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Maintenance Manual
1985 & ON MODEL 208 SERIES
Member of GAMA COPYRIGHT © 1995 CESSNA AIRCRAFT COMPANY WICHITA, KANSAS, USA D2078-25-13
1 AUGUST 1995 REVISION 25
1 MARCH 2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL REVISION SUMMARY 1.
2.
General A.
This section shows a table that gives operators and maintenance personnel a list of the changes that were made to different documents in the manual as part of the current revision.
B.
The table has three columns. The three columns are entitled: Chapter-Section-Subject, Document Title, and Action.
DeÞnition A.
Columns (1) Chapter-Section-Subject - This column gives the manual location for each document in the revision. (2) Document Title - This column gives the name of the document as it is given at the top of the actual document and in the Table of Contents. (3) Action - This column gives the step you must complete to include this revision in a paper copy of the manual. There are three different steps that can be given. The three steps are ADD, REPLACE, and REMOVE. NOTE:
B. 3.
This column does not apply to CD-ROM, DVD-ROM, or internet delivered publications.
Rows (1) Each row gives all the necessary data for one document that is part of the current revision.
Procedure A.
Find the manual location for each document in the revision as given by the data in the Chapter-SectionSubject column. NOTE:
For data about document page numbers and how to put them in the manual, refer to Introduction, Page Number System. Also, pages 1 - 99 are used for both "General", and "Description And Operation" documents.
B.
Make sure that the title of the document that you remove and/or the title of the document that you add agree with the data in the Document Title column of the table.
C.
Complete the step given in the Action column as directed below: (1) ADD - This step is for a new document that was not in the manual before. Put it in the applicable location. (2) REPLACE - This step is for an existing document that was changed in the current revision. Remove the existing document and put the revised one in its place. (3) REMOVE - This step is for an existing document that is no longer applicable. Remove it from the manual.
CHAPTER SECTION SUBJECT
DOCUMENT TITLE
ACTION
REVISION SUMMARY
REPLACE
INSPECTION PROGRAM CHANGE SUMMARY
REPLACE
PUBLICATION TITLE PAGE
REPLACE
Introduction
LIST OF EFFECTIVE PAGES
REPLACE
Introduction
TABLE OF CONTENTS
REPLACE
Introduction
ICA SUPPLEMENT LIST
ADD
REVISION SUMMARY © Cessna Aircraft Company
Page 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CHAPTER SECTION SUBJECT
DOCUMENT TITLE
ACTION
Introduction
LIST OF CHAPTERS
REMOVE
Introduction
LIST OF REVISIONS
REPLACE
Introduction
LIST OF PUBLICATIONS
REPLACE
Chapter 04
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 04
TABLE OF CONTENTS
REPLACE
04-00-00
AIRWORTHINESS LIMITATIONS
REPLACE
04-10-00
TYPICAL INSPECTION TIME LIMITS
REPLACE
04-10-01
SEVERE INSPECTION TIME LIMITS
REPLACE
04-11-00
REPLACEMENT TIME LIMITS
REPLACE
Chapter 05
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 05
TABLE OF CONTENTS
REPLACE
05-00-00
TIME LIMITS/MAINTENANCE CHECKS
REPLACE
05-10-00
INSPECTIONS
REPLACE
05-10-01
INSPECTION TIME LIMITS
REPLACE
05-11-00
COMPONENT TIME LIMITS
REPLACE
05-11-01
COMPONENT TIME LIMITS - RUSSIAN CERTIFIED AIRPLANES
REPLACE
05-13-00
SUPPLEMENTAL INSPECTION DOCUMENT
REPLACE
05-14-00
LISTING OF SUPPLEMENTAL INSPECTIONS
REPLACE
05-15-00
SCHEDULED INSPECTION PROGRAM
REPLACE
05-15-0A
INSPECTION DOCUMENT 0A
REPLACE
05-15-01
INSPECTION DOCUMENT 01
REPLACE
05-15-02
INSPECTION DOCUMENT 02
REPLACE
05-15-03
INSPECTION DOCUMENT 03
REPLACE
05-15-04
INSPECTION DOCUMENT 04
REPLACE
05-15-05
INSPECTION DOCUMENT 05
REPLACE
05-15-06
INSPECTION DOCUMENT 06
REPLACE
05-15-07
INSPECTION DOCUMENT 07
REPLACE
05-15-08
INSPECTION DOCUMENT 08
REPLACE
05-15-09
INSPECTION DOCUMENT 09
REPLACE
05-15-10
INSPECTION DOCUMENT 10
REPLACE
05-15-11
INSPECTION DOCUMENT 11
REPLACE
05-15-12
INSPECTION DOCUMENT 12
REPLACE
05-15-13
INSPECTION DOCUMENT 13
REPLACE
05-15-14
INSPECTION DOCUMENT 14
REPLACE
REVISION SUMMARY © Cessna Aircraft Company
Page 2 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CHAPTER SECTION SUBJECT
DOCUMENT TITLE
ACTION
05-15-15
INSPECTION DOCUMENT 15
REPLACE
05-15-16
INSPECTION DOCUMENT 16
REPLACE
05-15-17
INSPECTION DOCUMENT 17
REPLACE
05-15-18
INSPECTION DOCUMENT 18
REPLACE
05-15-19
INSPECTION DOCUMENT 19
REPLACE
05-15-20
INSPECTION DOCUMENT 20
REPLACE
05-15-21
INSPECTION DOCUMENT 21
REPLACE
05-15-22
INSPECTION DOCUMENT 22
REPLACE
05-15-MA
INSPECTION DOCUMENT MA
REPLACE
05-15-MB
INSPECTION DOCUMENT MB
REPLACE
05-15-MD
INSPECTION DOCUMENT MD
REPLACE
05-15-ME
INSPECTION DOCUMENT ME
REPLACE
05-15-MF
INSPECTION DOCUMENT MF
REPLACE
05-15-MG
INSPECTION DOCUMENT MG
REPLACE
05-15-MH
INSPECTION DOCUMENT MH
REPLACE
05-15-MI
INSPECTION DOCUMENT MI
REPLACE
05-15-MJ
INSPECTION DOCUMENT MJ
REPLACE
05-15-MK
INSPECTION DOCUMENT MK
REPLACE
05-15-ML
INSPECTION DOCUMENT ML
REPLACE
05-20-01
EXPANDED INSPECTION
05-50-00
UNSCHEDULED MAINTENANCE CHECKS
REPLACE
Chapter 06
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 06
TABLE OF CONTENTS
REPLACE
AIRPLANE ZONING - DESCRIPTION AND OPERATION
REPLACE
Chapter 12
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 12
TABLE OF CONTENTS
REPLACE
Chapter 12
LIST OF TASKS
REPLACE
ENGINE OIL SYSTEM - SERVICING
REPLACE
Chapter 26
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 26
TABLE OF CONTENTS
REPLACE
Chapter 26
LIST OF TASKS
REPLACE
PORTABLE FIRE EXTINGUISHING - DESCRIPTION AND OPERATION
REPLACE
Chapter 27
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 27
TABLE OF CONTENTS
REPLACE
06-20-01
12-11-02
26-20-00
ADD
REVISION SUMMARY © Cessna Aircraft Company
Page 3 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CHAPTER SECTION SUBJECT Chapter 27
DOCUMENT TITLE
ACTION
LIST OF TASKS
REPLACE
27-30-02
ELEVATOR TRIM - MAINTENANCE PRACTICES
REPLACE
27-30-02
ELEVATOR TRIM - INSPECTION/CHECK
REPLACE
Chapter 30
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 30
TABLE OF CONTENTS
REPLACE
Chapter 30
LIST OF TASKS
REPLACE
30-11-00
TKS ANTI-ICE SYSTEM - INSPECTION/CHECK
REPLACE
30-11-10
TKS ANTI-ICE FLUID TANK COMPONENTS - MAINTENANCE PRACTICES
REPLACE
30-11-11
TKS ANTI-ICE SYSTEM - MAINTENANCE PRACTICES
REPLACE
30-11-20
TKS ANTI-ICE LEADING EDGE POROUS PANEL - MAINTENANCE PRACTICES
REPLACE
30-11-20
TKS ANTI-ICE LEADING EDGE POROUS PANEL - ADJUSTMENT/TEST
REPLACE
30-11-30
TKS ANTI-ICE FLUID DISTRIBUTION SYSTEM - MAINTENANCE PRACTICES
REPLACE
Chapter 32
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 32
TABLE OF CONTENTS
REPLACE
Chapter 32
LIST OF TASKS
REPLACE
32-10-00
MAIN LANDING GEAR - MAINTENANCE PRACTICES
REPLACE
32-10-00
MAIN LANDING GEAR - INSPECTION/CHECK
REPLACE
32-40-00
WHEELS AND BRAKES - MAINTENANCE PRACTICES
REPLACE
Chapter 61
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 61
TABLE OF CONTENTS
REPLACE
Chapter 61
LIST OF TASKS
REPLACE
61-11-00
PROPELLER (MCCAULEY) - DESCRIPTION AND OPERATION
REPLACE
61-11-00
PROPELLER (MCCAULEY) - MAINTENANCE PRACTICES
REPLACE
61-11-00
DYNAMIC BALANCING (MCCAULEY) - ADJUSTMENT/TEST
REPLACE
61-11-00
PROPELLER (MCCAULEY) - INSPECTION/CHECK
REPLACE
Chapter 74
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 74
TABLE OF CONTENTS
REPLACE
SPARK IGNITERS - MAINTENANCE PRACTICES
REPLACE
Chapter 79
LIST OF EFFECTIVE PAGES
REPLACE
Chapter 79
TABLE OF CONTENTS
REPLACE
CHIP DETECTORS - MAINTENANCE PRACTICES
REPLACE
74-21-00
79-31-00
REVISION SUMMARY © Cessna Aircraft Company
Page 4 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION PROGRAM CHANGE SUMMARY 1.
General A.
2.
This section gives a list of changes that were made to the inspection program during Revision 25. This list gives operators and maintenance personnel a record of the changes to the inspection program and is for reference only.
Changes A.
Changes made to the Chapter 5 documents: (1) 05-00-00 - TIME LIMITS/MAINTENANCE CHECKS: (a) ClariÞed the 100-hour inspection requirements for airplanes operated for hire. (b) Changed Continuous Inspection Program references to Scheduled Inspection Programs. (2) 05-10-00 - INSPECTIONS: (a) Revised this document to specify components that are to be inspected in accordance with the manufacturers requirements. (3) 05-10-01 - INSPECTION TIME LIMITS: (a) This document was updated to reßect the changes listed in each of the inspection documents shown below. (4) 05-11-00 - COMPONENT TIME LIMITS: (a) Updated the vacuum system inspection information to refer to the manufacturer service letters. (b) Updated the propeller inspection information to refer to the correct manufacturer documents. (c) Added the overhaul requirements for propellers. (5) 5-15-00 SCHEDULED INSPECTION PROGRAM: (a) Changed Continuous Inspection Program references to Scheduled Inspection Programs. (6) 5-15-0A INSPECTION DOCUMENT 0A: (a) Added Item Code A052001 to this Inspection Document and tied it to Task 5-20-01-280. (7) 5-15-01 INSPECTION DOCUMENT 1: (a) Added Item Code B301102 and tied it to Task 30-11-00-721. (8) 5-15-12 INSPECTION DOCUMENT 12: (a) Added the 2661215-9 actuator to Task 27-30-02-641 for Item Code C273003. (9) 5-20-01 EXPANDED INSPECTIONS: (a) Added this document to the manual. (b) Added Inspection Task 05-20-01-280 for the Aircraft Records Check. (10) 5-50-00 UNSCHEDULED MAINTENANCE CHECKS: (a) ClariÞed the propeller inspections required after a lightning strike.
B.
Changes to Tasks (1) The list that follows contains tasks that were revised or added in this revision.
Task Number
Task Title
Action
5-20-01-280
Aircraft Records Check.
Added
27-30-02-641
Elevator Trim Tab Actuator (2661215-1 and 2661215-9) Lubrication.
30-11-00-721
Inboard TKS Wing Panel Pressurization Functional Check.
Added
32-10-00-220
Main Landing Gear Detailed Inspection.
Revised
32-10-00-221
Center-Spring and Main Gear-Spring Interface Area Special Detailed (Corrosion Inspection and Repair).
Revised
61-11-00-720
McCauley Propeller Functional Check.
Revised
Revised
INSPECTION PROGRAM CHANGE SUMMARY © Cessna Aircraft Company
Page 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
Page 1
Mar 1/2012
Pages 1-6
Aug 2/2004
Page 1
Mar 1/2012
SERVICE KIT LIST
Pages 1-6
Dec 1/2006
LIST OF PUBLICATIONS
Pages 1-7
Mar 1/2012
Page 1
DELETED
00-Title 00-List of Effective Pages 00-Record of Revisions 00-Record of Temporary Revisions 00-Table of Contents LIST OF REVISIONS INTRODUCTION ICA SUPPLEMENT LIST
LIST OF CHAPTERS
00 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Mar 1/2012
RECORD OF REVISIONS Revsion Number
Date Inserted
Date Removed
Page Number
Revsion Number
Date Inserted
Date Removed
Page Number
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS LIST OF REVISIONS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page 1 Page 1
INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cross-Reference Listing of Popular Name Versus Model Numbers and Serials . . . Coverage and Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Temporary Revisions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AeroÞche (MicroÞche) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compact Disc (CD-ROM) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Caravan Service Bulletins and Service Kits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Using the Maintenance Manual. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Effectivity Pages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Revision Filing Instructions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Identifying Revised Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Warnings, Cautions and Notes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cessna Propeller Aircraft Customer Care Supplies and Publications Catalog . . . . . Customer Comments on Manual . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page 1 Page 1 Page 1 Page 2 Page 2 Page 2 Page 2 Page 2 Page 3 Page 5 Page 5 Page 6 Page 6 Page 6 Page 6
ICA SUPPLEMENT LIST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page 1
SERVICE KIT LIST. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page 1
LIST OF PUBLICATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . List of Manufacturers Technical Publications Available Through Cessna . . . . . . . . . . Radio Manufacturer Manuals. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sigmatek Inc./ARC Avionics Manufacturer Manuals . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page Page Page Page
00 - CONTENTS © Cessna Aircraft Company
1 1 5 6
Page 1 of 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL LIST OF REVISIONS 1.
General A.
This Maintenance Manual includes the Þrst issue and the revisions listed in Table 1. The revisions must be included in this manual as they are supplied. Make sure the information in this manual is current and that the latest maintenance and inspections procedures are available.
Table 1. Original Issue 1 August 1995 Revision Number
B.
Date
Revision Number
Date
1
1 April 1996
2
3 September 1996
3
3 March 1997
4
2 September 1997
5
1 June 1998
6
1 March 1999
7
15 October 1999
8
1 March 2000
9
1 September 2000
10
1 March 2001
11
4 September 2001
12
3 December 2001
13
3 June 2002
14
2 September 2002
15
5 May 2003
16
3 November 2003
17
2 August 2004
18
3 January 2005
19
2 January 2006
20
1 December 2006
21
1 March 2008
22
1 April 2010
23
1 July 2010
24
1 June 2011
25
1 March 2012
FAA Approved Airworthiness Limitations are incorporated in this maintenance manual as Chapter 4. Revisions to Chapter 4 are dated as approved by the FAA. The revisions listed in Table 2 must be included in Chapter 4 as they are issued. Make sure the maintenance information required is current under Parts 43.16 and 91.409 of the Federal Aviation Regulations.
Table 2. Original Issue 8 May 1990 Revision Number
Date
Revision Number
Date
1
24 June 1993
2
22 August 1995
3
8 May 1998
4
25 May 2000
5
9 February 2001
6
26 November 2001
7
4 April 2002
8
16 August 2002
9
20 December 2002
10
11 April 2003
11
6 October 2003
12
17 June 2004
13
6 December 2004
14
16 December 2005
15
31 January 2008
16
15 January 2010
17
8 May 2010
18
25 May 2010
19
21 June 2010
20
5 October 2010
LIST OF REVISIONS © Cessna Aircraft Company
Page 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INTRODUCTION 1.
General
WARNING: All the inspection intervals, replacement time limits, overhaul time limits, inspection methods, life limits, cycle limits, etc., recommended by Cessna are based on the use of new, repaired, or overhauled Cessna approved parts. The data in Cessna’s maintenance manuals and parts catalogs is not applicable if the parts are designed, built, repaired, overhauled, and/or approved by entities other than Cessna. The purchaser is warned not to rely on such data for non-Cessna parts. The purchaser must get from the manufacturer and/or seller of non-Cessna parts, the inspection intervals, replacement time limits, overhaul time limits, inspection methods, life limits, cycle limits, etc., for all non-Cessna parts.
2.
A.
The procedures in this manual are based on data that is available at the time of publication. This manual is updated, supplemented, and will automatically change by all data that is issued in the Service Newsletters, Service Bulletins, Supplier Service Notices, Publication Changes, Revisions, Reissues and Temporary Revisions. The revisions become part of and are specifically incorporated in this publication. The user must know the latest changes to this publication through data available at the Cessna Authorized Service Stations or from the Cessna Product Support subscription services. Cessna Service Stations are supplied with a group of supplier publications. The supplier publications will give disassembly, overhaul, and parts breakdown data for some of the different supplier items. Suppliers publications are updated, supplemented, and specifically changed by the supplier issued revisions and service data which may be issued by Cessna. This will automatically change this publication and is communicated to the field by the Cessna Authorized Service Stations and/or by Cessna’s subscription services.
B.
The inspection, maintenance and parts requirements for Supplemental Type Certification (STC) installations are not included in this manual. When an STC installation is included on the airplane, the parts of the airplane affected by the installation must be checked. Do an inspection in accordance with the inspection program published by the owner of the STC. The STC installations may change systems interface, operating characteristics and component loads or stresses on adjacent structures. The Cessna-provided inspection criteria may not be correct for airplanes with STC installations.
C.
The revisions, temporary revisions and reissues can be purchased from your Cessna Service Station or directly from Cessna Propeller Product Support, Dept. 751, Cessna Aircraft Company, P.O. Box 7706, Wichita, Kansas 67277-7706.
D.
The data in this Maintenance Manual is applicable to all U.S. and Foreign Certified Model 208 airplanes. Data that applies to a particular country is identified in the chapter(s) affected.
E.
The Cessna Service Stations are supplied with all the supplemental maintenance data for this manual. They have the latest authoritative recommendations for servicing the Cessna airplanes. It is recommended that Cessna owners use the knowledge and experience of the Cessna Service Organization.
Cross-Reference Listing of Popular Name Versus Model Numbers and Serials A.
The airplanes are certified by the model number. Names are often used for marketing purposes. To supply the same method of referring to the airplanes, the model number will be used in this manual. The airplane name may be used to distinguish the different types of the same model. The following table supplies a list of the names, model numbers and serial numbers.
INTRODUCTION © Cessna Aircraft Company
Page 1 Aug 2/2004
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
NAME
MODEL
SERIALS BEGINNING
CARAVAN I
208
20800001
CARAVAN I
208B
208B0001
CARGOMASTER
208
20800113
SUPER CARGOMASTER
208B
208B0045
GRAND CARAVAN
208B
208B0214
3.
Coverage and Format A.
4.
Temporary Revisions A.
5.
Information that is available may be supplied by a temporary revision. This is used to supply, without delay, new data that will assist in maintaining safe flight/ground operations. The temporary revisions are numbered consecutively with the ATA chapter assignment. Page numbering uses the threeelement number which matches the maintenance manual. The temporary revisions are normally included in the maintenance manual at the next scheduled revision.
Aerofiche (Microfiche) A.
6.
The Cessna Model 208 Series 1985 & On Maintenance Manual, is prepared in accordance with the Air Transport Association (ATA) Specification Number 100 for Manufacturer's Technical Data dated February 1, 1983.
This maintenance manual is prepared for aerofiche presentation. A List of Chapters, which identifies the initial fiche/frame of each chapter section, has been assembled and included in the introduction. The List of Chapters is used to assist in the use of the aerofiche index. The List of Chapters is shown in the upper left frame of each aerofiche card.
Compact Disc (CD-ROM) A.
The following manuals and service publications are available on one CD-ROM (Compact Disc-Read Only Memory): NOTE: 1. 2. 3. 4. 5.
7.
The listed publications are kept up to date through routine revisions. Cessna Model 208 Series 1985 & On Maintenance Manual Structural Repair Manual Illustrated Parts Catalog Wiring Diagram Manual Avionic Installation Service/Parts Manual
Caravan Service Bulletins and Service Kits A.
Caravan Service Bulletins (CABs) and Service Kits (SKs) are supplied to inform and/or authorize modification to the airplane and/or system. As service kits are supplied, they will be included and appear in the Service Kit List, located previous to the introduction. The list of service kits uses five columns to list data. (1) Service Kit Number - The service kit number column identifies the kit by number. Service bulletins are numbered sequentially. (2) Title - The title column identifies the service kit by name. It is the same title displayed on page one of the service kit. (3) Service Bulletin Number - This column shows and refers to the CAB or SSP which refers to the service kit. (4) Service Kit Date - The service kit date column shows the initial date the kit became active.
INTRODUCTION © Cessna Aircraft Company
Page 2 Aug 2/2004
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (5)
8.
Manual Incorporation - The manual incorporation column shows if the service kit has been included in the maintenance manual (Incorp), if the service kit had no effect on the maintenance manual (No Effect), or if the service kit has not been included at the time of the revision (dashed lines).
Using the Maintenance Manual A.
Division of Subject Matter. (1) Cessna Model 208 Series 1985 & On Maintenance Manual is divided into four sections. The four sections are divided into chapters. Each chapter has its own effectivity page and table of contents. The following shows the manual divisions: (a) Section 1 - Airplane General
Chapter
Title
4
Airworthiness Limitations
5
Time Limits/Maintenance Checks
6
Dimensions and Areas
7
Lifting and Shoring
8
Leveling and Weighing
9
Towing and Taxiing
10
Parking, Mooring, Storage and Return to Service
11
Placards and Markings
12
Servicing (b)
Section 2 - Airframe Systems
Chapter
Title
20
Standard Practices - Airframe
21
Air Conditioning
22
Auto Flight
23
Communications
24
Electrical Power
25
Equipment/Furnishings
26
Fire Protection
27
Flight Controls
28
Fuel
30
Ice and Rain Protection
31
Indicating/Recording Systems
32
Landing Gear
33
Lights
34
Navigation
35
Oxygen
INTRODUCTION © Cessna Aircraft Company
Page 3 Aug 2/2004
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Chapter
Title
36
Pneumatic
37
Vacuum
38
Water/Waste (c)
Section 3 - Structures
Chapter
Title
51
Standard Practices and Structures - General
52
Doors
53
Fuselage
55
Stabilizers
56
Windows
57
Wings (d)
Section 4 - Power Plant
Chapter
Title
61
Propeller
71
Power Plant
73
Engine Fuel and Control
74
Ignition
76
Engine Controls
77
Engine Indicating
78
Exhaust
79
Oil
80
Starting B.
Page Numbering System. (1) The page numbering system used in the Cessna Model 208 Series 1985 & On Maintenance Manual has three-element numbers separated by dashes. The page number and date are located to the right of the three-element number on each page.
(2)
When the chapter/system element number is followed with zeros in the section/subsystem and subject/unit element number (21-00-00), the data is applicable to the entire system.
INTRODUCTION © Cessna Aircraft Company
Page 4 Aug 2/2004
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (3) (4) (5)
(6)
(7)
(8)
(9)
When the section/subsystem element number is followed with zeros in the subject/unit element number (21-51-00), the data is applicable to the subsystem within the system. Subject/unit element number is used to identify data applicable to units in the subsystems. The subject/unit element number progresses in sequence from the number -01 in accordance with the number of subsystem units requiring maintenance information. All system/subsystem/unit (chapter/section/subject) maintenance instructions are separated into specific types of data: 1. Description and Operation 2. Troubleshooting 3. Maintenance Practices Blocks of page numbers are used to identify the type of information: Page 1 through 99 - Description and Operation Page 101 through 199 - Troubleshooting Page 201 through 299 - Maintenance Practices The description and operation or troubleshooting information may not be necessary for simple units. For these units, the pages are omitted. When subtopics are small, they may be combined into a single topic titled Maintenance Practices. Maintenance Practices is a combination of subtopics. It may include Servicing, Removal/Installation, Adjustment/Test, Cleaning/Painting or Approved Repairs. Large subtopics may be treated as an individual topic. The list below shows the page numbering for individual topics: Page 301 through 399 - Servicing Page 401 through 499 - Removal/Installation Page 501 through 599 - Adjustment/Test Page 601 through 699 - Inspection/Check Page 701 through 799 - Cleaning/Painting Page 801 through 899 - Approved Repairs A typical page number:
(10) The illustrations are also included in the page block numbering system. For example, all illustrations in a Maintenance Practice section will begin with the number 2 (i.e., Figure 201, Figure 202, etc.). All illustrations within an Approved Repair section will begin with the number 8 (Figure 801, Figure 802, etc.). 9.
Effectivity Pages A.
10.
A List of Effective Pages is supplied at the beginning of each maintenance manual chapter. All pages in the specific chapter are listed in numerical order on the effectivity page(s) with the date of issue for each page.
Revision Filing Instructions A.
Regular Revision. (1) You can determine which pages to remove or insert into the maintenance manual by the effectivity page. The effectivity page lists the pages in sequence by the three-element number (chapter/section/subject) and then by page number. When two pages display the same
INTRODUCTION © Cessna Aircraft Company
Page 5 Aug 2/2004
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL three-element number and page number, the page with the most recent date is put into the maintenance manual. The date column on the corresponding chapter effectivity page must agree with the active page. B.
11.
12.
Identifying Revised Material A.
Additions or revisions to the text in an section will be identified by a revision bar in the left margin of the page and adjacent to the change.
B.
The minimum revisable unit in the Maintenance Manual is a subject (refer to Page Numbering System for definition of subject). All the pages in a subject will have the same date regardless if the data in that page has changed or not.
C.
When large changes are made to the text in an existing section, revision bars will be shown for the full length of the text.
D.
The illustrations shown in this manual may have hand indicators for revisions before March 1, 2000 or Revision 8, that direct attention to the change. Future revisions to the illustrations, beginning at Revision 8, will be shown by a revision bar along the entire vertical length of one side of illustration. No hand indicators will be shown.
Warnings, Cautions and Notes A.
13.
Warnings, cautions and notes are applicable to the instructions. These adjuncts to the text highlight or emphasize important data. (1) WARNING - This is applicable to the data that follows it. A warning puts attention to the use of the instructions, materials, methods, or limits that must be obeyed to prevent an injury or death. (2) CAUTION - This is applicable to the data that follows it. A caution puts attention to the use of the instructions, materials, methods, or limits which must be obeyed to prevent damage to the airplane or equipment. (3) NOTE - A NOTE follows the applicable instructions. It is for data only.
Cessna Propeller Aircraft Customer Care Supplies and Publications Catalog A.
14.
Temporary Revision. (1) File the temporary revisions in the applicable chapter(s) in accordance with the filing instructions shown on the first page of the temporary revision. (2) The rescission of a temporary revision is completed by including it into the maintenance manual or by a superseding temporary revision. A Record of Temporary Revisions is supplied in the Temporary Revision List. The temporary revision list is located before the Introduction-List of Effective Pages. A Manual Incorporation Date column on the Temporary Revision List page will indicate the date the Temporary Revision was incorporated, thus authorizing the rescission of the temporary revision.
A Cessna Propeller Aircraft Customer Care Supplies and Publications Catalog is available from a Cessna Service Station or directly from Cessna Propeller Product Support Dept. 751 Cessna Aircraft Company, P.O. Box 7706, Wichita, Kansas 67277-7706. The catalog lists all publications and Customer Care Supplies available from Cessna for prior year models as well as new products. To maintain this catalog in a current status, it is revised yearly and issued in paper form.
Customer Comments on Manual A.
Cessna Aircraft Company has worked to furnish you with an accurate, useful, up-to-date manual. This manual can be improved with your help. Please use the return card, supplied with your manual, to report any errors, discrepancies, and omissions in this manual as well as any general comments you want to make.
INTRODUCTION © Cessna Aircraft Company
Page 6 Aug 2/2004
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ICA SUPPLEMENT LIST ICA Supplement Number
Revision
Title
ICA Supplement Date
Manual Incorporation Date
ICA-208-2300001
B
Dual GMA1347 Audio Panel ICA Supplement
Jul 15/2011
Pending Incorporation
ICA-208-2600001 ICA-208-3000001
A
Cabin Fire Extinguisher Instl
Mar 23/2009
Mar 1/2012
A
Low Airspeed Awareness ICA
Oct 7/2008
Pending Incorporation
ICA-208-3000002
-
Low Airspeed Awareness ICA
Oct 21/2008
Pending Incorporation
ICA-208-3000003
B
TKS W/ Fairing ICA Supplement
Dec 3/2010
Pending Incorporation
ICA-208-3100001
B
L-3 Communications FA2100 CVDR
Jul 27/2010
Pending Incorporation
ICA-208-3500001
C
Portable Therapeutic Oxygen Supply Installation
Aug 1 20110
Pending Incorporation
ICA SUPPLEMENT LIST © Cessna Aircraft Company
Page 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SERVICE KIT LIST Service Kit Number
Title
SK208-1
Service Bulletin Number
Service Kit Date
Manual Incorporation
Oil Breather Line Improvement
CAB85-06
9/13/85
No Effect
SK208-2
Pilot and Copilot Seat Locking Pin
CAB85-02
8/09/85
3/01/99
SK208-3
Volt/Ammeter Switch Replacement
CAB85-11
11/08/85
No Effect
SK208-4
Improved Courtesy Light Installation
SSP85-09
12/01/85
No Effect
SK208-5
Optional RH Panel Improvement
SSP85-13
12/20/85
No Effect
SK208-6
Jackscrew Replacement (C/B Panel)
CAB85-12
11/08/85
No Effect
SK208-7
Inertial Separator Linkage Support Bracket Replacement
CAB85-14
12/06/85
No Effect
SK208-8
Fuel Transmitter and Gage ModiÞcation
CAB85-09
11/01/85
No Effect
SK208-9
Hourmeter Air Switch Installation
CAB85-16
12/20/85
3/01/99
SK208-10
Cargo Pod Water Drains Installation
CAB86-27
6/19/87
3/01/99
SK208-11
Electrical Bonding Ground Strap Installation
CAB86-21
7/25/86
3/01/99
SK208-12
Maximum Takeoff Weight Increase ModiÞcation
CAB85-13
11/15/85
3/01/99
SK208-13
208 Extended Vertical Tail Installation
CAB85-13
11/15/85
3/01/99
SK208-14
Fin/Rudder Hinge Bracket
CAB85-13
11/15/85
3/01/99
SK208-15
LH Instrument Panel ModiÞcation
CAB86-04
10/24/86
No Effect
SK208-16
Torque Indicator Installation Improvement
CAB85-16
12/13/85
3/01/99
SK208-17
Improved Drive For Standby Alternator
CAB86-18R1
12/07/90
3/01/99
SK208-18
Fuel Shutoff Valve Screw Replacement
CAB86-03
4/11/86
No Effect
SK208-19
Cowl Door Latch Pin ModiÞcation
CAB86-10
4/25/86
No Effect
SK208-20
Fuel Selector Off Warning System Installation
CAB86-08
10/10/86
3/01/99
SK208-21
Nose/Main Landing Gear Wheel and Bearing Improvements
CAB88-02
3/04/88
No Effect
SK208-22
Inertial Separator and Induction Air Plenum ModiÞcation
CAB87-12
9/18/87
No Effect
SK208-23
Improved Secondary Exhaust System
CAB86-32R1
12/21/90
3/01/99
SK208-24
External Start Contactor Suppressor
CAB86-24
8/29/86
3/01/99
SK208-25
Fuel Reservoir and Flapper Valve ModiÞcation
CAB87-13
10/16/87
3/01/99
SK208-26
Low Fuel Level Transmitter Nut Retention
CAB86-25
9/05/86
No Effect
SK208-27
Fuselage Sealing Procedures
CAB86-25
9/05/86
No Effect
SK208-28
Cancelled
-----
- - - -
- - - -
SK208-29
Improved Door Restraint Installation
CAB88-16
4/15/88
3/01/99
SERVICE KIT LIST © Cessna Aircraft Company
Page 1 Dec 1/2006
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Service Bulletin Number
Service Kit Date
Manual Incorporation
CAB88-05
3/18/88
No Effect
-----
- - - -
- - - -
Circuit Breaker Panel Cover Installation
CAB88-26
7/29/88
3/01/99
SK208-33
Circuit Breaker Panel Cover Installation
CAB88-26
7/29/88
3/01/99
SK208-34
Circuit Breaker Panel Cover Installation
CAB88-26
7/29/88
3/01/99
SK208-35
Inertial Separator/Induction Air Plenum ModiÞcation
CAB88-06
9/18/87
No Effect
SK208-36
Outboard Wing Fuel Drain Valve Installation
CAB88-06
3/18/88
3/01/99
SK208-37
Cargo Barrier Attachment Improvement
CAB88-18
5/06/88
3/01/99
SK208-38
Elevator Trim Chain Cover Assembly
CAB88-07
3/18/88
No Effect
SK208-39
Ground Wire For Windshield Anti-Ice Panel
CAB87-05
4/24/87
3/01/99
SK208-40
Large Oil Cooler Installation
CAB87-17 SNL90-07
4/23/93
3/01/99
SK208-41
Crew Step Structural Improvement
CAB89-14 & CAB89-12
6/02/89
- - - -
SK208-42
Radio Cover Watershield Installation and Cabin Top Sealing
CAB88-08
3/18/88
3/01/99
SK208-43
Cowl Scupper and Drain Installation
CAB88-41
12/09/88
3/01/99
SK208-44
Crew Step Structural Improvements
CAB89-14
6/02/89
3/01/99
SK208-45
Inertial Vane Separator Vane Improvements
CAB88-15
4/08/88
3/01/99
SK208-46
Partition Net Placard
CAB88-25
7/29/88
3/01/99
SK208-47
Cowl Door Rub Strips
CAB88-17
4/22/88
No Effect
SK208-48
Tow Limit ModiÞcation
CAB88-09
3/18/88
3/01/99
SK208-49
Pitot Static System Line Replacement
CAB88-36
11/18/88
3/01/99
SK208-50
Flap System Adjustable Interconnect Rod
CAB88-13
3/25/88
No Effect
SK208-51
Nose Gear Torque Link Replacement
CAB89-02R2
7/06/90
3/01/99
SK208-52
Wing Tank External Sump Installation
CAB88-23
7/29/88
3/01/99
SK208-53
Cancelled
- - - -
- - - -
- - - -
SK208-54
Static Source Selector Valve ModiÞcation
CAB88-34
11/11/88
No Effect
SK208-55
Exhaust Hanger ModiÞcation (Cargo Pod or Floats)
CAB88-20
5/31/88
3/01/99
SK208-56
Cancelled
- - - -
- - - -
- - - -
SK208-57
Fuel Pump Unit Drain Reservoir Installation
None
6/09/89
3/01/99
SK208-58
Pointer 3000-1 ELT Installation
CAB89-04
1/13/89
No Effect
SK208-59
Nose Gear Steering Bushing Replacement
CAB88-40
12/09/88
No Effect
Service Kit Number
Title
SK208-30
Control Column Aileron Cable Guard Improvement
SK208-31
Cancelled
SK208-32
SERVICE KIT LIST © Cessna Aircraft Company
Page 2 Dec 1/2006
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Service Bulletin Number
Service Kit Date
Manual Incorporation
Utility Seat Frame Padding Installation
CAB89-11
5/19/89
No Effect
SK208-61
Turbine Oil Hose Assembly Replacement
CAB89-01
1/06/89
3/01/99
SK208-62
Door Lock Pin Replacements
CAB89-03
1/13/89
No Effect
SK208-63
Wing Skin Stiffener
CAB88-32
10/14/88
No Effect
SK208-64
Bleed Air Pressure Relief Valve
CAB89-26 & CAB90-14
9/08/89
3/01/99
SK208-65
Bleed Air Pressure Relief Valve
CAB89-26 & CAB 90-14
9/08/89
3/01/99
SK208-66
Air Conditioner Compressor Bracket Replacements
CAB89-19
7/04/89
3/01/99
SK208-67
Nose Gear Fork Replacement - Extended Fork
None
6/20/89
3/01/99
SK208-68
Nose Gear Fork Replacement - Standard Fork
None
6/20/89
3/01/99
SK208-69
Enlarged Cargo Pod Heat Shield Installation
CAB89-30
11/03/89
- - - -
SK208-70
Airborne Bleed Air Valves RetroÞt Kit
CAB90-09
3/23/90
3/01/99
SK208-71
Inboard Flap Track and Roller Removal
CAB89-31
11/24/89
3/01/99
SK208-72
Upper Right Flap Bellcrank Support Replacement (EC25588)
CAB89-32R1
7/02/92
3/01/99
SK208-73
Cancelled
- - - -
- - - -
- - - -
SK208-74
Nose Gear Fork Tow Point
CAB90-05
3/02/90
3/01/99
SK208-75
Bulkhead Control Cable Cutout ModiÞcation and Floorboard Reinforcement
CAB90-21R2
6/21/91
- - - -
SK208-76
Bulkhead Control Cable Cutout ModiÞcation and Floorboard Reinforcement
CAB90-21
6/21/91
- - - -
None
Secondary Exhaust Duct Mid Support Removal
CAB90-27
10/26/90
11/03/03
SK208-77
Cancelled
-----
- - - -
- - - -
SK208-78
Air Conditioner Condenser Cowl Seal Replacement
CAB91-12
4/26/91
3/01/98
SK208-79
Improved Airspeed Warning Switch Installation
CAB90-32
12/07/90
3/01/98
SK208-80
675 SHP PT6A-114A Engine Installation
SNL90-07
10/12/90
3/01/98
SK208-81
Door Handle Plunger Replacement
CAB91-24
8/02/91
- - - -
SK208-82
Heater Valve Improvement
CAB91-26
8/02/91
3/01/98
SK208-83
Control Column Bearings and Aileron Cable Guard Replacements
CAB91-27
9/20/91
3/01/98
SK208-84
600 SHP PT6A-114 Engine Installation
SNL90-08
9/24/93
3/01/98
SK208-85
208A to 208 Cargomaster Conversion
None
11/06/90
- - - -
Service Kit Number
Title
SK208-60
SERVICE KIT LIST © Cessna Aircraft Company
Page 3 Dec 1/2006
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Service Bulletin Number
Service Kit Date
Manual Incorporation
Conversion of US 208 to Brazilian 208
None
11/06/90
- - - -
SK208-87
208 Cargo to Passenger Conversion
None
5/31/91
- - - -
SK208-88
Upper Rudder Hinge Bracket/Skin Reinforcement
CAB92-02
1/17/92
3/01/98
SK208-90
Firewall Fitting Reinforcement Installation
CAB91-09
7/02/92
3/01/98
SK208-91
Cancelled
- - - -
- - - -
- - - -
SK208-92
Flap Support ModiÞcation Installation
- - - -
8/28/92
3/01/98
SK208-93
Cancelled
- - - -
- - - -
- - - -
SK208-94
Engine Power Control Cable Replacement
CAB91-34
11/22/91
3/01/98
SK208-95
Engine Fuel Control Lever Replacement
CAB91-22
7/05/91
3/01/98
SK208-96
Cancelled
- - - -
- - - -
- - - -
SK208-97
Floorboard Reinforcement Installation
CAB93-03
4/30/93
- - - -
SK208-98
Flap Bellcrank Connecting Rod Clearance ModiÞcation
CAB92-15
7/24/92
- - - -
SK208-99
Lower Rudder Hinge Replacement
CAB92-18
10/16/92
3/01/98
SK208-100
Control Cable Clearance ModiÞcation
No Longer Available
- - - -
- - - -
SK208-101
Upper and Lower Crew Door Hinge Attach Repair
CAB93-05
4/30/93
3/01/99
SK208-102
Left Elevator Hinge Replacement
CAB93-12
6/30/95
3/01/99
SK208-103
Floorboard Reinforcement Installation
CAB93-03
3/12/93
- - - -
SK208-104
Upper and Center Rudder Hinge Bracket ModiÞcation
CAB93-06
4/09/93
3/01/99
SK208-105
Upper and Center Vertical Fin Rudder Hinge Replacement
CAB93-06
5/28/93
3/01/99
SK208-106
Right Elevator Hinge Replacement
CAB93-12
6/30/95
3/01/99
SK208-107
Elevator Center Pivot Arm Hinge Replacement
CAB93-12
8/27/93
3/01/99
SK208-108
Propeller Control Cable Replacement
CAB93-04
4/30/93
3/01/99
SK208-109
Flap Outboard Support Inspection and ModiÞcation
CAB93-11
8/27/93
- - - -
SK208-110
Low Fuel Level Switch Cover Replacement
CAB92-19
10/23/92
3/01/99
SK208-111
Throttle Control Quadrant Cover Doubler Installation
CAB93-09
7/23/93
3/01/99
None
Aileron Cable Seal Retainer Inspection
CAB93-13
8/27/93
No Effect
SK208-112
Bleed Air System Pressure Relief Valve Installation
CAB93-02
3/12/93
3/01/99
SK208-113
Windshield De-Ice ModiÞcation
CAB93-21
11/19/93
8/1/95
Service Kit Number
Title
SK208-86
SERVICE KIT LIST © Cessna Aircraft Company
Page 4 Dec 1/2006
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Service Kit Number
Title
Service Bulletin Number
Service Kit Date
Manual Incorporation
SK208-114
Cancelled
-----
- - - -
- - - -
SK208-115
Oxygen Lever Arm ModiÞcation
CAB93-19
10/22/93
8/1/95
SK208-116
Aileron Hinge Bolt Change
CAB93-18
10/08/93
- - - -
SK208-117
Crew Seat Vertical Adjust Bearing Replacement
CAB95-6
5/26/95
- - - -
SK208-118
Elevator Trim Chain Cover Installation
CAB94-09
3/25/94
8/1/95
SK208-119
Flap Switch Guard Installation/ModiÞcation
CA894-16
10/28/94
8/1/95
SK208-120
Cancelled
-----
- - - -
- - - -
SK208-121
Fuel Reservoir Structure ModiÞcation
CAB98-17
11/30/98
- - - -
SK208-122
Standby Alternator Drive Pulley With Drain and Oil Collection Can
CAB96-23
12/16/96
5/1/96
None
Nose Cap/Induction Inlet Support and Engine Flange Guard Installation
CAB95-1
None
8/1/95
SK208-123
Flap Bellcrank Attach Bolt Replacement
CAB95-11
6-30-95
8/1/95
None
Nose Landing Gear Spring Inspection and IdentiÞcation
CAB95-13
None
7/18/95
None
Engine Oil Placard
CAB96-9
11/09/96
5/01/96
None
Nose Gear Axle Spacer Replacement
CAB91-30
10/11/91
5/01/96
SK208-125
Ignitor Lead Cable Routing ModiÞcation
CAB96-11
3/01/96
- - - -
SK208-128
Avionics Cooling Fan Replacement
CAB96-10
2/09/96
- - - -
SK208-129
Engine Secondary Exhaust Duct Hanger ModiÞcation and Cowling Inspection
CAB00-8
5/29/00
11/03/03
SK208139A
Engine Secondary Exhaust Duct Hanger ModiÞcation
CAB00-9
9/29/00
11/03/03
SK208-141
Secondary Exhaust Duct Installation
- - - -
8/18/00
11/03/03
None
Yaw Damper AC Inverter Wiring Fuse Installation
CAB99-9
11/29/99
- - - -
SK208-142
Emergency Power Lever Shear Wire Installation
CAB01-15
12/17/01
12/17/01
None
Transcal Altitude Digitizer Wiring ModiÞcation
CAB01-16
None
- - - -
None
Flap System Inboard Forward Bellcrank Life Limit and Inspection
CAB02-1
None
11/03/03
None
Wing Spar Inspection and Inspection Plate ModiÞcation
CAB02-2
None
11/03/03
SK208-145
Plywood Floorboard Installation - 5/8 Inch
CAB02-3
2/25/02
11/03/03
None
Emergency Power Lever Control Cable Bracket Replacement
CAB02-4
None
11/03/03
None
Fuel-Vent Line Float Valves Inspection
CAB02-11
None
11/03/03
SERVICE KIT LIST © Cessna Aircraft Company
Page 5 Dec 1/2006
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Service Bulletin Number
Service Kit Date
Manual Incorporation
Flap System Inboard Forward Bellcrank Installation
CAB02-12
1/27/03
11/03/03
None
Engine Mount Bolts and Washers Inspection
CAB02-13
None
11/03/03
None
Flight Hourmeter Electrical Wiring ModiÞcation
CAB03-03
None
- - - -
None
Radar Indicator Inspection/ModiÞcation
CAB03-4
None
- - - -
None
Main Landing Gear Part Number and Serial IdentiÞcation Placard Installation
CAB03-7
None
- - - -
None
Flap System Bell Cranks Weld Inspection
CAB03-11
None
11/03/03
Service Kit Number
Title
SK208148A
SERVICE KIT LIST © Cessna Aircraft Company
Page 6 Dec 1/2006
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL LIST OF PUBLICATIONS 1.
List of Manufacturers Technical Publications Available Through Cessna A.
The following listed publications are necessary for the support of your Model 208. Contact the appropriate manufacturer to order the publications that are necessary for the support of your Model 208. NOTE:
The publications made by King Radio and Sigmatek Inc./ARC Avionics must be ordered directly from the manufacturer. The publication names, numbers and addresses are listed after this section.
Table 1. Chapter 21 - Air Conditioning Manufacturers Part Number
Publication Part Number
Publication Title
Manufacturer
Vent Blower
33E83-2
37E661-13
33E83 Blower Assembly Component Maintenance Manual (Vent Blower) with Illustrated Parts List
FL Aerospace Corp. Janitrol Aero Div. 4200 Surface Rd. Columbus, OH 43228
Flow Control Valve, and Temperature Control Valve
1H101-4, 1H102-2
CM-200-1
Airborne Component Overhaul Manual With Illustrated Parts List
Parker-HanniÞn Corp. Airborne Division 711 Taylor St. P.O. Box 4032 Elyria, OH 44036
Publication Part Number
Publication Title
Manufacturer
19000869-00
G1000 Caravan Line Maintenance Manual
Garmin International, Inc. 1200 E. 151st Street Olathe, Kansas 66062
19000303-72
Installation Manual
Garmin Inc.
Item
Cessna Part Number
Table 2. Chapter 22 - Auto Flight Item
Cessna Part Number
Manufacturers Part Number
GFC -700 Autopilot
Autopilot Servo
GSA 8X/GSM 85
International,
LIST OF PUBLICATIONS © Cessna Aircraft Company
Page 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 3. Chapter 24 - Electrical Power Item
Manufacturers Part Number
Publication Part Number
Publication Title
Manufacturer
DC to DC Power Converter
RG40
83706C
Removal/Installation Manual Model RG40 Regulated DC to DC Converter (FAA TSO-C71 Approved)
KGS Electronics, Inc. 418A E. Live Oak Ave. Arcadia, CA 91006-5690
Battery
G6381E
GSM68213
Gill Battery Service Manual
Teledyne Battery Products 840 W. Brockton Ave. P.O. Box 431 Redlands, CA 92373
Battery
30994-001
BA89-9/ 92-13
Battery Instruction Manual (NICAD) Marathon Battery Service Manual
Marathon Power Technologies Co. 8301 Imperial Dr. Waco, TX 76712-6588
Table 4.
Chapter 27 - Flight Controls Publication Part Number
Publication Title
Manufacturer
AT-RL1001-ICA
Rudder Gust Lock Kit Maintenance Manual with Illustrations
Aero Twin Inc. 2404 Merrill Field Dr. Anchorage, AK 99501
Cessna Part Number
Item
Rudder Gust Lock Kit
Cessna Part Number
Manufacturers Part Number
LIST OF PUBLICATIONS © Cessna Aircraft Company
Page 2 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 5. Chapter 30 - Deice Item
Manufacturers Part Number
Publication Part Number
Publication Title
Manufacturer
Propeller De-Ice Brush Block
3E2090-1
68-04714K13
Propeller Deice System Brush Block Assembly Overhaul Manual and Illustrated Parts Catalog
BF Goodrich Aerospace Div. Engineered Polymer Products 150 Division Dr. Wilmington, NC 28401
Propeller De-Ice
3E2205-4
68-04712D-13
Propeller Deice System Maintenance Manual
BF Goodrich Aerospace Div. Engineered Polymer Products 150 Division Dr. Wilmington, NC 28401
P70363680154
8304151-13
McCauley Electrothermal Deice Systems Service Parts Manual
McCauley Propeller Systems P.O. Box 7704 Wichita, KS 67277-7704
Cessna Part Number
Propeller De-Ice
9910587202
Pnuematic De-Ice Boots
BFG 30-10-31
Pneumatic De-Icer Installation, Maintenance and Repair Manual
Goodrich Corporation Aerospace Div. Ice Protection Systems 1555 Corporate Woods Parkway Uniontown, OH 44685
Pnuematic De-Ice Boots
BFG 30-10-70
Fastboot Pneumatic De-Icer Installation Instructions
Goodrich Corporation Aerospace Div. Ice Protection Systems 1555 Corporate Woods Parkway Uniontown, OH 44685
Publication Part Number
Publication Title
Manufacturer
19000869-00
G1000 Caravan Line Maintenance Manual
Garmin Inc.
Manufacturers Part Number
Publication Part Number
Publication Title
Manufacturer
1H75-14
CM200-1
Airborne Component Overhaul Manual with Illustrated Parts List
Parker-HanniÞn Corp. 711 Taylor St. P.O. Box 4032 Elyria, OH 44036
Table 6.
Chapter 34 - Navigation
Item
Cessna Part Number
Manufacturers Part Number
G1000 system components Table 7.
International,
Chapter 36 - Pneumatic
Item
Pressure Regulator
Cessna Part Number
LIST OF PUBLICATIONS © Cessna Aircraft Company
Page 3 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 8. Chapter 61 - Propellers Item
Manufacturers Part Number
Publication Part Number
Publication Title
Manufacturer
Propeller
HC-B3MN3/ M10083
135C-13
Composite Blade Inspection, Repair, and Overhaul Instructions
TRW Hartzell Propeller Division of TRW Inc. 1800 Covington Ave. Piqua, OH 45356
Propeller Spinner
D4898P D4899
127-13
Spinner Assembly Maintenance Instruction Guide
TRW Hartzell Propeller Division of TRW Inc. 1800 Covington Ave. Piqua, OH 45356
Propeller
MPC700
MPC700
McCauley MPC700 Propeller Overhaul Manual
McCauley Propeller Systems P.O. Box 7704 Wichita, KS 67277-7704
Propeller
MPC26
MPC26
MPC26 Owner/Operator Information Manual
McCauley Propeller Systems
Cessna Part Number
Table 9. Chapter 71 - Engine Item
Cessna Part Number
Manufacturers Part Number
Publication Part Number
Publication Title
Manufacturer
Engine 600SHP
991057 9-1
PT6A-114
3043512
Pratt & Whitney PT6A-114/116/135/135A Maintenance Manual
Pratt & Whitney Canada Inc. 100 Marie-Victorin Blvd. Longueuil, Quebec Canada J4G 1A1
Engine 600SHP
991057 9-1
PT6A-114
3043514
Pratt & Whitney PT6A-114/116/135/135A Illustrated Parts Catalog
Pratt & Whitney Canada Inc. 100 Marie-Victorin Blvd. Longueuil, Quebec Canada J4G 1A1
Engine 675SHP
991058 9-1
PT6A-114A
3043512
Pratt & Whitney PT6A114/114A/116/135/135A Maintenance Manual
Pratt & Whitney Canada Inc. 100 Marie-Victorin Blvd. Longueuil, Quebec Canada J4G 1A1
Engine 675SHP
991058 9-1
PT6A-114A
3043514
Pratt & Whitney PT6A114/114A/116/135/135A Illustrated Parts Catalog
Pratt & Whitney Canada Inc. 100 Marie-Victorin Blvd. Longueuil, Quebec Canada J4G 1A1
LIST OF PUBLICATIONS © Cessna Aircraft Company
Page 4 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 10. Chapter 80 - Starter/Generator Item
Publication Title
Manufacturers Part Number
Publication Part Number
Starter/ Generator
23081-22
23081023-3-13
Lear Siegler 23081 DC Starter/ Generator Overhaul Manual with Illustrated Parts List
Lucas Aerospace Power Equipment Corp. 777 Lena Drive Aurora, OH 44202
Starter/ Generator
23081-22
237006-13
Lucas Aerospace DC Generators and Starter-Generators Maintenance Manual
Lucas Aerospace Power Equipment Corp. 777 Lena Drive Aurora, OH 44202
Generator Control Unit Assembly
51539012N
51539012-13
Lear Siegler 51539-012 DC Generator Control Unit Overhaul Manual with Illustrated Parts List
Lucas Aerospace Power Equipment Corp. 777 Lena Drive Aurora, OH 44202
2.
Cessna Part Number
Manufacturer
Radio Manufacturer Manuals A.
The following publications must be ordered directly from Honeywell. Refer to the following list for address and telephone numbers:
UNITED STATES OPERATIONS Honeywell International, Inc., Aerospace Electronics Systems 1 Technology Center 23500 W. 105th Street Olathe, KS 66061 800-257-0726 - Direct line to the Product Support Operator for assistance from Customer Service, Product Specialists, Warranty, Field Engineering and Training. 913-782-0600 - 24 hour emergency service during nonworking hours. Telex: WUD (0) 4-229 Cable: KINGRAD EUROPEAN MARKETING OFFICE: Telephone: Switzerland 22 98 58 80 Telex: Switzerland 289445. Answer back: King CHGeneva, Switzerland Telephone: France 4 422 1747 Telex: France 150952. Answer back: KINGRAD Ivry-Le Temple MANUAL NUMBER
MANUAL TITLE
006-0180-01
Audio Panel/Marker Beacon Receiver Installation Manual (Type KMA-24)
006-5180-01
Audio Panel/Marker Beacon Receiver Overhaul/Maintenance Manual (Type KMA-24)
006-0179-03
VHF NAV/COMM Transceiver Installation Manual (Type KX-155/KX-165)
006-5179-03
Weather Radar Installation Manual (Type KX-155/KX-165)
LIST OF PUBLICATIONS © Cessna Aircraft Company
Page 5 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
MANUAL NUMBER
MANUAL TITLE
006-0191-01
Weather Radar Installation Manual (Type KWX-56, KI- 244, KA-126)
006-5193-01
Weather Radar Overhaul/Maintenance Manual (Type KWX-56, KI-244, KA-126)
006-0184-02
Automatic Direction Finder Installation Manual (Type KR-87, KI-227, KA-44B)
006-0534-01
Transponder Installation Manual (Type KT-79)
006-5534-00
Transponder Overhaul/Maintenance Manual (Type KT-79)
006-0176-03
DME/Master Indicator/Slave Indicator Installation Manual (Type KN-63/KDI-572/KDI-573/KDI-574)
006-5176-01
DME Overhaul/Maintenance Manual (Type KN-63)
006-5178-00
DME Indicator Overhaul/Maintenance Manual (Type KDI-572/KDI-573/KDI-574)
006-0193-01
Radio Magnetic Indicator Installation Manual (Type KNI-582)
006-5193-00
Radio Magnetic Indicator Installation Manual (Type KNI-582)
006-0137-03
VOR/LOC/GS Indicator Installation Manual (Type KI-202, KI-203, KI-204, KI-206, KI-207)
006-5137-03
VOR/LOC/GS Indicator Overhaul/Maintenance Manual (Type KI-202, KI-203, KI-204, KI-206, KI-207)
006-0185-00
Digital Area Navigation System Installation Manual (Type KNS-81)
006-5185-00
Digital Area Navigation System Overhaul/Maintenance Manual (Type KNS-81)
006-0152-02
Radar Altimeter Installation Manual (Type KRA-10A) (KI-250 Included)
006-5152-02
Radar Altimeter Overhaul/Maintenance Manual (Type KRA-10A) (KI-250 Included)
006-10536-02
Radar Altimeter Installation Manual (Type KRA-405B) (KNI-415/416 Included)
006-10557-0006
Autopilot Installation Manual (Type KFC225)
006-00702-0000
Flight Control and Avionics System Installation Maintenance Manual
006-0169-03
VHF Comm Transceiver Installation Manual (Type KY-196)
006-5169-03
VHF Comm Transceiver Overhaul/Maintenance Manual (Type KY-196)
006-0192-02
Radio Magnetic Indicator Installation Manual (Type KI-229)
006-5192-00
Radio Magnetic Indicator Overhaul/Maintenance Manual Type KI-229)
3.
Sigmatek Inc./ARC Avionics Manufacturer Manuals NOTE:
Following listed Sigmatek Inc./ARC publications must be ordered directly from Sigmatek Inc./ARC Avionics, Attn: ARC Order Entry, 1001 Industrial Rd., Augusta, Kansas 67010, USA. Telephone No. 316-775-1178.
MANUAL NUMBER
MANUAL TITLE
RP0070104-320
300/400 Series Navigation Indicators
RP0070104-310
400 Nav/Com (Type RT-485B) Service/Parts Manual
RP0070104-370
400 Marker Beacon (Type R-402) Service/Parts Manual
RP0070103-580
400 Glide Slope Receiver (Type R-443B) Service/Parts Manual
LIST OF PUBLICATIONS © Cessna Aircraft Company
Page 6 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
MANUAL NUMBER
MANUAL TITLE
RP0070104-140
400 RMI (Type IN-404A) Service/Parts Manual
RP0070103-610
400 ADF (Type R-546E) Service/Parts Manual
RP0070103-780
400 ADF (Type R446A) Service/Parts Manual
RP0070104-280
400/1000 DME (Type RN-477A) Service/Parts Manual
RP0070104-290
400/1000 RNAV (Type RN-479A) Service/Parts Manual
RP0070104-341
400B Autopilot/IFCS Vol. I (Type AF-550A & IF-550A)
RP0070104-342
400B Autopilot/IFCS Vol. II (Type AF-550A & IF-550A)
RP0070104-343
400B Autopilot/IFCS Vol. III (Type AF-550A & IF-550A)
RP0070104-344
400B Autopilot/IFCS Vol. IV (Type AF-550A & IF-550A)
RP0070104-440
400 Audio AmpliÞer (Type SDM-490A) Service/Parts Manual
RP0070104-447
300 DME (Type SDM-77A) Service/Parts Manual
RP0070104-330
300/400/800 Transponder (Type RT-359A & RT-459A) Service/Parts Manual
7010429
400 RNAV (Type RN-479A) Service/Parts Manual
D4551-13
400 Encoding Altimeter (Type EA-401A) Service/Parts Manual
D4551-13
800 Encoding Altimeter (Type EA-801A) Service/Parts Manual
D4556-13
800 Alerter (Type AA-801A) Service/Parts Manual
LIST OF PUBLICATIONS © Cessna Aircraft Company
Page 7 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CHAPTER
AIRWORTHINESS LIMITATIONS THE AIRWORTHINESS LIMITATIONS SECTION IS FAA APPROVED AND GIVES INSPECTIONS AND MAINTENANCE THAT ARE REQUIRED BY PARTS 43.16 AND 91.403 OF TITLE 14 OF THE CODE OF FEDERAL REGULATIONS, UNLESS AN ALTERNATIVE PROGRAM HAS BEEN FAA APPROVED.
REVISION 20
8 MAY 1990 5 OCTOBER 2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
4-00-00
Pages 1-2
Oct 5/2010
4-10-00
Pages 1-2
Jan 31/2008
4-10-01
Pages 1-2
Jan 15/2010
4-11-00
Pages 1-2
Jan 31/2008
04-Title 04-List of Effectivity Page 04-Record of Temporary Revisions 04-Contents
4 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Oct 5/2010
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS AIRWORTHINESS LIMITATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4-00-00 Page 1 4-00-00 Page 1 4-00-00 Page 1
TYPICAL INSPECTION TIME LIMITS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Schedule. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4-10-00 Page 1 4-10-00 Page 1 4-10-00 Page 1
SEVERE INSPECTION TIME LIMITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Schedule. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4-10-01 Page 1 4-10-01 Page 1 4-10-01 Page 1
REPLACEMENT TIME LIMITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Replacement Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4-11-00 Page 1 4-11-00 Page 1 4-11-00 Page 1
4 - CONTENTS © Cessna Aircraft Company
Page 1 of 1 Oct 5/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL AIRWORTHINESS LIMITATIONS 1.
Scope A.
2.
This chapter gives the mandatory replacement times and inspection intervals for components and structures that are life-limited. The section also gives the scheduled inspection requirements for structural and fatigue components that are considered a part of the certification process. NOTE:
The Airworthiness Limitations section is FAA-Approved and gives specified inspection and maintenance necessary under Parts 43.16 and 91.409 of Title 14 of the Code of Federal Regulations, unless an alternative program has been approved by the FAA.
NOTE:
For EASA certified airplanes, the Airworthiness Limitations section is applicable to airplanes with less than 50,000 flight hours. Flight beyond 50,000 flight hours is prohibited until new or revised EASA-approved Airworthiness Limitations are obtained.
NOTE:
For airplanes registered in the Ukraine, the supplemental inspections defined by the Listing of Supplemental Inspections (5-14-00) are mandatory. Extension of the thresholds and intervals of these supplemental inspections is prohibited.
NOTE:
For IAC AR certified airplanes, the Airworthiness Limitation section is applicable to airplanes with less than 50,000 flight hours. Flight beyond 50,000 flight hours is prohibited.
Definition A.
This chapter has three sections. (1) Typical Inspection Time Limits (4-10-00). This section gives the systems and components that must be inspected at specified intervals for typical operations. The intervals are the maximum time permitted between inspections. (2) Severe Inspection Time Limits (4-10-01). This section gives the systems and components that must be inspected at specified intervals for severe operations. The intervals are the maximum time permitted between inspections. (3) Replacement Time Limits (4-11-00) This section gives the life limited components which must be replaced at a specific time.
B.
Operational Inspection Times. (1) You must first find the category of your airplane's operation based on average flight length. (2) You must also find the number of hours and number of landings on the airplane, then find the average flight length based on the formulas found below. You must use whichever number is less.
Number of Flight Hours Number of Flights
=
Average Flight Length
=
Average Flight Length
or Number of Flight Hours Number of Landings (3) (4)
If the average flight length is less than or equal to thirty-five minutes, then you must use the inspection times found in section 4-10-01, Severe Inspection Time Limits. For airplanes with an average flight length greater than thirty-five minutes, you must find the severity of the operating environment. If the airplane operates thirty percent or more of its flight time in severe environments, you must use the severe operation inspection times found in section 4-10-01, Severe Inspection Time Limits. Examples of severe environments would
4-00-00 © Cessna Aircraft Company
Page 1 Oct 5/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
(5) (6)
include flight operations at low altitude (i.e., less than 5,000 ft. above ground level) such as pipeline patrol, sightseeing, training flights, traversing mountainous terrain or flying near coastal areas identified in section 51-12-00, Corrosion Severity Maps - Description and Operation. For all other operating environments, inspections should be conducted using the typical operation inspection times in section 4-10-00, Typical Inspection Time Limits. After the operating environment is known, make an airplane logbook entry that states which inspection schedule (Typical or Severe) is being used.
4-00-00 © Cessna Aircraft Company
Page 2 Oct 5/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL TYPICAL INSPECTION TIME LIMITS 1.
General A.
2.
Inspection time intervals for the components that follow, show the maximum intervals. Expanded inspection and maintenance check procedures are shown in the applicable sections throughout this manual.
Inspection Schedule A.
B.
Flight Controls (Chapter 27) (1) Flap Bell Crank (Part Numbers 2622281-2, -12) - Do the inspection at the first 4000 landings, then every 500 landings thereafter in accordance with the latest revision of CAB02-1. (2) Flap Bell Crank (Part Number 2692001-2) - Do the inspection at the first 4000 landings, then every 500 landings thereafter in accordance with the latest revision of CAB02-1. NOTE:
Total landings include the accumulated landings of 2622281-2 prior to modification by SK208-123 to the 2692001-2 configuration.
NOTE:
These components also have replacement time limits. Refer to 4-11-00, Replacement Time Limits for replacement criteria.
Landing Gear (Chapter 32) (1) Main Landing Gear Axles (Part Numbers 2641011-1, -3, -4) (NDI - Magnetic Particle Inspection) - Do the inspection at the first 5000 landings and every 1000 landings thereafter (Supplemental Inspection Number 32-10-01). NOTE:
These components also have replacement time limits. Refer to 4-11-00, Replacement Time Limits for replacement criteria.
C.
Fuselage (Part Numbers 2610000-1, -2 and 2610001-1, -2) (Chapter 53) (1) Fuselage to Strut Attach Fitting and Lugs (NDI - Eddy Current) - Do the inspection at the first 10,000 hours and every 2500 hours thereafter (Supplemental Inspection Number 53-20-07). (2) Fuselage to Wing Carry-Thru Attach Fitting (NDI - Eddy Current) - Do the inspection at the first 20,000 hours and every 5000 hours thereafter (Supplemental Inspection Number 53-20-02).
D.
Wing (Chapter 57) (1) Wing (Part Numbers 2622000-1, -2, -101, -102, -119, -120, -123, -124) (a) Center Flap Track (NDI - Eddy Current) - Do the inspection at the first 15,000 landings and every 3000 landings thereafter (Supplemental Inspection Number 57-50-01). (b) Inboard Flap Track (NDI - Eddy Current) - Do the inspection at the first 15,000 landings and every 3000 landings thereafter (Supplemental Inspection Number 57-50-01). (c) Outboard Flap Track (NDI - Eddy Current) - Do the inspection at the first 15,000 landings and every 3000 landings thereafter (Supplemental Inspection Number 57-50-01). (d) Front Spar Lower Cap Inspection Inboard of WS 141.20 (NDI - Eddy Current) - Do the inspection at the first 20,000 hours and every 5000 hours thereafter (Supplemental Inspection Number 57-20-02). (e) Rear Spar Lower Cap Inspection Inboard of WS 141.20 (NDI - Eddy Current) - Do the inspection at the first 20,000 hours and every 5000 hours thereafter (Supplemental Inspection Number 57-20-03). (f) Wing/Strut Attachment to Front Spar (NDI - Eddy Current) - Do the inspection at the first 20,000 hours and every 5000 hours thereafter (Supplemental Inspection Number 57-6002). (g) Wing to Carry-Thru Front Spar Attachment Fittings (NDI - Eddy Current) - Do the inspection at the first 20,000 hours and every 5000 hours thereafter (Supplemental Inspection Number 57-20-01). (h) Wing to Carry-Thru Rear Spar Attachment Fittings (NDI - Eddy Current) - Do the inspection at the first 20,000 hours and every 5000 hours thereafter. (Supplemental Inspection Number 57-20-01).
4-10-00 © Cessna Aircraft Company
Page 1 Jan 31/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (2)
Wing Strut (Part Numbers 2621000-5, -6, -11, -12, -19, -20, -21, 22) (a) Wing Strut Attach Fitting (NDI - Eddy Current) - Do the inspection at the first 10,000 hours and every 5000 hours thereafter (Supplemental Inspection Number 57-60-01).
4-10-00 © Cessna Aircraft Company
Page 2 Jan 31/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SEVERE INSPECTION TIME LIMITS 1.
General A.
2.
Inspection time intervals for the components that follow, show the maximum intervals. Expanded inspection and maintenance check procedures are shown in the applicable sections throughout this manual.
Inspection Schedule A.
B.
Flight Controls (Chapter 27) (1) Flap Bell Crank (Part Numbers 2622281-2, -12) - Do the inspection at the first 4000 landings, then every 500 landings thereafter in accordance with the latest revision of CAB02-1. (2) Flap Bell Crank (Part Number 2692001-2) - Do the inspection at the first 4000 landings, then every 500 landings thereafter in accordance with the latest revision of CAB02-1. NOTE:
Total landings include the accumulated landings of 2622281-2 prior to modification by SK208-123 to the 2692001-2 configuration.
NOTE:
These components also have replacement time limits. Refer to 4-11-00, Replacement Time Limits for replacement criteria.
Landing Gear (Chapter 32) (1) Main Landing Gear Axles (Part Numbers 2641011-1, -3, -4) (NDI - Magnetic Particle Inspection) - Do the inspection at the first 5000 landings, then every 1000 landings thereafter (Supplemental Inspection Number 32-10-01). NOTE:
These components also have replacement time limits. Refer to 4-11-00, Replacement Time Limits for replacement criteria.
C.
Fuselage (Part Numbers 2610000-1, -2 and 2610001-1, -2) (Chapter 53) (1) Fuselage to Strut Attach Fitting and Lugs (NDI - Eddy Current) - Do the inspection at the first 5000 hours (Supplemental Inspection Number 53-20-07), and then (a) Every 1200 hours thereafter for lugs with a nominal/standard bolt size (Part Number S346174). (b) Every 500 hours thereafter for lugs with a 1/64 inch oversize bolt (Part Number S3461-159). (c) Every 400 hours thereafter for lugs with a 1/32 inch oversize bolt (Part Number S3461-160). (2) Fuselage to Wing Carry-Thru Attach Fitting (NDI - Eddy Current) - Do the inspection at the first 20,000 hours, then every 5000 hours thereafter (Supplemental Inspection Number 53-20-02).
D.
Wing (Chapter 57) (1) Wing (Part Numbers 2622000-1, -2, -101, -102, -119, -120, -123, -124) (a) Center Flap Track (NDI - Eddy Current) - Do the inspection at the first 15,000 landings, then every 3000 landings thereafter (Supplemental Inspection Number 57-50-01). (b) Inboard Flap Track (NDI - Eddy Current) - Do the inspection at the first 15,000 landings, then every 3000 landings thereafter (Supplemental Inspection Number 57-50-01). (c) Outboard Flap Track (NDI - Eddy Current) - Do the inspection at the first 15,000 landings, then every 3000 landings thereafter (Supplemental Inspection Number 57-50-01). (d) Front Spar Lower Cap Inspection Inboard of WS 141.20 (NDI - Eddy Current) - Do the inspection at the first 20,000 hours, then every 5000 hours thereafter (Supplemental Inspection Number 57-20-02). (e) Rear Spar Lower Cap Inspection Inboard of WS 141.20 (NDI - Eddy Current) - Do the inspection at the first 20,000 hours, then every 5000 hours thereafter (Supplemental Inspection Number 57-20-03). (f) Wing/Strut Attachment to Front Spar (NDI - Eddy Current) - Do the inspection at the first 20,000 hours (Supplemental Inspection Number 57-60-02), and then 1 Every 5000 hours thereafter for lugs with a nominal/standard bolt size (Part Number S3461-77). 2 Every 4400 hours thereafter for lugs with a 1/64 inch oversize bolt (Part Number S3461-163).
4-10-01 © Cessna Aircraft Company
Page 1 Jan 15/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL 3
(2)
Every 3600 hours thereafter for lugs with a 1/32 inch oversize bolt (Part Number S3461-164). (g) Wing to Carry-Thru Front Spar Attachment Fittings (NDI - Eddy Current) - Do the inspection at the first 20,000 hours, then every 5000 hours thereafter (Supplemental Inspection Number 57-20-01). (h) Wing to Carry-Thru Rear Spar Attachment Fittings (NDI - Eddy Current) - Do the inspection at the first 20,000 hours, then every 5000 hours thereafter (Supplemental Inspection Number 57-20-01). Wing Strut (Part Numbers 2621000-5, -6, -11, -12, -19, -20, -21, -22) (a) Wing Strut Attach Fitting (NDI - Eddy Current) - Do the inspection at the first 5000 hours, then every 3600 hours thereafter (Supplemental Inspection Number 57-60-01).
4-10-01 © Cessna Aircraft Company
Page 2 Jan 15/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL REPLACEMENT TIME LIMITS 1.
General A.
2.
The following life-limited components are to be replaced at the specified time. It is recommended that the components be scheduled for replacement during the airplane's inspection interval coinciding with, or occurring just before, the expiration of the specified time limit. Procedures for replacement of the components are described in the applicable chapters in this Maintenance Manual.
Replacement Schedule A.
Flight Controls (Chapter 27) (1) Flap Bell Crank (Part Number 2622083-18) - Replace at every 2250 landings. (2) Flap Bell Crank (Part Number DDA00028-4) - Replace at every 2250 landings. (3) Flap Bell Crank (Part Numbers 2622281-2, -12) - Replace at every 7000 landings. (4) Flap Bell Crank (Part Number 2692001-2) - Replace at every 7000 landings. NOTE: (5) (6) (7)
B.
Total landings include the accumulated landings of 2622281-2 prior to modification by SK208-123 to the 2692001-2 configuration.
Flap Bell Crank (Part Numbers 2622311-7, -16) - Replace at every 40,000 landings. Flap Bell Crank (Part Number 2622311-7) attaching parts: Bearings (Part Number MS27641-5 or S3952-5) and Bolt (Part Number AN5-77) - Replace at every 10,000 landings. Flap Bell Crank (Part Number 2622311-16) attaching parts: Bearings (Part Number KP5A-H) and Bolt (Part Number AN5-77) - Replace at every 10,000 landings.
Landing Gear (Chapter 32) (1) Main Landing Gear NOTE:
(a) (b) (c) (d) (e) (f) (g)
Attaching hardware (bolts, bearings, bushings, and trunnion pins related to the installation of the components below) is to be replaced whenever the associated component is replaced.
Main Landing Gear Center Spring (Part Numbers 2641014-2, -3, -4, -5, -6, -7, -8, -9) Replace at every 31,500 landings. Main Landing Gear Trunnion Assembly (Part Numbers 2641012-1, -2, -8, -9, -13, -14) Replace at every 31,500 landings. Main Landing Gear Spring (Part Numbers 2641013-1, -2, -3, -4, -5, -6, -7, -8) - Replace at every 31,500 landings. Main Landing Gear Attach Pin (Part Numbers 2641008-1, -2, -200) - Replace at every 31,500 landings. Main Landing Gear Axles (Part Numbers 2641011-1, -3, -4) - Replace at every 10,000 landings. Main Landing Gear Axles (Part Number 2641011-5) - Replace at every 31,500 landings. Main Landing Gear Axle Fittings (Part Numbers 2641010-1, -3, -7) - Replace at every 31,500 landings.
4-11-00 © Cessna Aircraft Company
Page 1 Jan 31/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (2)
Nose Landing Gear NOTE:
(a)
(b)
Attaching hardware (bolts, bearings, bushings, and axle components related to the installation of the components below) is to be replaced whenever the associated component is replaced.
Nose Gear Drag Link Spring (Part Numbers 2643062-1, -2, -3) - Replace at every 15,000 landings. NOTE:
For nose gear drag link springs repaired per CAB96-24 or per Chapter 32, Nose Landing Gear - Cleaning/Painting with damage repaired between 0.050 inch and 0.062 inch (1.270 mm and 1.575 mm), the life limit is an additional 12,000 landings after repair, not to exceed 15,000 landings.
NOTE:
For nose gear drag link springs repaired per CAB96-24 or per Chapter 32, Nose Landing Gear - Cleaning/Painting with damage repaired between 0.063 inch and 0.075 inch (1.600 mm and 1.905 mm), the life limit is an additional 10,000 landings after repair, not to exceed 15,000 landings.
Nose Gear Assembly (Part Numbers 2643045, 2643100 and 2643095 Series Part Numbers) - Replace at every 40,000 landings. NOTE:
(c)
Support Assembly, Nose Gear Spring (Part Numbers 2643030, 2643055 and 2643099 Series Part Numbers) - Replace at every 40,000 landings. NOTE:
(d)
For an illustration of the nose gear spring support assembly to be replaced, refer to Chapter 32, section 32-20-00, Nose Landing Gear - Maintenance Practices.
Fork Assembly, Nose Gear Spring (Part Numbers 2643031-1, -7) - Replace at every 40,000 landings. NOTE:
C.
For an illustration of the nose gear assembly to be replaced, refer to Chapter 32, section 32-20-00, Nose Landing Gear - Maintenance Practices.
For an illustration of the nose gear spring fork assembly to be replaced, refer to Chapter 32, section 32-20-00, Nose Landing Gear - Maintenance Practices.
Oxygen (Chapter 35) (1) The airplane may be equipped with a two-port oxygen system incorporating a 50.67 cubic-footcapacity oxygen cylinder (Part Numbers C166001-1101 and C166001-1201) or a ten port oxygen system incorporating a 116.95 cubic-foot-capacity oxygen cylinder (Part Numbers C1660011102 and C166001-1103). Both cylinders have a life limit of 15 years.
4-11-00 © Cessna Aircraft Company
Page 2 Jan 31/2008
5
CHAPTER
TIME LIMITS/ MAINTENANCE CHECKS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT 05-Title 05-List of Effective Pages 05-Record of Temporary Revisions 05-Table of Contents 05-List of Tasks 5-00-00 5-10-00 5-10-01 5-11-00 5-11-01 5-13-00 5-14-00 5-15-00 5-15-0A 5-15-01 5-15-02 5-15-03 5-15-04 5-15-05 5-15-06 5-15-07 5-15-08 5-15-09 5-15-10 5-15-11 5-15-12 5-15-13 5-15-14 5-15-15 5-15-16 5-15-17 5-15-18 5-15-19 5-15-20 5-15-21 5-15-22 5-15-MA 5-15-MB 5-15-MD 5-15-ME 5-15-MF 5-15-MG 5-15-MH 5-15-MI
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CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS TIME LIMITS/MAINTENANCE CHECKS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-00-00 5-00-00 5-00-00 5-00-00
Page 1 Page 1 Page 1 Page 2
INSPECTIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Interval Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tasks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-10-00 5-10-00 5-10-00 5-10-00 5-10-00
Page 1 Page 1 Page 1 Page 3 Page 4
INSPECTION TIME LIMITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-10-01 Page 1 5-10-01 Page 1
COMPONENT TIME LIMITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Component Time Limit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-11-00 Page 1 5-11-00 Page 1 5-11-00 Page 1
COMPONENT TIME LIMITS - RUSSIAN CERTIFIED AIRPLANES . . . . . . . . . . . . . . . . . . . Component Time Limit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-11-01 Page 1 5-11-01 Page 1 5-11-01 Page 1
SUPPLEMENTAL INSPECTION DOCUMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Supplemental Inspection Document . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Principal Structural Elements. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Durability - Fatigue And Damage Tolerance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reporting - Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Applicability/Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PSE Details . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-13-00 5-13-00 5-13-00 5-13-00 5-13-00 5-13-00 5-13-00 5-13-00
LISTING OF SUPPLEMENTAL INSPECTIONS - DESCRIPTION AND OPERATION . . . Supplemental Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Supplemental Inspections to Task Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-14-00 Page 1 5-14-00 Page 1 5-14-00 Page 1
SCHEDULED INSPECTION PROGRAM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Purpose and Use . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Construction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Time Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Manual - Compact Disc - Read Only Memory (CD-ROM) . . . . . . . . . . . Inspection Guidelines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-00 5-15-00 5-15-00 5-15-00 5-15-00 5-15-00
INSPECTION DOCUMENT 0A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-0A Page 1 5-15-0A Page 1 5-15-0A Page 1
INSPECTION DOCUMENT 01 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-01 Page 1 5-15-01 Page 1 5-15-01 Page 1
INSPECTION DOCUMENT 02 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-02 Page 1 5-15-02 Page 1 5-15-02 Page 1
INSPECTION DOCUMENT 03 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-03 Page 1 5-15-03 Page 1 5-15-03 Page 1
INSPECTION DOCUMENT 04 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-04 Page 1 5-15-04 Page 1 5-15-04 Page 1
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MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 05 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-05 Page 1 5-15-05 Page 1 5-15-05 Page 1
INSPECTION DOCUMENT 06 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-06 Page 1 5-15-06 Page 1 5-15-06 Page 1
INSPECTION DOCUMENT 07 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-07 Page 1 5-15-07 Page 1 5-15-07 Page 1
INSPECTION DOCUMENT 08 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-08 Page 1 5-15-08 Page 1 5-15-08 Page 1
INSPECTION DOCUMENT 09 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-09 Page 1 5-15-09 Page 1 5-15-09 Page 1
INSPECTION DOCUMENT 10 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-10 Page 1 5-15-10 Page 1 5-15-10 Page 1
INSPECTION DOCUMENT 11 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-11 Page 1 5-15-11 Page 1 5-15-11 Page 1
INSPECTION DOCUMENT 12 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-12 Page 1 5-15-12 Page 1 5-15-12 Page 1
INSPECTION DOCUMENT 13 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-13 Page 1 5-15-13 Page 1 5-15-13 Page 1
INSPECTION DOCUMENT 14 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-14 Page 1 5-15-14 Page 1 5-15-14 Page 1
INSPECTION DOCUMENT 15 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-15 Page 1 5-15-15 Page 1 5-15-15 Page 1
INSPECTION DOCUMENT 16 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-16 Page 1 5-15-16 Page 1 5-15-16 Page 1
INSPECTION DOCUMENT 17 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-17 Page 1 5-15-17 Page 1 5-15-17 Page 1
INSPECTION DOCUMENT 18 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-18 Page 1 5-15-18 Page 1 5-15-18 Page 1
INSPECTION DOCUMENT 19 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-19 Page 1 5-15-19 Page 1 5-15-19 Page 1
INSPECTION DOCUMENT 20 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-20 Page 1 5-15-20 Page 1 5-15-20 Page 1
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MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 21 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-21 Page 1 5-15-21 Page 1 5-15-21 Page 1
INSPECTION DOCUMENT 22 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-22 Page 1 5-15-22 Page 1 5-15-22 Page 1
INSPECTION DOCUMENT MA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MA Page 1 5-15-MA Page 1 5-15-MA Page 1
INSPECTION DOCUMENT MB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MB Page 1 5-15-MB Page 1 5-15-MB Page 1
INSPECTION DOCUMENT MD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MD Page 1 5-15-MD Page 1 5-15-MD Page 1
INSPECTION DOCUMENT ME . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-ME Page 1 5-15-ME Page 1 5-15-ME Page 1
INSPECTION DOCUMENT MF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MF Page 1 5-15-MF Page 1 5-15-MF Page 1
INSPECTION DOCUMENT MG . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MG Page 1 5-15-MG Page 1 5-15-MG Page 1
INSPECTION DOCUMENT MH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MH Page 1 5-15-MH Page 1 5-15-MH Page 1
INSPECTION DOCUMENT MI . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MI Page 1 5-15-MI Page 1 5-15-MI Page 1
INSPECTION DOCUMENT MJ. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MJ Page 1 5-15-MJ Page 1 5-15-MJ Page 1
INSPECTION DOCUMENT MK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-MK Page 1 5-15-MK Page 1 5-15-MK Page 1
INSPECTION DOCUMENT ML. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Inspection Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-15-ML Page 1 5-15-ML Page 1 5-15-ML Page 1
EXPANDED INSPECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aircraft Records Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5-20-01 Page 1 5-20-01 Page 1 5-20-01 Page 1
UNSCHEDULED MAINTENANCE CHECKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Unscheduled Maintenance Checks DeÞned and Areas to be Inspected . . . . . . . . . .
5-50-00 Page 1 5-50-00 Page 1 5-50-00 Page 1
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 5-20-01-280
Aircraft Records Check
5-20-01 Page 1
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MODEL 208 MAINTENANCE MANUAL TIME LIMITS/MAINTENANCE CHECKS - GENERAL 1.
Scope A.
2.
This chapter gives the time limits and maintenance checks for the Model 208 and 208B airplanes. It is divided into several sections, each with a special purpose toward providing information necessary to establish inspection criteria. Refer to the Description section for detailed information concerning each of these sections. NOTE:
In accordance with Title 14 of the Code of Federal Regulations (CFR) 23.1529, Chapter 4 (Airworthiness Limitations) of this manual is published as a separate document. Refer to Chapter 4 for those components that have a mandatory inspection and components replacement schedule.
NOTE:
For EASA-certiÞed airplanes, the Chapter 4 Airworthiness Limitations section is applicable to airplanes with less than 50,000 ßight hours. Flight beyond 50,000 ßight hours is prohibited until new or revised FAA-approved Airworthiness Limitations are obtained.
NOTE:
The time limits and maintenance checks recorded in this chapter are the minimum requirements for airplanes operated under normal conditions. For airplanes that operate in areas of bad conditions can be found, such as, high salt coastal environments, areas of high heat and humidity, areas where industrial or other airborne pollutants are present, extreme cold, unimproved surfaces, etc., the time limits shall be changed as necessary.
NOTE:
It is recommended that all Cessna Model 208 owners participate in CESCOM (Computerized Maintenance Records System). This is a comprehensive system which gives an easy procedure to monitor and schedule inspections, Service Bulletins, Service Kits, Airworthiness Directives, and scheduled and unscheduled maintenance activities. For additional information on CESCOM, refer to Section 8 in the Pilot’s Operating Handbook, or the CESCOM Instruction Manual supplied with your airplane
B.
Chapter 4 of this manual is FAA approved and issued separately from the maintenance manual. Some inspection interval and life limit requirements of Chapter 4 possibly will not agree with the current Chapter 5. When there is a conßict between the two chapters, Chapter 4 requirements must always be followed. Chapter 5 requirements will be made to agree with Chapter 4 at the next revision to the manual.
C.
Inspection Documents that begin with the letter M are those inspections found in Chapter 4. These were added because there can be no grace period for these inspections.
Inspection Requirements A.
Two basic types of inspections are available as deÞned below: (1) As required by Title 14 of the Code of Federal Regulations Part 91.409 (a), all civil airplanes of U.S. registry must have a complete examination (Annual) each 12 calendar months. In addition to the required Annual inspection, airplanes operated commercially (for hire) must also have complete inspection each 100 hours of operation as required by Title 14 of the Code of Federal Regulations Part 91.409 (b). (a) If ßown for hire. An airplane operating in this category must have a complete airplane inspection each 1 100 hours. Refer to Inspection Document 0A. The Component Time Limits shall also be examined at each inspection interval to make sure the correct overhaul and replacement requirements are done at the speciÞed times. (b) If not ßown for hire. An airplane that operates in this category must have a complete airplane inspection 1 each 12 calendar months of operation (Annual). Refer to Inspection Document 0A. The Component Time Limits shall be examined at each inspection interval to make sure the correct overhaul and replacement requirements are done at the speciÞed times.
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MODEL 208 MAINTENANCE MANUAL (2) 3.
If an airplane is being operated under a CFR Part 135 CertiÞcate, the operator can choose to use an Approved Aircraft Inspection Program.
Description NOTE:
Given below is a detailed description and the purpose of each section of this chapter.
A.
Section 5-00-00, Time Limits/Maintenance Checks - General. This section gives a description and purpose of each section of this chapter.
B.
Section 5-10-01, Inspection Time Limits. (1) This section supplies a list, in chart format, of all of the inspection and service requirements which must be done. Each page has the six columns that follow: (a) Revision Status gives the date that an item was added, deleted or revised. A blank entry in this column shows no change was made since the reissue of this manual. (b) The ITEM CODE NUMBER column gives a seven-character, alphanumeric code that is related to each inspection. The item code number does not change. The alphanumeric code contains one letter and six numbers. The letter at the start of the code is A, B, C, or D. Refer to the list that follows for a description of the code: • The letter A shows that a visual inspection is necessary • The letter B shows that a functional check or an operational check is necessary • The letter C shows that a lubrication is necessary • The letter D shows that a clean, service, or replacement is necessary. (c) The Task column gives a short description of the maintenance item and are supplied in chapter order. (d) The Interval is an alphanumeric code character that shows the frequency of the item. The frequencies for each code are given in Chapter 5-10-00. (e) The CH SE SU is a reference to the applicable Inspection Document that currently has the inspection item. (f) Applicable Zone refers to the physical location(s) in the airplane where the item is. Most functional and operational tests do not give a zone, but a code which shows the special conditions required to do the test. The codes and conditions are as follows: ALL - This code and condition is applicable to the entire airplane. ENG - Airplane engine to be running. AUX - External source of electrical power. Airplane engine power sources shall not be used for these tests. NOTE:
(2) (3)
It is possible to do many of the tests in the AUX category with the airplane's battery power. However, it is not recommended because of the power drain on the battery.
BAT - These tests must be done with the component powered by the airplane’s battery or the batteries built into the individual component, like the ELT or other components with internal battery power. LAB - Is when special equipment is used which requires that the component be removed from the airplane and taken to a place equipped to do the check or calibration. FLT - The test is to be done during a ßight. The primary purpose of the Inspection Time Limits section is to give a complete list of all inspection items in an order that lets the information given previously be easily found. This section is not to be used as a method to examine the airplane. The Inspection Time Limits Chart shows the recommended intervals at which items are to be examined for normal use in average environmental conditions. Airplanes operated in very humid areas (tropics), or in very cold, damp climates, etc., can need more frequent inspections for wear, corrosion, and lubrication. When the airplane is used in these bad conditions, complete periodic
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MODEL 208 MAINTENANCE MANUAL inspections that agree with this chart at more frequent intervals until the operator can use Þeld experience to set his own inspection periods. The operator’s inspection intervals shall obey the inspection time limits shown in this manual except as given below: (a) Each inspection interval can have 10 more hours (if time controlled), or 30 more days (if date controlled) or can be done early any time before the regular interval as given below: If any inspection document is done late, the next inspection document in sequence 1 keeps a due point from the time the late inspection document was initially scheduled. If any inspection document is done early, 10 hours or less ahead of schedule, the 2 next inspection document due point can remain where initially set. If any inspection document is done early, more than 10 hours ahead of schedule, the 3 next inspection document due point must be rescheduled to set a new due point from the time it was done early. C.
Section 5-11-00, Component Time Limits. This section gives a list of overhaul or replacement intervals for components in chapter order. These requirements are not given in the 5-10-01 Inspection Time Limits section. The component overhaul or replacement criteria must be used to Þnd the correct action for the components in the list. These requirements must be worked into the scheduled inspection program to supply a complete inspection program.
D.
Section 5-15-00, Scheduled Inspection Program. (1) This section gives information about the scheduled inspection programs. (2) Each section of 5-15-XX is the Inspection Document that records the items to be examined for that given interval. The last two characters give the subject of the chapter/section/subject identiÞcation.
E.
Section 5-14-00, Listing of Supplemental Inspections (1) This section has a matrix or cross-reference table for the Supplemental Inspection Documents to the Task Inspection Documents.
F.
Section 5-50-00, Unscheduled Maintenance Checks. This section has the inspections and checks which can be required because of special or unusual circumstances and do not have regular repeated intervals to be done.
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MODEL 208 MAINTENANCE MANUAL INSPECTIONS 1.
General A.
2.
General inspection criteria. (1) When you do each of the speciÞed inspections in this chapter, you must do a general examination of the adjacent areas while access is available. These general visual examinations can help Þnd conditions which will need more maintenance procedures. (2) When you access an area, examine the wire bundles and make repairs as applicable. Make sure the wire bundles are not attached to hydraulic tubes or lines. (3) Inspection items are given for speciÞed components and systems. The inspection program must have professionalism and good judgment used by all inspection personnel. The technician must make sure that all components and systems are in good condition and kept to the highest safety standards. (4) If a component or system is moved or changed (because of maintenance done) after a required operational or functional test is done, then you must do the test again before the system or component is returned to service. Refer to the appropriate chapter in this Maintenance Manual for removal, installation, operational tests, and functional tests of components and/or systems. (5) Refer to Chapter 12, Servicing for information about the lubricant, lubrication points and the method of lubrication for items or components that are lubricated. (6) Refer to Chapter 6, Airplane Zoning - Description and Operation, for airplane zone deÞnition. (7) After you complete the applicable inspections, refer to the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual to do a preßight inspection. (8) Inspection Document intervals that begin with the letter M are those inspections that match Chapter 4. These were added because there is no grace period for these inspections. (9) TKS is a ßuid based anti-ice system.
Inspection Interval Requirements NOTE:
Inspection requirements for the engine and propeller are supplied by the component manufacturer. To make sure that the latest inspection requirements given by the manufacturer are done, refer to and do the requirements at the intervals published by the component manufacturer.
Inspection Interval
Inspection Document
Interval 0A gives a list of item(s), which are completed during the Annual inspection
5-15-0A
Interval 1C item(s), which are completed every 12 calendar months.
5-15-01
Interval 2C item(s), which are completed every 24 calendar months.
5-15-02
Interval 4C item(s), which are completed every 48 calendar months.
5-15-03
Interval 6C item(s), which are completed every 72 calendar months.
5-15-04
Interval 12C item(s), which are completed every 144 calendar months.
5-15-05
Interval 200hrs/1C item(s), which are completed every 200 Hours or 12 calendar months, whichever occurs Þrst.
5-15-06
Interval 1A/1C item(s), which are completed every 400 Hours or 12 calendar months, whichever occurs Þrst.
5-15-07
Interval 1A/2C item(s), which are completed every 400 Hours or 24 calendar months, whichever occurs Þrst.
5-15-08
Interval 2A/1C item(s), which are completed every 800 Hours or 12 calendar months, whichever occurs Þrst.
5-15-09
Interval 2A/2C item(s), which are completed every 800 Hours or 24 calendar months, whichever occurs Þrst.
5-15-10
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MODEL 208 MAINTENANCE MANUAL
Inspection Interval
Inspection Document
Interval 4A/2C item(s), which are completed every 1600 Hours or 24 calendar months, whichever occurs Þrst.
5-15-11
Interval 4A/5C item(s), which are completed every 1600 Hours or 60 calendar months, whichever occurs Þrst.
5-15-12
Interval AD item(s), which are completed at the Þrst 20,000 hours and every 5000 hours thereafter.
5-15-13
Interval AE item(s), which are completed at the Þrst 5000 hours and every 2500 hours thereafter.
5-15-14
Interval AF item(s), which are completed at the Þrst 7500 hours and every 2500 hours thereafter.
5-15-15
Interval AG item(s), which are completed at the Þrst 12,500 hours and every 2500 hours thereafter.
5-15-16
Interval AH item(s), which are completed at the Þrst 16,500 hours and every 5000 hours thereafter.
5-15-17
Interval AI item(s), which are completed at the Þrst 17,500 hours and every 1000 hours thereafter.
5-15-18
Interval AJ item(s), which are completed at the Þrst 25,000 landings and every 5000 landings thereafter.
5-15-19
Required 14CFR 91.207 interval item(s), which are completed every 12 calendar months (No grace period).
5-15-20
Required 14CFR 91.411 certiÞcation interval item(s), which are completed every 24 calendar months (No grace period).
5-15-21
Required 14CFR 91.413 certiÞcation interval item(s), which are completed every 24 calendar months (No grace period).
5-15-22
Interval MA item(s), which are completed at 10,000 hours and every 5000 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-MA
Interval MB item(s), which are completed at 5000 landings and every 1000 landings thereafter, up to 10,000 landings. Replace at 10,000 landings. (Chapter 4 requirement - No grace period)
5-15-MB
Interval MD item(s), which are completed at 15,000 landings and every 3000 landings thereafter. (Chapter 4 requirement - No grace period)
5-15-MD
Interval ME item(s), which are completed at 5000 hours and every 3600 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-ME
Interval MF item(s), which are completed at 20,000 hours and every 5000 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-MF
Interval MG item(s), which are completed at 5000 hours and every 1200 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-MG
Interval MH item(s), which are completed at 10,000 hours and every 2500 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-MH
Interval MI item(s), which are completed at 5000 hours and every 500 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-MI
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MODEL 208 MAINTENANCE MANUAL
Inspection Interval
Inspection Document
Interval MJ item(s), which are completed at 5000 hours and every 400 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-MJ
Interval MK item(s), which are completed at 20,000 hours and every 4400 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-MK
Interval ML item(s), which are completed at 20,000 hours and every 3600 hours thereafter. (Chapter 4 requirement - No grace period)
5-15-ML
A.
3.
Touch-and-go landings are to be considered identical to full-stop landings and must therefore be included in the count of accumulated landings for all inspections and maintenance. Both full-stop landings and touch-and-go landings must be tracked.
Tasks A.
The inspection tasks have more data than that given in Inspection Time Limits. The inspection tasks are identiÞed with ATA (chapter-section-subsection-function) numbers. Each task has different ATA numbers. The chapter-section-subsection digits give the location of the inspection task in the Maintenance Manual. The last three digits give the function of the inspection task. Refer to the task example.
B.
An example task number is given in the task example. The numbers that follow the word Task give the chapter, section, and subsection location of the inspection task. The Þrst two digits of the function gives the aircraft maintenance and task oriented support system (AMTOSS) code. Refer to Table 1 for a description of the function AMTOSS codes. The last digit of the function gives the task sequence number. If the same chapter-section-subsection and function number is used for a different task, the sequence number increments by one or more.
Table 1. Function AMTOSS Codes 10
CLEANING
64
Lubricating
11
Chemical
65
Fueling, Defueling
12
Abrasive
67
Disinfect, Sanitize
13
Ultrasonic
68
Drain Fluid
14
Mechanical
70
TESTING, CHECKING
15
Stripping
71
Operational
5-10-00 © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 1. Function AMTOSS Codes (continued)
4.
16
Miscellaneous Cleaning
72
Functional
17
Flushing
73
System
20
INSPECTION, CHECKS
74
Bite
21
General Visual
75
Special
22
Detailed Dimensional
76
Electrical
23
Penetrant
78
Pressure
24
Magnetic
79
Leak
25
Eddy Current
80
MISCELLANEOUS
26
X-Ray
81
Fault Isolation
27
Ultrasonic
82
Adjusting, Aligning, Calibration, Rigging
28
SpeciÞc, Special
87
Bleeding
29
Borescope
90
CHANGE, REMOVE, INSTALL
60
SERVICING, PRESERVING, LUBRICATING
96
Replace
61
Servicing
C.
Personnel that use this Maintenance Manual on compact disc-read only memory (CD-ROM) can link from a highlighted task number to the inspection task in the related chapter. You can Þnd a list of the task numbers in Inspection Time Limits.
D.
Supplemental Type CertiÞcate (STC) Installations (1) The necessary inspections for STC installations are not included in this manual. You must use the inspection program made by the owner of the STC to examine the components included in the STC installation. The inspections supplied by Cessna are not applicable for airplanes with STC installations since STC installations can change the system interface, operation property, component load, or stress on adjacent structures.
DeÞnitions A.
Lubrication or Servicing Task (1) Any act of lubrication or servicing for the purpose of maintaining inherent design capabilities.
B.
Visual Check (1) A visual check is an observation to determine that an item is fulÞlling its intended purpose. The check does not require quantitative tolerances. This is a failure Þnding task.
C.
General Visual (Surveillance) Inspection (1) A visual examination of an interior or exterior area, installation or assembly to detect obvious damage, failure, or irregularity. This level of inspection is made under normally available lighting conditions such as daylight, hangar lighting, ßashlight, or drop-light and may require removal or opening of access panels or doors. Stands, ladders, or platforms may be required to gain proximity to the area being checked.
D.
Detailed Inspection (1) An intensive visual examination of a speciÞc structural area, system, installation or assembly to detect damage, failure, or irregularity. Available lighting is normally supplemented with a direct source of good lighting at an intensity deemed appropriate by the inspector. Inspection aids such as mirrors, magnifying lenses, etc. may be used. Surface cleaning and elaborate access procedures may be required.
5-10-00 © Cessna Aircraft Company
Page 4 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL E.
Special Detailed Inspection (1) An intensive examination of a speciÞc item(s), installation or assembly to detect damage, failure, or irregularity. The examination is likely to make extensive use of specialized inspection techniques and/or equipment. Intricate cleaning and substantial access or disassembly procedure may be required.
F.
Operational Check (1) An operational check is a task to determine that an item is fulÞlling its intended purpose. The check does not require quantitative tolerances. This is a failure Þnding task.
G.
Functional Check (1) A functional check is a quantitative check to determine if one or more functions of an item perform within speciÞed limits.
H.
Restoration Task (1) That work necessary to return the item to a speciÞc standard. Since restoration may vary from cleaning or replacement of single parts up to a complete overhaul, the scope of each assigned restoration task has to be speciÞed.
I.
Discard Task (1) The removal from service of an item at a speciÞed life limit. Discard tasks are normally applied to so-called single parts such as cartridges, canisters, cylinders, engine disks, and safe-life structural members.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION TIME LIMITS 1.
Inspection Items
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
Added Mar 1/12
A052001
Aircraft Records Check Task 5-20-01280
Annual
05-15-0A
ALL
A110001
Interior and Exterior Placard and Decal Detailed Inspection Task 11-00-00-220
Annual
05-15-0A
ALL
D121001
Brake System Servicing Task 12-1001-610
Annual
05-15-0A
121
D121003
Shimmy Damper Servicing Task 1210-01-611
Annual
05-15-0A
710
C122101
Landing Gear Lubrication Task 12-2103-640
Annual
05-15-0A
700
C122103
Hartzell Propeller Lubrication Task 1221-04-640
Annual
05-15-0A
110
B212401
Avionics Cooling Fan Operational Check Task 21-24-00-710
Annual
05-15-0A
211 212
B215001
Compressor Drive Belt Functional Check Task 21-50-00-720
1A/2C
05-15-08
121 122
C221201
Autopilot Servos Lubrication Task 2212-00-640
2A/2C
05-15-10
226 232
B221201
Garmin Autopilot (GFC 700) Functional Check Task 22-12-00-720
2A/1C
05-15-09
226 232
B236001
Static Discharge System Functional Check Task 23-60-00-720
200 Hours/1C
05-15-06
343 375 376 571 671
B243201
Gill Flooded Lead-Acid Battery Functional Check (Capacity Check) Task 24-32-00-720
2A/1C
05-15-09
122
B243301
Concord Sealed Lead Acid Battery Functional Check (Capacity Check) Task 24-33-00-720
2A/1C
05-15-09
122
B243401
Marathon Ni-Cad Battery Functional Check (Capacity Check) Task 24-3400-720
200 Hours/1C
05-15-06
122
A243601
Standby Alternator Detailed Inspection Task 24-36-00-220
1A/1C
05-15-07
121
A245001
Power Distribution Boxes Detailed Inspection Task 24-50-00-220
2A/2C
05-15-10
121 122
A251000
Smoke Goggle General Visual Inspection Task 25-10-00-210
1C
05-15-01
801 802
A251001
Crew Seats Detailed Inspection Task 25-10-00-220
2A/2C
05-15-10
231 232
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Page 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
B251001
Inertia Reel Operational Check Task 25-10-00-710
1A/1C
05-15-07
221 232
A251003
Passenger Seats Detailed Inspection Task 25-21-00-220
2A/2C
05-15-10
231 232
A255101
Cargo Nets Detailed Inspection Task 25-51-00-220
200 Hours/1C
05-15-06
251 252 255 256 257 258
B255201
Cargo Pod Drains Operational Check Task 25-52-00-710
Annual
05-15-0A
901 902 903 904 905 906
A255201
Cargo Pod Zonal Inspection Task 2552-00-210
6C
05-15-04
901 902 903 904 905 906
B256001
ARTEX C406-2 Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-720
12 calendar months (required 14CFR 91.207)
05-15-20
220 311 312 340
B256003
ARTEX ME406 Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-721
12 calendar months (required 14CFR 91.207)
05-15-20
220 311 312 340
B256005
ARTEX C406-N Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-722
12 calendar months (required 14CFR 91.207)
05-15-20
220 311 312 340
B256007
Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-723
12 calendar months (required 14CFR 91.207)
05-15-20
220 311 312 340
A261001
Engine Fire Detection System General Visual Inspection Task 26-10-00-210
1A/1C
05-15-07
121 122
B262001
Portable Fire Extinguisher Functional Check (Weight Check) Task 26-20-00720
1C
05-15-01
215 216 251 252
B262003
Portable Fire Extinguisher Restoration (Hydrostatic Test) Task 26-20-00-780
12C
05-15-05
215 216 251 252
B262005
Portable Fire Extinguisher Restoration (Internal Inspection) Task 26-20-00290
6C
05-15-04
215 216 251 252
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
C270001
Flight Controls Lubrication Task 2700-00-640
Annual
05-15-0A
215 226 374 525 625
C271001
Aileron Trim System Lubrication Task 27-10-02-640
1A/1C
05-15-07
211 212 217 218 233 234 253 254 251 252 551 571 651 671
B271001
Spoiler System Functional Check Task 27-10-00-720
2A/2C
05-15-10
211 212 217 218 233 234 253 254 251 252 503 525 603 625
B271003
Aileron System Functional Check Task 27-10-00-721
4A/2C
05-15-11
211 212 217 218 233 234 253 254 251 252 503 525 603 625
C271003
Aileron Trim Tab Actuator (2660044-1) Lubrication Task 27-10-02-641
2A/2C
05-15-10
551 571 651 671
B271005
Aileron Trim Tab (Free Play) Functional Check Task 27-10-02-720
1A/1C
05-15-07
551 571 651 671
C271005
Aileron Trim Tab Actuator (2661615-1, 26616159, or 2661615-10) Lubrication Task 27-10-02-642
4A/5C
05-15-12
551 571 651 671
B272001
Rudder System Functional Check (Standard Rudder Installation) Task 27-20-00-720
4A/2C
05-15-11
211 212 213 214 217 218 233 234 253 254 257 258 311 312 320 341
C272001
Rudder Bar Bearings and Rudder Pedals Lubrication Task 27-20-00-640
4A/2C
05-15-11
211 212 213 214
216 373 503 603
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Page 3 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
Revised Mar 1/12
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
B272003
Rudder System Functional Check (Float Kit Installation) Task 27-20-00721
1C
05-15-01
211 213 217 233 253 257 311 320
212 214 218 234 254 258 312 341
B273001
Elevator System Functional Check Task 27-30-00-720
4A/2C
05-15-11
211 213 217 233 253 257 311 320 374 376
212 214 218 234 254 258 312 373 375
C273001
Elevator Trim Tab Actuator (26600171) Lubrication Task 27-30-02-640
2A/2C
05-15-10
371 372 375 376
C273003
Elevator Trim Tab Actuator (26612151 and 2661215-9) Lubrication Task 2730-02-641
4A/5C
05-15-12
371 372 375 376
B273003
Elevator Trim Tab (Free Play) Functional Check Task 27-30-02-720
1A/1C
05-15-07
371 372 375 376
B273101
Stall Warning System Operational Check Task 27-31-00-710
Annual
05-15-0A
211 212 503
C275001
Flap Tracks and Rollers Lubrication Task 27-50-00-640
Annual
05-15-0A
525 527 625 627
B275001
Flap System Functional Check Task 27-50-00-720
4A/2C
05-15-11
'251 252 511 611 525 625
A275001
Flap Actuator Mount Bracket Detailed Inspection Task 27-50-00-220
2A/1C
05-15-09
231 232
A275003
Flap Bellcrank Detailed Inspection Task 27-50-00-221
2A/1C
05-15-09
251 252 511 611 525 625
A281001
Fuel Filler Assembly Detailed Inspection Task 28-10-01-220
Annual
05-15-0A
521 621
B281001
Fuel Vent Line Float Valve Operational Check Task 28-10-03-710
200 Hours/1C
05-15-06
575 675
A281003
Fuel Storage System Detailed Inspection Task 28-10-01-221
4C
05-15-03
521 621
B282103
Firewall Fuel Shutoff Valve Control Operational Check Task 28-21-00-711
1A/1C
05-15-07
213 214 220
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
Added Mar 1/12
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
C282301
Wing Shutoff Valve Linkage Lubrication Task 28-23-00-640
1A/1C
05-15-07
231 232 511 611
B284101
Fuel Reservoir Warning System Operational Check Task 28-41-00-710
Annual
05-15-0A
ENG
B284103
Fuel Quantity and Low Fuel Warning Systems Functional Check Task 2841-00-720
2A/2C
05-15-10
AUX
B301001
Bleed Air Pressure Regulator Functional Check Task 30-10-00-720
1A/1C
05-15-07
122 AUX
B301003
Bleed Air Pressure Regulator Functional Check Task 30-10-00-720
200 Hours/1C
05-15-06
122 AUX
B301101
TKS Anti-Ice System Functional Check Task 30-11-00-720
1A/1C
05-15-07
AUX
B301102
Inboard TKS Wing Panel Pressurization Functional Check Task 30-11-00-721
1C
05-15-01
501, 601, AUX
B304001
Windshield Anti-Ice System Operational Check Task 30-40-00-710
1A/1C
05-15-07
AUX
A321001
Main Landing Gear Detailed Inspection Task 32-10-00-220
1C
05-15-01
721 722
A321003
Center-Spring and Main Gear-Spring Interface Area Special Detailed (Corrosion Inspection and Repair) Task 32-10-00-221
4C
05-15-03
721 722
A321005
Main Landing Gear Axle Special Detailed Inspection (SID 32-10-01) Task 32-10-00-240
MB
05-15-MB
721 722
A322001
Nose Landing Gear Detailed Inspection Task 32-20-00-220
1A/2C
05-15-08
710
B322001
Shimmy Damper Functional Check Task 32-20-02-720
Annual
05-15-0A
710
A324001
Brakes Detailed Inspection Task 3240-00-220
1C
05-15-01
721 722
B324001
Brakes Operational Check Task 3240-00-710
2A/2C
05-15-10
ENG
A324005
Main Landing Gear Wheels and Tires Detailed Inspection Task 32-40-00-222
1C
05-15-01
721 722
A324009
Nose Landing Tire Detailed 32-40-00-224
Gear Wheel and Inspection Task
1C
05-15-01
710
B332001
Passenger/Cargo Compartment Lighting Operational Check Task 33-20-00-710
2A/2C
05-15-10
AUX
5-10-01 © Cessna Aircraft Company
Page 5 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
B341101
Pitot Tube Heaters Operational Check Task 34-11-00-710
200 Hours/1C
05-15-06
AUX
B341103
Pitot/Static System Functional Check Task 34-11-00-720
24 calendar months (required 14 CFR 91.411)
05-15-21
AUX
B342101
Magnetic Compass Functional Check Task 34-21-00-720
1C
05-15-01
ENG
B345001
Transponder Functional Check Task 34-50-00-720
24 calendar months (required 14 CFR 91.413)
05-15-22
AUX
B350101
Oxygen System Operational Check Task 35-01-00-710
1C
05-15-01
231 251 255 311 801
A520001
Crew Doors Detailed Inspection Task 52-00-00-220
2A/2C
05-15-10
801 802
A520003
Passenger/Cargo Doors and Door Frames Detailed Inspection Task 52-00-00-221
2A/2C
05-15-10
255 256 257 258 803 804
A531001
External Fuselage Zonal Inspection Task 53-10-00-210
6C
05-15-04
ALL
A531003
Internal Cockpit Zonal Inspection Task 53-10-00-211
4C
05-15-03
211 213 215 217 231 233
A531004
Internal Cabin Zonal Inspection Task 53-10-00-212
6C
05-15-04
251 252 253 254 255 256 257 258 311 312
A531005
Fuselage to Strut Attach Fitting Lugs (Nominal Standard Bolt Size) (Typical Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 53-20-07-250
MH
05-15-MH
251 252 253 254
232 252 256 312 802
212 214 216 218 232 234
5-10-01 © Cessna Aircraft Company
Page 6 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
A531006
Fuselage to Strut Attach Fitting Lugs (Nominal Standard Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 53-20-07-251
MG
05-15-MG
251 252 253 254
A531007
Internal Tail Cone Zonal Inspection Task 53-10-00-213
4C
05-15-03
311 312 320 330
A531008
Fuselage Engine Mount Fittings Special Detailed Inspection (SID 53-10-01) Task 53-10-00-250
AD
05-15-13
121 122 130
A531009
Seat Rails and Attachment Structure Detailed Inspection (SID 53-10-07) Task 53-25-00-220
AF
05-15-15
231 232 233 234 251 252 253 254 255 256 257 258
A531010
Fuselage to Strut Attach Fitting Lugs (Oversize 1/64 - Inch Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 53-20-07-252
MI
05-15-MI
251 252 253 254
A531011
Fuselage to Strut Attach Fitting Lugs (Oversize 1/32- Inch Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 53-20-07-253
MJ
05-15-MJ
251 252 253 254
A531013
Empennage and Horizontal Stabilizer Zonal Inspection Task 53-10-00-214
6C
05-15-04
340 341 373 374
A532003
Cargo and Passenger Door Doublers Special Detailed Inspection (SID 5320-01) Task 53-10-00-251
AF
05-15-15
255 256 257 258 803 804
A532004
Lower Forward Carry-Thru Bulkhead Special Detailed Inspection (SID 5320-03) Task 53-10-00-253
AG
05-15-16
253 254
A532005
Main Landing Gear Fitting Special Detailed Inspection (SID 53-20-04) Task 53-10-00-254
AJ
05-15-19
253 254
A532006
Main Landing Gear Attach Fittings and Aft Carry-Thru Bulkhead (SID 53-2005) Task 53-10-00-255
AG
05-15-16
253 254
A532007
Fuselage to Wing Attach Fitting Lugs Special Detailed Inspection (SID 5320-02) Task 53-10-00-252
MF
05-15-MF
251 252 501 511 525 601 611 625
A532008
Firewall Brace and Doubler Assemblies Detailed Inspection (SID 53-20-11) Task 53-10-00-223
AD
05-15-13
121 122 130
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Page 7 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
A532009
Carry-Through Root Rib Detailed Inspection (SID 53-20-08) Task 53-10-00-220
AD
05-15-13
251 252 500 600
A532011
Crew Door Frames Detailed Inspection (SID 53-20-09) Task 53-10-00-221
AD
05-15-13
231 232 233 234 801 802
A532012
Passenger and Cargo Door Frames Detailed Inspection (SID 53-20-10) Task 53-10-00-222
AD
05-15-13
255 256 257 258 803 804
A532013
Bulkheads and Stiffeners Below the Seat Rail Attachments at FS 143.00 and FS 158.00 Detailed Inspection (SID 53-20-12) Task 53-25-00-221
AF
05-15-15
233 234
A532014
Stringers at Intersections with Forward and Aft Carry - Thru Bulkheads Detailed Inspection (SID 53-20-13) Task 53-10-00-224
AG
05-15-16
251 252
A532015
Fuselage Skin Doubler at Main Landing Gear Cutout Detailed Inspection (SID 53-20-14) Task 53-10-00-225
AE
05-15-14
253 254
A532016
Fuselage to Wing Carry-Thru Attach Fitting and Bulkhead Special Detailed Inspection (SID 53-20-06) Task 53-1000-256
AG
05-15-16
251 252
A535001
Fuselage to Horizontal Stabilizer Attach Fittings Special Detailed Inspection (SID 53-50-01) Task 53-10-00-257
AD
05-15-13
320 373 374
A535002
Vertical Stabilizer Attach Points Special Detailed Inspection (Typical Inspection Compliance) (SID 53-50-02) Task 53-10-00-258
AD
05-15-13
311 312 320 341
A535003
Vertical Stabilizer Attach Points Special Detailed Inspection (Severe Inspection Compliance) (SID 53-50-02) Task 53-10-00-259
AH
05-15-17
311 312 320 341
A551003
Horizontal Stabilizer Forward and Aft Attach Points Special Detailed Inspection (SID 55-10-01) Task 55-10-00-250
AD
05-15-13
373 374
A551005
Horizontal Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) (SID 55-10-02) Task 55-10-00-252
AI
05-15-18
373 374
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Page 8 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
A553001
Vertical Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) (SID 55-30-01) Task 55-30-00-250
AD
05-15-13
320 341
A553002
Vertical Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) (SID 55-30-01) Task 55-30-00-251
AH
05-15-17
320 341
A553004
Horizontal Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) (SID 55-10-02) Task 55-10-00-251
AD
05-15-13
373 374
B560001
Functional Check of the Windshield Task 56-00-01-720
2C
05-15-02
240
A564002
Windshield and Attachment Structure Detailed Inspection (SID 56-30-01) Task 56-00-01-220
AD
05-15-13
240
A570008
Wing Strut Fittings Special Detailed Inspection (Typical Inspection Compliance) (SID 57-60-01) Task 57-10-01-250
MA
05-15-MA
531 631
A570009
Wing Strut Fittings Special Detailed Inspection (Severe Inspection Compliance) (SID 57-60-01) Task 57-10-01-251
ME
05-15-ME
531 631
A570010
Front Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection (SID 57-20-02) Task 57-10-00-252
MF
05-15-MF
501 521 601 621
A570011
Rear Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection (SID 57-20-03) Task 57-10-00-253
MF
05-15-MF
521 525 621 625
A570012
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Typical Inspection Compliance) (SID 57-60-02) Task 57-10-01-252
MF
05-15-MF
531 631
A570013
Wing to Carry - Thru Front Spar Attachment Fittings Special Detailed Inspection (SID 57-20-01) Task 57-10-00-250
MF
05-15-MF
251 252
A570014
Wing to Carry - Thru Rear Spar Attachment Fittings Special Detailed Inspection (SID 57-20-01) Task 57-10-00-251
MF
05-15-MF
251 252
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
A570015
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Severe Inspection Compliance) (SID 57-60-02) Task 57-10-01-253
MF
05-15-MF
531 631
A570016
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/64 Inch Oversize Bolt Size) (Severe Inspection Compliance) (SID 57-60-02) Task 57-10-01-254
MK
05-15-MK
531 631
A570017
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/32 Inch Oversize Bolt Size) (Severe Inspection Compliance) (SID 57-60-02) Task 57-10-01-255
ML
05-15-ML
531 631
A571001
Wing Zonal Inspection Task 57-10-00210
4C
05-15-03
500 600
A575002
Center Flap Track and Inboard Flap Track Special Detailed Inspection (SID 57-50-01) Task 57-10-00-254
MD
05-15-MD
525 527 625 627
A575003
Outboard Flap Track Special Detailed Inspection (SID 57-50-01) Task 57-1000-255
MD
05-15-MD
525 527 625 627
B611001
Hartzell Propeller Functional Check Task 61-10-00-720
1A/1C
05-15-07
110
B611101
McCauley Propeller Functional Check Task 61-11-00-720
1A/1C
05-15-07
110
A710001
Engine Compartment Inspection Task 71-00-01-210
2C
05-15-02
130
A712001
Engine Mounts and Firewall Detailed Inspection Task 71-20-00-220
1A/1C
05-15-07
130
A712003
Engine Truss and Ring Assembly Special Detailed Inspection (SID 71-20-01) Task 71-20-00-240
AD
05-15-13
130
A714101
Engine Wash Ring, Air Plenum, and Thermocouple (T1) Detailed Inspection Task 71-41-00-220
1C
05-15-01
130
A716001
Inertial Air Separator Detailed Inspection Task 71-60-00-220
1A/1C
05-15-07
130
B761001
Engine Controls Functional Check Task 76-10-00-720
1A/1C
05-15-07
130 211 212 ENG
Zonal
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REVISION STATUS
ITEM CODE NUMBER
TASK
INTERVAL CH SE SU
ZONE
B761003
Emergency Power Lever Annunciator Light (EPL) Operational Check Task 76-10-01-710
Annual
05-15-0A
AUX
A781001
Primary and Secondary Exhaust Duct General Visual Inspection Task 78-1000-211
2A/2C
05-15-10
130
A801001
Starter-Generator (Part Number 23081 Series only) Detailed Inspection Task 80-10-00-220
1A/1C
05-15-07
130
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MODEL 208 MAINTENANCE MANUAL COMPONENT TIME LIMITS 1.
2.
Component Time Limit A.
All components not recorded here must be examined as given in other sections of this chapter and repaired, overhauled or replaced as required. Items shown here must be overhauled or replaced during the regular maintenance periods that are to be done nearest to the speciÞed limit. Also, those items that are underlined are recorded in Chapter 4, Airworthiness Limitations.
B.
The replacement life for each component recorded in this section applies to the part throughout its life on the original installation and on later installations. The life (number of hours or number of landings) must be recorded individually for these components and must stay with the component during removal. For example, if a component is removed for overhaul, it must be tagged with the life (number of hours or number of landings) to the date of removal and this tag must remain with the component throughout the overhaul process. (Overhaul of a component does not zero time the life of the component.) When received from overhaul and installed on an airplane, the life of the component must be recorded for continued accumulation toward the life-limit.
Schedule A.
Equipment and Furnishings (Chapter 25) (1) Emergency Locator Transmitter (ELT) Battery Pack - Replace at replacement date. Refer to Task 25-60-00-960. NOTE:
B.
Replace battery if transmitter has been in use for more than one cumulative hour or when 50 percent of the useful life of the battery has expired.
Flight Controls (Chapter 27) (1) Flap Bell Crank (Part Number 2622083-18) - Replace at every 2250 landings. (2) Flap Bell Crank (Part Number DDA00028-4) - Replace at every 2250 landings. (3) Flap Bell Crank (Part Numbers 2622281-2, -12) - Replace at every 7000 landings. (4) Flap Bell Crank (Part Number 2692001-2) - Replace at every 7000 landings. NOTE:
Total landings includes the accumulated landings of 2622281-2 prior to modiÞcation by SK208-123 to the 2692001-2 conÞguration.
(5) (6)
Flap Bell Crank (Part Numbers 2622311-7, -16) - Replace at every 40,000 landings. Flap Bell Crank (Part Number 2622311-7) attaching parts: Bearings (Part Number MS27641-5 or S3952-5) and Bolt (Part Number AN5-77) - Replace at every 10,000 landings. (7) Flap Bell Crank (Part Number 2622311-16) attaching parts: Bearings (Part Number KP5A-H) and Bolt (Part Number AN5-77) - Replace at every 10,000 landings. (8) Elevator Forward Pushrod Part Numbers 2613440-1, 2613414-1, and 2660034-1 - Replace at 9500 landings. (9) Elevator Forward Pushrod Part Numbers 2613440-3, 2613440-5, DDA05946-1 - Replace at 40,000 landings. (10) Elevator Aft Pushrod Part Numbers 2634009-1, 2634027-1, and 2634027-3 - Replace at 40,000 landings. C.
Ice and Rain Protection (Chapter 30) (1) TKS Metering Pumps Part Number 9514A-1- Replace every 5000 ßight hours.
D.
Landing Gear (Chapter 32) (1) Main Landing Gear NOTE:
(a)
Attaching hardware (bolts, bearings, bushings, and trunnion pins related to the installation of the components below) is to be replaced whenever the associated component is replaced.
Main Landing Gear Center Spring (Part Numbers 2641014-2, -3, -4, -5, -6 -7, -8, -9) Replace at every 31,500 landings.
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MODEL 208 MAINTENANCE MANUAL (b)
(2)
Main Landing Gear Trunnion Assembly (Part Numbers 2641012-1, -2, -8, -9, -13 -14) Replace at every 31,500 landings. (c) Main Landing Gear Spring (Part Numbers 2641013-1, -2, -3, -4, -5, -6, 7, -8) - Replace at every 31,500 landings. (d) Main Landing Gear Attach Pin (Part Numbers 2641008-1, -2, -200) - Replace at every 31,500 landings. (e) Main Landing Gear Axles (Part Numbers 2641011-1, -3, -4) - Replace at every 10,000 landings. (f) Main Landing Gear Axles (Part Numbers 2641011-5) - Replace at every 31,500 landings. (g) Main Landing Gear Axle Fittings (Part Numbers 2641010-1, -3, -7) - Replace at every 31,500 landings. Nose Landing Gear NOTE:
(a)
(b)
Attaching hardware (bolts, bearings, bushings, and axle components related to the installation of the components below) is to be replaced whenever the associated component is replaced.
Nose Gear Drag Link Spring (Part Numbers 2643062-1, -2, -3) - Replace at every 15,000 landings. NOTE:
For nose gear drag link springs repaired per CAB96-24 or per Chapter 32, Nose Landing Gear - Cleaning/Painting with damage repaired between 0.050 inch and 0.062 inch (1.270 mm and 1.575 mm), the life limit is an additional 12,000 landings after repair, not to exceed 15,000 landings.
NOTE:
For nose gear drag link springs repaired per CAB96-24 or per Chapter 32, Nose Landing Gear - Cleaning/Painting with damage repaired between 0.063 inch and 0.075 inch (1.600 mm and 1.905 mm), the life limit is an additional 10,000 landings after repair, not to exceed 15,000 landings.
Nose Gear Assembly (Part Numbers 2643045, 2643100 and 2643095 Series Part Numbers) - Replace at every 40,000 landings. NOTE:
(c)
Support Assembly, Nose Gear Spring (Part Numbers 2643030, 2643055 and 2643099 Series Part Numbers) - Replace at every 40,000 landings. NOTE:
(d)
For an illustration of the nose gear assembly to be replaced, refer to Chapter 32, section 32-20-00, Nose Landing Gear - Maintenance Practices.
For an illustration of the nose gear spring support assembly to be replaced, refer to Chapter 32, section 32-20-00, Nose Landing Gear - Maintenance Practices.
Fork Assembly, Nose Gear Spring (Part Numbers 2643031-1, -7) - Replace at every 40,000 landings. NOTE:
For an illustration of the nose gear spring fork assembly to be replaced, refer to Chapter 32, section 32-20-00, Nose Landing Gear - Maintenance Practices.
E.
Navigation (Chapter 34) (1) Pitot and static hoses - Replace after 10 years in service.
F.
Oxygen (Chapter 35) (1) The airplane may be equipped with a two-port oxygen system incorporating a 50.67 cubic-footcapacity oxygen cylinder (Part Numbers C166001-1101 and C166001-1201) or a ten port oxygen system incorporating a 116.95 cubic-foot-capacity oxygen cylinder (Part Numbers C1660011102 and C166001-1103). Both cylinders have a life limit of 15 years. Refer to Task 35-01-00960. (2) Oxygen Cylinder hydrostatic test intervals are determined by DOT-E 8162 or DOT-SP 8162. Refer to Table 1, Hydrostatic Test Intervals.
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MODEL 208 MAINTENANCE MANUAL
Table 1. Hydrostatic Test Intervals DOT Cylinder Marking
Last Retest
Next Retest
Subsequent Retest Interval
Refer to Task
E 8162
Before July 1, 2006
3 years after last hydrostatic test
5 years
Task 35-01-00-780
E 8162
On or After July 1, 2006
5 years after last hydrostatic test
5 years
Task 35-01-00-780
SP 8162
On or After July 1, 2006
5 years after last hydrostatic test
5 years
Task 35-01-00-780
(3)
Scott 359 Series Oxygen Mask All Components Including Regulator - Overhaul/Replacement 5 years. Refer to Task 35-15-00-960.
G.
Vacuum (Chapter 37) (1) Vacuum hoses - Replace after 10 years in service. (2) Vacuum System Central Air Filter Discard (C294502-0201 ) - Refer to Parker HanniÞn Airborne Services Service Letter 59B for interval, Refer to Task 37-10-00-960 for procedure. (3) Vacuum Relief Valve Filter Discard (C482001-0202) - Refer to Parker HanniÞn Airborne Services Service Letter 59B for interval, Refer to Task 37-10-00-961 for procedure.
H.
Propeller - Hartzell (Chapter 61) (1) Propeller - Overhaul, Refer to Hartzell Propeller Service Letter HC-SL-61-61Y, Revision 3 or later. (2) Governor (Woodward) - Overhaul, Refer to Pratt & Whitney Service Bulletin number 1703. (3) Overspeed Governor (Woodward) - Overhaul, Refer to Service Bulletin 33580.
I.
Propeller McCauley (Chapter 61) (1) Propeller - Overhaul, Refer to McCauley Propeller Service Bulletin 137AE or latest revision. (2) Governor (Woodward) - Overhaul, Refer to Pratt & Whitney Service Bulletin number 1703. (3) Overspeed Governor (Woodward) - Overhaul, Refer to Service Bulletin 33580.
J.
Power plant (Chapter 71) (1) Refer to Pratt & Whitney Maintenance Manual listed in the List of Publications in the front of this publication and Pratt & Whitney Service Bulletin No. PT6A-72-1703, Rev. No. 1 or later. NOTE:
Engine components, such as standby alternator, etc., should be inspected for condition at time of engine overhaul, as it may be cost effective to overhaul or replace marginal components at that time. A determination is to be made during engine overhaul such that if components have less hours in service than the engine, or have not accumulated sufÞcient hours for economic reasons, these components may not require overhaul or replacement concurrent with engine overhaul. It is recommended that the overhaul or replacement interval for these components not exceed the engine overhaul interval.
NOTE:
Inspect the engine compartment for structural damage when engine is removed for overhaul, and make the necessary repairs.
NOTE:
Inspect the engine exhaust as it may be cost effective to replace marginal components at engine overhaul.
NOTE:
Inspect electrical harnesses for damage which would be cost effective to replace at engine overhaul.
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MODEL 208 MAINTENANCE MANUAL (2)
(3)
K.
Fuel Hose: From Þrewall fuel Þlter to engine fuel heater. From fuel control unit motive ßow return to Þrewall Þtting - Replace rubber hoses, base number S2495 (Purchased through Cessna), every 5 years or 3600 hours, whichever occurs Þrst. Replace Teßon hoses, base number S2808 (Purchased through Cessna), every 10 years or at engine TBO, whichever occurs Þrst. Oil Hose: Oil cooler supply from engine external scavenge pump to oil cooler inlet. From oil cooler return outlet to engine oil tank. Torque indicating pressure hose. Engine oil pressure indicating hose from engine to Þrewall. Torque indicating vent hose - Replace Teßon engine compartment ßexible ßuid-carrying hoses, base number S2808/AE3663 (Purchased through Cessna), every 10 years or at engine TBO, whichever occurs Þrst.
Starter-Generator (Chapter 80) (1) Starter-Generator (Lear Siegler/Lucas Aerospace) (a) Except part number 23081-023A, - Overhaul every 1000 hours. Refer to Task 80-10-00960. (b) Part number 23081-023A - Overhaul/replace every 1200 hours. Refer to Task 80-10-00960. (2) Starter-Generator (Aircraft Parts Corp. - APC "XL") (a) Overhaul with genuine APC parts (SNL02-1) - Overhaul every 1600 hours. Refer to Task 80-10-00-960 (b) Original equipment starter-generators APC 200SGL119Q-2 - Overhaul every 2000 hours. Refer to Task 80-10-00-960 (c) Starter-generators 200SGL119Q-2RX units overhauled by Cessna/APC - Overhaul every 2000 hours. Refer to Task 80-10-00-960 (3) Starter/Generator (Aircraft Parts Corp. - APC part number 300SGL145Q) - Overhaul/ replace every 1000 hours. Refer to Task 80-10-00-960
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MODEL 208 MAINTENANCE MANUAL COMPONENT TIME LIMITS - RUSSIAN CERTIFIED AIRPLANES 1.
2.
Component Time Limit A.
All components not listed herein should be inspected as detailed elsewhere in this chapter and repaired, overhauled or replaced as required. Items shown here should be overhauled or replaced during regular maintenance periods falling due nearest to the specified limit.
B.
Replacement life of each component listed in this section applies to the part throughout its life on the original installation and on later installations. The life (number of hours or number of landings) must be recorded individually for these components and must remain with the component during removal. For example, if a component is removed for overhaul, it must be tagged with the life (number of hours or number of landings) to the date of removal and this tag must remain with the component throughout the overhaul process. (Overhaul of a component does not zero time the life of the component.) When received from overhaul and installed on an airplane, the life of the component must be recorded to allow continued accumulation toward the life limit.
Schedule A.
Flight Controls (Chapter 27) (1) Flap bellcrank Part Number 2622083-18 -- Replace at 1500 landings. (2) Flap bellcrank Part Number DDA00028-4 -- Replace at 1500 landings. (3) Flap bellcrank Part Number 2622281-2 -- Replace at 5000 landings. (4) Flap bellcrank Part Number 2622281-12 -- Replace at 5000 landings. (5) Flap bellcrank Part Number 2692001-2 -- Replace at 5000 landings. NOTE:
B.
Total landings include the accumulated landings of 2622281-2 prior to modification by SK208-123 to the 2692001-2 configuration.
Landing Gear (Chapter 32) (1) Main landing gear axle Part Numbers 2641011-1, -3 and -4 -- Replace at 10,000 landings (2) Main landing gear assembly Part Numbers 2641000 and 2641027 (all assembly dash numbers) -- Replace at 20,000 landings. (3) Nose landing gear drag link spring Part Numbers 2643062-1 and -2 -- Replace at 15,000 landings. (4) Nose landing gear assembly Part Numbers 2643000-1, -2, -3, -4 and -5 -- Replace at 20,000 landings.
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MODEL 208 MAINTENANCE MANUAL SUPPLEMENTAL INSPECTION DOCUMENT 1.
2.
Supplemental Inspection Document A.
Introduction (1) The Supplemental Structural Inspection Program for the Cessna Models 208 and 208B airplane is a result of the Models 208 and 208B airplane's current usage and state-of-the-art analysis, testing and inspection methods. Analysis methods include durability, fatigue, and damage tolerance assessments. A practical state-of-the-art inspection program is found for each Principal Structural Element (PSE). The FAA has defined a PSE in AC25.571: (a) A PSE is an element that contributes significantly to carrying flight, ground or pressurization loads, and whose integrity is essential in maintaining the overall structural integrity of the airplane. (2) The Supplemental Structural Inspection Program was made with the combined efforts of Cessna Aircraft Company and Model 208 and 208B operators. The inspection program is the current structural maintenance inspection, plus supplemental inspections for continued airworthiness of the airplane as years of service are collected. The primary function of the Supplemental Structural Inspection Program is to find fatigue damage which will increase with time. In addition to the supplemental inspections, we started a Corrosion Prevention and Control Program (CPCP) to prevent or control corrosion that can have an effect on the continued airworthiness of the airplane. (3) The Supplemental Structural Inspection Program is valid for airplanes with less than 50,000 hours. Beyond this the continued airworthiness of the airplane can no longer be assured. Retirement of the airframe is recommended when 50,000 flight hours have been accumulated.
B.
Function (1) The function of the Supplemental Structural Inspection Program is to find damage from fatigue, overload or corrosion through the use of the Nondestructive Inspections (NDI), and visual inspections. This Supplemental Inspection Document (SID) is only for primary and secondary airframe components. Engine, electrical items and primary and secondary systems are not included in this document. (a) A list is included to show the requirements for the SID program for primary and secondary airframe components. 1 The airplane maintenance agrees with Cessna's recommendations or the equivalent. 2 If the SID is for a specific part or component, you must examine and evaluate the surrounding area of the parts and equipment. If problems are found outside these areas, report them to Cessna Aircraft Company on a reporting form. Changes can then be made to SID program, if necessary. 3 The SID inspections are for all Cessna Models 208 and 208B airplanes. The inspection intervals are for unmodified airplanes, and represent the maximum approved inspection times. On airplanes that changed the airplane design, gross weight, or airplane performance, it can be necessary to do inspections more frequently. Examples of some Supplemental Type Certification (STC) installations, which will require modified inspection intervals include vortex generators and non-standard engines. The owner or the maintenance organization should contact the STC holder(s) or modification originator to get new FAA approved inspection information.
Principal Structural Elements A.
Principal Structural Elements Description (1) An airplane component is classified as a Principal Structural Element (PSE) if: (a) The component contributes significantly to carrying flight and ground loads. (b) And, if the component fails, it can result in a catastrophic failure of the airframe. (2) The monitoring of these PSEs is the main focus of this Supplemental Structural Inspection Program. (3) Typical examples of PSEs, as listed in FAA Advisory Circular 25.571, are listed in Table 1.
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Table 1. Typical Examples of Principal Structural Elements (PSEs) Wing and Empennage Control surfaces, flaps and their mechanical systems and attachments (hinges, tracks and fittings) Primary fittings Principal splices Skin or reinforcement around cutouts or discontinuities Skin-stringer combinations Spar caps Spar webs Fuselage Circumferential frames and adjacent skin Door frames Pilot window posts Bulkheads Skin and single frame or stiffener element around a cutout Skin or skin splices, under circumferential loads Skin or skin splices, under fore-and-aft loads Door skins, frames and latches Window frames Landing Gear and their Attachments Engine Support Structure and Mounts B.
3.
Selection Criteria (1) The factors used to find the PSEs in this document include: (a) Service Experience 1 Two sources of information were used to find the service discrepancies. a Cessna Service Bulletins and Service Information Letters issued to repair common service discrepancies were examined. b FAA Service Difficulty Records were examined. 2 The data collected was also used to find a component's susceptibility to corrosion or accidental damage as well as its inspectability. (b) Fatigue And Damage Tolerance Analysis 1 Fatigue and damage tolerance analyses were conducted for the critical areas of the PSEs. Details of these analyses are presented in Section 3, Durability - Fatigue And Damage Tolerance. (c) Testing 1 Test results from previous static tests and fatigue cyclic tests were reviewed to identify the critical areas of the PSEs.
Durability - Fatigue And Damage Tolerance A.
Airplane Usage (1) Airplane usage data for the SID program is from the analysis of the in-service use of the airplane. Operational data for the SID is from operator surveys and from CESCOM (Computerized Maintenance Program). This information was used to make the fatigue loads spectra.
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MODEL 208 MAINTENANCE MANUAL (2) (3)
Usage results from inspection of typical flight lengths, takeoff gross weights, payloads and fuel. One flight profile was made to represent each usage. Different usages are mixed to make a fatigue loads spectrum. The flight profile has the applicable facts of the flight. The profile gives the gross weight, payload, fuel, altitude, speed, distance, etc., necessary to find the pertinent flight and ground parameters to develop the fatigue loads. The flight is then divided into operational segments, where each segment represents the average values of the parameters (speed, payload, fuel, etc.) that are used to calculate the fatigue loads spectrum.
B.
Stress Spectrum (1) A fatigue loads spectrum, which is the gross area stress, was made for each PSE to be analyzed. This analysis is based on the usage-flight profile. (2) The spectrum represents the loading environments, they are: (a) Flight loads (gust and maneuver) (b) Landing impact (c) Balancing tail loads (d) Thrust loads (e) Ground loads (taxi, turning, landing, braking, pivoting, etc.), (f) Ground-air-ground cycles (3) The spectrum is representative of a flight-by-flight, cycle-by-cycle random loading sequence that gives a reflection of the applicable and important airplane response characteristics. (4) After an examination of the airplane usage data and the ways in which the surveyed airplanes were flown, two sets of stress profiles were developed, one for each flight profile. (a) The first flight profile represents a typical usage with an average flight length of approximately one hour. (b) The second flight profile represents severe usage, which includes flight lengths of less than 35 minutes, flights at low altitudes or operations in high humidity climates.
C.
Classification for Types of Operation (1) To find the frequency of inspections, first find the total number of hours and the total number of landings on the airplane, then find an average flight length over the life of the airplane, whichever is less.
(2) (3)
(4)
If the average flight length is less than or equal to 35 minutes, then use the Severe Flight Profile for inspections. For airplanes with an average flight length greater than 35 minutes you must determine the severity of the operating environment. If the airplane operates thirty percent or more of its flight time in severe environments, you must use the severe operation inspection times found in section 4-10-01, Severe Operation Inspection and 5-14-00, Listing Of Supplemental Inspections. Examples of severe environments would include flight operations at low altitude (i.e., less than 5,000 ft. above ground level) such as pipeline patrol, sight-seeing, training flights, traversing mountainous terrain or flying near coastal areas identified in section 51-11-00, Corrosion Severity Maps - Description and Operation. Airplanes operated in all other environmental conditions will use the typical operation inspection times found in 4-10-00, Typical Inspection Time Limits and 5-14-00,Listing Of Supplemental Inspections.
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MODEL 208 MAINTENANCE MANUAL (5) D.
Make an airplane logbook entry that gives the applicable inspection schedule after it has been found.
Damage Tolerance and Fatigue Assessments (1) The damage tolerance and fatigue analysis gives the information used to set the inspection frequency requirements for each PSE. The analysis includes the possible fatigue failure locations and types of damage. Required inspections are set using analytical results, available test data and service experience. The analysis includes standard fatigue analyses, the crack growth time history and the strength that remains. Linear elastic fracture mechanics are used for the damage tolerance analysis, while fatigue analyses use the "Palmgren-Miner" linear cumulative damage theory. (2) In the analysis, the possible structural condition areas related to the airplanes as they are used more frequently, are carefully examined. Examples include: (a) Large areas of structure working at the same stress level, which can start widespread fatigue damage. (b) A number of small (less than detection size) adjacent cracks suddenly connected into a long crack (as in a line of rivet holes). (c) A load supplied from adjacent damaged parts that can cause faster damage to nearby parts (the "domino" effect). (d) Failure of multiple load path structure at the same time (crack arrest structure). (3) Initial inspections of a selected area of structure are set by both crack growth and fatigue analysis results, as well as test results. (a) The structures which were found to be fail-safe, use the fatigue life to set the initial inspections. (b) The locations with long fatigue life, the maximum initial inspection has a limit of 20,000 flight hours. (c) Structure which was found to be fail-safe includes the Models 208 and 208B wing, fuselage and empennage. (4) For locations that were identified as very important by tests, the initial inspection time was set by crack growth. The crack growth for each PSE is calculated from the initial crack size co to the crack length at instability/failure, ccrit, because of limit load. The crack's growth during time is shown by crack length versus time in flight hours. All inspections that were done again and again were derived from crack growth.
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MODEL 208 MAINTENANCE MANUAL
4.
Reporting - Communications A.
Discrepancies (1) For the SID to continue to stay applicable, it is necessary to have a free flow of information between the operator, the FAA and Cessna Aircraft Company. The important information about the inspection results, repairs and modifications done must be supplied to Cessna Aircraft Company to let the effect of the recommended inspection procedures and inspection intervals to be calculated. (2) Also, the operator's inspections and reports can find items not included in the SID before. These items will be examined by Cessna Aircraft Company and will be added to the SID for all of the operators, if applicable. (3) Cessna Propeller Aircraft Product Support has a system to collect the reports. The applicable forms are included in this document. Copies of these forms are also available from a Cessna Service Station or Cessna Field Service Engineer.
B.
Discrepancy Reporting (1) Discrepancy reporting is very important to let the inspection start times and repeat times be adjusted as well as to add or remove PSEs. It can be possible to make the inspection methods, repairs, and modifications better for the PSEs from the reported data. (2) All cracks, multiple cut off fasteners, and corrosion found during the inspection must be reported to Cessna Aircraft Company within ten days. The PSE inspection results are to be reported on a form as shown on the pages that follow.
C.
Send the Discrepancy Form (1) Send all available data, which includes forms, repairs, photographs, sketches, etc., to: Cessna Aircraft Company Attn: SID Program Technical Support Services Dept. 751 Wichita, KS 67277 USA Fax: 316-942-9006 NOTE:
D.
5.
Cessna Aircraft Company Follow-Up Action (1) All SID reports will be examined to find if any of the steps are necessary: (a) Complete a check of the effect on the structural or operational condition. (b) Complete a check of other high-time airplanes to find if a service bulletin shall be issued. (c) Find if a reinforcement is required. (d) Change the SID if required.
Inspection Methods A.
6.
This system does not replace the normal channels to send information for items not included in the SID.
Selection of inspection method. (1) A very important part of the SID program is to select and evaluate the state-of-the-art nondestructive inspection (NDI) methods applicable to each PSE, and to find a minimum detectable crack length, cdet, for each NDI method. (2) The selection of NDI method uses crack direction, location, ccrit, part thickness and access as criteria. The size of the inspection task, human factors (such as qualifications of the inspection), equipment used, and access will effect an inspection. Visual, radiographic, liquid penetrant, eddy current and magnetic particle methods are used. A complete description of each of these methods is given in the Model 208 Nondestructive Testing Manual.
Applicability/Limitations A.
This SID is applicable to the Models 208 and 208B airplanes.
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MODEL 208 MAINTENANCE MANUAL B.
7.
STC Installations (1) The Cessna Models 208 and 208B airplanes can have modifications that were done by STCs by other organizations without Cessna Engineering approval. The inspection intervals given in this SID are for unchanged airplanes, and are the maximum approved inspection times. (2) On airplanes that changed the airplane design, gross weight, or airplane performance, it can be necessary to do inspections more frequently. (a) Examples of common STCs not applicable in this SID document include vortex generators and non-standard engines. The owner and/or maintenance organization must contact the STC holder(s) or modification originator to get new FAA approved inspection requirements. (3) The SID inspection times use total airframe hours/landings in service. If a specific airframe component has been replaced, the component is examined for the total component hours/landings requirements. However, any attachment structure that was not replaced when the component was replaced must be examined and use the total airframe hour/landings requirements.
PSE Details A.
Details (1) This section contains the important instructions selected by the rationale process described in Section 2, Principal Structural Elements. These items are considered important for continued airworthiness of the Cessna Models 208 and 208B airplane. Service Information Letters and Service Bulletins about the PSEs are available from Cessna Aircraft Company. (2) A summary of the PSEs is presented in the 5-14-00,Listing Of Supplemental Inspections. This can be used as a checklist by the operators. A summary of the inspections by flight hours is also given.
B.
Repairs, Alterations and Modifications (RAM) (1) Repairs, Alterations and Modifications (RAM) made to PSEs can have an effect on the inspection times and methods necessary for the SID. The flowchart in Figure 1 can be used to find if a new damage tolerance assessment and FAA approved supplemental inspection criteria are necessary. (2) For repairs that are not included in the recommendations of this SID document or the Model 208 Structural Repair Manual, contact Cessna Propeller Aircraft Product Support for assistance; (316) 517-5800 or Fax (316) 942-9006. (a) All repairs supplied by Cessna Aircraft Company since January 2003 meet the damage tolerance assessment requirements.
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Damage Tolerance Assessment Flowchart Figure 1 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL LISTING OF SUPPLEMENTAL INSPECTIONS - DESCRIPTION AND OPERATION 1.
2.
Supplemental Inspection Procedures A.
Each of the Supplemental Inspections listed in this section are now incorporated into Task based inspections. This document provides a cross reference between the Supplemental Inspections. The Supplemental Inspection Number corresponds to the section of the Model 208 Nondestructive Testing Manual section number and is also referenced in the tasks.
B.
Procedure (1) Each supplemental inspection is assigned an independent item code in Chapter 5 and Task number in the applicable ATA chapters. (2) The item codes are in Chapter 5, Inspection Time Limits and in the Inspection Documents. The item codes for the supplemental inspections below have not changed but for editorial reasons, the letter A was added. (3) Inspections that are also necessary for Chapter 4, Airworthiness Limitations, are done at intervals that start with M to help you keep records. The intervals for these inspections are specified in Chapter 4. It is necessary to complete the Chapter 4 inspections on or before the specified interval. Chapter 4 inspections do not have a grace period
Supplemental Inspections to Task Matrix
Table 1. Task Inspection to SID DETAILS FOUND IN TASK
SUPPLEMENTAL INSPECTION NUMBER
INSPECTION COMPLIANCE TITLE
INSPECTION DOCUMENT
ITEM CODE
Task 240
32-10-00-
32-10-01
Main Landing Gear Axle Special Detailed Inspection
05-15-MB
A321005
Task 250
53-10-00-
53-10-01
Fuselage Engine Mount Fittings Special Detailed Inspection
05-15-13
A531008
Task 251
53-10-00-
53-20-01
Cargo and Passenger Door Doublers Special Detailed Inspection
05-15-15
A532003
Task 250
53-20-07-
53-20-07
Fuselage to Strut Attach Fitting Lugs (Nominal Standard Bolt Size) (Typical Inspection Compliance) Special Detailed Inspection
05-15-MH
A531005
Task 251
53-20-07-
53-20-07
Fuselage to Strut Attach Fitting Lugs (Nominal Standard Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (Part Number S3461-74)
05-15-MG
A531006
Task 252
53-20-07-
53-20-07
Fuselage to Strut Attach Fitting Lugs (Oversize 1/64 - Inch Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (Part Number S3461-159)
05-15-MI
A531010
Task 253
53-20-07-
53-20-07
Fuselage to Strut Attach Fitting Lugs (Oversize 1/32- Inch Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (Part Number S3461-160)
05-15-MJ
A531011
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MODEL 208 MAINTENANCE MANUAL Table 1. Task Inspection to SID (continued) DETAILS FOUND IN TASK
SUPPLEMENTAL INSPECTION NUMBER
INSPECTION COMPLIANCE TITLE
INSPECTION DOCUMENT
ITEM CODE
Task 253
53-10-00-
53-20-03
Lower Forward Carry-Thru Bulkhead Special Detailed Inspection
05-15-16
A532004
Task 254
53-10-00-
53-20-04
Main Landing Gear Fitting Special Detailed Inspection
05-15-19
A532005
Task 255
53-10-00-
53-20-05
Main Landing Gear Attach Fittings and Aft Carry-Thru Bulkhead
05-15-16
A532006
Task 252
53-10-00-
53-20-02
Fuselage to Wing Attach Fitting Lugs Special Detailed Inspection
05-15-MF
A532007
Task 223
53-10-00-
53-20-11
Firewall Brace and Doubler Assemblies Detailed Inspection
05-15-13
A532008
Task 220
53-10-00-
53-20-08
Carry-Through Root Rib Detailed Inspection
05-15-13
A532009
Task 221
53-10-00-
53-20-09
Crew Door Frames Detailed Inspection
05-15-13
A532011
Task 222
53-10-00-
53-20-10
Passenger and Cargo Door Frames Detailed Inspection
05-15-13
A532012
Task 220
53-25-00-
53-10-07
Seat Rails and Attachment Structure Detailed Inspection
05-15-15
A531009
Task 221
53-25-00-
53-20-12
Bulkheads and Stiffeners Below the Seat Rail Attachments at FS 143.00 and FS 158.00 Detailed Inspection
05-15-15
A532013
Task 224
53-10-00-
53-20-13
Stringers at Intersections with Forward and Aft Carry - Thru Bulkheads Detailed Inspection
05-15-16
A532014
Task 225
53-10-00-
53-20-14
Fuselage Skin Doubler at Main Landing Gear Cutout Detailed Inspection
05-15-14
A532015
Task 257
53-10-00-
53-50-01
Fuselage to Horizontal Stabilizer Attach Fittings Special Detailed Inspection
05-15-13
A535001
Task 258
53-10-00-
53-50-02
Vertical Stabilizer Attach Points Special Detailed Inspection (Typical Inspection Compliance)
05-15-13
A535002
Task 259
53-10-00-
53-50-02
Vertical Stabilizer Attach Points Special Detailed Inspection (Severe Inspection Compliance)
05-15-17
A535003
Task 250
55-10-00-
55-10-01
Horizontal Stabilizer Forward and Aft Attach Points Special Detailed Inspection
05-15-13
A551003
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MODEL 208 MAINTENANCE MANUAL Table 1. Task Inspection to SID (continued) DETAILS FOUND IN TASK
SUPPLEMENTAL INSPECTION NUMBER
INSPECTION COMPLIANCE TITLE
INSPECTION DOCUMENT
ITEM CODE
Task 251
55-10-00-
55-10-02
Horizontal Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance)
05-15-13
A553004
Task 252
55-10-00-
55-10-02
Horizontal Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance)
05-15-18
A551005
Task 256
53-10-00-
53-20-06
Fuselage to Wing Carry-Thru Attach Fitting and Bulkhead Special Detailed Inspection
05-15-16
A532016
Task 250
55-30-00-
55-30-01
Vertical Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance)
05-15-13
A553001
Task 251
55-30-00-
55-30-01
Vertical Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance)
05-15-17
A553002
Task 220
56-00-01-
56-30-01
Windshield and Attachment Structure Detailed Inspection
05-15-13
A564002
Task 250
57-10-00-
57-20-01
Wing to Carry - Thru Front Spar Attachment Fittings Special Detailed Inspection
05-15-MF
A570013
Task 251
57-10-00-
57-20-01
Wing to Carry - Thru Rear Spar Attachment Fittings Special Detailed Inspection
05-15-MF
A570014
Task 252
57-10-00-
57-20-02
Front Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection
05-15-MF
A570010
Task 253
57-10-00-
57-20-03
Rear Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection
05-15-MF
A570011
Task 254
57-10-00-
57-50-01
Center Flap Track and Inboard Flap Track Special Detailed Inspection
05-15-MD
A575002
Task 255
57-10-00-
57-50-01
Outboard Flap Track Special Detailed Inspection
05-15-MD
A575003
Task 250
57-10-01-
57-60-01
Wing Strut Fittings Special Detailed Inspection (Typical Inspection Compliance)
05-15-MA
A570008
Task 251
57-10-01-
57-60-01
Wing Strut Fittings Special Detailed Inspection (Severe Inspection Compliance)
05-15-ME
A570009
Task 252
57-10-01-
57-60-02
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Typical Inspection Compliance)
05-15-MF
A570012
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MODEL 208 MAINTENANCE MANUAL Table 1. Task Inspection to SID (continued) DETAILS FOUND IN TASK
SUPPLEMENTAL INSPECTION NUMBER
INSPECTION COMPLIANCE TITLE
INSPECTION DOCUMENT
ITEM CODE
Task 253
57-10-01-
57-60-02
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Severe Inspection Compliance) (Part Number S3461-77)
05-15-MF
A570015
Task 254
57-10-01-
57-60-02
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/64 Inch Oversize Bolt Size) (Severe Inspection Compliance) (Part Number S3461-163)
05-15-MK
A570016
Task 255
57-10-01-
57-60-02
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/32 Inch Oversize Bolt Size) (Severe Inspection Compliance) (Part Number S3461-164)
05-15-ML
A570017
Task 240
71-20-00-
71-20-01
Engine Truss and Ring Assembly Special Detailed Inspection
05-15-13
A712003
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MODEL 208 MAINTENANCE MANUAL SCHEDULED INSPECTION PROGRAM 1.
Purpose and Use A.
As the person who has control of the airworthiness of the airplane, the owner or operator must use only qualiÞed personnel to do maintenance on the airplane.
B.
Title 14 of the Code of Federal Regulations (1) The Cessna Scheduled Inspection Program will help the owner or operator to do the intent of Title 14 CFR Part 91.409 paragraph (a), (b) and (c). (a) Except as provided in paragraph (c) of Title 14 CFR Part 91.409, no person may operate an aircraft unless, within the preceding 12 calendar months, it has had: An annual inspection in accordance with Title 14 CFR Part 43 of this chapter and has 1 been approved for return to service by a person authorized by Title 14 CFR Part 43.7 of this chapter; or An inspection for the issuance of an airworthiness certiÞcate in accordance with Title 2 14 CFR Part 21 of this chapter. NOTE:
(b)
(c)
2.
No inspection performed under paragraph (b) of Title 14 CFR Part 91.409 may be substituted for any inspection required by this paragraph unless it is performed by a person authorized to perform annual inspections and is entered as an “Annual” inspection in the required maintenance records.
Except as provided in paragraph (c) of Title 14 CFR Part 91.409, no person may operate an aircraft carrying any person (other than a crew member) for hire, and no person may give ßight instruction for hire in an aircraft which that person provides, unless within the preceding 100 hours of time in service the aircraft has received an annual or 100-hour inspection and been approved for return to service in accordance with Title 14 CFR Part 43 of this chapter or has received an inspection for the issuance of an airworthiness certiÞcate in accordance with Title 14 CFR Part 21 of this chapter. The 100-hour limitation may be exceeded by not more than 10 hours while en route to reach a place where the inspection can be done. The excess time used to reach a place where the inspection can be done must be included in computing the next 100 hours of time in service. Paragraphs (a) and (b) of Title 14 CFR Part 91.409 do not apply to: An aircraft that carries a special ßight permit, a current experimental certiÞcate, or a 1 light-sport or provisional airworthiness certiÞcate. 2 An aircraft inspected in accordance with an approved aircraft inspection program under Title 14 CFR Part 125 or 135 of this chapter and so identiÞed by the registration number in the operations speciÞcations of the certiÞcate holder having the approved inspection program. An aircraft subject to the requirements of paragraph (d) or (e) of Title 14 CFR Part 3 91.409.
Construction A.
The Scheduled inspection program includes all of the inspections for the Model 208, 208 Cargomaster, 208B, 208B Super Cargomaster and 208B passenger airplanes and is recommended by Cessna Aircraft Company.
B.
A Scheduled maintenance inspection program schedule applicable for your airplane must be selected as early as possible. The inspection program timing begins from the Date of Airworthiness. Inspections can begin soon after delivery.
C.
Operators can use their own inspection method, but it will not be given in this manual. However, these inspection programs must be monitored by the operator.
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MODEL 208 MAINTENANCE MANUAL D.
The inspections are done at subsequent intervals that are related to hours, calendar months, years, or in accordance with (IAW) the manufacturer's instructions NOTE:
3.
E.
Intervals for some inspections are given by the manufacturer of the component that is examined. The applicable manufacturer's manual gives the data necessary to Þnd those intervals. Refer to the Introduction, List of Publications.
F.
The inspection program is divided into inspections documents for each of the required inspection intervals.
Inspection Time Limitations A.
Any inspection time limit found in Chapter 5 and also required in Chapter 4 shall not be extended. Any inspection required by the Code of Federal Regulations (CFR) shall not be extended.
B.
Inspection item time limits can be extended for maintenance scheduling purposes only as provided below. NOTE:
(1) (2) (3) (4) C. 4.
5.
A calendar month starts on the Þrst day of the month. You must complete the inspections on or before the last day of the month for their related calendar month interval.
The Annual inspection is due on the last day of the calendar month in which it was previously completed twelve months before and cannot be extended. The 100-hour inspection can only be extended up to 10 hours while en route to reach a place where the inspection can be done. Refer to Title 14 CFR Part 91.409 paragraph (a) and (b).
Inspection items that use ßight hours with calendar limits can be extended up to a maximum of 30 ßight hours or two calendar months beyond the time when the inspection is due.. Inspection items that use ßight hour limits can be extended up to a maximum of 30 ßight hours beyond the due time. Inspection items that use calendar limits can be extended up to two calendar months beyond the due date. The part of the allowable extension that is used does not need to be deducted from the subsequent due point.
Any inspection program based upon the intervals of the items in this chapter, or more frequent intervals, is acceptable.
Maintenance Manual - Compact Disc - Read Only Memory (CD-ROM) A.
The maintenance manual on Compact Disc - Read Only Memory (CD-ROM) contains a search feature. With this, any item code can be found, quickly and entirely. Along with the item code, task information about the inspection and zone is given. A task number is supplied with the items which give special inspection instructions. When a task number is clicked on, the link will show the task information from the appropriate section of the manual.
B.
A print task function selection is available when that button is set with the cursor. The task text and any linked documents, such as illustrations, are printed together as one document.
C.
The end of each task is clearly shown with the message "End of Task".
Inspection Guidelines. A.
The Inspection Documents are to be used as a recommended inspection outline. Detailed information of systems and components in the airplane will be found in various chapters of this Maintenance Manual and the applicable vendor publications. It is recommended that you refer to the applicable portion of this manual for service instructions and installation instructions and to the vendor’s data or publications speciÞcations for torque values, clearances, settings, tolerances, and other requirements.
B.
DeÞnitions and procedures (1) On Condition is deÞned as follows: The necessary inspections and/or checks to make sure that a failure of the component will not occur before the next scheduled inspection.
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MODEL 208 MAINTENANCE MANUAL (2) (3)
(4) (5) (6) (7)
Condition is deÞned as follows: Inspect for, but not limited to, cleanliness, cracks, deformation, corrosion, wear, and loose or missing fasteners. MOVABLE PARTS: Inspect for lubrication, servicing, make sure the part is tight, binding, more than normal wear, safetying, correct operation, correct adjustment, correct travel, cracked Þttings, condition of the hinges, defective bearings, cleanliness, corrosion, deformation, sealing, and tension. FLUID LINES AND HOSES: Inspect for leaks, cracks, bulging, collapsed, twisted lines/hoses, dents, kinks, chaÞng, proper radius, security, discoloration, bleaching, deterioration, and proper routing; rubber hoses for hardness or ßexibility and metal lines for corrosion. METAL PARTS: Make sure the installation is correct and tight, and that there are no cracks and/or metal distortion. WIRING: Inspect for correct and tight installation, chaÞng, burning, arcing, defective insulation, loose or broken terminals, heat deterioration, and corroded terminals. STRUCTURAL BOLTS: Inspect for correct torque. Obey the applicable torque values. Refer to Bolt Torque Data during installation or when visual inspection shows the need for a torque check. NOTE:
(8) (9)
C.
The torque values that are listed are not to be used for the measurement of tightness of installed parts while they are in service.
FILTERS, SCREENS, AND FLUIDS: Make sure the Þlters and screens are replaced at the required interval. Make sure you use clean ßuids and that the Þlters or screens are kept clean. Make sure the system checks (operation or function) that need electrical power are done with 28.5 Volts, +0.25 or -0.25 Volts, bus voltage. This will make sure all components operate at their operational voltage.
Airplane Þle. (1) Miscellaneous data, information, and licenses are a part of the airplane Þle. Make sure that the documents listed in this section are up-to-date and obey the current 14 CFR. If other documents and data are needed for other nations, owners of exported airplanes should talk with their aviation ofÞcials to get their individual requirements. (a) To be displayed in the airplane at all times: Standard Airworthiness CertiÞcate (FAA Form 8100-2). 1 2 Aircraft Registration CertiÞcate (FAA Form 8050-3). Aircraft Radio Station License (Federal Communication Commission Form 556 if 3 transmitter is installed). 4 Radio Telephone Station License (Federal Communication Commission Form 409 if Flitefone Radio Telephone is installed). (b) To be in the airplane at all times: 1 Weight and Balance Data Sheets and associated papers (all copies of the Repair and Alteration Form, FAA Form 337, are applicable). Equipment List. 2 3 Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. (c) To be made available upon request: Airplane, Engine, and Propeller Logbooks. 1
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MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 0A
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 0A gives a list of item(s), which are completed during the Annual inspection interval.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A052001
Aircraft Records Check Task 5-20-01-280
ALL
A110001
Interior and Exterior Placard and Decal Detailed Inspection Task 11-00-00-220
ALL
D121001
Brake System Servicing Task 12-10-01-610
121
D121003
Shimmy Damper Servicing Task 12-10-01-611
710
C122101
Landing Gear Lubrication Task 12-21-03-640
700
C122103
Hartzell Propeller Lubrication Task 12-21-04-640
110
B212401
Avionics Cooling Fan Operational Check Task 2124-00-710
211 212
B255201
Cargo Pod Drains Operational Check Task 25-5200-710
901 902 903 904 905 906
C270001
Flight Controls Lubrication Task 27-00-00-640
215 226 374 525 625
MECH INSP REMARKS
216 373 503 603
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MODEL 208 MAINTENANCE MANUAL
ITEM CODE NUMBER
TASK
ZONE
B273101
Stall Warning System Operational Check Task 2731-00-710
211 212 503
C275001
Flap Tracks and Rollers Lubrication Task 27-50-00640
525 527 625 627
A281001
Fuel Filler Assembly Detailed Inspection Task 28-1001-220
521 621
B284101
Fuel Reservoir Warning System Operational Check Task 28-41-00-710
ENG
B322001
Shimmy Damper Functional Check Task 32-20-02720
710
B761003
Emergency Power Lever Annunciator Light (EPL) Operational Check Task 76-10-01-710
AUX
MECH INSP REMARKS
*** End of Inspection Document 0A Inspection Items ***
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MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 01
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 01 gives a list of item(s), which are completed at every 12 calendar months.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A251000
Smoke Goggle General Visual Inspection Task 2510-00-210
801 802
B262001
Portable Fire Extinguisher Functional Check (Weight Check) Task 26-20-00-720
215 216 251 252
B272003
Rudder System Functional Check Installation) Task 27-20-00-721
211 213 217 233 253 257 311 320
B301102
Inboard TKS Wing Panel Pressurization Functional Check Task 30-11-00-721
501, 601, AUX
A321001
Main Landing Gear Detailed Inspection Task 32-1000-220
721 722
A324001
Brakes Detailed Inspection Task 32-40-00-220
721 722
A324005
Main Landing Gear Wheels and Tires Detailed Inspection Task 32-40-00-222
721 722
(Float
Kit
MECH INSP REMARKS
212 214 218 234 254 258 312 341
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MODEL 208 MAINTENANCE MANUAL
ITEM CODE NUMBER
TASK
ZONE
A324009
Nose Landing Gear Wheel and Tire Detailed Inspection Task 32-40-00-224
710
B342101
Magnetic Compass Functional Check Task 34-2100-720
ENG
B350101
Oxygen System Operational Check Task 35-01-00710
231 232 251 252 255 256 311 312 801 802
A714101
Engine Wash Ring, Air Plenum, and Thermocouple (T1) Detailed Inspection Task 71-41-00-220
130
MECH INSP REMARKS
*** End of Inspection Document 01 Inspection Items ***
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MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 02
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 02 gives a list of item(s), which are completed at every 24 calendar months.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B560001
Functional Check of the Windshield Task 56-00-01720
240
A710001
Engine Compartment Zonal Inspection Task 71-0001-210
130
MECH INSP REMARKS
*** End of Inspection Document 02 Inspection Items ***
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MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 03
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 03 gives a list of item(s), which are completed at every 48 calendar months.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A281003
Fuel Storage System Detailed Inspection Task 2810-01-221
521 621
A321003
Center-Spring and Main Gear-Spring Interface Area Special Detailed (Corrosion Inspection and Repair) Task 32-10-00-221
721 722
A531003
Internal Cockpit Zonal Inspection Task 53-10-00-211
211 213 215 217 231 233
A531007
Internal Tail Cone Zonal Inspection Task 53-10-00213
311 312 320 330
A571001
Wing Zonal Inspection Task 57-10-00-210
500 600
MECH INSP REMARKS
212 214 216 218 232 234
*** End of Inspection Document 03 Inspection Items ***
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MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 04
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 04 gives a list of item(s), which are completed at every 72 calendar months.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A255201
Cargo Pod Zonal Inspection Task 25-52-00-210
901 902 903 904 905 906
B262005
Portable Fire Extinguisher Restoration (Internal Inspection) Task 26-20-00-290
215 216 251 252
A531001
External Fuselage Zonal Inspection Task 53-10-00210
ALL
A531004
Internal Cabin Zonal Inspection Task 53-10-00-212
251 252 253 254 255 256 257 258 311 312
A531013
Empennage and Horizontal Inspection Task 53-10-00-214
340 341 373 374
Stabilizer
Zonal
MECH INSP REMARKS
*** End of Inspection Document 04 Inspection Items ***
5-15-04 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 05
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 05 gives a list of item(s), which are completed at every 144 calendar months.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B262003
Portable Fire Extinguisher Restoration (Hydrostatic Test) Task 26-20-00-780
215 216 251 252
MECH INSP REMARKS
*** End of Inspection Document 05 Inspection Items ***
5-15-05 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 06
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 06 gives a list of item(s), which are completed at every 200 Hours or 12 calendar months, whichever occurs Þrst.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B236001
Static Discharge System Functional Check Task 2360-00-720
343 375 376 571 671
B243401
Marathon Ni-Cad Battery Functional (Capacity Check) Task 24-34-00-720
122
A255101
Cargo Nets Detailed Inspection Task 25-51-00-220
251 252 255 256 257 258
B281001
Fuel Vent Line Float Valve Operational Check Task 28-10-03-710
575 675
B301003
Bleed Air Pressure Regulator Functional Check Task 30-10-00-720
122 AUX
B341101
Pitot Tube Heaters Operational Check Task 34-1100-710
AUX
Check
MECH INSP REMARKS
*** End of Inspection Document 06 Inspection Items ***
5-15-06 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 07
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 07 gives a list of item(s), which are completed at every 400 Hours or 12 calendar months, whichever occurs Þrst.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A243601
Standby Alternator Detailed Inspection Task 24-3600-220
121
B251001
Inertia Reel Operational Check Task 25-10-00-710
221 232
A261001
Engine Fire Detection System General Visual Inspection Task 26-10-00-210
121 122
C271001
Aileron Trim System Lubrication Task 27-10-02-640
211 217 233 253 251 551 651
B271005
Aileron Trim Tab (Free Play) Functional Check Task 27-10-02-720
551 571 651 671
B273003
Elevator Trim Tab (Free Play) Functional Check Task 27-30-02-720
371 372 375 376
B282103
Firewall Fuel Shutoff Valve Control Operational Check Task 28-21-00-711
213 214 220
MECH INSP REMARKS
212 218 234 254 252 571 671
5-15-07 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
ITEM CODE NUMBER
TASK
ZONE
C282301
Wing Shutoff Valve Linkage Lubrication Task 28-2300-640
231 232 511 611
B301001
Bleed Air Pressure Regulator Functional Check Task 30-10-00-720
122 AUX
B301101
TKS Anti-Ice System Functional Check Task 30-1100-720
AUX
B304001
Windshield Anti-Ice System Operational Check Task 30-40-00-710
AUX
B611001
Hartzell Propeller Functional Check Task 61-10-00720
110
B611101
McCauley Propeller Functional Check Task 61-1100-720
110
A712001
Engine Mounts and Firewall Detailed Inspection Task 71-20-00-220
130
A716001
Inertial Air Separator Detailed Inspection Task 7160-00-220
130
B761001
Engine Controls Functional Check Task 76-10-00720
130 211 212 ENG
A801001
Starter-Generator (Part Number 23081 Series only) Detailed Inspection Task 80-10-00-220
130
MECH INSP REMARKS
*** End of Inspection Document 07 Inspection Items ***
5-15-07 © Cessna Aircraft Company
Page 2 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 08
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 08 gives a list of item(s), which are completed at every 400 Hours or 24 calendar months, whichever occurs Þrst.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B215001
Compressor Drive Belt Functional Check Task 2150-00-720
121 122
A322001
Nose Landing Gear Detailed Inspection Task 32-2000-220
710
MECH INSP REMARKS
*** End of Inspection Document 08 Inspection Items ***
5-15-08 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 09
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 09 gives a list of item(s), which are completed at every 800 Hours or 12 calendar months, whichever occurs Þrst.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B221201
Garmin Autopilot (GFC 700) Functional Check Task 22-12-00-720
226 232
B243201
Gill Flooded Lead-Acid Battery Functional Check (Capacity Check) Task 24-32-00-720
122
B243301
Concord Sealed Lead Acid Battery Functional Check (Capacity Check) Task 24-33-00-720
122
A275001
Flap Actuator Mount Bracket Detailed Inspection Task 27-50-00-220
231 232
A275003
Flap Bellcrank Detailed Inspection Task 27-50-00221
251 252 511 611 525 625
MECH INSP REMARKS
*** End of Inspection Document 09 Inspection Items ***
5-15-09 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 10
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 10 gives a list of item(s), which are completed at every 800 Hours or 24 calendar months, whichever occurs Þrst.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
C221201
Autopilot Servos Lubrication Task 22-12-00-640
226 232
A245001
Power Distribution Boxes Detailed Inspection Task 24-50-00-220
121 122
A251001
Crew Seats Detailed Inspection Task 25-10-00-220
231 232
A251003
Passenger Seats Detailed Inspection Task 25-2100-220
231 232
B271001
Spoiler System Functional Check Task 27-10-00720
211 217 233 253 251 503 603
C271003
Aileron Trim Tab Actuator (2660044-1) Lubrication Task 27-10-02-641
551 571 651 671
C273001
Elevator Trim Tab Actuator (2660017-1) Lubrication Task 27-30-02-640
371 372 375 376
MECH INSP REMARKS
212 218 234 254 252 525 625
5-15-10 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
ITEM CODE NUMBER
TASK
ZONE
B284103
Fuel Quantity and Low Fuel Warning Systems Functional Check Task 28-41-00-720
AUX
B324001
Brakes Operational Check Task 32-40-00-710
ENG
B332001
Passenger/Cargo Compartment Operational Check Task 33-20-00-710
AUX
A520001
Crew Doors Detailed Inspection Task 52-00-00-220
801 802
A520003
Passenger/Cargo Doors and Door Frames Detailed Inspection Task 52-00-00-221
255 256 257 258 803 804
A781001
Primary and Secondary Exhaust Duct General Visual Inspection Task 78-10-00-211
130
Lighting
MECH INSP REMARKS
*** End of Inspection Document 10 Inspection Items ***
5-15-10 © Cessna Aircraft Company
Page 2 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 11
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 11 gives a list of item(s), which are completed at every 1600 Hours or 24 calendar months, whichever occurs Þrst.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B271003
Aileron System Functional Check Task 27-10-00721
211 217 233 253 251 503 603
212 218 234 254 252 525 625
B272001
Rudder System Functional Check (Standard Rudder Installation) Task 27-20-00-720
211 213 217 233 253 257 311 320
212 214 218 234 254 258 312 341
C272001
Rudder Bar Bearings and Rudder Pedals Lubrication Task 27-20-00-640
211 212 213 214
MECH INSP REMARKS
5-15-11 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
ITEM CODE NUMBER
TASK
ZONE
B273001
Elevator System Functional Check Task 27-30-00720
211 213 217 233 253 257 311 320 374 376
B275001
Flap System Functional Check Task 27-50-00-720
'251 252 511 611 525 625
MECH INSP REMARKS
212 214 218 234 254 258 312 373 375
*** End of Inspection Document 11 Inspection Items ***
5-15-11 © Cessna Aircraft Company
Page 2 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 12
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 12 gives a list of item(s), which are completed at every 1600 Hours or 60 calendar months, whichever occurs Þrst.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
C271005
Aileron Trim Tab Actuator (2661615-1, 2661615- 9, or 2661615-10) Lubrication Task 27-10-02-642
551 571 651 671
C273003
Elevator Trim Tab Actuator (2661215-1 2661215-9) Lubrication Task 27-30-02-641
371 372 375 376
and
MECH INSP REMARKS
*** End of Inspection Document 12 Inspection Items ***
5-15-12 © Cessna Aircraft Company
Page 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 13
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 13 gives a list of item(s), which are completed at the Þrst 20,000 hours and every 5000 hours thereafter.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A531008
Fuselage Engine Mount Fittings Special Detailed Inspection (SID 53-10-01) Task 53-10-00-250
121 122 130
A532008
Firewall Brace and Doubler Assemblies Detailed Inspection (SID 53-20-11) Task 53-10-00-223
121 122 130
A532009
Carry-Through Root Rib Detailed Inspection (SID 53-20-08) Task 53-10-00-220
251 252 500 600
A532011
Crew Door Frames Detailed Inspection (SID 53-2009) Task 53-10-00-221
231 232 233 234 801 802
A532012
Passenger and Cargo Door Frames Detailed Inspection (SID 53-20-10) Task 53-10-00-222
255 256 257 258 803 804
A535001
Fuselage to Horizontal Stabilizer Attach Fittings Special Detailed Inspection (SID 53-50-01) Task 53-10-00-257
320 373 374
A535002
Vertical Stabilizer Attach Points Special Detailed Inspection (Typical Inspection Compliance) (SID 53-50-02) Task 53-10-00-258
311 312 320 341
MECH INSP REMARKS
5-15-13 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
ITEM CODE NUMBER
TASK
ZONE
A551003
Horizontal Stabilizer Forward and Aft Attach Points Special Detailed Inspection (SID 55-10-01) Task 5510-00-250
373 374
A553001
Vertical Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) (SID 55-30-01) Task 55-30-00-250
320 341
A553004
Horizontal Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) (SID 55-10-02) Task 55-10-00-251
373 374
A564002
Windshield and Attachment Structure Detailed Inspection (SID 56-30-01) Task 56-00-01-220
240
A712003
Engine Truss and Ring Assembly Special Detailed Inspection (SID 71-20-01) Task 71-20-00-240
130
MECH INSP REMARKS
*** End of Inspection Document 13 Inspection Items ***
5-15-13 © Cessna Aircraft Company
Page 2 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 14
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 14 gives a list of item(s), which are completed at the Þrst 5000 hours and every 2500 hours thereafter.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A532015
Fuselage Skin Doubler at Main Landing Gear Cutout Detailed Inspection (SID 53-20-14) Task 53-10-00225
253 254
MECH INSP REMARKS
*** End of Inspection Document 14 Inspection Items ***
5-15-14 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 15
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 15 gives a list of item(s), which are completed at the Þrst 7500 hours and every 2500 hours thereafter.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A531009
Seat Rails and Attachment Structure Detailed Inspection (SID 53-10-07) Task 53-25-00-220
231 233 251 253 255 257
A532003
Cargo and Passenger Door Doublers Special Detailed Inspection (SID 53-20-01) Task 53-10-00251
255 256 257 258 803 804
A532013
Bulkheads and Stiffeners Below the Seat Rail Attachments at FS 143.00 and FS 158.00 Detailed Inspection (SID 53-20-12) Task 53-25-00-221
233 234
MECH INSP REMARKS
232 234 252 254 256 258
*** End of Inspection Document 15 Inspection Items ***
5-15-15 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 16
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 16 gives a list of item(s), which are completed at the Þrst 12,500 hours and every 2500 hours thereafter.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A532004
Lower Forward Carry-Thru Bulkhead Special Detailed Inspection (SID 53-20-03) Task 53-10-00253
253 254
A532006
Main Landing Gear Attach Fittings and Aft CarryThru Bulkhead (SID 53-20-05) Task 53-10-00-255
253 254
A532014
Stringers at Intersections with Forward and Aft Carry - Thru Bulkheads Detailed Inspection (SID 53-20-13) Task 53-10-00-224
251 252
A532016
Fuselage to Wing Carry-Thru Attach Fitting and Bulkhead Special Detailed Inspection (SID 53-20-06) Task 53-10-00-256
251 252
MECH INSP REMARKS
*** End of Inspection Document 16 Inspection Items ***
5-15-16 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 17
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 17 gives a list of item(s), which are completed at the Þrst 16,500 hours and every 5000 hours thereafter.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A535003
Vertical Stabilizer Attach Points Special Detailed Inspection (Severe Inspection Compliance) (SID 53-50-02) Task 53-10-00-259
311 312 320 341
A553002
Vertical Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) (SID 55-30-01) Task 55-30-00-251
320 341
MECH INSP REMARKS
*** End of Inspection Document 17 Inspection Items ***
5-15-17 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 18
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 18 gives a list of item(s), which are completed at the Þrst 17,500 hours and every 1000 hours thereafter.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A551005
Horizontal Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) (SID 55-10-02) Task 55-10-00-252
373 374
MECH INSP REMARKS
*** End of Inspection Document 18 Inspection Items ***
5-15-18 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 19
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 19 gives a list of item(s), which are completed at the Þrst 25,000 landings and every 5000 landings thereafter.
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A532005
Main Landing Gear Fitting Special Detailed Inspection (SID 53-20-04) Task 53-10-00-254
253 254
MECH INSP REMARKS
*** End of Inspection Document 19 Inspection Items ***
5-15-19 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 20
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 20 gives a list of required 14CFR 91.207 interval item(s), which are completed every 12 calendar months (No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B256001
ARTEX C406-2 Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-720
220 311 312 340
B256003
ARTEX ME406 Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-721
220 311 312 340
B256005
ARTEX C406-N Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-722
220 311 312 340
B256007
Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 Emergency Locator Transmitter (ELT) Functional Check Task 25-60-00-723
220 311 312 340
MECH INSP REMARKS
*** End of Inspection Document 20 Inspection Items ***
5-15-20 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 21
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 20 gives a list of required 14 CFR 91.411 interval item(s), which are completed every 24 calendar months (No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B341103
Pitot/Static System Functional Check Task 34-11-00720
AUX
MECH INSP REMARKS
*** End of Inspection Document 21 Inspection Items ***
5-15-21 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT 22
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document 20 gives a list of required 14 CFR 91.413 interval item(s), which are completed every 24 calendar months (No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
B345001
Transponder Functional Check Task 34-50-00-720
AUX
MECH INSP REMARKS
*** End of Inspection Document 22 Inspection Items ***
5-15-22 © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MA
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MA gives a list of item(s), which are completed at 10,000 hours and every 5000 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A570008
Wing Strut Fittings Special Detailed Inspection (Typical Inspection Compliance) (SID 57-60-01) Task 57-10-01-250
531 631
MECH INSP REMARKS
*** End of Inspection Document MA Inspection Items ***
5-15-MA © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MB
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MB gives a list of item(s), which are completed at 5000 landings and every 1000 landings thereafter, up to 10,000 landings. Replace at 10,000 landings (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A321005
Main Landing Gear Axle Special Detailed Inspection (SID 32-10-01) Task 32-10-00-240
721 722
MECH INSP REMARKS
*** End of Inspection Document MB Inspection Items ***
5-15-MB © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MD
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MD gives a list of item(s), which are completed at 15,000 landings and every 3000 landings thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A575002
Center Flap Track and Inboard Flap Track Special Detailed Inspection (SID 57-50-01) Task 57-10-00254
525 527 625 627
A575003
Outboard Flap Track Special Detailed Inspection (SID 57-50-01) Task 57-10-00-255
525 527 625 627
MECH INSP REMARKS
*** End of Inspection Document MD Inspection Items ***
5-15-MD © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT ME
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document ME gives a list of item(s), which are completed at 5000 hours and every 3600 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A570009
Wing Strut Fittings Special Detailed Inspection (Severe Inspection Compliance) (SID 57-60-01) Task 57-10-01-251
531 631
MECH INSP REMARKS
*** End of Inspection Document ME Inspection Items ***
5-15-ME © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MF
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MF gives a list of item(s), which are completed at 20,000 hours and every 5000 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A532007
Fuselage to Wing Attach Fitting Lugs Special Detailed Inspection (SID 53-20-02) Task 53-10-00252
251 252 501 511 525 601 611 625
A570010
Front Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection (SID 57-20-02) Task 57-10-00252
501 521 601 621
A570011
Rear Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection (SID 57-20-03) Task 57-10-00253
521 525 621 625
A570012
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Typical Inspection Compliance) (SID 57-60-02) Task 57-10-01-252
531 631
A570013
Wing to Carry - Thru Front Spar Attachment Fittings Special Detailed Inspection (SID 57-20-01) Task 5710-00-250
251 252
MECH INSP REMARKS
5-15-MF © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
ITEM CODE NUMBER
TASK
ZONE
A570014
Wing to Carry - Thru Rear Spar Attachment Fittings Special Detailed Inspection (SID 57-20-01) Task 5710-00-251
251 252
A570015
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Severe Inspection Compliance) (SID 57-60-02) Task 57-10-01-253
531 631
MECH INSP REMARKS
*** End of Inspection Document MF Inspection Items ***
5-15-MF © Cessna Aircraft Company
Page 2 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MG
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MG gives a list of item(s), which are completed at 5000 hours and every 1200 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A531006
Fuselage to Strut Attach Fitting Lugs (Nominal Standard Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 53-20-07-251
251 252 253 254
MECH INSP REMARKS
*** End of Inspection Document MG Inspection Items ***
5-15-MG © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MH
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MH gives a list of item(s), which are completed at 10,000 hours and every 2500 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A531005
Fuselage to Strut Attach Fitting Lugs (Nominal Standard Bolt Size) (Typical Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 53-20-07-250
251 252 253 254
MECH INSP REMARKS
*** End of Inspection Document MH Inspection Items ***
5-15-MH © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MI
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MI gives a list of item(s), which are completed at 5000 hours and every 500 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A531010
Fuselage to Strut Attach Fitting Lugs (Oversize 1/ 64 - Inch Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 5320-07-252
251 252 253 254
MECH INSP REMARKS
*** End of Inspection Document MI Inspection Items ***
5-15-MI © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MJ
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MJ gives a list of item(s), which are completed at 5000 hours and every 400 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A531011
Fuselage to Strut Attach Fitting Lugs (Oversize 1/ 32- Inch Bolt Size) (Severe Inspection Compliance) Special Detailed Inspection (SID 53-20-07) Task 5320-07-253
251 252 253 254
MECH INSP REMARKS
*** End of Inspection Document MJ Inspection Items ***
5-15-MJ © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT MK
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document MK gives a list of item(s), which are completed at 20,000 hours and every 4400 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A570016
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/64 Inch Oversize Bolt Size) (Severe Inspection Compliance) (SID 57-60-02) Task 57-10-01-254
531 631
MECH INSP REMARKS
*** End of Inspection Document MK Inspection Items ***
5-15-MK © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INSPECTION DOCUMENT ML
1.
2.
Date:
_______________
Registration Number:
_______________
Serial Number:
_______________
Total Time:
_______________
Description A.
Inspection Document ML gives a list of item(s), which are completed at 20,000 hours and every 3600 hours thereafter (Chapter 4 requirement - No grace period).
B.
Inspection items are given in the sequence of the zone in which the inspection is completed. A description of the inspection, as well as the Item Code Number are supplied for cross-reference to section 5-10-01. Frequently, tasks give more information about each inspection. These tasks are found in the individual chapters of this manual.
C.
The right portion of each page gives space for the mechanic's and inspector's initials and remarks. You can use copies of these pages as a checklist while you complete the tasks in this Inspection Document.
General Inspection Criteria A.
As you complete each of the inspection tasks in this Inspection Document, examine the adjacent area while access is available to Þnd conditions that need more maintenance.
B.
If it is necessary to replace a component or to make a change to a system while you complete a task, do the task again before the system or component is returned to service.
C.
Inspection Kits are available for some Inspection Documents. They supply consumable materials used to complete the inspection item(s) given for the interval. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Service Kit List to Þnd applicable part numbers.
ITEM CODE NUMBER
TASK
ZONE
A570017
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/32 Inch Oversize Bolt Size) (Severe Inspection Compliance) (SID 57-60-02) Task 57-10-01-255
531 631
MECH INSP REMARKS
*** End of Inspection Document ML Inspection Items ***
5-15-ML © Cessna Aircraft Company
Page 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL EXPANDED INSPECTION 1.
General A.
This section provides a more detailed description of the inspection items listed in the Inspection Time Limits section. It provides more information on what to inspect and how to inspect an item. The information in this section is arranged in chapter order.
Task 5-20-01-280 2.
Aircraft Records Check A.
General (1) This task gives the procedures to do a special inspection of the aircraft records for the airplane.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do An Aircraft Records Check Special Inspection. (1) Inspect aircraft records to make sure that all applicable Cessna Service Letters, Cessna Service Bulletins and Supplier Service Bulletins are complied with. (2) Inspect aircraft records to make sure that all applicable Airworthiness Directives and Federal Aviation regulations are complied with. (3) Inspect aircraft records to make sure that all logbook entries required by the Federal Aviation Regulations are complied with.
E.
Restore Access (1) None End of task
5-20-01 © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL UNSCHEDULED MAINTENANCE CHECKS 1.
2.
General A.
During operation, the airplane may be subjected to: (1) Hard landings. (2) Overspeed. (3) Severe air turbulence or severe maneuvers. (4) Towing with a large fuel unbalance or high drag/side loads due to ground handling. (5) Lightning Strike.
B.
When any of these conditions are reported by the ßight crew, a visual inspection of the airframe and speciÞc inspections of components and areas involved must be accomplished.
C.
The inspections are performed to determine and evaluate the extent of damage in local areas of visible damage, and to the structure and components adjacent to the area of damage.
D.
If foreign object damage is encountered (suspected or actual), a visual inspection of the airplane must be accomplished before airplane is returned to service.
Unscheduled Maintenance Checks DeÞned and Areas to be Inspected A.
Hard/Overweight Landings. (1) A hard landing is any landing made at what is believed to be an excessive sink rate. Closely related to hard landings, is an overweight landing, which is deÞned as landing the airplane at any gross weight which exceeds maximum gross landing weight as speciÞed in the FAA Approved Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. NOTE: (2)
B.
If the hard/overweight landing is combined with high drag/side loads, additional checks are required.
Hard or overweight landing check. (a) Landing gear. Main gear struts - Inspect for security of attachment and permanent set. 1 Main gear attachments and supporting structure - Inspect for security, loose or failed 2 fasteners, and any evidence of structural damage. Main gear spring - Inspect for gear spread or wing low. 3 Nose gear trunnion supports and attaching structure - Inspect for security, loose or 4 failed fasteners, and any evidence of structural damage. 5 Nose gear attachments and supporting structure - Inspect for security, loose or failed fasteners, and any evidence of structural damage. (b) Wings. 1 Wing surface and lift strut - Inspect for skin buckles, loose or failed fasteners, and security of attach Þttings and fuel leaks Trailing edge - Inspect for any deformation affecting normal ßap operation. 2
Overspeed. (1) Any time an airplane has exceeded one or both of the following: (a) Airplane overspeed exceeding placard speed limits of ßaps. (b) Airplane overspeed exceeding design speeds. (2) Airplanes equipped with an airspeed exceedance device capable of recording an airspeed exceedence with accompanying time duration: (a) For a recorded airspeed above 175 knots in smooth air, with duration greater than 5 seconds; or any airspeed above 181 knots, perform the speciÞed overspeed inspection. (3) Overspeed (airspeed) check. (a) Fuselage. Windshield and Windows - Inspect for buckling, dents, loose or failed fasteners, and 1 any evidence of structural damage. All hinged doors - Inspect hinges, hinge attach points, latches and attachments, and 2 skins for deformation and evidence of structural damage.
5-50-00 © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (b) (c)
(d)
Cowling. 1 Skins - Inspect for buckling, cracks, loose or failed fasteners, and indications of structural damage. Stabilizers. Stabilizers - Inspect skins, hinges and attachments, movable surfaces, mass balance 1 weights, and attaching structure for cracks, dents, buckling, loose or failed fasteners, and evidence of structural damage. Wings. Flaps - Inspect for skin buckling, cracks, loose or failed fasteners, attachments and 1 structure for damage. Fillets and fairings - Inspect for buckling, dents, cracks, and loose or failed fasteners. 2
C.
Severe air turbulence or severe maneuvers. (1) May be deÞned as atmospheric conditions producing violent buffeting of airplane. Severe maneuvers can be deÞned as any maneuvers exceeding Pilot's Operating Handbook and FAA Approved Airplane Flight Manual limits. (2) Severe turbulence and/or maneuvers checks. (a) Stabilizers. 1 Horizontal stabilizer hinge Þttings, actuator Þttings and stabilizer center section Inspect for security, loose or failed fasteners, and any evidence of structural damage. Vertical stabilizer - Inspect for evidence of structural damage, skin buckles and 2 security at primary attachments in tailcone, loose or failed fasteners, damage to hinges and actuator Þttings. 3 Elevator and rudder balance weight supporting structure - Inspect for security, loose or failed fasteners, and evidence of structural damage. (b) Wing. Wing to body strut Þttings and supporting structure - Inspect for security, loose or 1 failed fasteners, and evidence of structural damage. 2 Trailing Edge - Inspect for any deformation affecting normal operation of ßap and aileron.
D.
Lightning strike. (1) If ßown through an electrically stressed region of the atmosphere where electrical discharges are transferred from cloud to cloud and from cloud to earth, the airplane may become a part of this discharge path. During a lightning strike, the current enters the airplane at one point and exits at another, usually at opposite extremities. The wing tips, nose and tail sections are the areas where damage is most likely to occur. Burning and/or eroding of small surface areas of the skin and structure may be detected during inspection. In most cases, the damage is obvious. In some cases, however, hidden damage may result. The purpose of the lightning strike inspection is to locate any damage that may have occurred to the airplane before returning it to service. (2) Lightning strike check. (a) As the following checks are performed, complete Lightning Strike/Static Discharge Incident Reporting Form. Completed form must be mailed to Cessna Contract Services, P.O. Box 7706, Wichita, KS 67277 Attn: Manager Contract Services. (b) Communications. Antennas - Inspect all antennas for evidence of burning or eroding. If damage is 1 noted, perform functional check of affected system. (c) Navigation. Radar reßector, feed horn, motor box assembly and mounting structure. Inspect for 1 damage. If damage is noted, perform a bench check of system. If superÞcial pitting or burning of mount structure only is noted, perform a functional check of radar system. Glideslope antenna - Inspect for burning and pitting. If damage is noted, perform a 2 functional check of glideslope system. Compass - Compass should be considered serviceable if the corrected heading is 3 within plus or minus 10 degrees of heading indicated by the remote compass system. If remote compass is not within tolerance, remove, repair or replace. (d) Fuselage. Skin - Inspect surface of fuselage skin for evidence of damage. 1 2 Tailcone - Inspect tailcone and static dischargers for damage.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (e) (f)
(g)
(h) E.
Stabilizers. 1 Inspect surfaces of stabilizers for evidence of damage. Wings. 1 Skins - Inspect for evidence of burning and eroding. Wing tips - Inspect for evidence of burning and pitting. 2 3 Flight surfaces and hinging mechanisms - Inspect for burning and pitting. 4 Radome - Inspect for evidence of burning or eroding. Propeller. Propeller - Inspect for evidence of burns or arcing on the blades or hub. Remove 1 from service and have inspected at an authorized repair facility if there are signs of damage present. Powerplant Refer to the engine manufacturer's maintenance manual, unscheduled inspection. 1
Foreign object damage. (1) Damage to the airplane engine may be caused by the ingestion of slush, by a bird strike or by any other foreign object while operating the airplane on the ground or in normal ßight. Damage may also be caused by tools, bolts, nuts, washers, rivets, rags or pieces of safety-wire left in the engine inlet duct during maintenance operations. The purpose of the foreign object damage inspection is to locate any damage prior to repairing or returning the airplane to service. (2) Safety precautions should be taken to prevent foreign objects from coming in contact with the airplane during towing and at all times when airplane is not in service. To prevent dirt and foreign objects damage, the engines should be provided with suitable covers. When there is wind and dust conditions, the covers should be installed as soon as practicable following engine shutdown and engine cooling. (3) The aerodynamic cleanliness level (degree of surface smoothness), contributes to performance capabilities of the airplane. It is important that the high cleanliness level be maintained. (4) Contour and waviness distortion of the aerodynamic surface may be developed in the course of normal operation or by improper handling during maintenance operations. Doors and access panels are susceptible to waviness through rough handling. Care should be exercised in the handling of these items. (5) Foreign object damage check. (a) Landing gear. 1 Fairings - Inspect for dents, cracks, misalignment, and indication of structural damage. (b) Fuselage. 1 Skin - Inspect forward and belly areas for dents, punctures, cracks, and any evidence of damage. (c) Cowling. Skins - Inspect for dents, punctures, loose or failed fasteners, cracks or indications 1 of structural damage. (d) Stabilizers. Leading edge skins - Inspect for dents, cracks, scratches, and any evidence of 1 structural damage. Surface de-ice boots - Inspect for cuts, punctures, or tears. 2 (e) Windows. Windshield - Inspect for chipping, scratches, and cracks. 1 (f) Wings. Leading edge skins - Inspect for dents, cracks, punctures, and evidence of possible 1 structural damage. Radome - Inspect for dents, cracks, punctures, scratches, etc. 2 3 Surface de-ice boots - Inspect for cuts, punctures, or tears. (g) Engine. Air inlet section - Inspect for dents, cracks, scratches, punctures, blood and feathers. 1 2 Propeller - Inspect for nicked, bent, broken, cracked, or rubbing blades.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL F.
High drag/side loads due to ground handling. (1) High drag/side load condition shall be deÞned to exist whenever the airplane skids or overruns from the prepared surface onto an unprepared surface, or landings short of prepared surface, or makes a landing which involves the blowing of tires or skids on a runway to the extent that the safety of the airplane was in question. This covers takeoff and landings or unusual taxi conditions. (2) High drag/side loads due to ground handling check. (a) Landing gear. Main gear and fairings - Inspect for loose or failed fasteners, buckling, security, 1 cracks, and evidence of structural damage. Nose gear and fairing - Inspect for loose or failed fasteners, cracks, steering cables 2 tension, security, buckling, and evidence of structural damage. (b) Wings. 1 Wing to fuselage attach Þttings and attaching structure - Inspect for security, loose or failed fasteners, and evidence of structural failure.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
5-50-00 © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
5-50-00 © Cessna Aircraft Company
Page 6 Mar 1/2012
6
CHAPTER
DIMENSIONS AND AREAS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
6-00-00
Page 1
Aug 1/1995
6-00-01
Pages 1-18
Apr 1/2010
6-00-02
Pages 1-9
Aug 1/1995
6-00-03
Pages 1-4
Aug 1/1995
6-20-01
Pages 1-5
Mar 1/2012
6-20-02
Pages 1-28
Apr 1/2010
06-Title 06-List of Effective Pages 06-Record of Temporary Revisions 06-Table of Contents
06 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS DIMENSIONS AND AREAS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6-00-00 Page 1 6-00-00 Page 1 6-00-00 Page 1
AIRPLANE DIMENSIONS AND AREAS - DESCRIPTION AND OPERATION . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Dimensions and Areas - 208 and 208 Cargomaster . . . . . . . . . . . . . . . . . . . . . . . . . . . . Dimensions and Areas - 208B, 208B Super Cargomaster and 208B Passenger . . .
6-00-01 Page 1 6-00-01 Page 1 6-00-01 Page 1 6-00-01 Page 7
AIRPLANE STATIONS - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6-00-02 Page 1 6-00-02 Page 1 6-00-02 Page 1
MAJOR STRUCTURAL MEMBERS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6-00-03 Page 1 6-00-03 Page 1
AIRPLANE ZONING - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6-20-01 Page 1 6-20-01 Page 1 6-20-01 Page 1
ACCESS PLATES AND PANELS IDENTIFICATION - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Access Plates and Panel IdentiÞcation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6-20-02 6-20-02 6-20-02 6-20-02
06 - CONTENTS © Cessna Aircraft Company
Page 1 Page 1 Page 1 Page 1
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL DIMENSIONS AND AREAS - GENERAL 1.
2.
Scope A.
This chapter includes illustrations and statistical information concerning the Model 208 airplane. Provided are the overall airplane dimensions, surface areas, station locations, location of major structural members, access plates, panels, floorboards, fairings and airplane zoning.
B.
Dimensions and measurements are presented to aid the operator and/or maintenance personnel in ground handling the airplane and locating components.
Definition A.
Airplane Dimensions and Areas. (1) The section on airplane dimensions and areas provides airplane dimensions and identifies areas of the airplane.
B.
Airplane Stations. (1) The section on stations provides illustrations to identify reference points on the airplane along a three axis division.
C.
Major Structural Members. (1) This section provides illustrations of major structural members.
D.
Airplane Zoning. (1) This section provides illustrations of all airplane zones.
E.
Access Plates/Panels. (1) This section provides numbering of all plates and panels based on specific airplane zones.
6-00-00 © Cessna Aircraft Company
Page 1 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL AIRPLANE DIMENSIONS AND AREAS - DESCRIPTION AND OPERATION 1.
2.
General A.
This section identifies dimensions and areas of the airplane. Dimensions are selected for pertinent information of measurements to assist operators, maintenance personal and/or ground handling personnel. Dimensions for the 208 and 208 Cargomaster are presented separately from dimensions of the 208B, 208B Super Cargomaster and 208B Passenger. Refer to the respective charts below.
B.
Airplane areas are illustrated after Dimensions and Areas information. Refer to Figure 1.
C.
Airplane dimensions are illustrated following airplane areas. Refer to Figure 2.
Dimensions and Areas - 208 and 208 Cargomaster
AIRPLANE OVERALL Length (Overall)
37.58 Feet
Height (Maximum)
14.83 Feet
Wing Span (Overall)
52.16 Feet
Propeller Diameter (Hartzell)
100.0 Inches
Propeller Diameter (McCauley)
106.0 Inches
Propeller Ground Clearance (Nose tire inflated and nose strut fully extended 4.50 inches) (Hartzell)
14.52 Inches
Propeller Ground Clearance (Nose tire inflated and nose strut fully extended 4.50 inches) (McCauley)
11.53 Inches
Landing Gear Track Width (centerline to centerline MLG tire): Standard Tires
11.66 Feet
Optional Tires
11.66 Feet
Wheelbase (At static empty weight)
11.62 Feet
FUSELAGE Cabin Width (Maximum sidewall to sidewall)
62.00 Inches
Cabin Height (Floorboard to headliner)
51.00 Inches
Cabin Volume (Including rear baggage area)
341.4 Cubic Feet
Cargo Pod
83.7 Cubic Feet
WINGS Span
52.16 Feet
Area (Includes cabin top)
279.40 Square Feet
Chord Length and Root W.S. 35.00
77.995 Inches
Mean Aerodynamic Chord (W. S. 141.45)
66.398 Inches
Wing Station 308.00
48.108 Inches
Projected Tip (W.S. 310.00)
47.892 Inches
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MODEL 208 MAINTENANCE MANUAL
NACA Airfoil Designation and Root W.S. 35.00 Wing Station 308.00
NACA 23017.424 NACA 23012 +2.62 Degrees
Incidence and Root W.S. 35.00 Mean Aerodynamic Chord (at W.S. 141.45)
+1.707 Degrees
Wing Station 308.00
+0.608 Degrees
Leading Edge Fuselage Station and Root W.S. 35.00
151.73 Inches
Mean Aerodynamic Chord W.S. 141.45
157.57 Inches
Wing Station 308.00
166.67 Inches +3.135 Degrees
Sweep Angles and Leading Edge Front Spar at 20 Percent Chord
+1.888 Degrees
25 Percent Chord
+1.56 Degrees
50 Percent Chord
0.0 Degrees
Rear Spar at 57.5 Percent Chord
-0.47 Degrees
Trailing Edge
-3.135 Degrees
Wing Loading: Landplane at 7300 Pounds
26.1 Pounds per Square Foot
Landplane at 8000 Pounds
28.6 Pounds per Square Foot
Floatplane and Amphibian at 7600 Pounds
27.2 Pounds per Square Foot
FLAPS Type
Single-Slotted (Type B Nose)
Span (Percent of 310 Inches)
62.9 Percent
Span
16.25 Feet
Inboard Location
W.S. 33.50
Outboard Location
W.S. 228.50
Chord - Projected (Inboard 22.666 Inches)
29 Percent Wing Chord
Chord - Projected (Outboard 16.476 Inches)
29 Percent Wing Chord
Front Spar Location (Percent Flap Chord)
20 Percent
Rear Spar Location (Percent Flap Chord)
65 Percent
Tracks Inboard
W.S. 53.158
Center
W.S. 126.658
Outboard
W.S. 214.456
Area (Both)
52.79 Square Foot
AILERONS
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Type and 1/4 Chord
Round Nose with Aft Set Hinge Line
Span (Percent of 310 Inches Measured at Hinge Line)
25.387 Percent
Span (Measured at Hinge Line)
6.558 Feet
Inboard Location (Hinge line at ends of aileron)
W.S. 228.924
Outboard Location (Hinge line at ends of aileron)
W.S. 307.25
Hinge Line Location (Percent Wing Chord): Outboard W.S. 230.063
77.51 Percent
Inboard W.S. 304.553
78.08 Percent
Chord Length (Percent Wing Chord): Inboard Hinge Point
29.27 Percent
Outboard Hinge Point
29.95 Percent
Aft of Hinge Line (Includes Trim Tab)
6.34 Square Feet
Trim Tab
1.10 Square Feet
Forward of Hinge Line
1.99 Square Feet
Aileron
8.33 Square Feet
Total (Both Ailerons)
16.66 Square Feet
Area
SLOT LIP SPOILERS Span (Percent of 310 Inches)
18.54 Percent
Span
57.47 Inches (4.79 Feet)
Inboard Location
W.S. 170.72
Outboard Location
W.S. 228.19
Chord: Inboard
5.82 Inches along upper surface
Outboard
5.18 Inches along upper surface
Trailing Edge Location: Inboard
80.82 Percent
Outboard
80.71 Percent
Area (Both)
4.39 Square Feet
STRUTS Span (Attach Hole to Attach Hole)
104.354 Inches
Chord Length
8.882 Inches
Airfoil Section Designation
NACA 0033.3 (Modified)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
HORIZONTAL TAIL Root
MAC
Tip
Stabilizer Station (S.S.)
0.00
56.00
123.00
Leading Edge Fuselage Station
394.270
400.600
408.174
NACA Airfoil
0012
Incidence
-0.75 Degrees
0010 -0.75 Degrees
-0.75 Degrees -2.50 Degrees
Elevator Balance Horn (Ref. Stabilizer/Elevator Chord Line) Chord Length
52.00 Inches
41.984 Inches
30.00 Inches
Front Spar (Percent Chord)
20.00 Percent
20.00 Percent
20.00 Percent
Rear Spar (Percent Chord)
61.538 Percent
64.170 Percent
Elevator Hinge Line (Percent Chord)
66.346 Percent
68.650 Percent
Trim Tab Hinge Line (Percent Chord)
86.597 Percent
87.040 Percent
87.139 Percent
W.L.
B.L.
Span: Stabilizer
20.50 Feet
Elevator
9.65 Feet
Trim Tab
5.11 Feet
Aspect Ratio
6.00
Taper Ratio
0.577
Dihedral
0.000 Degrees
Elevator Trim Tab: Inboard Location
S.S. 44.12
Outboard Location
S.S. 105.460
Sweep Angles: Leading Edge
+6.450 Degrees
Front Spar
+4.419 Degrees
25 Percent Chord
+3.88 Degrees
Rear Spar
+0.539 Degrees
Elevator Hinge Line
0.00 Degrees
Trim Tab Hinge Line
-2.411 Degrees
Trailing Edge
-3.765 Degrees
Elevator Hinge Locations:
F.S.
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MODEL 208 MAINTENANCE MANUAL
Inboard
428.750
132.625
0.000
Center
428.750
132.625
55.000
Outboard
428.750
132.625
114.026
VERTICAL TAIL
Root
MAC
Tip
Vertical Tail Waterline (Refer to Note 1)
127.50
168.605
225.50
Leading Edge Fuselage Station
385.000
406.776
436.916
NACA Airfoil
0012
Incidence
0.0 Degrees
0.0 Degrees
0.0 Degrees
Chord Length
72.50 Inches
52.682 Inches
25.251 Inches
Front Spar (Percent Chord)
20.00 Percent
20.00 Percent
20.00 Percent
Rear Spar (Percent Chord)
63.45 Percent
50.209 Percent
Rudder Hinge Line (Percent Chord)
67.862 Percent
55.634 Percent
Span:
98.00 Inches
Aspect Ratio
1.772
Taper Ratio
0.401
009.8
Sweep Angles: Leading Edge
+27.91 Degrees
Front Spar
+23.43 Degrees
25 Percent Chord
+22.26 Degrees
Rear Spar
+10.45 Degrees
Rudder Hinge Line
+9.437 Degrees
Trailing Edge
+2.726 Degrees
Dorsal Fin
6.66 Square Feet
CONTROL SURFACE TRAVELS/CABLE TENSION SETTINGS AILERONS Aileron Up Travel
25 Degrees, +4 or -0 Degrees
Aileron Down Travel
16 Degrees, +1 or -0 Degree
Aileron Cable Tension: Fuselage Loop
20 Pounds, +5 or -5 Pounds
Wing Loop
40 Pounds, +5 or -5 Pounds
Control Wheel Interconnect
30 Pounds, +5 or -5 Pounds
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Roll Autopilot Cable: KFC-225
12 Pounds, +2 or -2 Pounds
GFC-700
12 Pounds, +2 or -2 Pounds
Aileron Friction Band: Without Autopilot
6 Pounds maximum
With Autopilot
8 Pounds maximum
Aileron Trim Tab: Right (Up)
15 Degrees, +2 or -2 Degrees
Right (Down)
15 Degrees, +2 or -2 Degrees
Aileron Trim Tab Cable Tension
3 Pounds maximum
Aileron (Right and Left) Servo Tab: Servo Tab Up
50 Percent of Aileron Travel, +1 or -1 Degree
Servo Tab Down
50 Percent of Aileron Travel, +1 or -1 Degree
RUDDER Rudder (Landplane Only): Maximum Right Rudder Travel
25 Degrees, +2 or -2 Degrees
Maximum Left Rudder Travel
25 Degrees, +2 or -2 Degrees
Rudder Cable Tension
30 Pounds, +5 or -5 Pounds
Rudder (Floatplane Only): Maximum Right Rudder Travel
23 Degrees, +2 or -0 Degrees
Maximum Left Rudder Travel
23 Degrees, +2 or -0 Degrees
Rudder Cable Tension
30 Pounds, +5 or -5 Pounds
Yaw Autopilot Cable: KFC-225
20 Pounds, +5 or -5 Pounds
GFC-700
20 Pounds, +5 or -5 Pounds
ELEVATORS Elevator Up Travel
25 Degrees, +2 or -2 Degrees
Elevator Down Travel
20 Degrees, +2 or -2 Degrees
Elevator Cable Tension
60 Pounds, +5 or -5 Pounds
Elevator Trim Tab: Tab Up (Refer to Note 2)
15 Degrees, +2 or -2 Degrees
Tab Down (Refer to Note 2)
15 Degrees, +2 or -2 Degrees 20 Pounds, +5 or -5 Pounds
Elevator Trim Tab Cable Tension Pitch Autopilot Cable:
20 Pounds, +5 or -5 Pounds
KFC-225
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20 Pounds, +5 or -5 Pounds
GFC-700 Elevator Friction Bands: Without Autopilot
15 Pounds maximum
With Autopilot
20 Pounds maximum
Flap Setting: 0 Degrees (Refer to Note 3)
0 Degrees, +0 or -0 Degrees
10 Degrees (Refer to Note 3)
10 Degrees, +1 or -2 Degrees
20 Degrees (Refer to Note 3)
20 Degrees, +2 or -2 Degrees
30 Degrees (Refer to Note 3)
30 Degrees, +1 or -2 Degrees
Flap Cable Tension: Flaps In Up Position
35 Pounds, +5 or -5 Pounds
Between 0 and 10 Degrees
15 Pounds, +5 or -5 Pounds
SLOT LIP SPOILERS Spoiler Up
40 Degrees, +5 or -5 Degrees
Spoiler Down
0 Degrees, +0 or -5 Degrees
NOTE 1: Adjust the spoiler pushrod as necessary to give a 0.01 to 0.03 inch clearance between the spoiler trailing edge and the top of the flap surface at the minimum clearance position. NOTE 2: Model 208 vertical tail tip waterline is 217.50 prior to Serial 20800029 unless modified by SK208-13. Dimensions marked are for extended tail only. NOTE 3: Maximum allowable servo with full elevator travel must not exceed 1.0 degree. NOTE 4: Left and right flap extension to be symmetrical within 0.5 degree at all positions. 3.
Dimensions and Areas - 208B, 208B Super Cargomaster and 208B Passenger
AIRPLANE OVERALL Length (Overall)
41.62 Feet
Height (Maximum)
15.46 Feet
Wing Span (Overall)
52.16 Feet
Propeller Diameter (Hartzell)
100.0 Inches
Propeller Diameter (McCauley)
106.0 Inches
Propeller Ground Clearance (Nose tire inflated and nose strut fully extended 3.625 inches) (Hartzell)
14.28 Inches
Propeller Ground Clearance (Nose tire inflated and nose strut fully extended 3.625 inches) (McCauley)
11.29 Inches
Landing Gear Track Width (centerline to centerline MLG tire):
11.66 Feet
Wheelbase (At static empty weight)
13.29 Feet
FUSELAGE Cabin Width (Maximum sidewall to sidewall)
62.00 Inches
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Cabin Height (Floorboard to headliner)
51.00 Inches
Cabin Volume (Including rear baggage area)
427 Cubic Feet
Cargo Pod
111.5 Cubic Feet
WINGS Span
52.16 Feet
Area (Includes cabin top)
279.40 Square Feet
Chord Length and Root W.S. 35.00
77.995 Inches
Mean Aerodynamic Chord (at W.S. 141.45)
66.40 Inches
Wing Station 308.00
48.111 Inches
Projected Tip (W.S. 310.00)
47.892 Inches
NACA Airfoil Designation and Root W.S. 35.00 Wing Station 308.00
NACA 23017.424 NACA 23012
Incidence Root W.S. 35.00
+2.62 Degrees
Mean Aerodynamic Chord (at W.S. 141.45)
+1.707 Degrees
Wing Station 308.00
+0.608 Degrees
Leading Edge Fuselage Station Root W.S. 35.00
171.73 Inches
Mean Aerodynamic Chord W.S. 141.52
177.57 Inches
Wing Station 308.00
186.67 Inches
Sweep Angles Leading Edge
+3.135 Degrees
Front Spar at 20 Percent Chord
+1.888 Degrees
25 Percent Chord
+1.56 Degrees
50 Percent Chord
0.0 Degrees
Rear Spar at 57.5 Percent Chord
-0.47 Degrees
Trailing Edge
-3.135 Degrees
Wing Loading at 8750 Pounds
31.3 Pounds per Square Foot
FLAPS Type
Single-Slotted (Type B Nose)
Span (Percent of 310 Inches)
62.9 Percent
Span
16.25 Feet
Inboard Location
W.S. 33.50
Outboard Location
W.S. 228.50
Chord - Projected (Inboard 22.666 Inches)
29 Percent Wing Chord
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Chord - Projected (Outboard 16.476 Inches)
29 Percent Wing Chord
Front Spar Location (Percent Flap Chord)
20 Percent
Rear Spar Location (Percent Flap Chord)
65 Percent
Tracks: Inboard
W.S. 53.158
Center
W.S. 126.658
Outboard
W.S. 214.456
Area (Both)
52.79 Square Foot
AILERONS Type and 1/4 Chord
Round Nose with Aft Set Hinge Line
Span (Percent of 310 Inches Measured at Hinge Line)
25.387 Percent
Span (Measured at Hinge Line)
6.558 Feet
Inboard Location (Hinge line at ends of aileron)
W.S. 228.924
Outboard Location (Hinge line at ends of aileron)
W.S. 307.625
Hinge Line Location (Percent Wing Chord): Outboard W.S. 230.063
77.51 Percent
Inboard W.S. 304.553
78.08 Percent
Chord Length (Percent Wing Chord): Inboard Hinge Point
29.3 Percent
Outboard Hinge Point
29.9 Percent
Aft of Hinge Line (Includes Trim Tab)
6.34 Square Feet
Trim Tab
1.10 Square Feet
Forward of Hinge Line
1.99 Square Feet
Aileron
8.33 Square Feet
Total (Both Ailerons)
16.66 Square Feet
Area
SLOT LIP SPOILERS Span (Percent of 310 Inches)
18.54 Percent
Span
57.47 Inches (4.79 Feet)
Inboard Location
W.S. 170.72
Outboard Location
W.S. 228.19
Chord: Inboard
5.82 Inches along upper surface
Outboard
5.18 Inches along upper surface
Trailing Edge Location: Inboard
80.82 Percent
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Outboard
80.71 Percent
Area (Both)
4.39 Square Feet
STRUTS Span (Attach Hole to Attach Hole)
104.354 Inches
Chord Length
8.882 Inches
Airfoil Section Designation
NACA 0033.3 (Modified)
HORIZONTAL TAIL Root
MAC
Tip
Stabilizer Station (S.S.)
0.00
56.00
123.00
Leading Edge Fuselage Station
394.270
400.600
408.174
NACA Airfoil
0012
Incidence
0.0 Degrees
0010 0.0 Degrees
Elevator Balance Horn (Ref. Stabilizer/Elevator Chord Line)
0.0 Degrees -2.50 Degrees
Chord Length
52.00 Inches
41.984 Inches
30.00 Inches
Front Spar (Percent Chord)
20.00 Percent
20.00 Percent
20.00 Percent
Rear Spar (Percent Chord)
61.538 Percent
64.170 Percent
Elevator Hinge Line (Percent Chord)
66.346 Percent
68.650 Percent
Trim Tab Hinge Line (Percent Chord)
86.597 Percent
87.040 Percent
87.139 Percent
Span: Stabilizer
20.50 Feet
Elevator
9.65 Feet
Trim Tab
5.11 Feet
Aspect Ratio
6.00
Taper Ratio
0.577
Dihedral
0.000 Degrees
Elevator Trim Tab: Inboard Location
S.S. 44.12
Outboard Location
S.S. 105.460
Sweep Angles: Leading Edge
+6.450 Degrees
Front Spar
+4.419 Degrees
25 Percent Chord
+3.88 Degrees
Rear Spar
+0.539 Degrees
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Elevator Hinge Line
0.00 Degrees
Trim Tab Hinge Line
-2.411 Degrees
Trailing Edge
-3.765 Degrees
Elevator Hinge Locations:
F.S.
W.L.
B.L.
Inboard
476.768
132.240
0.000
Center
476.768
132.240
55.000
Outboard
476.768
132.240
114.026
VERTICAL TAIL
Root
MAC
Tip
Vertical Tail Waterline
127.50
168.605
225.50
Leading Edge Fuselage Station
385.000
406.776
436.916
NACA Airfoil
0012
Incidence
0.0 Degree
0.0 Degree
0.0 Degree
Chord Length
72.50 Inches
52.682 Inches
25.251 Inches
Front Spar (Percent Chord)
20.00 Percent
20.00 Percent
20.00 Percent
Rear Spar (Percent Chord)
63.45 Percent
50.209 Percent
Rudder Hinge Line (Percent Chord)
67.862 Percent
55.634 Percent
Span:
98.00 Inches
Aspect Ratio
2.005
Taper Ratio
0.401
009.8
Sweep Angles: Leading Edge
+27.91 Degrees
Front Spar
+23.43 Degrees
25 Percent Chord
+22.26 Degrees
Rear Spar
+10.45 Degrees
Rudder Hinge Line
+9.437 Degrees
Trailing Edge
+2.726 Degrees
Dorsal Fin
6.66 Square Feet
CONTROL SURFACE TRAVELS/CABLE TENSION SETTINGS AILERONS Aileron Up Travel
25 Degrees, +4 or -0 Degrees
Aileron Down Travel
16 Degrees, +1 or -0 Degree
Aileron Cable Tension: Fuselage Loop
20 Pounds, +5 or -5 Pounds
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Wing Loop
40 Pounds, +5 or -5 Pounds
Control Wheel Interconnect
30 Pounds, +5 or -5 Pounds
Roll Autopilot Cable: KFC-225
12 Pounds, +2 or -2 Pounds
GFC-700
12 Pounds, +2 or -2 Pounds
Aileron Friction Band: Without Autopilot
6 Pounds maximum
With Autopilot
8 Pounds maximum
Aileron Trim Tab: Right (Up)
15 Degrees, +2 or -2 Degrees
Right (Down)
15 Degrees, +2 or -2 Degrees 3 Pounds maximum
Aileron Trim Tab Cable Tension Aileron (Right and Left) Servo Tab: Servo Tab Up
50 Percent of Aileron Travel, +1 or -1 Degree
Servo Tab Down
50 Percent of Aileron Travel, +1 or -1 Degree
RUDDER Rudder: Maximum Right Rudder Travel
25 Degrees, +2 or -2 Degrees
Maximum Left Rudder Travel
25 Degrees, +2 or -2 Degrees
Rudder Cable Tension
30 Pounds, +5 or -5 Pounds
Yaw Autopilot Cable: KFC-225
20 Pounds, +5 or -5 Pounds
GFC-700
20 Pounds, +5 or -5 Pounds
ELEVATORS Elevator Up Travel
25 Degrees, +2 or -2 Degrees
Elevator Down Travel
20 Degrees, +2 or -2 Degrees
Elevator Cable Tension
60 Pounds, +5 or -5 Pounds
Elevator Trim Tab: Tab Up (Refer to Note 4)
15 Degrees, +2 or -2 Degrees
Tab Down (Refer to Note 4)
15 Degrees, +2 or -2 Degrees
Elevator Trim Tab Cable Tension
20 Pounds, +5 or -5 Pounds
Pitch Autopilot Cable: KFC-225
20 Pounds, +5 or -5 Pounds
GFC-700
20 Pounds, +5 or -5 Pounds
Elevator Friction Bands:
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Without Autopilot
15 Pounds maximum
With Autopilot
20 Pounds maximum
Flap Setting: 0 Degrees (Refer to Note 5)
0 Degree, +0 or -0 Degrees
10 Degrees (Refer to Note 5)
10 Degrees, +1 or -2 Degrees
20 Degrees (Refer to Note 5)
20 Degrees, +2 or -2 Degrees
30 Degrees (Refer to Note 5)
30 Degrees, +1 or -2 Degrees
Flap Cable Tension: Flaps In Up Position
35 Pounds, +5 or -5 Pounds
Between 0 and 10 Degrees
15 Pounds, +5 or - 5 Pounds
SLOT LIP SPOILERS Spoiler Up
40 Degrees, +5 or -5 Degrees
Spoiler Down
0 Degrees, +0 or -5 Degrees
NOTE:
Adjust the spoiler pushrod as necessary to give a 0.01 to 0.03 inch clearance between the spoiler trailing edge and the top of the flap surface at the minimum clearance position.
NOTE:
Left hand and right hand flap extension to be symmetrical within 0.5 degree at all positions.
NOTE:
Maximum allowable servo with full elevator travel must not exceed 1.0 degree.
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Airplane Areas Figure 1 (Sheet 1)
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Airplane Dimensions Figure 2 (Sheet 1)
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Airplane Dimensions Figure 2 (Sheet 2)
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Airplane Dimensions Figure 2 (Sheet 3)
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Airplane Dimensions Figure 2 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL AIRPLANE STATIONS - DESCRIPTION AND OPERATION 1.
2.
General A.
The airplane is divided into reference points in inches. These reference points provide a means of quickly identifying the location of components. Three axes are used as reference points.
B.
Illustrations are provided to aid the maintenance technician in locating various stations and reference dimensions. Refer to Figure 1.
Description A.
Abbreviations and Terminology. (1) FS - FUSELAGE STATION is a horizontal reference designation starting in front of the nose of the airplane. (2) CFS - CANTED FUSELAGE STATIONS are stations tilted at an angle to waterlines; greater or less than 90 degrees. (3) WS - WING STATIONS are measured outboard from center of fuselage to wing tip, along wing. (4) WL - WATERLINE is a vertical reference designation measured parallel to ground. (5) BL - BUTTOCK LINE is a horizontal reference designation starting at airplane centerline. Right or left is added to indicate direction from airplane centerline (RBL, LBL). (6) CBL - CANTED BUTTOCK LINES are lines tilted at an angle to fuselage centerline. (7) RSS - REAR SPAR STATIONS are horizontal reference designations perpendicular to rear spar. (8) NACA - NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS mathematically constructed airfoils made available to the public by NACA. (9) SS - STABILIZER STATIONS are measured outboard from center of fuselage, along stabilizer to stabilizer tip. (10) CWS - CANTED WING STATIONS are stations tiled at an angle (not parallel) relative to centerline of wing. (11) CSS - CANTED STABILIZER STATIONS are stations tilted at an angle (not parallel) relative to centerline of stabilizer. (12) MAC - MEAN AERODYNAMIC CHORD is a mathematically intermediate chord located between root and tip chords.
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Airplane Station Diagrams Figure 1 (Sheet 1)
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Airplane Station Diagrams Figure 1 (Sheet 2)
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Airplane Station Diagrams Figure 1 (Sheet 3)
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Airplane Station Diagrams Figure 1 (Sheet 4)
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Airplane Station Diagrams Figure 1 (Sheet 5)
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Airplane Station Diagrams Figure 1 (Sheet 6)
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Airplane Station Diagrams Figure 1 (Sheet 7)
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Airplane Station Diagrams Figure 1 (Sheet 8)
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MODEL 208 MAINTENANCE MANUAL MAJOR STRUCTURAL MEMBERS - DESCRIPTION AND OPERATION 1.
General A.
The location of some major structural members is provided as a convenient reference point for locating other components. For an illustration of these major structural members, refer to Figure 1.
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Major Structural Members Locations Figure 1 (Sheet 1)
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Major Structural Members Locations Figure 1 (Sheet 2)
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Major Structural Members Locations Figure 1 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL AIRPLANE ZONING - DESCRIPTION AND OPERATION 1.
2.
General A.
The Model 208 is divided into numbered zones to provide a method for locating components. The zones are identiÞed by a three-digit number as shown in the example below. Each digit designates a zone category: major, submajor or subdivision.
B.
Major Zones. (1) 100 - FS 100.0 forward side of Þrewall and forward. (2) 200 - FS 100.0 aft side of Þrewall to FS 308.00 (208), FS356.00 (208B). (3) 300 - FS 308.00 (208), FS356.00 (208B) to end of airplane. (4) 500 - Left wing. (5) 600 - Right wing. (6) 700 - Landing gear. (7) 800 - Cabin and cargo doors. (8) 900 - Cargo pod (if applicable).
Description A.
Airplane zones may be utilized for location of work areas and components prior to beginning maintenance or service tasks on the airplane. Airplane zones are used in this manual to locate items such as placards and markings displayed on interior and exterior surfaces of the airplane. For a breakdown of the airplane zones, refer to Figure 1.
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Airplane Zones Figure 1 (Sheet 1)
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Airplane Zones Figure 1 (Sheet 2)
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Airplane Zones Figure 1 (Sheet 3)
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Airplane Zones Figure 1 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL ACCESS PLATES AND PANELS IDENTIFICATION - DESCRIPTION AND OPERATION 1.
General A.
2.
The airplane access plates cover holes that are small. Access panels cover access areas that are larger in comparison with access plates. Any access cover with a hinge is identified as a door.
Description NOTE:
3.
Access doors are on the left side of the cargo pod, if installed, and on the left and right side of the engine cowling. Refer to Chapter 52, General for location of other doors.
A.
Plates (Refer to Figure 1 thru Figure 9). (1) Access plates are on the lower surface of each wing, the bottom of the fuselage, the right side of the empennage, the side of the vertical stabilizer, and on the bottom of each horizontal stabilizer.
B.
Panels (Refer to Figure 1 thru Figure 9). (1) Access panels are on the bottom of the fuselage, the bottom of the empennage, and if installed, on the left side of the cargo pod.
Access Plates and Panel Identification A.
Access plates and panels have identifiers with a series of numbers and letters. These identifiers give the zone, the panel identifier, and the location in the zone. (1) Zones are identified by a three-number sequence that is shown in Airplane Zoning - Description and Operation. (2) The Primary identifiers follow a three-number sequence, with the first plate/panel identified as “A,” the second plate/panel identified as “B” and so on. (3) Locators follow the primary identifier and identify the top, bottom, left, right or internal location of the plate/panel.
B.
As an example, access plate 521AB identifies: (1) Zone Location (521). (2) The first panel in the zone (panel A). (3) Orientation of the panel (B for bottom).
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Fuselage Access Plates/Panels Identification Figure 1 (Sheet 1)
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Table 1. Fuselage Access Plates/Panels Identification. Panel
Equipment Located In Area (Refer to Figure 1)
253AC
Fuel Sump Tank
312AR
Oxygen filler Valve
Table 2. Model 208 Floorboard Access Plates/Panels Identification. Panel
Equipment Located In Area (Refer to Figure 2)
211AC
Heater Duct and Rudder Trim
211AL
Fuel Control Valve
211BL
Rudder Torque Tube
211CL
Rudder Pedal Linkage
211DL
Structure
211EL
Control Column, Aileron Cables and Pulleys
212AR
Rudder Pedal Linkage
212BR
Rudder Torque Tube
212CR
Rudder Torque Tube, TKS Tubing
212DR
Structure
212ER
Structure
212FR
Control Column, Aileron Cables and Pulleys, TKS Tubing
215AL
Structure
216AR
Structure, TKS Tubing
216AC
Rudder Pulleys
216BC
Rudder Pulleys
231AL
Heater Duct
231BL
Aileron Pulleys, Elevator Bell Crank, Rudder and Elevator Trim Cables and Transponder Antenna
231CL
Heater Duct
231DL
Elevator and Elevator Trim Cables, DME Antenna, and TKS tank vent line, TKS Tubing
232AR
Heater Duct
232BR
No. 2 Glide Slope Receiver and RMI Dynaverter, AHRS No. 2, TKS tank vent line
232AC
Elevator Bell Crank, Rudder and Elevator Trim
232CR
Heater Duct, TKS Tubing
232DR
Autopilot Roll Actuator/Servor, TKS propeller proportioning unit (if installed), TKS Tubing
232BC
Rudder Pulleys, TKS level sender, filler hose
251AL
Rudder, Elevator an Elevator Trim Cables, TKS Tubing, TKS Filler Tubing
251BL
Fuel Lines and Heating Ducts, TKS Tubing, TKS Filler Tubing
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MODEL 208 MAINTENANCE MANUAL Table 2. Model 208 Floorboard Access Plates/Panels Identification. (continued) Panel
Equipment Located In Area (Refer to Figure 2)
251CL
Aileron Cables and Fuel Lines, TKS Tubing, TKS Filler Tubing
251DL
Fuel Lines and Heating Ducts, TKS Tubing
251EL
Fuel Lines, TKS Tubing
251FL
Rudder, Elevator an Elevator Trim Cables
251GL
Heating Ducts, TKS Tubing
251HL
TKS Filler Tube
252AR
Electrical Wiring and Heating Ducts, TKS Tubing
252BR
Aileron Cable, TKS Propeller Proportioning Unit (If Installed), TKS Tubing
252CR
Electrical Wiring and Heating Ducts
252DR
Structure
252ER
Structure
252FR
Electrical Wiring and Heating Ducts
252GR
TKS Tubing
255AL
Rudder, Elevator an Elevator Trim Cables
255BL
Heating Ducts, TKS Tubing
255CL
Structure
255DL
Rudder, Elevator an Elevator Trim Cables
255EL
Heating Ducts, TKS Tubing
255FL
Structure
255GL
Rudder, Elevator an Elevator Trim Cables
255HL
Structure, TKS Tubing
255JL
Rudder, Elevator an Elevator Trim Cables
255KL
Structure, TKS Tubing
255LL
Rudder, Elevator an Elevator Trim Cables
255ML
Structure, TKS Tubing
255NL
Rudder, Elevator an Elevator Trim Cables
255PL
Structure, TKS Tubing
255QL
Rudder, Elevator an Elevator Trim Cables
255RL
Structure, TKS Tubing
255SL
Rudder, Elevator an Elevator Trim Cables
255TL
Structure, TKS Tubing
256AR
Structure
256BR
Electrical Wiring and Heating Ducts
256CR
Structure
6-20-02 © Cessna Aircraft Company
Page 4 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 2. Model 208 Floorboard Access Plates/Panels Identification. (continued) Panel
Equipment Located In Area (Refer to Figure 2)
256DR
Structure
256ER
Electrical Wiring and Heating Ducts
256FR
Structure
256GR
Structure
256HR
Electrical Wiring
256JR
Structure
256KR
Electrical Wiring
256LR
Structure
256MR
Electrical Wiring
256NR
Structure
256PR
Electrical Wiring
256QR
Structure
256RR
Electrical Wiring
256SR
Structure
256TR
Electrical Wiring
6-20-02 © Cessna Aircraft Company
Page 5 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Model 208 Floorboard Access Plates/Panels Identification Figure 2 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 6 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Model 208 Cargomaster Access Plates/Panels Identification Figure 3 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 7 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 3. Model 208 Cargomaster Access Plates/Panels Identification. Panel
Equipment Located In Area (Refer to Figure 3)
211AC
Heater Duct and Rudder Trim
211AL
Fuel Control Valve
211BL
Rudder Torque Tube
211CL
Rudder Pedal Linkage
211DL
Structure
211EL
Control Column, Aileron Cables and Pulleys
212AR
Rudder Pedal Linkage
212BR
Rudder Torque Tube
212CR
Rudder Torque Tube, TKS Tubing
212DR
Structure
212ER
Structure
212FR
Control Column, Aileron Cables and Pulleys, TKS Tubing
215AL
Structure
216AR
Structure, TKS Tubing
216AC
Rudder Pulleys
216BC
Rudder Pulleys
231AL
Heater Duct
231BL
Aileron Pulleys, Elevator Bell Crank, Rudder and Elevator Trim Cables and Transponder Antenna
231CL
Heater Duct
231DL
Elevator and Elevator Trim Cables, DME Antenna TKS Tank Vent Line
232AR
Heater Duct, TKS Tubing
232BR
No. 2 Glide Slope Receiver and RMI Dynaverter, AHRS No. 2, TKS vent line
232AC
Elevator Bell Crank, Rudder and Elevator Trim
232CR
Heater Duct, TKS Tubing
232DR
Autopilot Roll Actuator/Servor, TKS propeller proportioning unit (if installed), TKS Tubing
232BC
Rudder Pulleys, TKS level sender, TKS Tubing
251AL
Rudder, Elevator an Elevator Trim Cables, TKS Tubing, TKS Filler Tubing
251BL
Fuel Lines and Heating Ducts, TKS Tubing, TKS Filler Tubing
251CL
Aileron Cables and Fuel Lines, TKS Tubing
251DL
Fuel Lines and Heating Ducts, TKS Tubing
251EL
Fuel Lines, TKS Tubing
251FL
Rudder, Elevator an Elevator Trim Cables
251GL
Heating Ducts, TKS Tubing
6-20-02 © Cessna Aircraft Company
Page 8 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 3. Model 208 Cargomaster Access Plates/Panels Identification. (continued) Panel
Equipment Located In Area (Refer to Figure 3)
251HL
TKS Tubing
252AR
Electrical Wiring and Heating Ducts
252BR
Aileron Cable, TKS Propeller Proportioning Unit (If Installed), TKS Tubing
252CR
Electrical Wiring and Heating Ducts
252DR
Structure
252ER
Structure
252FR
Electrical Wiring
252GR
TKS Tubing
255AL
Rudder, Elevator an Elevator Trim Cables
255BL
Heating Ducts, TKS Tubing
255CL
Structure
255DL
Rudder, Elevator an Elevator Trim Cables
255EL
Heating Ducts, TKS Tubing
255FL
Structure
255GL
Rudder, Elevator an Elevator Trim Cables
255HL
Structure, TKS Tubing
255JL
Rudder, Elevator an Elevator Trim Cables
255KL
Structure, TKS Tubing
255LL
Rudder, Elevator an Elevator Trim Cables
255ML
Structure, TKS Tubing
255NL
Rudder, Elevator an Elevator Trim Cables
255PL
Structure, TKS Tubing
255QL
Rudder, Elevator an Elevator Trim Cables
255RL
Structure, TKS Tubing
255SL
Rudder, Elevator an Elevator Trim Cables
255TL
Structure, TKS Tubing
256AR
Structure
256BR
Electrical Wiring
256CR
Structure
256DR
Structure
256ER
Electrical Wiring
256FR
Structure
256GR
Structure
256HR
Electrical Wiring
256JR
Structure
6-20-02 © Cessna Aircraft Company
Page 9 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 3. Model 208 Cargomaster Access Plates/Panels Identification. (continued) Panel
Equipment Located In Area (Refer to Figure 3)
256KR
Electrical Wiring
256LR
Structure
256MR
Electrical Wiring
256NR
Structure
256PR
Electrical Wiring
256QR
Structure
256RR
Electrical Wiring
256SR
Structure
256TR
Electrical Wiring
6-20-02 © Cessna Aircraft Company
Page 10 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Model 208B and 208B Super Cargomaster Panels Figure 4 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 11 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 4. Model 208B and 208B Super Cargomaster Panels. Panel
Equipment Located In Area (Refer to Figure 4)
211AC
Heater Duct and Rudder Trim
211AL
Fuel Control Valve
211BL
Rudder Torque Tube
211CL
Rudder Pedal Linkage
211DL
Structure
211EL
Control Column, Aileron Cables and Pulleys
212AR
Rudder Pedal Linkage
212BR
Rudder Torque Tube
212CR
Rudder Torque Tube, TKS Tubing
212DR
Structure
212ER
Structure
212FR
Control Column, Aileron Cables and Pulleys, TKS Tubing
215AL
Structure
216AR
Structure, TKS Tubing
216AC
Rudder Pulleys
216BC
Rudder Pulleys
231AL
Heater Duct
231BL
Aileron Pulleys, Elevator Bell Crank, Rudder and Elevator Trim Cables and Transponder Antenna
231CL
Heater Duct
231DL
Elevator and Elevator Trim Cables and DME Antenna
232AR
Heater Duct, TKS Tubing
232BR
No. 2 Glide Slope Receiver and RMI Dynaverter, AHRS No. 2
232AC
Elevator Bell Crank, Rudder and Elevator Trim
232CR
Heater Duct, TKS Tubing
232DR
Autopilot Roll Actuator/Servor, TKS Tubing
232BC
Rudder Pulleys
251AL
Rudder, Elevator and Elevator Trim Cables
251BL
Rudder, Elevator and Elevator Trim Cables
251CL
Fuel Lines
251DL
Aileron Cables and fuel Lines
251EL
Rudder, Elevator and Trim Cables, TKS Tubing
251FL
Fuel Lines, TKS Tubing
251GL
Structure
6-20-02 © Cessna Aircraft Company
Page 12 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 4. Model 208B and 208B Super Cargomaster Panels. (continued) Panel
Equipment Located In Area (Refer to Figure 4)
251HL
Fuel Lines, TKS Tubing
251JL
Structure, TKS Tubing
251KL
Fuel Lines, TKS Tubing
251LL
Structure, TKS Tubing
251ML
Rudder, Elevator and Elevator Trim Cables
251NL
Structure, TKS Tubing
251PL
TKS Tubing
252AR
Electrical Wiring and Heating Ducts
252BR
Aileron Cable
252CR
Electrical Wiring and Heating Ducts
252DR
Structure, TKS Propeller Proportioning Unit (If Installed), TKS Tubing
252ER
Structure,TKS Propeller Proportioning Unit (If Installed), TKS Tubing
252FR
Electrical Wiring
252GR
Electrical Wiring, TKS Tubing
252HR
Structure, TKS Tubing
252JR
Electrical Wiring
252KR
Structure
252LR
Structure
252MR
Electrical Wiring
252NR
TKS Tubing
255AL
Rudder, Elevator an Elevator Trim Cables
255BL
Heating Ducts, TKS Tubing
255CL
Structure
255DL
Rudder, Elevator an Elevator Trim Cables
255EL
Heating Ducts
255FL
Structure
255GL
Rudder, Elevator an Elevator Trim Cables
255HL
Structure, TKS Tubing
255JL
Rudder, Elevator an Elevator Trim Cables
255KL
Structure
255LL
Rudder, Elevator an Elevator Trim Cables, TKS Tubing
255ML
Structure
255NL
Rudder, Elevator an Elevator Trim Cables
255PL
Structure, TKS Tubing
255QL
Rudder, Elevator an Elevator Trim Cables
6-20-02 © Cessna Aircraft Company
Page 13 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 4. Model 208B and 208B Super Cargomaster Panels. (continued) Panel
Equipment Located In Area (Refer to Figure 4)
255RL
Structure
255SL
Rudder, Elevator an Elevator Trim Cables, TKS Tubing
255TL
Structure
255UL
Structure, TKS Tubing
255VL
Rudder, Elevator and Elevator Trim Cables
255WL
Structure, TKS Tubing
255XL
Rudder, Elevator and Elevator Trim Cables
255YL
Structure, TKS Tubing
256AR
Structure
256BR
Electrical Wiring
256CR
Structure
256DR
Structure
256ER
Electrical Wiring
256FR
Structure
256GR
Structure
256HR
Electrical Wiring
256JR
Structure
256KR
Electrical Wiring
256LR
Structure
256MR
Electrical Wiring
256NR
Structure
256PR
Electrical Wiring
256QR
Structure
256RR
Electrical Wiring
256SR
Structure
256TR
Electrical Wiring
256VR
Structure
256WR
Electrical Wiring
256XR
Structure
256YR
Electrical Wiring
255AAL
Structure, TKS Tubing
255ABL
Rudder, Elevator an Elevator Trim Cables, TKS Tubing
255ACL
Structure
256ZR
Structure
6-20-02 © Cessna Aircraft Company
Page 14 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 4. Model 208B and 208B Super Cargomaster Panels. (continued) Panel
Equipment Located In Area (Refer to Figure 4)
256AAR
Electrical Wiring
256ABR
Structure
256ACR
Electrical Wiring
6-20-02 © Cessna Aircraft Company
Page 15 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Model 208B Passenger Figure 5 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 16 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 5. Model 208B Passenger. Panel
Equipment Located In Area (Refer to Figure 5)
211AC
Heater Duct and Rudder Trim
211AL
Fuel Control Valve
211BL
Rudder Torque Tube
211CL
Rudder Pedal Linkage
211DL
Structure
211EL
Control Column, Aileron Cables and Pulleys
212AR
Rudder Pedal Linkage
212BR
Rudder Torque Tube
212CR
Rudder Torque Tube
212DR
Structure
212ER
Structure
212FR
Control Column, Aileron Cables and Pulleys
215AL
Structure
216AR
Structure
216AC
Rudder Pulleys
216BC
Rudder Pulleys
231AL
Heater Duct
231BL
Aileron Pulleys, Elevator Bell Crank, Rudder and Elevator Trim Cables and Transponder Antenna
231CL
Heater Duct
231DL
Elevator and Elevator Trim Cables and DME Antenna
232AR
Heater Duct
232BR
No. 2 Glide Slope Receiver and RMI Dynaverter, AHRS No. 2
232AC
Elevator Bell Crank, Rudder and Elevator Trim
232CR
Heater Duct
232DR
Autopilot Roll Actuator/Servor
232BC
Rudder Pulleys
251AL
Rudder, Elevator an Elevator Trim Cables
251BL
Rudder, Elevator an Elevator Trim Cables
251CL
Fuel Lines and Heating Ducts
251DL
Aileron Cables and Fuel Lines
251EL
Rudder, Elevator an Elevator Trim Cables
251FL
Fuel Lines and Heating Ducts
251GL
Structure
6-20-02 © Cessna Aircraft Company
Page 17 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 5. Model 208B Passenger. (continued) Panel
Equipment Located In Area (Refer to Figure 5)
251HL
Fuel Lines and Heating Ducts
251JL
Structure
251KL
Fuel Lines and Heating Ducts
251LL
Structure
251ML
Rudder, Elevator an Elevator Trim Cables
251NL
Structure and Heating Duct
252AR
Structure
252BR
Electrical Wiring and Heating Ducts
252CR
Structure, TKS Propeller Proportioning Unit (If Installed)
252DR
Structure, TKS Propeller Proportioning Unit (If Installed)
252ER
Electrical Wiring and Heating Ducts, TKS Propeller Proportioning Unit (If Installed)
252FR
Aileron Cable
252GR
Electrical Wiring and Heating Ducts
252HR
Structure
252JR
Electrical Wiring and Heating Ducts
252KR
Structure
252LR
Structure
252MR
Electrical Wiring and Heating Ducts
255AL
Rudder, Elevator an Elevator Trim Cables
255BL
Structure and Heating Ducts
255CL
Structure
255DL
Rudder, Elevator an Elevator Trim Cables
255EL
Structure and Heating Ducts
255FL
Structure
255GL
Rudder, Elevator an Elevator Trim Cables
255HL
Structure and Heating Ducts
255JL
Structure
255KL
Rudder, Elevator an Elevator Trim Cables
255LL
Structure and Heating Ducts
255ML
Structure
255NL
Rudder, Elevator an Elevator Trim Cables
255PL
Structure and Heating Ducts
255QL
Structure
255RL
Rudder, Elevator an Elevator Trim Cables
255SL
Structure
6-20-02 © Cessna Aircraft Company
Page 18 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 5. Model 208B Passenger. (continued) Panel
Equipment Located In Area (Refer to Figure 5)
255TL
Rudder, Elevator an Elevator Trim Cables
255UL
Structure
255VL
Rudder, Elevator an Elevator Trim Cables
255WL
Structure
255XL
Rudder, Elevator an Elevator Trim Cables
255YL
Structure
255ZL
Rudder, Elevator an Elevator Trim Cables
255AAL
Structure
255ABL
Rudder, Elevator an Elevator Trim Cables
255ACL
Structure
256AR
Structure
256BR
Electrical Wiring and Heating Ducts
256CR
Structure
256DR
Structure
256ER
Electrical Wiring and Heating Ducts
256FR
Structure
256GR
Structure
256HR
Electrical Wiring and Heating Ducts
256JR
Structure
256KR
Structure
256LR
Electrical Wiring and Heating Ducts
256MR
Structure
256NR
Structure
256PR
Electrical Wiring and Heating Ducts
256QR
Structure
256RR
Structure
256SR
Electrical Wiring
256TR
Structure
256UR
Electrical Wiring
256VR
Structure
256WR
Electrical Wiring
256XR
Structure
256YR
Electrical Wiring
256ZR
Structure
6-20-02 © Cessna Aircraft Company
Page 19 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 5. Model 208B Passenger. (continued) Panel
Equipment Located In Area (Refer to Figure 5)
256AAR
Electrical Wiring
256ABR
Structure
256ACR
Electrical Wiring
6-20-02 © Cessna Aircraft Company
Page 20 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Panels Figure 6 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 21 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 6.
Panels.
Panel
Equipment Located In Area (Refer to Figure 6)
226A
Elevator Trim
226B
Aileron and Rudder Trim
226C
Electric Elevator Trim
226D
Electric Elevator Trim
6-20-02 © Cessna Aircraft Company
Page 22 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Left Lower Wing Panels Figure 7 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 23 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 7.
Left Lower Wing Panels.
Panel
Equipment Located In Area (Refer to Figure 7)
501AB
Cabin Air Inlet Duct
501BB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
501CB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
501DB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
501EB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
503AB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed), TKS Proportioning Unit (if installed)
503BB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed), TKS Proportioning Unit (if installed)
503CB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
503DB
Aileron Cables, Pulleys and Bell Crank
503EB
Aileron and Spoiler Bell Cranks
503FB
Electrical Wiring
503GB
Electrical Wiring
503HB
Electrical Wiring
503JB
Electrical Wiring
511AB
Flow Control Valves, Pressure Switches, Fuel Level Sensor
521AB
Fuel Tank
521BB
Fuel Tank
521CB
Fuel Tank
521DB
Fuel Tank
521EB
Fuel Tank
523AB
Magnetometer
523BB
Fuel Level Sensor
525AB
Flap Bell Crank
525BB
Flap Interconnect Rod
525CB
Flap Bell Crank
525DB
Flap Interconnect Rod
525EB
Flap Cables
525FB
Flap Cables
525GB
Flap Cable and Pulley
551AB
Aileron Trim Actuator
575AB
Fuel Vent
6-20-02 © Cessna Aircraft Company
Page 24 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Right Lower Wing Panels Figure 8 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 25 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 8.
Right Lower Wing Panels.
Panel
Equipment Located In Area (Refer to Figure 8)
601AB
Cabin Air Inlet Duct
601BB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
601CB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
601DB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
601EB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
603AB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed), TKS Proportioning Unit (if installed)
603BB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed), TKS Proportioning Unit (if installed)
603CB
Aileron Cables, Electrical Wiring and Deice Tubing, (if installed)
603DB
Aileron Cables, Pulleys and Bell Crank
603EB
Aileron and Spoiler Bell Cranks
603FB
Electrical Wiring
603GB
Electrical Wiring
603HB
Electrical Wiring
603JB
Electrical Wiring
611AB
Air Conditioner Evaporator, Fuel Tank Shutoff Valves, Fuel Level Sensor
621AB
Fuel Tank
621BB
Fuel Tank
621CB
Fuel Tank and Strut Fitting
621DB
Fuel Tank
621EB
Fuel Tank
623AB
Flux Gate, magnetometer
623BB
Fuel Level Sensor
625AB
Flap Bell Crank
625BB
Flap Interconnect Rod
625CB
Flap Bell Crank
625DB
Flap Interconnect Rod
625EB
Flap Cables
625FB
Flap Cables
625GB
Flap Cable and Pulley
651AB
Aileron Trim Actuator
675AB
Fuel Vent
6-20-02 © Cessna Aircraft Company
Page 26 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Aft Fuselage, Horizontal and Vertical Stabilizer Panels Figure 9 (Sheet 1)
6-20-02 © Cessna Aircraft Company
Page 27 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Table 9. Aft Fuselage, Horizontal and Vertical Stabilizer Panels. Panel
Equipment Located In Area (Refer to Figure 9)
320A 330A
Elevator Bell Crank and Rudder Cables, TKS Proportioning Units (IF Installed), TKS Low Pressure Switches (IF Installed) Rudder Gustlock
340A
Emergency Locator Transmitter
341A
Deice Line (IF Installed)
341B
Antenna Harness
341C
TKS Deice Line (IF Installed)
343A
Rudder Balance Weight
373AL
Elevator Trim Actuator
373BL
Deice Line (IF Installed)
374AR
Elevator Trim Actuator
374BR
Deice Line (IF Installed)
6-20-02 © Cessna Aircraft Company
Page 28 Apr 1/2010
7
CHAPTER
LIFTING AND SHORING
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
7-00-00
Page 1
Aug 1/1995
7-10-00
Pages 201-212
Aug 1/1995
7-10-01
Pages 201-204
Aug 1/1995
7-20-00
Page 201
Aug 1/1995
07-Title 07-List of Effective Pages 07-Record of Temporary Revisions 07-Table of Contents
07 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jan 3/2005
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS LIFTING AND SHORING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7-00-00 7-00-00 7-00-00 7-00-00
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JACKING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Jacking Procedures For Airplanes Without Cargo Pods . . . . . . . . . . . . . . . . . . . . . . . . . Recommended Jacking Procedures For Airplanes With Cargo Pod . . . . . . . . . . . . . . Alternate Jacking Procedure For Airplanes With Cargo Pod . . . . . . . . . . . . . . . . . . . . . Jacking Individual Main Gear Wheels Using Integral Jack Pads . . . . . . . . . . . . . . . . .
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EMERGENCY LIFTING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation For Lifting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Lifting Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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SHORING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shoring Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL LIFTING AND SHORING - GENERAL 1.
Scope A.
2.
Definition A.
3.
This chapter describes both standard and emergency procedures used to lift the airplane off the ground.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief de finition of the sections incorporated in this chapter is as follows: (1) The section on jacking provides normal procedures and techniques used to jack the airplane off the ground. (2) The section on emergency lifting provides procedures, techniques and fabrication information needed to lift the airplane by overhead means. (3) The section on shoring provides information regarding construction and location of shoring devices.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Main/Nose Gear Jack
2-169-10
Cessna Aircraft Company Cessna Parts Distribution Department 701, CPD 25800 East Pawnee Road Wichita, KS 67218-5590 (P. O. Box 7704 Wichita, KS 67277-7704)
To jack airplane.
Nose Gear Jack
02-0222-0100
Tronair South 1740 Eber Rd. Holland, OH 43528
To jack nose gear.
Jack Pad
J-1370
Tronair
Used in conjunction with Tronair nose gear jack.
Tail Stand
2-531-01
Cessna Aircraft Company
To provide support for tailcone.
Wing Jack
Support Systems Corp. 4216 N. Carmichael Court Montgomery, AL 36106
Used to jack wings in alternate lifting method.
Wing Jack Pad
Locally Fabricated
Used in conjunction with alternate lifting method.
Spreader Jig
Locally Fabricated
Used for emergency lifting.
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MODEL 208 MAINTENANCE MANUAL JACKING - MAINTENANCE PRACTICES 1.
2.
General A.
The entire airplane may be lifted by using jacks. Jack placement is dependent on specific model and cargo pod configuration. The recommended method of jacking the airplane utilizes integral jack points located on the bottom of the fuselage area. An alternate method is also provided which uses tall jacks in conjunction with fabricated jack pads under the wings to raise the airplane.
B.
Airplane jacking is used to aid in removal/installation of the landing gear and anytime the airplane must be supported off the floor. When jacking, observe the following notes: The airplane may be jacked with full fuel tanks.
NOTE:
When possible, the airplane should be on a level surface. The jacking site should be protected from the wind, preferably inside a hangar.
NOTE:
In some instances, it may be necessary to use optional hoisting rings for the initial lift; to be followed up with jacks.
NOTE:
Jacks should be used in conjunction with wing and fuselage shoring.
Tools, Equipment and Materials A.
3.
NOTE:
Refer to Lifting and Shoring - General for a list of required tools, equipment and material.
Description and Operation A.
Two jack points are provided on the underside of the fuselage for jacking the main landing gear. These jack points may be used singularly to lift one wheel, or in conjunction with the other two jack points to lift the entire airplane. (1) For 208 and 208 Cargomaster airplanes without a cargo pod, jack points are located at each aft main landing gear support (FS 207.44, BL 23.77 left and right). (2) For 208 and 208 Cargomaster airplanes with a cargo pod, jack points are located at each aft main landing gear fitting using the outboard cap attach bolt (FS 207.44, BL 31.125 left and right). (3) For 208B airplanes, jack points are located at FS 227.44, BL 31.125 left and right. NOTE:
B.
Main landing gear to fuselage fairings must be removed to access jack points on all airplanes with cargo pods.
One jack point is provided for the nose wheel. The location of this jack point is common for all 208 Models. This jack point is located on the centerline of the airplane, aft of the nose gear at the firewall (FS 100.00).
CAUTION: A tail stand must be used when servicing airplane inside tail section. Ensure the tail stand is strong enough to support the airplane. C.
Airplanes may be equipped with two additional jack points for changing the main landing gear tires. These integral jack points are located on the aft side of main landing gear axle fittings. NOTE:
4.
These jack points are standard equipment on 208 serial 20800061 and On; on 208 Cargomaster serial 20800113 and On; and on 208B Models serial 208B0001 and On.
Jacking Procedures For Airplanes Without Cargo Pods A.
Jacking Instructions. NOTE: (1)
This is the recommended jacking procedure for airplanes without cargo pods.
Ensure static ground wire is attached to airplane.
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Jack Points Airplanes Without Cargo Pods Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4)
Position jacks beneath each jack point. Refer to Figure 201 for specific jack point locations. Ensure that jack base is level and jack cylinder is vertical at start of jacking operation. Raise jacks simultaneously, keeping airplane level until tires are clear of ground. NOTE:
(5) (6) (7) (8) 5.
Raise airplane no more than required for maintenance being performed.
Position tail stand under tail tie-down for stability. On completion of maintenance, remove tail stand. To lower airplane; lower jacks simultaneously. Remove jacks.
Recommended Jacking Procedures For Airplanes With Cargo Pod A.
Jacking Instructions.
CAUTION: Do not use cargo pod structure for jacking or blocking surface. (1) (2) (3) (4)
Ensure static ground wire is attached to airplane. Remove main gear to fuselage fairing. Refer to Chapter 32, Main Landing Gear - Maintenance Practices. Position jacks at left and right main landing gear attach trunnion bearing caps (forward outboard bolt heads). Refer to Figure 202 for specific jack point locations. Position jack at nose gear on centerline of airplane (FS 100.00).
CAUTION: Jack base must be level and jack cylinder vertical at start of jacking operations. (5)
Raise jacks simultaneously, keeping airplane level. NOTE:
(6) (7) (8) 6.
Raise jacks no more than required for maintenance being performed.
Position tail stand under tail tie-down. On completion of maintenance, remove jack stand and lower jacks simultaneously. Remove jacks and ground wire.
Alternate Jacking Procedure For Airplanes With Cargo Pod A.
Jacking Instructions.
CAUTION: Do not use cargo pod structure for jacking or blocking surface. (1) (2) (3) (4) (5)
Ensure static ground wire is attached to airplane. Position jack under tail tie-down ring. Position a jack at WS 141.2 or WS 155.9 on the front spar rivet line of each wing. Refer to Figure 203 for specific jack point locations. Fabricate jack pads. Refer to Figure 204. Place jack pads on jacks.
CAUTION: Wing jacks must be equipped with jack pads to protect wing structure. CAUTION: Jack base must be level and jack cylinder vertical at start of jacking operation. (6)
Raise three jacks simultaneously, keeping airplane level. NOTE:
Raise airplane no more than required for maintenance being performed.
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Jack Points for Airplanes With Cargo Pods Figure 202 (Sheet 1)
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Jack Points for Airplanes With Cargo Pods Figure 202 (Sheet 2)
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Jack Points for Airplanes With Cargo Pods Figure 202 (Sheet 3)
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Jack Points for Airplanes With Cargo Pods Figure 202 (Sheet 4)
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Alternate Jack Points for Airplanes With Cargo Pods Figure 203 (Sheet 1)
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Alternate Jack Points for Airplanes With Cargo Pods Figure 203 (Sheet 2)
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Alternate Jack Points for Airplanes With Cargo Pods Figure 203 (Sheet 3)
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Fabrication of Wing Jack Pad Figure 204 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL When airplane has been raised to a sufficient height, position a jack stand under nose gear jack point for stability. (8) On completion of maintenance, remove jack stand from nose. (9) Lower all jacks simultaneously. (10) Remove jacks and ground wire.
(7)
7.
Jacking Individual Main Gear Wheels Using Integral Jack Pads A.
Jacking Instructions. (1) Position wheel chocks at wheels that will not be jacked. (2) Remove optional brake fairing if installed. Refer to Chapter 32, Wheels and Brakes Maintenance Practices. (3) Position jack beneath integral main landing gear jack pad. (4) Ensure that jack base is level and jack cylinder is vertical at start of jacking operation. NOTE: (5) (6) (7)
Raise wheel no more than required for maintenance being performed.
On completion of maintenance, lower and remove jack. Install brake fairing. Refer to Chapter 32, Wheels and Brakes - Maintenance Practices. Remove wheel chocks.
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MODEL 208 MAINTENANCE MANUAL EMERGENCY LIFTING - MAINTENANCE PRACTICES 1.
General A.
2.
In some instances (i.e., off-runway landing, collapsed gear, etc.) it may be necessary to use overhead means to lift (hoist) the airplane. This maintenance practice provides information needed to fabricate a spreader jig, attach jig to airplane and lift airplane using overhead means.
Preparation For Lifting A.
The following procedures shall be accomplished prior to emergency lifting (hoisting). (1) Defuel airplane. Refer to Chapter 28, Fuel System - Maintenance Practices. (2) Remove cargo from airplane. (3) Remove all antennas that will interfere with hoisting. Refer to Model 208 Avionics Manual. (4) Secure propeller with propeller anchor assembly. Refer to Chapter 10, Mooring - Maintenance Practices. (5) Check hoist cables assembly for frayed or deteriorated cables and loose bolts or fittings. (6) Ensure cables are not twisted or kinked. (7) Ensure hoist(s) has a minimum capacity of 10 tons. (8) If hoisting airplane in a hangar, ensure that hanger structure will support 10 tons. (9) Cover windshield with padded blankets. (10) Check center of gravity. Airplane should be ballasted as required to locate the center of gravity between hoisting attach points.
WARNING: Do not attempt to raise the airplane with a sling if airplane is equipped with a float or amphibian kit. 3.
Emergency Lifting Procedures A.
Lifting Procedures (Refer to Figure 201 and Figure 202). (1) Fabricate a spreader jig. Refer to Figure 201 for fabrication details. NOTE: (2) (3)
Remove wing-to-fuselage fairing strips. Refer to Chapter 57, Wings - Removal/Installation. Attach spreader jig to chain at four points using bolt, washer and nut. Refer to Figure 202. NOTE:
(4) (5)
A spreader jig is used to ensure that only vertical force is applied to hoisting rings.
Length of chain below spreader jig should be kept at a minimum.
Attach chain to hoisting rings at four points using snap hook. Attach long ropes to tie-down fittings at wing and tail. Use these ropes to stabilize and guide airplane during lifting and lowering.
CAUTION: Raise airplane slowly to assure airplane stability and safety during emergency lifting operations. (6) (7) (8) (9)
Raise airplane enough to place jacks under fuselage jack points. Remove emergency lifting devices. Reinstall wing-to-fuselage fairing strips. Refer to Chapter 57, Wings - Removal/Installation. On completion of maintenance, lower and remove jacks.
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Spreader Jig Fabrication Figure 201 (Sheet 1)
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Spreader Jig Fabrication Figure 201 (Sheet 2)
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Spreader Jig Installation Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL SHORING - MAINTENANCE PRACTICES 1.
2.
Description and Operation A.
Shoring the airplane is accomplished by using contour boards. The boards can be fabricated locally using plywood side by side until a thickness of 2.0 inches is obtained, and contouring to fit the lower surface of the wing or fuselage. The contour boards should be padded with 0.5 inch felt and covered with canvas duck or equivalent.
B.
Fuselage jacks should be used in conjunction with wing and fuselage shoring.
Shoring Locations A.
Refer to the following wing and fuselage stations for locations where contour boards may be used for shoring purposes.
COMPONENT TO BE SHORED
PERMISSABLE SHORING LOCATIONS (BY WING OR FUSELAGE STATION)
Wing
Wing Station 141.20 Wing Station 155.90
Fuselage
Fuselage Station 100.00 Fuselage Station 166.48 (Primary shoring location) Fuselage Station 168.70 (Primary shoring location) Fuselage Station 194.80 (Primary shoring location) Fuselage Station 208.00 (Primary shoring location) Fuselage Station 234.00 Fuselage Station 284.00
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8
CHAPTER
LEVELING AND WEIGHING
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
8-00-00
Page 1
Aug 1/1995
8-20-00
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Aug 1/1995
08-Title 08-List of Effective Pages 08-Record of Temporary Revisions 08-Table of Contents
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MODEL 208 MAINTENANCE MANUAL
CONTENTS LEVELING AND WEIGHING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEVELING - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Longitudinal Leveling Using Fuselage Leveling Points . . . . . . . . . . . . . . . . . . . . . . . . . . Longitudinal Leveling Using Pilot’s Seat Rails. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lateral Leveling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL LEVELING AND WEIGHING - GENERAL 1.
2.
Scope A.
This chapter provides information necessary to properly level the airplane for any maintenance, overhaul or major repairs which might become necessary.
B.
For information on airplane weighing procedures, refer to the Pilot’s Operating Handbook.
Tools, Equipment and Material
NAME
NUMBER
Spirit Level
3.
MANUFACTURER
USE
Commercially available
Bubble level used to level airplane.
Definition A.
Leveling. (1) The section on leveling provides maintenance practices and instructions for longitudinal and lateral leveling of the airplane. This leveling is accomplished using a spirit level of at least 18 inches in length.
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MODEL 208 MAINTENANCE MANUAL LEVELING - MAINTENANCE PRACTICES 1.
General A.
To obtain longitudinal leveling indication, points are provided on left side of airplane fuselage. (1) 208 and 208 Cargomaster leveling points are as follows: FS 209.00, WL 97.50 and FS 227.00, WL 97.50. (2) 208B, 208B Super Cargomaster and 208B Passenger leveling points are as follows: FS 239.05, WL 97.50 and FS 272.13, WL 97.50.
B.
Longitudinal leveling of the airplane for weighing will require that the main landing gear be supported by stands, blocks, etc. on the main gear scales to a position at least four inches higher than the nose gear as it rests on an appropriate scale. This initial elevated position will compensate for the difference in waterline station between the main and nose gear so that final leveling can be accomplished solely by deflating the nose gear tire. NOTE:
2.
C.
The airplane can also be leveled longitudinally by raising or lowering the airplane at the jack points.
D.
Lateral leveling indication is obtained inside airplane by placing a spirit level directly on seat rails just aft of crew doors (removing carpet if necessary), allowing level to be observed from outside of airplane.
Longitudinal Leveling Using Fuselage Leveling Points A.
3.
Leveling Procedures (Refer to Figure 201). (1) Remove screws at longitudinal leveling points located on left side of fuselage. (2) Obtain two screws of sufficient length to provide resting points for level. (3) Install screws at longitudinal leveling points on fuselage. (4) Position a spirit level on screws. (5) Observe level. (6) To level airplane longitudinally, de flate nose gear tire to properly center bubble in level.
Longitudinal Leveling Using Pilot's Seat Rails. A.
4.
Since the nose gear strut on this airplane contains an oil snubber for shock absorption rather than an air/oil shock strut, it cannot be deflated to aid in airplane leveling.
Leveling Procedures (Refer to Figure 201). (1) Move pilot's seat to the most forward position. (2) Place level on top of (and parallel to) seat rail, just aft of pilot’s seat. (3) Observe level. (4) To level airplane longitudinally, de flate nose gear tire to properly center bubble in level.
Lateral Leveling A.
Leveling Procedures (Refer to Figure 201). (1) Place a spirit level directly on seat rails just aft of crew doors, removing carpet if necessary. (2) Observe level. (3) To level airplane laterally, deflate main gear tire to properly center bubble in level.
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Leveling Points Figure 201 (Sheet 1)
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Leveling Points Figure 201 (Sheet 2)
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9
CHAPTER
TOWING AND TAXIING
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
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CONTENTS TOWING AND TAXIING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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TOWING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear Towing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cold Weather Towing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL TOWING AND TAXIING - GENERAL 1.
2.
Scope A.
This chapter describes towing procedures for movement of the airplane on the ground.
B.
For taxiing procedures, refer to the Pilot’s Operating Handbook.
Definition A.
Towing. (1) The section on towing describes those procedures and cautions applicable to all model 208 airplanes. Refer to Towing - Maintenance Practices for complete instructions.
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MODEL 208 MAINTENANCE MANUAL TOWING - MAINTENANCE PRACTICES 1.
General A.
Towing the airplane is accomplished through the nose gear axle, using a yoke-type tow bar (standard equipment in the airplane).
CAUTION: Ensure all external equipment is disconnected from the airplane. Do not push or pull on control surfaces or propeller when maneuvering airplane. CAUTION: Use the yoke-type towbar (standard equipment in the airplane) to tow the airplane. This is the approved method to tow the airplane. Other tow methods could cause structural damage to the airplane. B.
During the towing operation, the maximum nose gear turning angle should not be exceeded on either side of center. Exceeding the angle will damage the nose gear. During nose wheel towing, all turning is accomplished through the tow bar. NOTE:
C.
2.
The nose gear is equipped with stop blocks and markings which provide an indication that the turning limits have been met.
A qualiÞed ground crew member should be stationed in the pilot’s seat during all phases of the towing operation to watch for hazardous conditions. This ground crew member can also stop the airplane if the tow bar breaks or becomes uncoupled. In congested areas, wing and/or tail walkers should be used to ensure adequate clearance between airplanes, adjacent equipment and structure.
Precautions A.
Observe the following cautions prior to towing the airplane.
CAUTION: Do not exceed 50 degrees turning limitation. CAUTION: The maximum nose gear towing/turning angle limit is 50 degrees either side of center. Forcing the nose gear beyond towing limits will result in damage to the nose gear, shimmy damper and structure. If turn limits are exceeded, an inspection of the nose gear assembly and nose gear wheel well structure must be performed. CAUTION: The parking brake, rudder gust locks, wheel chocks, static ground cable and mooring cable should be released or removed before towing. Failure to do so could result in structural damage to the airplane. 3.
Nose Gear Towing A.
Towing Instructions for Airplanes with a Lord Shimmy Dampener (Refer to Figure 201). (1) Insert tow bar into nose wheel axle.
CAUTION: Ensure all external equipment is disconnected from the airplane. Do not push or pull on control surfaces or propeller when maneuvering airplane. (2) (3) (4) (5) (6)
Station person in pilot's seat to assist with braking of airplane. Station persons at wing struts for pushing airplane. Remove wheel chocks, static ground cables and mooring cables. Release parking brake. Release rudder gust lock.
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Turning Radius Figure 201 (Sheet 1)
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Turning Radius Figure 201 (Sheet 2)
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Turning Radius Figure 201 (Sheet 3)
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Turning Radius Figure 201 (Sheet 4)
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If area is congested, station wing walkers and tail walkers around the airplane to ensure adequate clearance between airplanes and adjacent equipment or structures.
CAUTION: Do not exceed 50 degrees turning limitation. (8) (9) 4.
When towing is complete, center nose wheel, engage parking brake, chock wheels and apply gust locks as required. Disconnect tow bar.
Cold Weather Towing A.
During winter months and cold weather operations, maintenance personnel and ground support personnel must be aware of the following concerns and safety requirements: (1) Reduced visibility; (2) Poor traction; and (3) Increased stopping distance.
CAUTION: Dry snow provides better traction than wet snow. Wet snow thaws and refreezes, creating hazardous driving conditions while towing airplanes. Heavy trafÞc and exhaust from parked vehicles can warm and thaw ice and snow on ramps, causing the ramp surface to become wet and slippery. CAUTION: Brakes applied suddenly or too hard may cause the towing vehicle to jackknife. On hard packed snow, apply brakes until wheels begin to slide, then release brake pressure slightly to reduce speed and maintain control of vehicle. CAUTION: Use proper towing vehicle with chains installed, if required, and proper tow bar. Make gradual starts and turns, steering smoothly. Traction can be reduced with fast starts which may cause towing vehicle wheels to spin. CAUTION: When on a slick ramp, position airplane so it will not be required to make sharp turns during taxi. Position airplane directly on taxiway to minimize turns and allow lower power settings which reduce blowing snow and foreign object damage. B.
Towing Instructions (Refer to Figure 201). (1) Insert tow bar into nose wheel axle.
CAUTION: Ensure all external equipment is disconnected from the airplane. Do not push or pull on control surfaces or propeller when maneuvering airplane. (2) (3)
Station person in pilot's seat to assist with braking of airplane. Station persons at wing struts for pushing airplane. NOTE:
(4) (5) (6)
Chocks may be frozen to the ground. If chocks are frozen to the ground, check wheels to ensure they are not frozen to the ground.
Remove wheel chocks, static ground cables and mooring cables. Ensure wheels are not frozen to parking surface. Release rudder gust lock.
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MODEL 208 MAINTENANCE MANUAL (7)
If area is congested, station wing walkers and tail walkers around airplane to ensure adequate clearance between airplanes and adjacent equipment or structures.
CAUTION: Do not exceed 50 degrees turning limitation. (8)
When towing is complete, center nose wheel, chock wheels and apply gust locks, as required.
CAUTION: Using chocks on ice may cause them to slide. Ensure chocks are Þrmly positioned and tied together. CAUTION: Do not set parking brake during cold weather, as accumulated moisture may freeze brakes. (9)
Disconnect tow bar.
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CHAPTER
PARKING AND MOORING
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
10-00-00
Pages 1-2
Aug 1/1995
10-10-00
Pages 201-202
Jun 3/2002
10-11-00
Pages 201-205
Sep 1/2004
10-20-00
Pages 201-210
Aug 1/1995
10-30-00
Pages 201-203
Jun 1/1998
10-Title 10-List of Effective Pages 10-Record of Temporary Revisions 10-Table of Contents
10 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS PARKING, MOORING, STORAGE AND RETURN TO SERVICE - GENERAL . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10-00-00 10-00-00 10-00-00 10-00-00
Page 1 Page 1 Page 1 Page 2
PARKING - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Parking Instructions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cold Weather Parking Instructions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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STORAGE - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flyable Storage and Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Temporary Storage and Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Indefinite Storage and Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MOORING - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Temporary and Mild Weather Mooring. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Long Term and Severe Weather Mooring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Temporary Stake Tie-Down Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Permanent Ground Anchor Construction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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RETURN TO SERVICE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flyable Storage Return to Service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Temporary Storage Return to Service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Indefinite Storage Return to Service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cold Soaked Airplane Return to Service. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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10 - CONTENTS © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL PARKING, MOORING, STORAGE AND RETURN TO SERVICE - GENERAL 1.
Scope A.
2.
This chapter describes and provides maintenance instructions for parking, storing, mooring and returning the airplane to service.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
Wheel Chocks
MANUFACTURER
USE
Available Commercially
To chock landing wheels.
Propeller Anchor Assembly
2606037-1
Cessna Aircraft Company Cessna Parts Distribution P.O. Box 949 Wichita, KS 67201
To secure propeller on airplanes thru 20800127 and thru 208B0054.
Propeller Anchor Assembly
2684002-1
Cessna Aircraft
To secure propeller on airplanes 20800128 and On, and 208B0055 and On.
Induction Air Inlet Cover
2601046-1
Cessna Aircraft
To prevent entry of moisture and/or foreign particles on airplanes thru 20800127 and thru 208B0054.
Induction Air Inlet Cover
2684001-1
Cessna Aircraft
To prevent entry of moisture and/or foreign particles on airplanes 20800128 and On, and 208B0055 and On.
Oil Cooler Inlet Cover
2601046-2
Cessna Aircraft
To prevent entry of moisture and/or foreign particles on airplanes thru 20800127 and thru 208B0054.
Oil Cooler Inlet Cover
2684001-2
Cessna Aircraft
To prevent entry of moisture and/or foreign particles on airplanes 20800128 and On, and 208B0055 and On.
Cessna Aircraft
To prevent entry of moisture and/or foreign particles in pitot tubes.
Pitot Tube Cover
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Static Ground Cable
Available Commercially
To static ground airplane.
Rope (0.375 inch diameter minimum or equivalent) (Refer to Note 1)
Available Commercially
To tie down wing and tail.
Available Commercially
To coat ni-cad battery intercell hardware during storage procedures.
Petroleum, Technical
Federal Specification VVP236
NOTE 1: Tie down ropes should be capable of resisting a pull of approximately 3,000 pounds. Nylon, polypropylene or dacron ropes are preferred over manila (hemp) rope. Manila rope has the disadvantage of shrinking when wet, is subject to rot and mildew and has considerably less tensile strength than synthetic fiber ropes, therefore requiring a larger diameter cord (0.5625 inch minimum).Manila rope and steel cable can fail due to chafing. Synthetic fiber ropes resist snapping because of greater elasticity than manila rope and steel cable. Manila rope demonstrates a greater elasticity than steel cable.Beginning with airplanes 20800077 and On, a system of tie down straps is offered that can be utilized to tie down the airplane and to tie down cargo within the airplane.There are two basic categories for tie down kits. The standard configuration with a 3,000 pound rating may be attached to any tie down point in the airplane. The heavy duty configuration with a 5,000 pound rating may only be attached to the aft passenger seat tracks. Any of the belt assemblies with a minimum rating of 3,000 pounds may be used to tie down the airplane. Refer to the 208 Parts Catalog for order information. 3.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief definition of the sections incorporated in this chapter is as follows: (1) The section on parking describes methods, procedures and precautions used when parking the airplane. (2) The section on storage provides information on recommended storage procedures. Recommendations vary with the length of time the airplane is to be stored. (3) The section on mooring describes procedures and equipment used to moor the airplane. (4) The section on return to service describes procedures used when returning the airplane to service from short or long term storage.
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MODEL 208 MAINTENANCE MANUAL PARKING - MAINTENANCE PRACTICES 1.
2.
General A.
The procedures that follow are instructions to park the airplane.
B.
The airplane must be moored if high winds are possible or anytime the airplane remains outside for extended periods of time. Refer to Mooring - Maintenance Practices, and refer to Storage Maintenance Practices for detailed instructions for short term or long term storage.
Parking Instructions A.
Hard Surface and Sod.
CAUTION: Any time the airplane has a heavy load, the wheel print pressure (pressure of the airplane wheels upon the contact surface of the parking area or runway) will be extremely high, and surfaces such as hot asphalt or damp sod may not correctly support the weight of the airplane. Caution must be taken to avoid the parking or moving the airplane on those surfaces. (1) (2)
Set the airplane on a level surface facing into the wind. Set the brake to park and chock the main gear wheels.
CAUTION: Do not set the brake in park during cold weather when moisture can freeze brakes, or when brakes are overheated. (3) (4) 3.
Install the control column lock. Set the rudder gust lock. Refer to the Pilot’s Operating Handbook and the FAA Approved Airplane Flight Manual for instructions.
Cold Weather Parking Instructions A.
Maintenance practices cover procedures used to park the airplane during low ambient air temperatures to -40°F (-40°C). NOTE:
Engine starting is possible at temperatures of -31°F (-35°C) and above. Hangar use is recommended at temperatures below this.
B.
Hangar use is recommended when ice, snow or heavy frost is possible. The use of a hangar is more economical and environmentally better than the use of a deicing service. If no hangar is available, be alert for snow, ice or frost on the wings and fuselage.
C.
Be alert for ice formation when you move the airplane from a warm hangar to snow conditions. When falling snow melts on the warm airplane skin, it may freeze, requiring deicing of the airplane.
D.
When you park the airplane for extended periods of time, make sure all water and galley liquids are removed.
CAUTION: Possible fuel expansion and overflow can result when the airplane fuel tanks are filled in cold temperatures, then the airplane is moved to a warm hangar. A fire hazard can result. CAUTION: Do not set the brake in park during cold weather. Moisture can freeze the brakes. E.
Set the airplane on a level surface facing into the wind or inside the hangar.
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CAUTION: When wheel chocks are used on the ice, the ice may cause them to move. Make sure the chocks are firmly set in position and tied together. F.
Chock the main gear wheels.
G.
Install covers for the engine, pitot and windshield.
H.
Remove the oxygen masks and personal gear.
I.
Measure the tire pressure with a tire gage. Make sure the inflation of the tires is correct. NOTE:
J.
The tires will have a lower pressure in cold weather.
Remove the battery from the airplane. NOTE:
If the battery is left in the airplane, regular maintenance will be required to prevent any discharge from the battery. If the battery is removed from the airplane, measure the battery charge.
K.
Install the control column lock.
L.
Set the rudder gust lock in position. Refer to the Pilot’s Operating Handbook and the FAA Approved Airplane Flight Manual for instructions.
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MODEL 208 MAINTENANCE MANUAL STORAGE - MAINTENANCE PRACTICES 1.
2.
3.
General A.
This maintenance practice provides instructions for flyable storage, temporary storage and indefinite storage. Included in these instructions are inspection criteria to be used during storage. The following definitions apply to storage times: (1) Flyable Storage - Flyable storage is defined as a maximum of 28 days nonoperational storage and/or the first 25 hours of intermittent engine operation. (2) Temporary Storage - Temporary storage is defined as a maximum of 90 days nonoperational status. (3) Indefinite Storage - Indefinite storage is defined as more than 90 days nonoperational status.
B.
Airplanes in storage should be returned to service using steps which are detailed in Return To Service - Maintenance Practices.
Flyable Storage and Inspection A.
Storage Procedures. (1) If the airplane will be out of service for five days or more, disconnect battery; and, as necessary, clean and coat the intercell hardware with a light coat of neutral nonconductive grease, such as petroleum jelly, to prevent corrosion. (2) If battery is left in airplane, regular servicing will be required to prevent discharge. If battery is removed from the airplane check it regularly for state of charge. (3) After two weeks, rotate airplane tires to prevent flat areas. Mark tires with tape to ensure tire is placed at a minimum of 90 degrees from previous position. (4) Do not set parking brake if a long period of inactivity is anticipated, as brake seizing can result. (5) Airplanes in storage 0 to 7 days - Engine may be left in an inactive state, with no preservation protection, provided engine is sheltered, humidity is not excessively high and engine is not subjected to extreme temperature changes that would produce condensation. (6) Airplanes in storage 8 to 28 days - Engines inactive for up to 28 days require no preservation, provided all engine openings are sealed off and relative humidity in engine is maintained at less than 40 percent. Humidity control is maintained by placing desiccant bags and humidity indicator on wooden racks in engine primary exhaust duct. Suitable windows must be provided in exhaust closure to facilitate observation of humidity indicators. (7) Ensure fuel bays are full of fuel.
B.
Inspection During Storage. (1) Inspection shall be carried out after 14 days of storage. If the relative humidity (as indicated on the humidity indicator) is less than 40 percent, no further action is required. If humidity indicated exceeds 40 percent, the desiccant bags must be replaced by freshly activated desiccant bags.
Temporary Storage and Inspection A.
Storage Procedures. NOTE:
(1) (2)
The airplane is constructed of corrosion- resistant alclad aluminum, which will last indefinitely under normal conditions if kept clean. However, these alloys are subject to oxidation. The first indication of corrosion on unpainted surfaces is in the form of white deposits or spots. On painted surfaces, paint is discolored or blistered. Storage in a dry hangar is essential to good preservation and should be procured, if possible. Varying conditions will alter measures of preservation, but under normal conditions in a dry hangar, and for storage periods not to exceed 90 days, the following methods of treatment are suggested:
Clean and wax airplane thoroughly. Lubricate all airframe items.
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MODEL 208 MAINTENANCE MANUAL (3)
Remove battery and store in a cool, dry place; service battery periodically and charge as required. NOTE:
(4) (5) (6) (7) (8) (9)
The airplane battery serial number is recorded in airplane equipment list. To assure accurate warranty records, battery should be reinstalled in same airplane from which it was removed. If battery is returned to service in a different airplane, appropriate record changes must be made and notification sent to the Cessna Warranty Administration Department.
Clean any oil or grease from tires, and coat tires with a tire preservative. Cover tires to protect against grease or oil. Either block up fuselage to relieve pressure on tires or rotate wheels every two weeks to prevent flat areas on tires. Mark tires with tape to ensure tire is placed approximately 90 degrees from previous position. Do not set the parking brake as brake seizing can result. Close fuel supply firewall shutoff valve. Disconnect fuel inlet line to oil-to-fuel heater and connect suitable oil supply line to oil-to-fuel heater fuel inlet. Blank off disconnected fuel supply line. Disconnect fuel line at inlet to flow divider to prevent oil from entering fuel manifold, and loosen line as required to permit drainage into a suitable container. NOTE:
An engine treated in accordance with the following may be considered being protected against normal atmospheric corrosion for a period not to exceed 90 days.
NOTE:
Engine preservation carried out during temporary or indefinite storage should be recorded in the engine logbook and on tags secured to the engine.
CAUTION: Under no circumstances should preservative oil be sprayed into the compressor or exhaust ports of the engine. Dirt particles deposited on blades and vanes during engine operation will adhere and alter the airfoil shape, adversely affecting compressor efficiency. CAUTION: Extreme care must be taken to prevent foreign material from being drawn into engine fuel system. Equipment must be supplied with suitable filters no coarser than 10 micron rating. CAUTION: Under no circumstances permit preservative oil to enter engine where it may come in contact with thermocouple probe assembly. Oil contamination of probes may cause complete failure of thermocouple system. (10) Supply preservative oil (MIL-PRF-6081, Grade 1010) at 5 to 25 PSIG pressure. temperature is at least 16°C (60°F).
Ensure
CAUTION: Observe starter motor operating limits (refer to Pilot's Operating Handbook and Approved Airplane Flight Manual). (11) With ignition switch in NORMAL position, IGN circuit breaker pulled, and fuel condition lever in HIGH IDLE position, and power control lever to MAX, carry out normal motoring run until all preservative oil is displaced. During run, power control lever should be moved from MAX to IDLE and returned to MAX, and fuel condition lever from HIGH IDLE to CUTOFF and returned to HIGH IDLE to displace fuel from system. (12) After motoring run, check to see if preservative oil is coming from opened fuel line. If not, repeat motoring cycle until preservative oil flows from opened fuel line.
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MODEL 208 MAINTENANCE MANUAL (13) Return power control lever to IDLE, and fuel condition lever to CUTOFF. Reconnect fuel supply line to oil-to-fuel heater and fuel line to inlet of flow divider. (14) Install all plugs, caps and covers over all openings to prevent entry of foreign material and accumulation of moisture. Install desiccant bags and humidity indicators. (15) Ensure fuel bays are full of fuel. B.
4.
Inspection During Storage. (1) Inspection shall be carried out every 14 days if the airplane is stored outside, or every 30 days if the airplane is stored inside. If the relative humidity (as indicated on the humidity indicator) is less than 40 percent, no further action is required. If humidity indicated exceeds 40 percent, the desiccant bags must be replaced by freshly activated desiccant bags. (2) Drain any accumulated moisture and contamination from all fuel drains every 30 days. Refer to Chapter 12, Fuel - Servicing. (3) Check fuel additive concentration every 30 days using a differential refractometer. Refer to the Pilot’s Operating Handbook and Approved Airplane Flight Manual for allowable concentration ranges. (a) If concentration falls below acceptable range, airplane must be defueled and refueled.
Indefinite Storage and Inspection A.
Storage Procedures. NOTE:
(1) (2) (3)
The airplane is constructed of corrosion- resistant alclad aluminum, which will last indefinitely under normal conditions if kept clean. However, these alloys are subject to oxidation. The first indication of corrosion on unpainted surfaces is in the form of white deposits or spots. On painted surfaces, paint is discolored or blistered. Storage in a dry hangar is essential to good preservation and should be procured, if possible. Varying conditions will alter measures of preservation, but under normal conditions in a dry hangar, and for storage periods greater than 90 days, the following methods of treatment are suggested:
Clean and wax airplane thoroughly. Lubricate all airframe items. Remove battery and store in a cool, dry place; service battery periodically and charge as required. NOTE:
(4) (5) (6) (7) (8) (9)
The airplane battery serial number is recorded in airplane equipment list. To assure accurate warranty records, battery should be reinstalled in same airplane from which it was removed. If battery is returned to service in a different airplane, appropriate record changes must be made and notification sent to the Cessna Warranty Administration Department.
Clean any oil or grease from tires, and coat tires with a tire preservative. Cover tires to protect against grease or oil. Either block up fuselage to relieve pressure on tires, or rotate wheels every two weeks to prevent flat areas on tires. Mark tires with tape to ensure tire is placed approximately 90 degrees from previous position. Do not set the parking brake as brake seizing can result. Close fuel supply firewall shutoff valve. Disconnect fuel inlet line to oil-to-fuel heater and connect suitable oil supply line to oil-to-fuel heater fuel inlet. Blank off disconnected fuel supply line. Disconnect fuel line at inlet to flow divider to prevent oil from entering fuel manifold, and loosen line as required to permit drainage into a suitable container. NOTE:
An engine treated in accordance with the following may be considered being protected against normal atmospheric corrosion for a period not to exceed 90 days.
NOTE:
Engine preservation carried out during temporary or indefinite storage should be recorded in the engine log book and on tags secured to the engine.
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CAUTION: Under no circumstances should preservative oil be sprayed into the compressor or exhaust ports of the engine. Dirt particles deposited on blades and vanes during engine operation will adhere and alter the airfoil shape, adversely affecting compressor efficiency. CAUTION: Extreme care must be taken to prevent foreign material from being drawn into engine fuel system. Equipment must be supplied with suitable filters no coarser than 10 micron rating. CAUTION: Under no circumstances permit preservative oil to enter engine where it may come in contact with thermocouple probe assembly. Oil contamination of probes may cause complete failure of thermocouple system. (10) Supply preservative oil (MIL-PRF-6081, Grade 1010) at 5 to 25 PSIG pressure. temperature is at least 16°C (60°F).
Ensure
CAUTION: Observe starter motor operating limits (refer to Pilot's Operating Handbook and Approved Airplane Flight Manual). (11) With ignition switch in NORMAL position, IGN circuit breaker pulled, and fuel condition lever in HIGH IDLE position, and power control lever to MAX, carry out normal motoring run until all preservative oil is displaced. During run, power control lever should be moved from MAX to IDLE and returned to MAX, and fuel condition lever from HIGH IDLE to CUTOFF and returned to HIGH IDLE to displace fuel from system. (12) After motoring run, check to see if preservative oil is coming from opened fuel line. If not, repeat motoring cycle until preservative oil flows from opened fuel line. (13) Return power control lever to IDLE, and fuel condition lever to CUTOFF. Reconnect fuel supply line to oil-to-fuel heater and fuel line to inlet of flow divider. (14) Place suitable container under engine and remove drain plugs from oil tank and accessory gearbox, and chip detector from propeller reduction gearbox. (15) With drains open, motor engine with starter (ignition NORMAL and IGN circuit breaker pulled) to permit scavenge pumps to clear engine, indicated by cessation of steady stream of oil from drains. To prevent excessive operation with limited lubrication, limit rotation to shortest possible time to accomplish complete draining. (16) Remove oil filter element and allow oil to drain. Refer to Pratt and Whitney Engine Maintenance Manual for procedures. (17) Allow oil to drain from engine to a slow drip (approximately one-half hour), then reinstall oil filter element and chip detector and close drains. (18) Remove cover plates from pads of accessory drives, and spray exposed surfaces and gear shafts with engine lubricating oil (Exxon Turbo Oil 2380 or equivalent). Replace cover plates. (19) Install plugs, caps, and covers over all openings to prevent entry of foreign material and accumulation of moisture. NOTE:
If engine is to remain in airplane, place desiccant bags on wooden racks in inlet and exhaust ducts.
(20) Tag oil filler cap with date of preservation, and enter date and type of preservation in engine log book. (21) Install humidity indicator in air inlet end and in exhaust end of engine compartment. Cover with suitable airtight moisture barrier. Provide inspection windows at each end for observation of humidity indicators. (22) Install all plugs, caps and covers over all openings to prevent entry of foreign material and accumulation of moisture. Install desiccant bags and humidity indicators. (23) Ensure fuel bays are full of fuel.
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MODEL 208 MAINTENANCE MANUAL B.
Inspection During Storage. (1) Inspection shall be carried out every 14 days if the airplane is stored outside, or every 30 days if the airplane is stored inside. If the relative humidity (as indicated on the humidity indicator) is less than 40 percent, no further action is required. If humidity indicated exceeds 40 percent, the desiccant bags must be replaced with freshly activated desiccant bags. (2) Drain any accumulated moisture and contamination from all fuel drains every 30 days. Refer to Chapter 12, Fuel - Servicing. (3) Check fuel additive concentration every 30 days using a differential refractometer. Refer to the Pilot’s Operating Handbook and Approved Airplane Flight Manual for allowable concentration ranges. (a) If concentration falls below acceptable range, airplane must be defueled and refueled.
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MODEL 208 MAINTENANCE MANUAL MOORING - MAINTENANCE PRACTICES 1.
General A.
Mooring procedures must be utilized when the airplane is to be parked for an extended period of time or during existing or expected bad weather.
CAUTION: Any time the airplane is loaded heavily, the footprint pressure (pressure of the airplane wheels upon the contact surface of the parking area or runway) will be extremely high, and surfaces such as hot asphalt or damp sod may not adequately support the weight of the airplane. Precautions should be taken to avoid airplane parking or movement on such surfaces.
2.
B.
The best protection against storm damage is to fly the airplane out of the impending storm area, provided there is sufficient time. The next best procedure is to secure the airplane in a storm-proof hangar or shelter. The last alternative is to adequately tie down the airplane.
C.
Three fixed mooring points are provided on the airplane. Two are located on the underside of the wings at the wing-strut intersect, and the third is located on the underside of the tailcone. On the Model 208, the tail skid serves as the mooring point; on the Cargomaster, 208B Super Cargomaster and 208B Passenger, a ring is furnished.
Temporary and Mild Weather Mooring A.
Mooring Procedures (Refer to Figure 201). (1) Position airplane on level surface headed into wind. (a) In fixed parking areas, use ground anchor points which are located outboard or aft of airplane mooring points. It may be necessary to use two parking spaces to get adequate spacing between ground anchor points.
CAUTION: Do not set parking brake during cold weather, when accumulated moisture may freeze brakes, or when brakes are overheated. (2) (3) (4)
Set parking brake or chock main gear wheels. Install control column lock. Set rudder gust lock in accordance with the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual.
CAUTION: Never attach mooring lines directly to struts. Use designated tie down rings to prevent possible damage to struts. (5)
Connect mooring lines to mooring rings and tail skid. A tie-down rope requires using a secure antislip knot such as the bowline or square knot. Refer to Figure 203 for antislip knot configurations. NOTE:
(6)
(7) (8)
During existing or expected gusty or high wind conditions, mooring lines should have slack taken out of them to prevent excessive movement of airplane resulting in high shock load on airplane and moorings.
Install the following protective covers (as required) to prevent entry of foreign material: (a) Induction air inlet cover. (b) Pitot tube cover. (c) Bypass air outlet cover. (d) Oil cooler air inlet cover. Secure propeller with propeller anchor assembly. Attach static ground cable securely to the tie-down ring on the wing and the ground anchor.
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Temporary and Mild Weather Mooring Figure 201 (Sheet 1)
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Temporary and Mild Weather Mooring Figure 201 (Sheet 2)
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3.
Long Term and Severe Weather Mooring A.
Mooring Procedures (Refer to Figure 202). (1) Position airplane on level surface headed into wind. (a) In fixed parking areas, use multiple ground anchor points for each mooring point on the airplane. Ensure that all ground anchor points are outboard (or aft) of airplane mooring points. It may be necessary to use two parking spaces to get adequate spacing between ground anchor points.
CAUTION: Do not set parking brake during cold weather, when accumulated moisture may freeze brakes, or when brakes are overheated. (2) (3) (4)
Set parking brake or chock main gear wheels. Install control column lock. Set rudder gust lock in accordance with the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual.
CAUTION: Never attach mooring lines directly to struts. Use designated tiedown rings to prevent possible damage to struts. (5)
Connect mooring lines to airplane mooring points using tie-down ropes. A tie-down rope requires using a secure antislip knot such as the bowline or square knot. Refer to Figure 203 for antislip knot configurations. NOTE:
(6)
(7) (8) 4.
During existing or expected gusty or high wind conditions, mooring lines should have slack taken out of them to prevent excessive movement of airplane resulting in high shock load on airplane and moorings.
Install the following protective covers (as required) to prevent entry of foreign material: (a) Induction air inlet cover. (b) Pitot tube cover. (c) Bypass air outlet cover. (d) Oil cooler air inlet cover. Secure propeller with propeller anchor assembly. Attach static ground cable securely to the tie-down ring on the wing and the ground anchor.
Temporary Stake Tie-Down Installation A.
Temporary Tie-Down Procedures on Sod Surfaces (Refer to Figure 204).
CAUTION: Stake driven tie- downs will often pull out when the ground becomes soaked from heavy and torrential rains. Wooden stakes also suffer from rot and reduced resiliency when subjected to severe moisture conditions. (1) (2)
Drive metal anchor stakes into ground to provide an approximate 45-degree angle between airplane mooring points and anchor stakes. Use multiple anchor stakes for each airplane mooring point. Refer to Figure 202 for approximate location of anchor stakes in relation to airplane mooring points.
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Long Term and Severe Weather Mooring Figure 202 (Sheet 1)
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Long Term and Severe Weather Mooring Figure 202 (Sheet 2)
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Tying Antislip Knots Figure 203 (Sheet 1)
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Anchor Fabrication Figure 204 (Sheet 1)
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Anchor Fabrication Figure 204 (Sheet 2)
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CAUTION: Never attach mooring lines directly to struts. Use designated tiedown rings to prevent possible damage to struts. (3)
Connect mooring lines to airplane mooring points using tie-down ropes. A tie-down rope requires using a secure antislip knot such as the bowline or square knot. Refer to Figure 203 for antislip knot configurations. NOTE:
5.
During existing or expected gusty or high wind conditions, mooring lines should have slack taken out of them to prevent excessive movement of airplane resulting in high shock load on airplane and moorings.
Permanent Ground Anchor Construction A.
Anchor Fabrication (Refer to Figure 204). NOTE: (1) (2) (3)
Tie down anchors should provide a minimum holding power of approximately 3,000 pounds.
Determine placement of anchors according to mooring requirements. (a) Three anchors are required for temporary and mild weather mooring. Refer to Figure 201 for approximate anchor locations. Six anchors are required for long term and severe weather mooring. Refer to Figure 202 for approximate anchor locations. Fabricate and embed ground anchors according to type of surface being used. Refer to Figure 204 for construction details.
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MODEL 208 MAINTENANCE MANUAL RETURN TO SERVICE - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Maintenance practices covers procedures used to return an airplane to service following storage. Procedures used to return an airplane to service are dependent upon how long the airplane has been in storage. This section provides instructions for airplanes in storage from 0 to 7 days (flyable storage), airplanes in storage from 8 to 28 days (flyable storage), airplanes in storage 90 days or less (temporary storage), and airplanes in storage for more than 90 days (indefinite storage).
Flyable Storage Return to Service A.
Airplanes in Storage From 0 to 7 Days. (1) Connect battery. (2) Perform a thorough preflight inspection.
B.
Airplanes in Storage From 8 to 28 Days. (1) Connect battery. (2) Remove desiccant bags and humidity indicator. (3) Ensure all previously sealed engine openings are reopened and unobstructed. (4) Perform a thorough preflight inspection.
Temporary Storage Return to Service A.
Airplanes in Storage 90 Days or Less. (1) Install and connect battery. (2) Remove all installed plugs, caps and covers. (3) Remove desiccant bags and humidity indicator. (4) Fill the engine oil tank with approved oil. Refer to Chapter 12, Replenishing - Description and Operation for approved oil.
CAUTION: Do not mix different brands, viscosities or types of oil when changing or replenishing oil between oil changes (refer to SB72-1: P & WC SBL 2001). NOTE:
The lubricating oil system does not require any depreservation procedures.
CAUTION: Under no circumstances permit preservative oil to enter engine where it may come into contact with thermocouple probe assembly. Oil contamination of probes may cause complete failure of indicating system. (5) (6) (7) (8)
Disconnect fuel line at flow divider inlet, then loosen line, as required, to permit preservative oil drainage into a suitable container. Connect airplane fuel supply. Open the fuel supply firewall shutoff valve. With ignition switch in NORMAL position and IGN circuit breakers disengaged, displace preservative oil from fuel system as follows. (a) Place power control lever to MAX position and fuel condition lever to HIGH IDLE. (b) Turn fuel boost pump ON.
CAUTION: Observe starter motor operating limits. Refer to the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. (c)
Perform normal motor run, during which time, move power control lever to IDLE and return to MAX, fuel condition lever to CUTOFF and return to HIGH IDLE, until clean fuel commences to flow from drain.
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MODEL 208 MAINTENANCE MANUAL Reconnect fuel inlet line to flow divider, tighten all connections, torque to 90 to 100 inch-pounds, and safety wire. Refer to Chapter 20, Safetying - Maintenance Practices. (10) Return power control lever to IDLE and fuel condition lever to CUTOFF. (11) Check percent of anti-icing additive in fuel using a differential refractometer. Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual for specific concentration levels. Defuel and refuel airplane if concentration levels are below acceptable levels. (12) Check brake fluid reservoir for proper fluid level. Refer to Chapter 12, Hydraulic Fluid - Servicing.
(9)
4.
Indefinite Storage Return to Service A.
Airplanes in Storage Longer Than 90 Days. (1) Install and connect battery. (2) Remove all installed plugs, caps and covers. (3) Remove desiccant bags and humidity indicator.
CAUTION: Do not mix different brands, viscosities or types of oil when changing or replenishing oil between oil changes (refer to SB72-1: P & WC SBL 2001). (4)
Fill the engine oil tank with approved oil. Refer to Chapter 12, Replenishing - Description and Operation. NOTE:
The lubricating oil system does not require any depreservation procedures.
CAUTION: Under no circumstances permit preservative oil to enter engine where it may come into contact with thermocouple probe assembly. Oil contamination of probes may cause complete failure of indicating system. (5) (6) (7) (8)
Disconnect fuel line at flow divider inlet, then loosen line, as required, to permit drainage of preservative oil into a suitable container. Connect airplane fuel supply. Open the fuel supply firewall shutoff valve. With ignition switch in NORMAL position and IGN circuit breakers disengaged, displace preservative oil from fuel system as follows. (a) Place power control lever to MAX position and fuel condition lever to HIGH IDLE. (b) Turn fuel boost pump ON.
CAUTION: Observe starter motor operating limits (refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual). (c)
Perform normal motor run, during which time, move power control lever to IDLE and return to MAX, fuel condition lever to CUTOFF and return to HIGH IDLE, until clean fuel commences to flow from drain. (9) Reconnect fuel inlet line to flow divider, tighten all connections, torque to 90 to 100 inch-pounds, and safety wire. Refer to Chapter 20, Safetying - Maintenance Practices. (10) Return power control lever to IDLE and fuel condition lever to CUTOFF. (11) Check percent of anti-icing additive in fuel using a differential refractometer. Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual for specific concentration levels. Defuel and refuel airplane if concentration levels are below acceptable levels. (12) Check brake fluid reservoir for proper fluid level. Refer to Chapter 12, Hydraulic Fluid - Servicing.
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5.
Cold Soaked Airplane Return to Service A.
Airplanes remaining at temperatures below -40°F(- 40°C) for two hours or longer are classified as cold soaked airplanes. This maintenance practice covers those procedures required to return a cold soaked airplane to service.
B.
Aircraft preheating. (1) Warm cabin and cockpit to prevent instrument fogging and window condensation following crew and passenger boarding. (2) Tank and under wing fuel drains must be drained frequently and thoroughly, as extremely cold temperatures reduce fuel/water solubility and super cools water particles in the fuel, increasing possibility of fuel system icing. (3) Apply heat to the engine, cabin and cockpit. A ground external power unit and/or preheat should be used for starting.
C.
Preheat the aircraft. (1) Preheat the engine by installing the engine oil cooler cover and directing hot air through the induction air inlet. (2) Preheat the cockpit and cabin by routing heater hoses through the pilot’s side window and the aft passenger/cargo door. (3) Retrieve a fuel sample and inspect for water globules. Sample until a clean fuel sample (free of water) has been attained. NOTE:
D.
It is possible for water to settle in the sump and freeze, obstructing the drain. Should this occur, apply heat until fuel flows freely. Maintain heat after fuel flow begins to ensure all particles have melted. Collect drainage in a clear clean container and inspect for water globules.
After the engines are running, allow sufficient time for instruments and avionics to warm prior to taxi.
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CHAPTER
PLACARDS AND MARKINGS
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CONTENTS PLACARDS AND MARKINGS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Interior and Exterior Placard and Decal Detailed Inspection . . . . . . . . . . . . . . . . . . . . .
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LIST OF TASKS 11-00-00-220
Interior and Exterior Placard and Decal Detailed Inspection
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MODEL 208 MAINTENANCE MANUAL PLACARDS AND MARKINGS - INSPECTION/CHECK 1.
General A.
This section has information about the interior and the exterior placards inspection.
Task 11-00-00-220 2.
Interior and Exterior Placard and Decal Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the placards, decals, and markings on the airplane.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do a Interior and Exterior Placard and Decal Detailed Inspection. (1) Examine the interior of the airplane, including the nose and aft baggage areas, for the installation of all necessary placards, decals, and markings. (a) For the necessary placards, decals, and markings, refer to the Model 208, Illustrated Parts Catalog or the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. (2) Examine the exterior of the airplane for the installation of all required placards, decals, and markings. (a) For the necessary placards, decals, and markings, refer to the Model 208, Illustrated Parts Catalog or the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. (3) Examine the airplane identification plate. (a) The identification plate is found on the left side of the stinger, Zone 330. Refer to the Model 208, Illustrated Parts Catalog and Chapter 6, Airplane Zoning - Description and Operation or the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual.
E.
Restore Access (1) None End of task
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12 CHAPTER
SERVICING
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CONTENTS SERVICING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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REPLENISHING - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Capacity Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Approved Fuel Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Oil Capacity Table. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SpeciÞed Synthetic Lubricating Oil Table (Ambient Temperature Above 0F). . . . . . . SpeciÞed Synthetic Lubricating Oil Table (Ambient Temperature 0F or Below) . . . . TKS Anti-Ice Fluid. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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HYDRAULIC FLUID - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic Brake System Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shimmy Damper Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear Strut Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL AND ENGINE OIL - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page 1 Page 1 Page 1 Page 1 Page 1 Page 5 Page 5 Page 6 Page 6
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FUEL - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety and Maintenance Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anti-Ice Additive as a Biocide . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aviation Fuel Additives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Checking Fuel in Wing Tank. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Defueling. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Purging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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ENGINE OIL SYSTEM - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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OXYGEN SYSTEM - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Charging Oxygen System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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TIRES - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-13-01 Page 301 12-13-01 Page 301 12-13-01 Page 301 12-13-01 Page 301
FLOODED LEAD-ACID BATTERY - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-14-01 Page 301 12-14-01 Page 301 12-14-01 Page 301 12-14-01 Page 301
SEALED LEAD-ACID BATTERY - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-14-02 Page 301 12-14-02 Page 301 12-14-02 Page 301 12-14-02 Page 301
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Page 1 of 3 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-14-03 Page 301 12-14-03 Page 301 12-14-03 Page 301 12-14-03 Page 301
VACUUM SYSTEM CENTRAL AIR FILTER - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-15-01 Page 301 12-15-01 Page 301 12-15-01 Page 301
FREON AIR CONDITIONING - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Evacuating. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Discharging/Charging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check Compressor Oil Level With Compressor on Airplane . . . . . . . . . . . . . . . . . . . . . Functional Leak Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-16-01 Page 301 12-16-01 Page 301 12-16-01 Page 301 12-16-01 Page 302 12-16-01 Page 303 12-16-01 Page 306 12-16-01 Page 307 12-16-01 Page 310
R134A AIR CONDITIONING - SERVICING (Airplanes 20800274 and On and 208B0655 and On) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Evacuating. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Discharging/Charging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check Compressor Oil Level With Compressor on Airplane . . . . . . . . . . . . . . . . . . . . . Functional Leak Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-16-02 Page 301 12-16-02 Page 301 12-16-02 Page 301 12-16-02 Page 302 12-16-02 Page 303 12-16-02 Page 303 12-16-02 Page 303 12-16-02 Page 307 12-16-02 Page 308
NOSE GEAR SHOCK STRUT - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Service Nose Gear Shock Strut . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-17-01 Page 301 12-17-01 Page 301 12-17-01 Page 301
SCHEDULED SERVICING - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-20-00 Page 1 12-20-00 Page 1 12-20-00 Page 1
LUBRICANTS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication Service Notes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Recommended Lubricants Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-21-00 Page 12-21-00 Page 12-21-00 Page 12-21-00 Page
BATTERY RECEPTACLE - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-21-01 Page 301 12-21-01 Page 301 12-21-01 Page 301
FLIGHT CONTROLS - SERVICING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Slot Lip Spoiler Pushrod Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Control Assembly Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication of Elevator Bell Cranks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication of Control Tube and Control Column Bearings . . . . . . . . . . . . . . . . . . . . . . Rudder Bar Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Flap Actuator Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Bell Crank Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-21-02 Page 301 12-21-02 Page 301 12-21-02 Page 303 12-21-02 Page 304 12-21-02 Page 304 12-21-02 Page 304 12-21-02 Page 304 12-21-02 Page 304 12-21-02 Page 304
LANDING GEAR - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Landing Gear Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-21-03 Page 301 12-21-03 Page 301 12-21-03 Page 301
12 - CONTENTS © Cessna Aircraft Company
1 1 1 2
Page 2 of 3 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL PROPELLER (HARTZELL) - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hartzell Propeller Lubrication. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-21-04 Page 301 12-21-04 Page 301 12-21-04 Page 301
ENGINE CONTROL ROD ENDS - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-21-05 Page 301 12-21-05 Page 301
EXTERNAL - CLEANING/PAINTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preventive Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Approved Windshield/Window Products . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning Windshield and Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Waxing and Polishing Windshield and Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Rain Repellent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aluminum Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Painted External Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Compressor Wash . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Compartment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wheels. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Deice Boots . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Panels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-22-01 Page 701 12-22-01 Page 701 12-22-01 Page 701 12-22-01 Page 701 12-22-01 Page 702 12-22-01 Page 703 12-22-01 Page 703 12-22-01 Page 703 12-22-01 Page 703 12-22-01 Page 703 12-22-01 Page 704 12-22-01 Page 704 12-22-01 Page 704 12-22-01 Page 704 12-22-01 Page 705 12-22-01 Page 705
INTERIOR - CLEANING/PAINTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Airplane Interior Cleaning Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-23-01 Page 701 12-23-01 Page 701 12-23-01 Page 701
UNSCHEDULED SERVICING - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Extreme Weather Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ground Power Receptacle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-30-00 12-30-00 12-30-00 12-30-00 12-30-00
DEICING/ANTI-ICING - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Deicing/Anti-Icing Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-31-00 Page 1 12-31-00 Page 1 12-31-00 Page 2
DEICING/ANTI-ICING - SERVICING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Approved Products . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Deicing/Anti-Icing Precautions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Type I Deicing Preparations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Type II, Type III and Type IV Anti-Ice Preparations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Deicing Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Do an Anti-Ice Procedure on the Airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Post-Application Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Post-Flight Clean Up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wheel Brake Deicing Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Deicing Boot Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Deicing Boot Preservation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-31-00 Page 301 12-31-00 Page 301 12-31-00 Page 301 12-31-00 Page 301 12-31-00 Page 306 12-31-00 Page 306 12-31-00 Page 307 12-31-00 Page 308 12-31-00 Page 308 12-31-00 Page 309 12-31-00 Page 309 12-31-00 Page 309 12-31-00 Page 309
TKS ANTI-ICE SYSTEM - SERVICING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Deicing Porous Panel Care. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
12-31-10 Page 301 12-31-10 Page 301 12-31-10 Page 301 12-31-10 Page 304
12 - CONTENTS © Cessna Aircraft Company
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Page 3 of 3 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 12-10-01-610
Hydraulic Brake System Servicing
12-10-01 Page 301
12-10-01-611
Shimmy Damper Servicing
12-10-01 Page 301
12-21-03-640
Landing Gear Lubrication
12-21-03 Page 301
12-21-04-640
Hartzell Propeller Lubrication
12-21-04 Page 301
12 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SERVICING - GENERAL 1.
Scope A.
2.
This chapter provides instructions for the replenishment of fluids, scheduled and unscheduled servicing applicable to the entire airplane. Personnel shall observe safety precautions pertaining to the individual servicing application.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief description of each section follows. (1) The section on replenishing is subdivided into categories to group servicing information such as systems requiring hydraulic fluid or compressed gas. A brief description of the subdivision subjects follows. (a) Replenishing charts for the liquids most commonly used to service the airplane are grouped together to aid maintenance personnel in servicing. (b) The subdivision of fuel and oil provides maintenance personnel with general servicing procedures. Safety precautions and servicing procedures required by federal and local regulations may supersede the procedures described. (c) The subject on hydraulic fluid servicing provides servicing procedures for the airplane hydraulic brake system, nose gear shimmy damper and nose gear strut. (d) The remaining subject subdivisions provide service information on either a system, an assembly or a component. (2) The section on scheduled servicing includes lubrication information, external cleaning and internal cleaning. The section is subdivided to provide individual system, assembly or component service information. (3) The section on unscheduled servicing provides information on deicing an airplane or portions of an airplane.
12-00-00 © Cessna Aircraft Company
Page 1 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL REPLENISHING - DESCRIPTION AND OPERATION 1.
General A.
2.
3.
This section gives maintenance personnel the service information to replenish fuel, oil, and anti-ice ßuid.
Description A.
For an illustration of service points located on the airplane, refer to Figure 1. This illustration can be used in conjunction with replenishing tables to help maintenance technicians in servicing the airplane.
B.
The following tables are provided to establish replenishment capacities of various systems: (1) Fuel Capacity (Table 1) (2) Approved Fuels (Table 2) (3) Engine Oil Capacity (Table 3) (4) Synthetic Lubricating Oil For Ambient Temperature Above 0°F (Table 4) (5) Synthetic Lubricating Oil For Ambient Temperature 0° or Below (Table 5) (6) Approved TKS Anti-Ice Fluid (Table 6)
Fuel Capacity Table A.
The following table lists airplane fuel capacity. NOTE:
Total fuel and usable fuel quantities are based on 6.75 pounds per gallon.
WARNING: Only aviation grade fuels are approved for use. Table 1. Fuel Capacity SYSTEM
U.S.
IMPERIAL
METRIC
Fuel Capacity (Beginning with Airplanes 20800130 and 208B0089 when modiÞed per SK208-52)
335.6 Gallons
279.45 Gallons
1270.3 Liters
Fuel Capacity (Airplanes Thru 20800129 and 208B0088 when not modiÞed per SK208-52)
335.0 Gallons
278.95 Gallons
1268.0 Liters
Usable Fuel
332.0 Gallons
276.25 Gallons
1257 Liters
4.
Approved Fuel Table A.
The following table lists approved fuels for use in the airplane.
CAUTION: Aviation gasoline may be used for a maximum of 150 hours between engine overhauls; or a mixture of one part aviation gasoline and three parts of Jet A, Jet A-1 or JP-5 may be used for a maximum of 450 hours between engine overhauls. NOTE:
Fuels must comply with Pratt & Whitney Engine Service Bulletin number 1244 and all supplements and revisions.
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Airplane Service Points Figure 1 (Sheet 1)
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Airplane Service Points Figure 1 (Sheet 2)
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Airplane Service Points Figure 1 (Sheet 3)
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Table 2. Approved Fuels (NOTE 1) TYPE OF FUEL
SPECIFICATION
TYPE
Jet A
ASTM-D-1655
Jet A-1
ASTM-D-1655
Jet B
ASTM-D-1655
JP-1
MIL-L-5616
JP-4
MIL-T-5624
(NATO F-40)
JP-5
MIL-T-5624
(NATO F43 or F44)
JP-8
MIL-T-83133A
(Russian Type)
RT
GOST-10227-86
(Russian Type)
TS-1
GOST-10227-86
(Russian Type)
Alternate Emergency Fuel (Refer to Caution)
All grades of military and commercial aviation gasoline.
NOTE 1: Fuel used must contain anti-icing fuel additive in compliance with MIL-DTL-27686 (EGME), MIL-DTL85470, (DIEGME), or Phillips PFA 55 MB. 5.
Engine Oil Capacity Table A.
The following table lists oil capacity for the airplane.
WARNING: The U. S. Environmental Protection Agency advises mechanics and other workers who handle oil to minimize skin contact with used oil and promptly remove used oil from skin. In a laboratory study, mice developed skin cancer after skin was exposed to used engine oil twice a week without being washed off, for most of their span. Substances found to cause cancer in laboratory animals may also cause cancer in humans. Table 3. Engine Oil Capacity SYSTEM
U.S. QUARTS
IMPERIAL QUARTS
METRIC LITERS
Oil Capacity (total with Þlter, oil cooler and cooler hoses)
14.0 quarts
11.66 quarts
13.25 liters
Oil Tank Capacity
9.5 quarts
7.92 quarts
8.99 liters
6.
SpeciÞed Synthetic Lubricating Oil Table (Ambient Temperature Above 0°F) A.
The following table lists approved synthetic lubricating oil for ambient temperatures above 0°F (-18°C).
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CAUTION: Do not mix viscosities. CAUTION: Because of engine manufacturer recommendations, only change from an existing lubricant formulation to a “third generation” lubricant formulation when an engine is new or freshly overhauled. For additional information on use of third generation oils, refer to Engine Manufacturer’s Pertinent Oil Service Bulletins. Table 4. Synthetic Lubricating Oils for Ambient Temperature Above 0°F BRAND
SUPPLIER
Engine Oil
Refer to the Pratt & Whitney Maintenance Manual, P/N 3043512, Chapter 72-00 for approved oil types. (NOTE 1)
NOTE 1: Engine lubricating oils must comply with Pratt & Whitney Engine Service Bulletin number 1001 and all supplements or revisions. Do not mix brands unless speciÞcally approved. 7.
SpeciÞed Synthetic Lubricating Oil Table (Ambient Temperature 0°F or Below) A.
The following table lists approved synthetic lubricating oil for ambient temperatures 0°F (-18°C) or lower.
CAUTION: Do not mix viscosities. CAUTION: Because of engine manufacturer recommendations, only change from an existing lubricant formulation to a “third generation” lubricant formulation when an engine is new or freshly overhauled. For additional information on use of third generation oils, refer to Engine Manufacturer’s Pertinent Oil Service Bulletins. Table 5. Synthetic Lubricating Oils for Ambient Temperature 0°F or Below BRAND (TYPE II)
SUPPLIER
Engine Oil
Refer to the Pratt & Whitney Maintenance Manual, P/N 3043512, Chapter 72-00 for approved oil types. (NOTE 1)
NOTE 1: Engine lubricating oils must comply with Pratt & Whitney Engine Service Bulletin number 1001 and all supplements or revisions. Do not mix brands unless speciÞcally approved. 8.
TKS Anti-Ice Fluid A.
Designations and manufacturers of approved TKS anti-ice ßuid are shown in the table that follows: NOTE:
Ice protection ßuid must meet the AL5 (DTD406B) speciÞcations. You can mix ßuids that meet these speciÞcations in the anti-ice ßuid tank in any proportion. Refer to Figure 1.
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WARNING: Do not use automobile anti-freeze ßuids because they can contain additives that are harmful to the membranes in the porous panels or to other system components. Also, you are not permitted to use thickened deice ßuid for runway or parked aircraft deicing. Table 6. TKS Anti-Ice Fluid SYSTEM Windshield and Surface Reservoir
U.S. GALLONS 20.92
IMPERIAL GALLONS 17.45
METRIC LITERS 79.067
NAME, NUMBER OR TYPE AL 5 Canyon Industries P.O. Box 26447 Tempe, AZ 85282 AEROSHELL COMPOUND 07 Shell Oil Company One Shell Plaza Houston, TX 77001 BP AERO DEICING 2 BP Oil Limited BP House Victoria Street London SWIE 5NJ UK AVL-TKS Aviation Laboratories 5401 Mitchelldale B6 Houston, TX 77092
NOTE 1: All anti-icing ßuids that are in accordance with British Deicing Fluid SpeciÞcation DTD 406B (NATO Symbol S-745) are approved.
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MODEL 208 MAINTENANCE MANUAL HYDRAULIC FLUID - SERVICING 1.
General A.
This section gives procedures to fill with fluid, the components that follow: hydraulic brake system, nose gear shimmy damper, and nose gear strut.
Task 12-10-01-610 2.
Hydraulic Brake System Servicing NOTE:
The hydraulic brake system uses two brake cylinders found forward of the pilot's rudder pedals. A hydraulic fluid reservoir on the engine side of the firewall supplies fluid to both cylinders. Bleed the brake system when there is a spongy response to brake pedals. Refer to Chapter 32, Wheels and Brakes - Maintenance Practices.
CAUTION: Make sure to release the parking brake before the start of any servicing of the master cylinder. This will release the pressure in the system. Service the Hydraulic Brake Fluid Reservoir. (1) Open the upper left cowling door. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. (2) Remove the filler cap from the brake fluid reservoir. (3) Visually do a check of the fluid level in the reservoir. (a) If the reservoir level is approximately half full or less, fill the reservoir with MIL-PRF-5606 hydraulic fluid to within 0.75 inch of the vent hole that is 0.098 inch in diameter. (4) Install the filler cap on the reservoir. (5) Close the upper left cowling door. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. End of task A.
Task 12-10-01-611 3.
Shimmy Damper Servicing NOTE:
A.
There are two different types of shimmy dampers that have different servicing procedures. The two types of shimmy dampers are Cessna and Lord. The Lord shimmy damper does not have field servicing procedures.
Fill the Cessna Shimmy Damper with fluid. NOTE:
(1) (2) (3) (4) (5) (6) (7) (8)
The shimmy damper barrel is filled with MIL-PRF-5606 hydraulic fluid. A filler plug is on the top of the damper barrel. For servicing instructions that include more than filling the shimmy damper with fluid, refer to Shimmy Damper - Maintenance Practices.
Remove the nose gear fairings to get access to the shimmy damper. Refer to Nose Gear Fairing - Maintenance Practices. Remove the left upper cowling door and the left lower cowl panel to get access to the shimmy damper. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Remove the safety wire from the filler plug. Remove the filler plug from the shimmy damper. Visually do a check of the position of the piston. If it is necessary to add fluid to the shimmy damper, then fill the shimmy damper with MIL-PRF5606 hydraulic fluid. Install the filler plug in the shimmy damper. Install the safety wire on the filler plug.
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Install the left upper cowling door and the left lower cowl panel. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (10) Install the nose gear fairings. Refer to Nose Gear Fairing - Maintenance Practices. B.
Service the Lord Shimmy Damper. (1) The Lord Shimmy Damper is sealed and not serviceable. End of task 4.
Nose Gear Strut Servicing A.
Fill the nose gear strut with MIL-PRF-5606 hydraulic fluid. For servicing instructions, refer to Nose Gear Shock Strut - Servicing.
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MODEL 208 MAINTENANCE MANUAL FUEL AND ENGINE OIL - DESCRIPTION AND OPERATION 1.
General A.
2.
Description A.
3.
4.
This portion of the section on replenishing provides maintenance personnel with servicing procedures on the airplane fuel system and the engine oil system.
This portion of replenishing section is subdivided into fuel system and engine oil system. A brief description of the system follows. (1) The fuel system servicing procedures include adding fuel, mixing anti-icing additives to the fuel, checking anti-icing concentration in fuel tanks, defueling procedures and purging fuel storage areas. (2) The engine oil system servicing procedures provides information on adding oil to the engine, draining oil from the engine and descriptive information on synthetic turbine engine oil.
Fuel Precautions A.
Safety Precautions. (1) The safety precautions on fueling and defueling may be superseded by local directives. However, following is a typical list of precautions. (a) Ground, by designated grounding cables, the fueling and/or defueling vehicle to the airplane. Also, a static ground device shall contact the fueling or defueling vehicle and ground. (b) Fire fighting equipment shall be immediately available. (c) Wear proper clothing. Do not wear clothing that has a tendency to generate static electricity such as nylon 1 or synthetic fabrics. Do not wear metal taps on shoes when working in areas where fuel fumes may 2 accumulate at ground level. (d) The airplane shall be in a designated fuel loading or unloading area. (e) High wattage, pulse transmitting avionics equipment shall not be operated in the immediate vicinity.
B.
Maintenance Precautions. (1) Use designated equipment for fuel loading and unloading to prevent contamination. (2) Use proper procedures when adding fuel inhibitors. (3) Use specified type of fuel.
Oil Precautions A.
Maintenance Precautions. (1) Use specified synthetic turbine engine oil. (2) Use servicing procedures; do not overfill, do not mix manufacturers brand of oils.
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MODEL 208 MAINTENANCE MANUAL FUEL - SERVICING 1.
General A.
Fuel. (1) The fuel used in the airplane must have an anti-icing additive incorporated, or must be added to the fuel when the tanks are Þlled. The fuel may contain anti-icing and biocidal additives, if desired.
CAUTION: Lack of anti-icing additive may cause fuel Þlter or line icing and subsequently engine ßameout. Lack of anti-icing additive may also cause growth of fungi in the fuel tanks. (2) (3)
2.
When preblended fuel is available, add fuel to tanks as described. The wing fuel tank capacities and acceptable fuel speciÞcations are shown in Replenishing - Description and Operation. Mixing of anti-icing additives is accomplished as fuel is added to the tank. Procedures for mixing are identical for both the left and right wing tanks.
B.
Fuel Tanks. (1) An area of each wing, from WS 53.00 to WS 214.30, is sealed to form an integral fuel bay. Fuel bays must be Þlled after each ßight to lessen the possibility of condensation in fuel bays and lines. The fuel Þller caps are located on top of the wings, forward of the spoilers. For additional information on fuel tanks, refer to Chapter 28, Fuel Storage - Description and Operation. (2) Plugs or caps must be installed on lines, hoses, and Þttings to prevent thread damage, residual fuel drainage, and contamination.
C.
Fuel Drains. (1) Fuel drains are provided at various locations throughout the fuel system for drainage of water and sediment from the fuel system. For fuel drain locations and maintenance of fuel system, refer to Chapter 28, Fuel - General. To activate drain valves, a fuel sampler cup screwdriver is provided. The fuel must be checked before the Þrst ßight of the day and each refueling. (2) Place fuel sampler cup to valve; depress valve with Phillips head end of rod protruding from cup.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following listed items:
NAME
NUMBER
MANUFACTURER
USE
HI-FLO Prist (Aerosol Cans)
PFA-55MB (MILDTL-27686)
PPG Industries, Inc 1 Gateway Center Suite 6 South Pittsburgh, PA 15222
Fuel system icing inhibitor.
LO-FLO Prist (Aerosol Cans)
PFA-55MB (MILDTL-27686)
PPG Industries, Inc.
Fuel system icing inhibitor.
Prist (Bulk)
PFA-55MB (MILDTL-27686)
PPG Industries, Inc.
Fuel additive (to be used with proportioner PRB-101).
Prist Hi-Flash
MIL-DTL-85470
PPG Industries, Inc.
Fuel system icing inhibitor.
Proportioner (Dispensers)
PRB-101
Quannah Corp. 6713 Pharoah Drive Corpus Christi, TX 78412
To dispense fuel inhibitors.
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NAME
NUMBER
MANUFACTURER
USE
Anti-Ice Concentration Test Kit
CJMD 128-002
Cessna Aircraft Co. Citation Marketing Div. Department 579 P.O. Box 7706 Wichita, KS 67277
To test MIL-DTL-27686 anti-ice additive to fuel concentration.
Anti-Ice Concentration Test Kit
HB-P-C B/2
Gammon Tech Product 2300 Hwy 34 Manasquan, NJ 08736
To test anti-ice additive to fuel concentration in fuel (will test both MIL-DTL-27686 and MIL-DTL-85470).
Anti-Static Additive
Shell ASA3
Royal Lubricants Co. P. O. Box 518 Hanover, NJ 07936
Antistatic fuel additive.
Anti-Static Additive
Dupont Stadis 450
Dupont
Antistatic fuel additive.
Biocidal Protection Additive
Sohio Biobor JF
Sohio Engineered Materials Co. Refractories Division 3425 Hyde Park Blvd. P. O. Box 664 Niagra Falls, NY 14302
Biocidal protection additive.
Biocidal Protection Additive
KATHON FP 1.5
Fuel Quality Services, Inc . 309 Mohawk Drive Rotterdam Junction , NY 12150
Biocidal protection additive.
Anti-Icing Additive
Fluid I
Anti-Icing Additive
Anti-Icing Additive
Fluid I-M
Anti-Icing Additive
Anti-Static Additive
Sigbol
Anti-Static Additive
NOTE:
3.
When using RT and TS-1 fuels with Anti-Icing additives Fluid I and Fluid I-M, maintenance requirements of Pratt and Whitney Canada Service Bulletin SB1244, Revision 17 or latest revision must be obeyed.ye
Safety and Maintenance Precautions A.
Safety Precautions.
WARNING: During all fuel system servicing procedures, Þre Þghting equipment must be available. Two ground wires from tiedown rings on the airplane to approved ground stakes must be used to prevent accidental disconnection of one ground wire. Make sure battery switch is turned off, unless otherwise speciÞed. (1)
(2)
Ground the fueling/defueling equipment (vehicle or fuel hydrant equipment) to the airplane with designated grounding cable(s). Make sure the fueling/defueling equipment is grounded to an approved static ground. Ground the airplane to an approved static ground with grounding cable. Ground fuel nozzle to the tie down lug on the wing strut. Ground airplane as follows: (a) Ground airplane Þrst. (b) Ground vehicle (or hose cart) to the same ground as the airplane. (c) Bond vehicle (or hose cart) to airplane. (d) Bond refuel nozzle to airplane. Make sure Þre Þghting equipment is positioned and immediately available.
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6) B.
4.
Do not wear clothing that has a tendency to generate static electricity such as nylon or synthetic fabrics. Do not wear metal taps on shoes. The airplane must be in a designated fuel loading/unloading area. High wattage, pulse transmitting avionics equipment must not be operated in the vicinity of the fueling/defueling operation.
Maintenance Precautions. (1) Use designated equipment for fuel loading/unloading to prevent contamination. (2) Due to the chemical composition of anti-ice additive, improper blending of fuel and anti-icing additive may cause the deterioration of the integral fuel tanks interior Þnish, thus promoting corrosion. It is very important that proper anti-ice additive blending procedures be followed. (3) Use authorized type of fuel and anti-ice additive. (4) During defueling, Make sure anti-ice additive blended fuel and unblended fuel are not mixed.
Anti-Ice Additive as a Biocide A.
In addition to preventing icing in fuel tanks, anti-ice additive effectively controls the growth of bacterial and fungal microorganisms which can form in fuel storage tanks. (1) Bacterial and fungal microorganisms multiply where water and fuel interface. Because the weather, temperature and climate differ where a particular airplane is based and operated, the amount of water condensation in the fuel tank varies. (2) Microbiological contamination can be an expensive and potentially dangerous condition. This type of contamination is related to water which gravitates to low points in fuel reservoirs and is not circulated or removed. Airborne spores Þnd their way into the fuel tanks and migrate to the water which they utilize as a growth medium while feeding off the hydrocarbon fuel. The Þrst indication of microbiological contamination is a light grayish slime. Heavy contamination will be a thick grey, Þbrous formation which may contain black masses of decay products. If the contamination is left unchecked, it can eventually move as a mass and block the fuel system and/or cause corrosion. (3) Examination of the fuel tank for bacterial and fungal microorganisms requires opening areas of the fuel tank and checking where trapped water may exist, such as the lower corners near wing ribs. Also check internal screens at ßapper valve openings into the sump area for bacterial and fungal microorganisms which have formed a mass and may be caught on the screen during their movement.
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5.
Aviation Fuel Additives
WARNING: Ethylene Glycol Monomethyl Ether (EGME) MIL-DTL-27686 anti-ice additive is toxic and dangerous to health when it is breathed and/or absorbed into the skin. When you put an anti-ice additive into fuel in an unventilated area, use appropriate personal protective equipment such as eye goggles/shield, a respirator with organic vapor cartridges, non-absorbing gloves, and other personal protective equipment to protect skin from the anti-ice additive. If anti-ice additive enters the eyes, ßush them with water and contact a physician immediately. CAUTION: Diethylene Glycol Monomethyl Ether (DIEGME) MIL-DTL-85470 is slightly toxic if you swallow it. It can cause eye redness, swelling, and irritation. It is also combustible. Before you use this material, refer to all safety information on the container. Make sure the additive is directed into the ßowing fuel stream. Start the additive ßow after the fuel ßow starts and stop the additive ßow before the fuel ßow stops. Do not allow concentrated additive to contact the coated interior of the fuel tank or the airplane painted surface. A.
When servicing fuel with anti-icing additives containing ethylene glycol monomethyl ether (EGME, MIL-DTL-27686) or diethylene glycol monomethyl ether (DIEGME, MIL-DTL-85470), remember that they are harmful if inhaled, swallowed or absorbed through the skin, and will cause eye irritation. Also, they are combustible. Before using this material, refer to all safety information on the container.
B.
EGME is toxic under sustained exposure environments. When inhaled, EGME is primarily a central nervous system depressant, although various animal studies have revealed that acute inhalation overexposure may cause kidney injury. The primary symptoms of inhalation overexposure in conÞned or poorly ventilated areas include headache, drowsiness, blurred vision, weakness, lack of coordination, tremor, unconsciousness and even death. When ingested (swallowed) in massive doses, EGME is reported to exhibit a narcotic action, but at lower dosage levels, death is delayed and is accompanied by lung edema (excessive serious ßuid in lungs), slight liver injury and marked kidney injury. EGME is only mildly irritating to the eyes and skin; however, it can be readily absorbed through the skin in toxic amounts. Symptoms of overexposure due to skin absorption are essentially the same as those outlined for inhalation.
C.
In cases of acute exposure, DIEGME is an eye and mucous membrane irritant, a nephrotoxin and central nervous system depressant. It is toxic by skin absorption. Inhalation may cause irritation to mucous membranes, although, due to it’s low volatility, this is not an extreme hazard at room temperature or below. If DIEGME contacts the eye, it may cause pain and transient injury. It is absorbed through the skin in toxic amounts.
D.
In the event EGME or DIEGME contact is experienced, the following emergency and Þrst aid procedures must be used. (1) If EGME or DIEGME is inhaled, remove person to fresh air. If the person is not breathing, give artiÞcial respiration, preferably mouth-to-mouth; however, if breathing is difÞcult, administer oxygen. Always call a physician. (2) If ingested (swallowed), drink large quantities of water. Then induce vomiting by placing a Þnger far back into the throat. Contact a physician immediately. If vomiting cannot be induced, take victim immediately to the hospital or a physician. If victim is unconscious or in convulsions, take victim immediately to the hospital or a physician. Do not induce vomiting or give anything by mouth to an unconscious person.
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MODEL 208 MAINTENANCE MANUAL (3)
E.
6.
If eye or skin contact is experienced, ßush with plenty of water (use soap and water for skin) for at least 15 minutes while removing contaminated clothing and shoes. Call a physician. Thoroughly wash contaminated clothing and shoes before reuse.
Additional antistatic and biocidal protection may be provided using approved products. Refer to the Tools, Equipment and Materials section for approved manufacturers. Refer to the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual for concentration levels of these products.
Fuel Loading
CAUTION: Make sure the proper grade and type of fuel is used to service the airplane. Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual for a list of approved fuels. A.
Approved fuels for the Model 208 airplane may or may not contain an anti-ice additive. The additive incorporates a biocidal chemical which prevents fungi/living organisms in the fuel storage reservoirs. Mixing anti-ice additive and fuel during refueling involves the utilization of an aerosol can or proportioner. Refer to Tools, Equipment and Materials for approved mixing devices.
B.
Mixing Icing Inhibitor Procedures. NOTE: (1) (2)
Equivalent procedures may be substituted.
When using proportioner follow directions provided. When using aerosol cans, utilize the following procedures. (a) Insert the fueling nozzle and fuel additive nozzle into the fuel Þller.
WARNING: Ethylene Glycol Monomethyl Ether (EGME) MIL-DTL-27686 anti-ice additive is toxic and dangerous to health when it is breathed and/or absorbed into the skin. When you put an anti-ice additive into fuel in an unventilated area, use appropriate personal protective equipment such as eye goggles/shield, a respirator with organic vapor cartridges, non-absorbing gloves, and other personal protective equipment to protect skin from the anti-ice additive. If anti-ice additive enters the eyes, ßush them with water and contact a physician immediately. CAUTION: Diethylene Glycol Monomethyl Ether (DIEGME) MIL-DTL85470 is slightly toxic if you swallow it. It can cause eye redness, swelling, and irritation. It is also combustible. Before you use this material, refer to all safety information on the container. Make sure the additive is directed into the ßowing fuel stream. Start the additive ßow after the fuel ßow starts and stop the additive ßow before the fuel ßow stops. Do not allow concentrated additive to contact the coated interior of the fuel tank or the airplane painted surface. (b)
C.
Start refueling; then, direct the fuel additive into the fuel stream to blend the additive simultaneously with the fuel as it Þlls the tank. The additive concentration range must be maintained in accordance with instructions in the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual.
Tank Filling Procedure.
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WARNING: Perform fuel loading in area which permits free movement of Þre equipment. WARNING: Make sure that the fuel supply unit is grounded, and that the ground to airplane is connected. (1) (2) (3)
Connect fueling nozzle ground to the tie down lug on airplane strut. Place a protective pad on the wing adjacent to the fuel Þller and remove the Þller cap. Service as follows: (a) If the turbine fuel has the fuel system icing inhibitor added, Þll wing tanks. (b) If the turbine fuel does not have fuel system icing inhibitor added, select an inhibitor. Refer to Tools, Equipment and Materials for approved icing inhibitors; and add in accordance with Mixing Icing Inhibitor Procedures.
CAUTION: Make sure Þller cap is secured. (4) 7.
Remove fuel nozzle; remove protective pad; disconnect fueling nozzle ground; install fuel Þller cap.
Checking Fuel in Wing Tank A.
Fuel Samples. (1) Sampling of fuel, and checking and draining sediment from the tanks, are the main purposes of the poppet-type drain valves installed on the lower side of the fuel tank. The valves are installed mainly in the vicinity of the fuel tank sump area. (2) The poppet-type valve is a spring-loaded poppet, housed in the drain valve body. The poppet is spring-loaded in the closed position. A slot in the end of the poppet allows for screwdriver operation. To operate the valve, depress the slot end to open valve and rotate to lock the valve to the open position. Depress, rotate and release, slot end will also set the valve to a closed position. (3) During cold weather, if more than one hour elapses between removal of airplane from a heated shelter and takeoff, all fuel sumps must be drained through the drain valves during the preßight inspection. Enough fuel must be drained from each drain point to make sure that the fuel is free from water and/or other contaminants. At least 30 minutes must elapse between fueling and checking for contamination. The fuel must be drained into a clear, clean container suitable to permit a careful visual examination for water and other contaminants. To aid in distinguishing water from fuel, add one or two drops of water soluble food coloring in the container prior to draining fuel samples. The food coloring will mix readily with water but not with fuel.
B.
MIL-DTL-27686 Anti- Ice Additive Concentration Check using CJMD128-002 Anti-Ice Concentration test kit. NOTE:
Refer to Tools, Equipment and Materials list in the test kit.
NOTE:
When you add anti-icing additive to fuel which does not contain the additive and/or to determine if the anti-icing additive concentration has fallen outside the limits speciÞed in the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual, do the following check.
(1) (2) (3) (4) (5) (6)
Get a sample of the fuel to be tested. Fill a beaker with approximately 250 ml of water (tap water is acceptable). Put the beaker on a hot plate and bring to a full boil. Attach the repeating pipet Þller to the 10-ml transfer pipet and adjust to the 10 ml mark. Transfer 10-ml of distilled water into a clean ampoule. Transfer 10-ml of the fuel test sample to the ampoule with the 10-ml of water. Screw the cap on the ampoule and shake the tube for two minutes (use the timer). NOTE:
This extracts the anti-ice additive from the fuel.
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Allow the ampoule to stand undisturbed for at least two minutes (use the timer). NOTE:
The fuel and water/anti-ice additive will separate into two separate phases or layers.
CAUTION: The contents of the ampoule are a strong acid. Do not let the contents touch your skin. If they do touch the skin, wash skin with running water for 15 minutes and get immediate medical attention. (8) (9) (10) (11) (12) (13) (14)
Snap off the top of the glass ampoule and empty the potassium dichromate/sulfuric acid solution into a clean ampoule. Do not discard the empty ampoule. With a clean pipet, add a few drops (not over 2 ml) of distilled water to the ampoule. Empty the rinse solution into the ampoule with the acid. Discard the empty ampoule. Attach the repeating pipet Þller to the 5 ml transfer pipet and adjust to the 5 ml mark. Carefully withdraw 5 ml of the bottom (water/anti-ice additive) phase from the ampoule of fuel and water. Make sure that none of the fuel phase is transferred. Empty the pipet with the water/anti-ice additive into the ampoule with the potassium dichromate/ sulfuric acid solution. Thoroughly mix the acid-water solution by swirling it carefully. Do not cap the ampoule. Immediately place the ampoule in the boiling water bath (beaker on the hot plate) for 10 minutes, +30 or -30 seconds, using the timer for control. Acid-water solution may chemically react, which will create erroneously high results. NOTE:
If the acid-water solution cannot be immediately placed in the boiling water, it must be maintained in an ice water bath until just prior to heating. Otherwise, the acid-water solution may chemically react and, as a result, will create erroneously high results.
(15) Remove the ampoule from the bath and allow to cool gradually to room temperature. (16) Transfer the reaction solution from the ampoule into a clean 10 ml sample cell. Fill to the 10 ml mark. (17) Fill the second sample cell with 10 ml of distilled water. (18) Insert the sample cell containing the reaction solution into the right opening of the optical comparator. (19) Insert the remaining sample cell into the left opening. (20) Hold the optical comparator lens approximately 10 inches from the eye. Do not make the mistake of placing the eye close to the lens. Face the backplate of the optical comparator directly toward any indirect outdoor (natural) lighting (northern exposure is best). Take care that no shadows fall on the backplate, as this causes uneven illumination of the observation Þelds. Do not prolong the observations for more than 10 to 15 seconds. Let the eyes rest between observations, preferably by viewing a gray or green surface. (21) Slowly rotate the color disk so that one color standard after another is brought into the observation Þeld until the nearest color match is obtained. Read the concentration in either the upper or lower openings at the right side of the optical comparator. If the color of the test solution falls between the two standards, for example between 0.06 and 0.08, report the concentration as 0.07 percent. (22) Record the results of the above test as the volume percent of anti-ice additive to the nearest 0.01 percent. (23) If concentration is not within the limits speciÞed in the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual, defuel airplane and refuel with properly mixed fuel. (24) Dispose of the acid solution by diluting the acid into a beaker almost Þlled with tap water. The diluted solution may then be safely poured down a drain. Flush for a few seconds with tap water. C.
Anti-Ice Additive Concentration Check using HB-P-C B/2 Anti-Icing Additive Test Kit. NOTE: (1) (2)
Refer to Tools, Equipment and Materials list in the test kit.
Perform check in accordance with instruction supplied with test equipment. Verify that anti-icing additive concentration is within the limits speciÞed in the Pilot’s Operating Handbook and Approved Airplane Flight Manual.
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MODEL 208 MAINTENANCE MANUAL
8.
Defueling A.
9.
Defueling Fuel Bays. (1) Ground airplane to suitable ground or stake. (2) Make sure battery switch is turned OFF. (3) Turn fuel selector valves OFF. (4) Remove Þller cap(s) from tank(s) to be defueled; insert defueling nozzle. (5) Remove as much fuel as possible with defueling nozzle. (6) Remove drain valves from bottom of fuel tank and drain remaining fuel.
Purging A.
Fuel Bay Purging.
WARNING: Purge fuel tank(s) with argon or carbon dioxide before you repair leaks. This will minimize the possibility of and explosion. Use a portable vapor detector to determine when it is safe to repair fuel tank(s). (1) (2) (3) (4) (5)
Ground airplane to suitable ground or stake. Make sure battery is disconnected from electrical system. Drain all fuel from tank(s) as outlined in Defueling above. Remove access door and place inert gas supply hose in fuel tank. Allow gas to ßow into tank until fuel vapor cannot be detected. Non-sparking tools must be used to make repairs (air motors, plastic hammers and scrapers, etc.).
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MODEL 208 MAINTENANCE MANUAL ENGINE OIL SYSTEM - SERVICING 1.
General A.
For the engine oil system servicing procedures, use the Pratt and Whitney Engine Maintenance Manual. Refer to the Introduction, Supplier Publication List.
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MODEL 208 MAINTENANCE MANUAL OXYGEN SYSTEM - SERVICING 1.
General A.
An access plate is provided on right side of fuselage to provide access to oxygen system filler valve assembly.
B.
The optional oxygen system consists of a 50.67 or 116.95 cubic-foot capacity oxygen cylinder, altitudecompensating regulator (Model 208 and Model 208B), filler valve, pressure lines, and ten outlets (Model 208 only), two outlets (Model 208 and Model 208B), thirteen outlets (Model 208B Passenger), and oxygen masks and line assemblies as required for each system. Masks are color coded with a blue band adjacent to the plug in fitting. Pilot’s mask is equipped with a microphone, keyed by a button on pilot’s control wheel. The oxygen system pressure gage is located in the overhead console. NOTE:
2.
Some cargo airplanes may also be equipped with quick-don masks.
Charging Oxygen System
WARNING: Oil, grease, or other lubricants in contact with high pressure oxygen create a serious fire hazard. Such contact should be avoided. Do not permit smoking or open flame in or near airplane while work is performed on oxygen systems. A.
Charge Oxygen Cylinder. (1) Remove access panel 312AR. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. (2) Remove oxygen filler valve dust cap. (3) Connect charging cylinder line from oxygen service cart to filler valve.
WARNING: Ground airplane and servicing equipment before charging oxygen system. (4)
Do not attempt to charge oxygen cylinder if servicing equipment fittings or filler valve are corroded or contaminated. If in doubt, clean with stabilized trichlorethylene and let air-dry. Do not allow solvent to enter any internal parts.
CAUTION: A cylinder which is completely empty may be contaminated. The regulator and cylinder assembly must then be disassembled, inspected and cleaned by an FAA approved facility before filling. (5) (6)
If cylinder is completely empty, do not charge. Cylinder must be removed, inspected and cleaned. Slowly open valve on oxygen service cart and charge airplane oxygen bottle to correct pressure. Refer to Table 301 or Table 302 for charging pressures at different temperatures.
Table 301. MIL-O-27210, Type 1 Oxygen Cylinder Fill Pressure for Various Fahrenheit Temperatures Stabilized Temperature °F
Fill Pressure PSIG
Stabilized Temperature °F
Fill Pressure PSIG
Stabilized Temperature °F
Fill Pressure PSIG
-50
1242
20
1569
90
1892
-40
1289
30
1616
100
1937
-30
1336
40
1662
110
1983
-20
1383
50
1708
120
2029
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MODEL 208 MAINTENANCE MANUAL Table 301. MIL-O-27210, Type 1 Oxygen Cylinder Fill Pressure for Various Fahrenheit Temperatures (continued) Stabilized Temperature °F
Fill Pressure PSIG
Stabilized Temperature °F
Fill Pressure PSIG
Stabilized Temperature °F
Fill Pressure PSIG
-10
1430
60
1754
130
2074
0
1477
70
1800
140
2120
10
1523
80
1846
150
2165
Table 302. MIL-O-27210, Type 1 Oxygen Cylinder Fill Pressure for Various Celsius Temperatures Stabilized Temperature °C
Fill Pressure kPa
Stabilized Temperature °C
Fill Pressure kPa
Stabilized Temperature °C
Fill Pressure kPa
-40
8838
-6
11228
28
13365
-38
8989
-4
11359
30
13486
-36
9139
-2
11489
32
13606
-34
9287
0
11619
34
13726
-32
9434
2
11747
36
13845
-30
9579
4
11876
38
13964
-28
9723
6
12003
40
14083
-26
9865
8
12130
42
14201
-24
10006
10
12255
44
14318
-22
10146
12
12381
46
14436
-20
10285
14
12506
48
14553
-18
10423
16
12630
50
14670
-16
10559
18
12754
52
14786
-14
10695
20
12878
54
14903
-12
10830
22
13000
56
15018
-10
10963
24
13122
-8
11096
26
13244
(7) (8) (9)
Shut off oxygen at charging cylinder and disconnect line. Install dust cap. Reinstall access panel 312AR. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation.
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MODEL 208 MAINTENANCE MANUAL TIRES - SERVICING 1.
2.
General A.
Servicing the tire by maintaining correct inflation pressure is the most important job in any tire preventative maintenance program. Improper in flation pressure causes uneven tread wear. (1) Under inflation, indicated by excessive wear in the shoulder area, is particularly damaging. It increases the chance of bruising sidewalls and shoulders against rim flanges. In addition, it shortens tire life by permitting excessive heat buildup. (2) Over inflation is indicated by excessive wear in the center of the tire. This condition reduces traction, increases tire growth and makes treads more susceptible to cutting.
B.
Servicing the tire(s) requires maintenance personnel to handle compressed gas. Observe safety precautions.
Tools, Equipment and Materials NOTE:
NAME
Equivalent substitutes may be used for the following listed item: NUMBER
Nitrogen 3.
MANUFACTURER
USE
Available Commercially
Inflate tires.
Servicing A.
Safety Precautions.
WARNING: Introducing relatively cooler nitrogen into a tire that is hot or when the brakes are hot may cause the tire to burst. (1)
Allow the tire and brake to cool before attempting to service.
WARNING: The tendency of a bursting tire is to rupture along the bead. Standing in any position in front of either bead area could cause injury if the tire should burst. (2)
Stand at a 90-degree angle to the axle along the centerline of the tire during servicing.
CAUTION: Applying a tire sealant on the tire may cause wheel corrosion. (3) B.
Follow all local safety and technical directives while servicing tires.
Procedures. (1) Check tire pressure regularly. (a) Tire pressure should be checked when tire is cold (at least 2 or 3 hours after flight) with an accurate gage (preferably the more precise dial type) on a regular basis. Tire pressure should be checked prior to each flight when practicable. NOTE:
Keeping tires at correct inflation pressures is the most important factor in any preventive maintenance program. The problems caused by underinflation can be particularly severe. Underinflation produces uneven tread wear and shortens tire life because of excessive flex heating. Over-traction makes the tread more susceptible to cutting and increases stress on the wheels. It is recommended that only dry nitrogen be used for tire inflation. Nitrogen will not sustain combustion and will reduce degradation of the inner-liner material due to oxidation.
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MODEL 208 MAINTENANCE MANUAL (b)
(2)
When checking tire pressure, examine tires for wear, cuts, and bruises. Remove oil, grease and mud from tires with soap and water. Use recommended tire pressure. Consult the table below. NOTE:
Recommended tire pressures should be maintained, especially in cold weather. Any drop in temperature of the air inside a tire causes a corresponding drop in air pressure.
NOTE:
If tires freeze to parking ramp in cold weather, use hot air or water to free them before moving airplane.
NOTE:
Inaccurate tire pressure gages are a major cause of improper inflation pressures.
Main Gear Tire Type:
Pressure
6.50 x 10, 8-ply rated tire
83 PSI
8.50 x 10, 8-ply rated tire
48 to 52 PSI
8.50 x 10, 8-ply rated tire (Model 208B Passenger)
53 to 57 PSI
29 x 11-10, 10-ply rated tire
35 to 45 PSI
Main Gear Tire Type - Amphibian: 6.00 x 6, 8-ply rated tire
40 to 50 PSI
Nose Gear Tire Type: 6.50 x 8, 8-ply rated tire
53 to 63 PSI
22 x 8.00-8, 6-ply rated tire
30 to 42 PSI
Nose Gear Tire Type - Amphibian: 5.00 x 5, 10-ply rated tire (3)
Adjusting for temperature. (a) When tires will be subjected to ground temperature changes in excess of 50°F (possibly due to flight from one climate to another), inflation pressures should be adjusted for the worst case prior to takeoff. The minimum required in flation must be maintained for the cooler climate. Pressure 1 can be adjusted in the warmer climate. 2 Before returning to the cooler climate, adjust inflation pressure for the lower temperature. (b) An ambient temperature change of 5°F produces approximately one percent pressure change. NOTE:
(4)
55 to 65 PSI
Excessive inflation pressure should never be bled off from hot tires. All adjustments to inflation pressure should be performed on tires cooled to ambient temperature.
Cold weather servicing. (a) Check tires for excessive deflation. NOTE:
(b) (c) (d)
Tire air pressure will decrease somewhat as the temperature drops, but excessive deflation could indicate cold weather leakage at the air valve. Avoid unnecessary pressure checks.
If it is necessary to pressure check tires in cold climates, always apply heat to air valves and surrounding areas before unseating valves. Continue application of heat during reinflation to ensure air valve seal flexibility when valve closes. Do not allow tires to stand in snow soaked with jet fuel, or on fuel covered ramp areas.
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MODEL 208 MAINTENANCE MANUAL (e)
(5)
If tires become frozen to parking ramp, use hot air or water to melt ice bond before attempting to move airplane. Cold pressure setting. The following recommendations apply to cold inflation pressure settings: (a) “Minimum pressure” for safe airplane operation is the cold in flation pressure necessary to support the operational loads as determined by the formula under “Unloaded Inflation” or as specified by the airframe manufacturer. (b) The loaded inflation must be specified four percent higher than the unloaded inflation. (c) A tolerance of -0 to +5 percent of the minimum pressure is the recommended operating range. (d) If in-service pressure is checked and found to be less than the minimum pressure, the following table should be consulted. An in-service tire is defined as a tire installed on an operating airplane.
TIRE PRESSURE
RECOMMENDED ACTION
100 to 95 percent of service pressure
Reinflate to specified service pressure.
95 to 85 percent of service pressure
Reinflate and record in log book. Remove tire if pressure loss is greater than five percent and reoccurs within 24 hours.
85 percent or less
Remove tire from airplane (Refer to NOTE 1)
NOTE 1: Any tire removed because of low in flation pressure should be inspected to verify that the casing has not sustained internal degradation. If it has sustained internal degradation, the tire should be scrapped. (6)
(7)
Mounted tube-type tires. (a) A tube-type tire that has been freshly mounted and installed should be closely monitored during the first week of operation, ideally before every takeoff. Air trapped between the tire and the tube at the time of mounting could seep out under the beard, through sidewall vents or around the valve stem, resulting in an under-inflated assembly. Tire stretch. (a) The initial stretch or growth of a tire results in a pressure drop after mounting. Consequently, tires should not be placed in service until the have been inflated a minimum of 12 hours, pressures rechecked, and tires reinflated if necessary.
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MODEL 208 MAINTENANCE MANUAL FLOODED LEAD-ACID BATTERY - SERVICING 1.
General A.
2.
3.
Some airplanes may be equipped with a 24.0 VDC, 45 ampere-hour lead-acid battery.
Description A.
The battery is comprised of lead compound plates immersed in a diluted solution of sulfuric acid and water (electrolyte). Each cell has a nominal voltage of approximately 2.0 volts when fully charged. The cells are connected in series.
B.
The battery is equipped with overboard vent lines which connect to the vent fittings on the battery case. (1) The battery case is located on the right side of the forward firewall. This case incorporates integral firewall hinges which allow the battery (and case) to swing out from the firewall.
Servicing A.
Servicing of the flooded lead-acid battery is limited to adding distilled water to the individual cells and cleaning the battery box. For battery removal/installation, inspection, charging and cleaning procedures, refer to Chapter 24, Flooded Lead-Acid Battery - Maintenance Practices. (1) Adding Electrolyte.
WARNING: Do not allow lead acid deposits to come in contact with skin or clothing. Serious burns may result unless the affected area is washed immediately with soap and water. Clothing will be ruined upon contact with battery acid. CAUTION: Do not add any type of battery rejuvenator to the electrolyte. When electrolyte has been spilled from a battery, the electrolyte balance may be adjusted by following instructions in the gill battery maintenance manual for lead-acid batteries. Remove flooded lead-acid battery. Refer to Chapter 24, Flooded Lead-Acid Battery Maintenance Practices. (b) Remove battery caps. (c) Observe electrolyte level in cells. Electrolyte level should be even with the horizontal baffle plate or split ring at bottom of the filler neck. (d) If fluid level is low, add distilled water to obtain proper level. (e) Replace battery caps. (f) Inspect and clean battery box as outlined in step (2). (g) Reinstall battery to airplane. Refer to Chapter 24, Flooded Lead-Acid Battery Maintenance Practices. Cleaning The Battery Box. (a) The battery box should be inspected and cleaned periodically. Use a strong solution of baking soda and water to clean the battery box. Remove hard deposits with a brush and flush thoroughly with clean water. The box should be thoroughly dried before installing battery. (a)
(2)
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MODEL 208 MAINTENANCE MANUAL SEALED LEAD-ACID BATTERY - SERVICING 1.
General A.
2.
3.
Some airplanes may incorporate a sealed lead-acid battery. This 24.0 VDC battery is maintenance free and is rated at 40 ampere-hours.
Description A.
The battery is a recombinant gas (RG) absorbed electrolyte battery. Because the electrolyte is absorbed in glass mat (AGM) separators, no leakage will occur even if the case is cracked or damaged through mishandling.
B.
The battery is equipped with overboard vent lines which connect to the vent fittings on the battery case. (1) The battery case is located on the right side of the forward firewall. This case incorporates integral firewall hinges which allows the battery (and case) to swing out from the firewall.
Servicing A.
This sealed battery, when installed in the airplane, requires no routine maintenance or servicing. However, after time limits have been reached as defined in Chapter 5, Inspection Time Limits, a capacity check shall be performed to ensure continuing airworthiness beyond this point. NOTE:
B.
Batteries are serviced and charged at the factory prior to shipment. However, batteries in storage should be boost charged every 90 days.
For charging instructions and battery capacity checks, refer to Chapter 24, Sealed Lead- Acid Battery - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - SERVICING 1.
General A.
2.
3.
Airplanes may incorporate a nickel-cadmium (Ni-Cad) battery. This 25.2 VDC battery is rechargeable and is rated at 40 ampere-hours.
Description A.
The electrolyte used in a ni-cad battery is a solution of distilled water and potassium hydroxide. The electrolyte is used only as a conductor and does not react with plates. The negative plates in the battery are cadmium-oxide and the positive plates are nickel-oxide. During charging, all oxygen is driven out of negative plates and only metallic cadmium remains. The oxygen dispelled from negative plates is picked up by positive plates to form nickel dioxide. Toward the end of charging process, electrolyte will gas due to electrolysis taking place in electrolyte. A slight amount of gassing is necessary to completely charge the battery.
B.
During discharge, reverse chemical action takes place. The negative plates gradually gain back oxygen, as positive plates lose oxygen. Due to this interchange of oxygen, chemical energy of plates is converted into electrical energy and electrolyte is absorbed by plates. For this reason, the level of electrolyte should be check only when battery is fully charged.
Servicing A.
Servicing of the ni-cad battery is limited to adding distilled, deionized or demineralized water to the individual cells. For battery and cell removal/installation, refer to Chapter 24, Ni-Cad Battery Removal/Installation.
WARNING: The electrolyte used in nickel-cadmium batteries is a caustic solution of potassium hydroxide. Serious burns will result if it comes in contact with any part of the body. Use rubber gloves, rubber apron and protective goggles when handling this solution. If electrolyte gets on skin, wash affected areas thoroughly with water, and neutralize with three-percent acetic acid, vinegar or lemon juice. If electrolyte gets into eyes, flush with water and get immediate medical attention. WARNING: Rings, metal watchbands and other metallic jewelry should be removed before working around the battery. Should such metallic objects contact intercell connectors of opposing polarity, they may fuse themselves to the connectors and cause severe skin burns. CAUTION: Tools or equipment used for servicing lead-acid batteries shall not be used for servicing Ni-Cad batteries. Ni-Cad batteries should be completely removed from lead-acid battery service area. The slightest acid contamination will deteriorate Ni-Cad batteries. (1)
Adding Fluid To Individual Cells. (a) Remove battery. Refer to Chapter 24, Ni-Cad Battery - Removal/Installation. (b) Remove battery caps.
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MODEL 208 MAINTENANCE MANUAL (c)
(d) (e)
Check electrolyte level in cells and adjust as required. Electrolyte level of fully charged ni-cad battery should be 0.250 inch above top of 1 plates immediately after charging, and approximately 0.125 inch above plates two hours after charging. NOTE:
Electrolyte level in ni- cad batteries can be accurately checked only when battery is fully charged. Ensure battery is fully-charged before checking electrolyte level.
NOTE:
Use only distilled, deionized or demineralized water. Batteries are easily contaminated through the use of tap water, which contains minerals, chlorines, softening agents and other foreign material.
Replace battery caps. Reinstall battery to airplane. Refer to Chapter 24, Ni-Cad Battery - Removal/Installation.
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MODEL 208 MAINTENANCE MANUAL VACUUM SYSTEM CENTRAL AIR FILTER - SERVICING 1.
General A.
The vacuum system central air filter keeps dust and dirt from entering the vacuum operated instruments.
CAUTION: Do not operate vacuum system with filter removed or vacuum line disconnected, as dust and other foreign matter may enter the system and damage the vacuum operated instruments. B.
Refer to Chapter 5, Inspection Time Limits for filter inspection intervals. Replace filter element when damaged and whenever it becomes sufficiently clogged to cause suction gage reading to drop below 4.5 inches Hg (mercury).
CAUTION: Smoking during system operation will cause premature filter clogging. 2.
Servicing A.
Remove Air Filter (Refer to Figure 301). (1) Unscrew bolt and washer from bottom of central air filter. (2) Remove central air filter from filter bracket. (3) Inspect for damage, deterioration and contamination. Clean or replace as required.
B.
Install Air Filter (Refer to Figure 301 ). (1) Seat central air filter up and into filter bracket. (2) Secure central air filter to filter bracket using bolt and washer. (3) Check central air filter for unobstructed flow. A properly functioning filter should allow a reading of at least 4.5 inches Hg (mercury) on the instrument panel suction gage.
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MODEL 208 MAINTENANCE MANUAL
Vacuum System Central Air Filter Servicing Figure 301 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FREON AIR CONDITIONING - SERVICING 1.
General A.
2.
Airplane may be equipped with Freon air conditioning system. This servicing procedure provides instructions for system discharging, evacuating, charging, leak testing and checking compressor oil level with compressor installed on airplane.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following listed items:
NAME
NUMBER
MANUFACTURER
USE
Charging Manifold and Hose Assembly
CMN-4-3
Superior Supply Company 215 Laura Wichita, KS 67211
To charge system.
Refrigerant Can Adapter
CT-3
Superior Supply Co.
To charge system.
Halogen Leak Detector
RLD-G19G-1
Superior Supply Co.
To detect Freon leaks.
Commercially Available
To evacuate system.
Vacuum Pump (Capable of 28 to 29 inches of Mercury) Two (2) Dial-Type Thermometers (2.0 inch diameter. Dial range 0° to 120°F)
C668507-0101, -0103 or -0104
Cessna Aircraft Cessna Parts Distribution Department 701, CPD 25800 East Pawnee Road Wichita, KS 67218- 5590
To check temperature differential.
Refrigerant
R-12
Commercially Available
To charge system.
Commercially Available
To check for leaks in system.
Dry Nitrogen Refrigerant Oil (500 viscosity minimum)
Capella WF 100
Texaco Oil Company Box 7483 White Plains, NY 10650
To lubricate compressor, fittings and O-rings.
Refrigerant Oil (500 viscosity minimum)
Suniso 5GS
Sun Oil Company 1801 Market Street Philadelphia, PA 19103
To lubricate compressor, fittings and O-rings.
Suds Spray Leak Detector
Type F
Winton Products Co. P. O. Box 36332 Charlotte, NC 28236
To check for leaks in charged system.
Teflon Tape
0.50 inch wide
Commercially Available
To wrap lines.
Compressor Oil Dipstick
32447
Commercially Available
To check compressor oil level.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
Machinists Universal Level Robin Air Charging Station
13105
Freon Recovery System 3.
MANUFACTURER
USE
Commercially Available
To check mounting angle of compressor on airplane.
Robinair Mfg. Corp. 1224 Robinair Way Montpelier, OH 43543-1933
To charge system.
Commercially Available
To capture Freon from system.
General Precautions A.
Freon Handling. NOTE:
Effect of Montreal Protocol and U. S. Environmental Protection Agency’s Clean Air Act of 1990 is to ban unnecessary release of CFC-12 refrigerant, also known as R12 into the atmosphere. In compliance with the preceding, Cessna Aircraft recommends refrigerant be captured and recycled. For additional information, refer to Federal Clean Air Act, EPA 40 CFR Part 82.
WARNING: Liquid R12 at normal atmospheric pressure and temperature, will freeze anything it contacts. Eyes are especially susceptible to damage. Safety glasses are absolute minimum protection and shall be worn at all times when servicing Freon system. WARNING: Do not attempt to treat yourself should any liquid refrigerant get into eyes. Follow these instructions: do not rub eye. Splash large quantities of cool water into eye to raise temperature. Apply a few drops of mineral oil to eye to wash it, followed by a weak solution of boric acid to flush out all of the oil. Seek aid of a doctor immediately. (1) (2) B.
Observe safety precautions when handling refrigerant or servicing and performing maintenance on air conditioning system. Use of protective clothing, gloves and goggles will protect the skin and eyes.
General System Notes. NOTE:
C.
Cleanliness is of utmost importance to avoid system contamination and useless wear to compressor and other equipment items. All plumbing and hoses shall be cleaned and capped after fabrication and shall remain capped during storage and installation until hooked up to their mating components. All valves shall also be capped with clean caps or plugs. During time components are open, extreme care shall be exercised to assure that no contaminating matter enters parts or system. Receiver/dryer is easily contaminated with moisture from atmosphere. All care shall be exercised to prevent moisture from entering receiver/dryer.
Removing Hoses Under Pressure.
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WARNING: Do not remove hoses under pressure. This procedure will result in the release of refrigerant into the atmosphere. Removing hoses under pressure may also result in personal injury if hose ends are not restrained. D.
Use of Intense Heat.
WARNING: To avoid explosion, never weld, use a blow torch, steam clean, bake aircraft finish or use excess amounts of heat on or in immediate area of any part of the air conditioning system or refrigerant supply tank, while they are closed to atmosphere, charged or not. Although R12 gas, under normal conditions, is nonpoisonous, discharge of refrigerant gas near a flame can produce a very poisonous gas (Phosgene). This gas will also attack all bright metal surfaces. WARNING: Do not use a flame-type leak detector because of fire hazard on airplanes and production of minor amounts of phosgene gas. Inhaling WARNING: Do not smoke in vicinity of refrigerant discharge. refrigerant through burning tobacco will produce a poisonous gas like an open flame. E.
Use of Nitrogen. NOTE:
4.
All nitrogen pressure checks are to be made only with regulated nitrogen.
System Evacuating A.
Evacuation of System (Refer to Figure 301 and Figure 302). NOTE:
Understand and follow all safety precautions prior to evacuating system.
NOTE:
Charging manifold and hoses must be free of contamination.
(1) (2) (3) (4) (5) (6)
Gain access to servicing valves, located between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Connect charging manifold hoses to low and high pressure servicing valves. Connect center hose from charging manifold to vacuum pump. Open valve on vacuum pump and both hand valves on charging manifold. Turn on vacuum pump to begin evacuation of system. Continue evacuation a minimum of 30 minutes after compound (suction) gage has reached 27.0 to 29.0 inches of mercury vacuum. NOTE:
(7) (8)
27.0 to 29.0 inches of mercury vacuum value is for sea level condition. For each 1,000 feet increase in field elevation, decrease value 1.0 inch of mercury.
Close both hand valves. Turn vacuum pump off.
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Typical Air Conditioning Servicing Ports Figure 301 (Sheet 1)
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Charging Manifold Connections Figure 302 (Sheet 1)
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5.
System Discharging/Charging A.
Discharging System (Refer to Figure 301 and Figure 302). NOTE:
Understand and follow all safety precautions prior to discharging system.
NOTE:
Charging manifold and hoses must be free of contamination.
(1) (2) (3) (4) (5) (6) B.
Gain access to servicing valves, located between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Connect charging manifold hoses to low and high pressure servicing valves. Connect Freon reclaimer in accordance with manufacturer’s reclaimer instructions. Capture Freon in accordance with manufacturer’s reclaimer instructions. Close hand valves when Freon has been reclaimed. Install access panel 232AC.
Charging System (Refer to Figure 301, Figure 302 and Figure 304). (1) Gain access to servicing valves, between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Connect charging manifold hoses to low and high pressure servicing valves. Refer to Figure 301 and Figure 302. (3) Evacuate system. Refer to System Evacuation procedures. (4) Open valve on R12 container. (5) Open center hose at manifold gage set for five seconds to purge hose of air. The system is now purged and under a vacuum.
CAUTION: Do not start engine at this time. Do not turn air conditioner on. (6)
Open high side manifold hand valve; observe low side gage. Close high side hand valve. (a) If low side gage does not change from a vacuum condition to a pressure condition. System blockage is indicated. Re-evacuate system and proceed with following steps. (7) If blocked, correct blockage, (8) Ensure both high and low side manifold hand valves are closed. (9) Using a qualified assistant, start engine and adjust to ground idle. (10) Adjust air conditioning controls for maximum cooling, and blower on high speed. (11) Keep refrigerant cylinder in upright position.
CAUTION: If container is turned upside down, refrigerant will become liquid. Liquid refrigerant entering compressor may cause serious damage. (12) Open low side manifold hand valve and allow a 7.5 pound R12 refrigerant charge in the gaseous state to enter the system. NOTE:
A fully charged system will take between 7.5 and 8.5 pounds of R12 refrigerant.
(a) Continue charging until sight glass is clear and free of frothing or bubbles. (13) Verify gages display approximately 40.0 PSI on low pressure gage and 170.0 to 210.0 PSI on high pressure gage. Higher readings indicate overcharged or blocked system. (a) For pressure at various temperatures, refer to Figure 304. (b) If high surges of pressure on high and low gages are indicated during charging, receiver/ dryer may be contaminated or expansion valve blocked. Refer to Chapter 21, Freon Air Conditioning - Troubleshooting.
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MODEL 208 MAINTENANCE MANUAL (14) Bubbles in sight glass of stabilized system (five minutes minimum operation) constituting a foam or frothing action indicate an undercharge. NOTE:
Occasional individual bubbles passing through sight glass should be considered a clear sight glass and full charge.
(a)
(15) (16) (17) (18) (19) (20) 6.
If frothing continues, open suction valve on charging manifold to allow approximately two ounces of refrigerant into system. Wait two minutes and recheck sight glass. Clear glass indicates full charge, frothy glass indicates low charge. Repeat two ounce charge and two minute wait until sight glass is clear. Close low and high pressure valves or charging manifold. Close valve at charging cylinder. Shut down engine. After system has stabilized for five minutes, remove hose fittings from high and low pressure service valves. Use a clean cloth wrapped around knurled fitting to prevent refrigerant from getting on skin and to catch excess oil. Replace caps on service valves. Reinstall access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation.
Check Compressor Oil Level With Compressor on Airplane A.
Checking Compressor Oil Level (Refer to Figure 303). (1) Discharge system. Refer to System Discharging/Charging. (2) Remove oil filler plug. Look through oil filler hole and rotate clutch front plate to position internal components. NOTE: (3) (4) (5) (6) (7)
This step is necessary to clear dipstick of internal parts, allowing insertion to full depth.
Center parts as they are moving to rear of compressor (discharge stroke). Determine mounting angle of compressor from horizontal. Position a machinists universal level, across flat surfaces of compressor’s two front mounting ears, center bubble, and read angle to closest degree. Insert dipstick to STOP position. Fabricate dipstick using fabrication instructions in Figure 303. (a) Ensure point of angle is to right if mounting angle of compressor is to the left. (b) Ensure bottom surface of angle is flush with surface of oil filler hole. Remove dipstick and count increments of oil. Each mark equals one fluid ounce with compressor in level position. Determine correct oil level for mounting angle of compressor. If increments on dipstick do not match table, add or subtract oil to midrange value. For example, if mounting angle of compressor is ten degrees and dipstick is three degrees, add oil in one fluid ounce increments until 7 is read on dipstick. Refer to Table 301. NOTE:
Too much oil in system will greatly reduce the efficiency of system.
Table 301. Oil Level for Mounting Angle. MOUNTING ANGLE (IN DEGREES)
ACCEPTABLE OIL LEVEL (IN INCREMENTS)
0
3 to 5
10
4 to 6
20
5 to 7
30
6 to 8
40
7 to 9
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Checking Compressor Oil Level Figure 303 (Sheet 1)
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Checking Compressor Oil Level Figure 303 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL Table 301. Oil Level for Mounting Angle. (continued)
(8)
50
8 to 10
60
8 to 10
90
8 to 10
Install oil filler plug. Ensure sealing O-ring is not twisted. Torque from 6.0 to 9.0 foot-pounds.
CAUTION: Do not overtighten oil plug to stop a leak. Remove plug and install a new O-ring. 7.
Functional Leak Check A.
Air Conditioning System Leak Check. (Refer to Figure 305).
WARNING: Use halogen tester in a well ventilated area to prevent any concentration of poisonous gas being produced. Do not breathe fumes. NOTE:
This test assumes system is not charged. If system is charged, it must be discharged before proceeding.
Gain access to servicing valves, between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Connect charging manifold to high pressure and low pressure service valves. (3) Connect vacuum pump to center charging manifold. (4) Turn on vacuum pump and open both hand valves (high and low pressure) on charging manifold. (5) Allow pump to run for 15 minutes. (6) Shut off both hand valves on charging manifold. Open valve on R12 container. Open both hand valves on charging manifold and allow one pound of Freon into system. (7) Shut off both hand valves on charging manifold and disconnect vacuum pump. Connect a dry, high-pressure regulated dry nitrogen source at center connection on charging manifold. (8) Slowly open hand valve (low pressure) on charging manifold, while watching high pressure gage. service system from 175 to 200 PSIG. Close hand valve (low pressure) on charging manifold and hose, and disconnect nitrogen supply. (9) Record pressure reading from gage and ambient temperature only after airplane has been in hanger a minimum of two hours to allow temperature to stabilize. (10) After a minimum of 24 hours, record system pressure and temperature. (a) Depending on final temperature being above or below initial temperature, Refer to charts from Figure 305, add or subtract the delta P to initial pressure reading. (b) Compare adjusted initial pressure with final pressure. If system loss is greater than 10 PSI, a system leak is present and must be corrected. (c) Locate leak using halogen. (d) Repair leak and rerun pressure check with Suds Spray. (1)
NOTE:
B.
All instructions for Suds Spray Leak Check are identical to preceding instructions except all fittings shall be sprayed with suds instead of using Halogen leak detector. Refer to steps 7.A.(10) (a) thru 7.A.(10) (c).
Install Access Panel. (1) Install access panel 232AC between pilot and copilot seat, Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
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Pressure Versus Temperature Chart Figure 304 (Sheet 1)
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Pressure Versus Temperature Chart Figure 304 (Sheet 2)
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Temperature and Pressure Chart Figure 305 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL R134A AIR CONDITIONING - SERVICING (Airplanes 20800274 and On and 208B0655 and On) 1.
General A.
2.
The airplane may be equipped with R134a air conditioning system. This servicing procedure provides instructions for system discharging, evacuating, charging, leak testing and checking compressor oil level with compressor installed on airplane.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following listed items:
NAME
NUMBER
MANUFACTURER
USE
Continuity Tester
VOM
Commercially Available
To troubleshoot electrical system.
Mobile Oil Corp. Interstate 10 Beaumont, TX 77704
To lubricate compressor, fittings and O-rings.
Ester Refrigerant Oil Leak Detector, Electronic
16500
Robinair Mfg. Corp. 1224 Robinair Way Montpelier, OH 43543-1933
Used to detect refrigerant leaks.
Leak Detector, Suds Spray
Type F
Winton Products Box 36332 Charlotte, NC
Used to detect refrigerant leaks.
Thermometer, Digital Type, 0 to 150 degrees
Commercially Available
Used to check temperature differential.
Dry Nitrogen Regulated, Dry
Commercially Available
Used to check for leaks in system.
Refrigerant
R134a
Commercially Available
Used to charge system.
Robinair Recovery and Recycling Cart
P/N 34700 P/N 34560 with Flushing Kit
Robinair Mfg. Corp. 1224 Robinair Way Montpelier, OH 43543-1933
Used to recover, recycle, recharge and flush with optional flushing kit.
Teflon Tape
0.50 inch wide
Commercially Available
Used to wrap lines.
Tensiometer
Gates 150 (17599H)
Gates Rubber Co. 2707 W. Douglas Wichita, KS 67213
Used to check belt tension.
Compressor Oil Dipstick
32447
Sanden International 10710 Sanden Drive Dallas, TX 75328
Used to check compressor oil level.
Acme Union, 1/4 inch
Commercially Available
Used to connect regulated nitrogen.
Machinists Universal Level
Commercially Available
Used to check mounting angle of compressor on airplane.
Power Supply 28 VDC, 200 Amp minimum.
Commercially Available
Used to provide power.
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3.
General Precautions A.
R134a Refrigerant Handling
WARNING: Observe safety precautions when handling refrigerant or servicing and performing maintenance on air conditioning system. Care must be taken to minimize release of refrigerant into the atmosphere. The Environmental Protection Agency (EPA) requires recycling/recovery of 134A as of 11/15/95. All reclamation and recovery equipment must be EPA and UL listed. Use R134A reclamation system per manufacturers instructions, whenever evacuating system. WARNING: Liquid refrigerants, at normal atmospheric pressure and temperature, will expand and absorb heat. As a result, it will freeze anything it contacts. Use of protective clothing, gloves, and goggles will protect skin and eyes. Eyes are especially susceptible to damage, so safety glasses are an absolute minimum protection. Goggles are the preferred method of protection and should be worn at all times when servicing system. WARNING: Should any liquid get into eyes, follow these instructions: do not rub eye. Splash large quantities of cool water into eye to raise temperature. Apply a few drops of mineral oil to eye to wash it, followed by a weak solution of boric acid to flush out all of the oil. Seek aid of a doctor immediately. Do not attempt to treat yourself. (1) (2) B.
Observe safety precautions when handling refrigerant or servicing and performing maintenance on air conditioning system. Use of protective clothing, gloves and goggles will protect the skin and eyes.
General System Notes. NOTE:
C.
Cleanliness is of utmost importance to avoid system contamination and useless wear to compressor and other equipment items. All plumbing and hoses shall be cleaned and capped after fabrication and shall remain capped during storage and installation until hooked up to their mating components. All valves shall also be capped with clean caps or plugs. During time components are open, extreme care shall be exercised to assure that no contaminating matter enters parts or system. Receiver/dryer is easily contaminated with moisture from atmosphere. All care shall be exercised to prevent moisture from entering receiver/dryer.
Removing Hoses Under Pressure.
WARNING: Do not remove hoses under pressure. This procedure will result in release of refrigerant into the atmosphere. Removing hoses under pressure may also result in personal injury if hose ends are not restrained. D.
Use of Intense Heat.
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WARNING: To avoid explosion, never weld, use a blow torch, or use excessive amounts of heat on or in immediate area of any part of the air conditioning system or a refrigerant supply tank, while they are close to atmosphere, charged or not. E.
Use of Nitrogen. NOTE:
4.
5.
All nitrogen pressure checks are to be made only with regulated nitrogen.
Compressor Lubrication NOTE:
Compressors are shipped from factory with the PAG oil. The original oil must be drained from the compressor and then replaced with the ester based refrigerant oil.
NOTE:
Do not, at any time, leave oil container standing open. Keep cap on tight, as exposed refrigerant oil absorbs moisture rapidly.
NOTE:
Oil is needed to lubricate the seals, gaskets and other parts of the compressor. A small amount of oil is circulated through the system with the refrigerant and is necessary to keep the expansion valve functioning properly. The 208 airplane system requires 11 fluid ounces of oil.
System Evacuating A.
Evacuation of System (Refer to Figure 301 and Figure 302). NOTE:
Understand and follow all safety precautions prior to discharging system.
NOTE:
Charging manifold and hoses must be free of contamination.
(1) (2) (3) (4) (5) (6) (7)
Gain access to service valves, located between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Remove cap from service valve on high pressure side of system and connect high pressure hose (red) of the charging manifold. Remove cap from the service valve on low pressure side of system and connect the low pressure hose (blue) of the charging manifold. Connect vacuum hose (yellow) to center of charging manifold. Open high and low pressure valves on charging manifold and valve to the vacuum pump. Start vacuum pump to begin evacuation of system. Observe low pressure gage needle. It should pull down into a slight vacuum. After about five minutes, low pressure gage should be below 20.00 inches of mercury and the high pressure gage should be slightly below the zero index on the gage. NOTE:
(8) (9) 6.
If high pressure does not drop below zero, a blocked line is indicated. If blocked, then stop evacuation and repair or remove obstruction.
Operate vacuum pump for 15 minutes and observe gage. Gage should be down to 27 inches of mercury. Hold vacuum for at least 30 minutes and observe gage. If the gage rises it indicates a leak that will need repair. If no leak is indicated, charge the system Refer to step 6.B.
System Discharging/Charging A.
Discharging System (Refer to Figure 301 and Figure 302). NOTE:
Understand and follow all safety precautions prior to discharging system.
NOTE:
Charging manifold and hoses must be free of contamination.
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R134a Air Conditioning Servicing Ports Figure 301 (Sheet 1)
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R134a Charging Manifold Connections Figure 302 (Sheet 1)
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WARNING: Extreme care must be taken when discharging refrigerant from system. Discharging must be done very slowly. Gloves and goggles should be worn for protection. (1) (2) (3) (4) (5) (6) (7) (8) B.
Gain access to servicing valves, located between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Connect charging manifold hoses to low and high pressure service valves. Use hose from center port of charging manifold to discharge refrigerant into a container to catch oil. Open high pressure manifold valve slowly to bleed off refrigerant and oil into container without splattering. As pressure drops, valve may be opened for faster discharge. Close valve when pressure drops below 5-10 psi so no air enters system, unless a component must be replaced which necessitates complete bleed down. Observe amount of oil caught in container. If amount is approximately a tablespoon or less, disregard loss. If it is more than one tablespoon, add like amount in compressor at fill nut. Install access panel 232AC .
Charging System (Refer to Figure 301, Figure 302 and Figure 304). NOTE: (1)
Refer to paragraphs on handling refrigerant and removing hoses under pressure for safety precautions. Charging manifold and hoses must be free of contamination.
Gain access to servicing valves, between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. NOTE:
(2)
Ensure system has been evacuated per step 5.A, attaining 25-27 in. Hg within 15 minutes and maintaining vacuum for 30 minutes after vacuum pump is shut off. If either criteria is not met, a leak is indicated and system must be leak checked per step 8.A. before being re-evacuated.
Following the servicing instructions, replace any oil that was vented during the leak test procedure. NOTE:
Do not add more oil than was vented. Too much oil in the system can deteriorate cooling performance of evaporators. If compressor was drained, refer to Chapter 21, R134a Air Conditioning - Maintenance Practices, for compressor lubrication.
WARNING: Do not allow a slug of liquid into system by tilting refrigerant container on its side, because it could damage the system (3) (4) (5)
Connect charging manifold hoses to low and high pressure servicing valves. Refer to Figure 301 and Figure 302. Place refrigerant container on scales. Connect R134a container to center of manifold and open valve on R134a container. Open center hose at manifold gage set for five seconds to purge hose of air. System is now purged and under a vacuum.
CAUTION: Do not start engine at this time. Do not turn air conditioner on. (6) (7) (8) (9)
Ensure both high and low side manifold hand valves are closed. Using a qualified assistant, start engine and adjust to ground idle. Adjust air conditioning controls for maximum cooling, and blower on HIGH setting. With refrigerant in an upright position, open low side manifold and allow 5.5 pounds of R134a gas to enter system. A normal fully charged system will require a charge between 5.5 and 7 pounds of refrigerant.
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WARNING: Do not overcharge system or component or system damage may occur. (10) With system fully charged and operating, observe the suction and discharge pressure. Pressure will vary with air temperature. Refer to Figure 304. (11) Allow system to operate for 10 minutes and then shut down. NOTE:
After shutdown, both suction and discharge pressures will immediately start equalizing. Pressures should be equal after a maximum of six minutes.
(12) Turn off system and replace caps on service valves. (13) Reinstall access 232AC . Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. 7.
Check Compressor Oil Level With Compressor on Airplane A.
Checking Compressor Oil Level (Refer to Figure 303). (1) Discharge system. Refer to System Discharging/Charging. (2) Remove oil filler plug. Look through oil filler hole and rotate clutch front plate to position internal components. NOTE: (3) (4) (5) (6) (7)
This step is necessary to clear dipstick of internal parts, allowing insertion to full depth.
Center parts as they are moving to rear of compressor (discharge stroke). Determine mounting angle of compressor from horizontal. Position machinists universal level, across flat surfaces of compressors two front mounting ears, center bubble, and read angle to closest degree. Insert dipstick to STOP position. Fabricate dipstick using fabrication instructions in Figure 303. (a) Ensure point of angle is to right if mounting angle of compressor is to the left. (b) Ensure bottom surface of angle is flush with surface of oil filler hole. Remove dipstick and count increments of oil. Each mark equals one fluid ounce with compressor in level position. Determine correct oil level for mounting angle of compressor. If increments on dipstick do not match table, add or subtract oil to midrange value. For example, if mounting angle of compressor is ten degrees and dipstick is three degrees, add oil in one fluid ounce increments until 7 is read on dipstick. Refer to Table 301. NOTE:
Too much oil in system will greatly reduce the efficiency of system.
Table 301. Oil Level for Mounting Angle MOUNTING ANGLE (IN DEGREES)
ACCEPTABLE OIL LEVEL (IN INCREMENTS)
0
3 to 5
10
4 to 6
20
5 to 7
30
6 to 8
40
7 to 9
50
8 to 10
60
8 to 10
90
8 to 10
(8)
Install oil filler plug. Ensure sealing O-ring is not twisted. Torque from 6.0 to 9.0 foot-pounds.
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CAUTION: Do not overtighten oil plug to stop a leak. Remove plug and install a new O-ring. 8.
Functional Leak Check A.
Air Conditioning System Leak Check (Refer to Figure 305). NOTE:
This test assumes system is not charged. If system is charged, it must be discharged before proceeding.
NOTE:
The refrigeration system must be checked after assembly and prior to charging the system if vacuum cannot be held during evacuation. Due to installation of the aircraft interior, it may be necessary to leak test the refrigerant lines aft of the firewall before the entire system is checked. Refrigerant lines penetrating the firewall must be securely plugged or capped to leak test aft portion of aircraft.
(1)
If a leak is indicated during evacuation, relieve the vacuum in system and disconnect the service cart. If desired, allow approximately 1/4 lb. of R134a to enter system per the service cart instructions before relieving the vacuum from the system in order to enable the use of an electronic leak detector during the leak check.
CAUTION: Service cart must be disconnected before the system is pressurized with nitrogen. If it is not disconnected, service cart could be damaged. CAUTION: Do not use any leak dye in the system, or damage to expansion valves or compressor may result. (2) (3) (4)
Gain access to servicing valves, between pilot and copilot seat, by removing access panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Connect regulated dry nitrogen to the high pressure side of the system. Open valve at service coupling end of hose and slowly regulate nitrogen into system to pressure of 350-400 psig maximum. Allow the high and low sides to stabilize to the same pressure.
CAUTION: Regulate nitrogen into system slowly to maintain control of pressure. With high pressures obtainable from nitrogen bottles, over pressurizing system is possible which would result in system damage. (5)
Apply suds leak detector to all system connections. An electronic leak detector can be used if 1/4 lb. of R134a was introduced into the system per step 8.A.(1). NOTE:
(6)
Repair leaks and rerun leak test. The high pressure gage may also be monitored for continuing drop in system pressure indicating a leak still existed. Note slight variations in pressure may be caused by temperature changes. NOTE:
B.
The electronic leak detector is very sensitive to refrigerant leaks. Residual refrigerant vapors in the area may set off the instrument prematurely. Therefore, be sure the area is clear of all vapors. A high pressure air purge (blow gun) usually clears an area of residual vapors.
Do not over tighten plumbing connections. Stripped threads may result.
After all leaks are repaired, vent nitrogen pressure by closing valve on high pressure side service coupler and adjusting the regulator on nitrogen to 0 psig, allowing hose to service coupler to vent. Disconnect high pressure hose at adapter tee and slowly open valve on high side service coupler to allow the nitrogen to vent from system.
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R134a Compressor Oil Level Check Figure 303 (Sheet 1)
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R134a Compressor Oil Level Check Figure 303 (Sheet 2)
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Pressure Versus Temperature Chart Figure 304 (Sheet 1)
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R134a Temperature and Pressure Chart Figure 305 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL C.
Install Access Panel. (1) Install access panel 232AC . between pilot and copilot seat. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
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MODEL 208 MAINTENANCE MANUAL NOSE GEAR SHOCK STRUT - SERVICING 1.
General A.
2.
The servicing of the nose gear shock strut is limited to adding hydraulic fluid (MIL-PRF-5606) to the strut.
Service Nose Gear Shock Strut A.
Servicing Procedures (Refer to Figure 301). NOTE:
No air pressure is required in the strut.
(1) (2) (3) (4) (5) (6)
Lift the nose of the airplane. Refer to Chapter 7, Jacking - Maintenance Practices. Make sure that the strut is fully extended. Remove the AN913-1 plug from the upper forward side of the strut. Install an AN816-2D or equivalent 0.125 inch NPT fitting in the upper strut. Attach the hydraulic fluid filler tube to the fitting. Use a pressure feed system (hand pump, pressure bottle, etc.) to fill the strut with MIL-PRF-5606 (red) hydraulic fluid. (7) Remove the pump and let the fluid go back through the filler tube to the supply. (8) Apply shop air or other air pressure (70.0 PSIG maximum) to the filler tube to blow excess fluid from the strut (internal standpipe installed in strut). (9) Remove the filler tube and fitting from the upper forward side of the strut. (10) Install the AN913-1 plug in the upper forward side of the strut. (11) Remove the airplane from the jack. Refer to Chapter 7, Jacking - Maintenance Practices.
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Nose Gear Shock Strut Servicing Figure 301 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL SCHEDULED SERVICING - DESCRIPTION AND OPERATION 1.
General A.
2.
This section provides instructions necessary to carry out scheduled servicing as well as internal/external cleaning. It also includes instructions for lubricating specific points identified in periodic inspection and/or preventive maintenance programs. This section does not include lubrication procedures required for the accomplishment of maintenance practices.
Description A.
This section is subdivided to provide maintenance personnel with charts, text and illustrations to prevent confusion. Also included in this section is a table containing a list of lubricants. (1) The subdivisions are separated according to airplane systems. This aids maintenance personnel in locating service information.
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MODEL 208 MAINTENANCE MANUAL LUBRICANTS - DESCRIPTION AND OPERATION 1.
2.
General A.
This part of the scheduled servicing section helps the operator find the recommended lubricants. For the best results and continued trouble-free service, use only clean and approved lubricants.
B.
For a list of recommended lubricants, refer to Recommended Lubricants Table.
Lubrication Service Notes A.
Lubricant Application. (1) Keep lubricants and dispensing equipment clean. NOTE: (2) (3) (4) (5) (6) (7) (8)
Cleanliness is necessary for good lubrication.
Use only one lubricant in a grease gun or oil can. Keep lubricants in tightly closed containers in a protected area. Before you apply the lubricant, use clean, dry cloths to clean the grease fittings and areas to be lubricated. To lubricate vented bearings, push grease into the fitting until the old grease is removed from the bearing. Use the special grease gun adapter, Alemite Part Number 318049, or an equivalent, to lubricate the flush-type grease fittings (NAS516). After any lubrication procedure, clean the unwanted lubricant from all but the actual friction area parts. All sealed or prepacked antifriction bearings are lubricated with MIL-PRF-23827 grease by the manufacturer. NOTE:
Unless otherwise specified, no more lubrication is necessary for these bearings.
(9) Do not put oil on antifriction bearings. (10) Do not let spray from steam or chemical cleaning touch the antifriction bearings. (a) When it is necessary to clean the exterior bearing surfaces, clean with a cloth moist with Federal Specification PD-680 solvent. (11) The porous, sintered-type friction bearings are lubricated before delivery. (a) To extend its service life, use an oil can and randomly apply some general purpose oil (MIL-PRF-7870) to the bearings. (12) Use a general purpose oil (MIL-PRF-7870) to lubricate the unsealed pulley bearings, rod ends, pivot end hinge points and any other friction point where lubrication is necessary. (13) Rub paraffin wax on the seat rails to make it easier to move the seats fore and aft. (14) Do not lubricate the roller chains or cables except under sea coast conditions. Clean them with a clean, dry cloth. (15) Use a general purpose oil (MIL-PRF-7870) to lubricate the unsealed pulley bearings, pivot points and any other friction point during the installation. (16) With the exception of the control rod ends for the engine and propeller and the rod ends of the hanger assemblies of the secondary exhaust duct, use a general purpose oil (MIL-PRF-7870) to lubricate the rod ends. NOTE:
The rod ends of the hanger assemblies of the secondary exhaust duct are prelubricated. The heat given off by the secondary exhaust duct will burn off the lubrication. This is a normal occurrence.
(a)
Lubricate the control rod ends for the engine and propeller. Refer to the Pratt and Whitney Maintenance Manual, Chapter 72-00. (17) Use (PG) powdered graphite (SS-G-659) to lubricate all piano hinges when the assembly is installed.
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MODEL 208 MAINTENANCE MANUAL
3.
Recommended Lubricants Table A.
The table that follows helps maintenance personnel in lubrication procedures. NOTE:
Equivalent substitutes can be used for the following items:
Table 1. Recommended Lubricants SYMBOL
PROCUREMENT SPECIFICATION
DESCRIPTION AND USE
PRODUCT PART NUMBER
VENDOR
GR
MIL-PRF81322
Grease, wide temperature range.
Mobil grease 28
Mobil Oil Corp. 150 E. 42nd Street New York, NY 10017
Royco 22C
Royal Lubricants Co., Inc. River Road East Hanover, NJ 07936
Aeroshell grease 22
Shell Oil Co. One Shell Plaza Houston, TX 77001
GWB
None
Wheel Bearings
Mobil Aviation Grease, SHC 100.
Exxon Mobil Corp. 5959 Las Colinas Blvd. Irving, TX 75039
GH
MIL-PRF23827
Grease, aircraft and instrument, gear and actuator screw.
Southwest Grease 16215
Southwest Petro-Chem, Inc. Division - Witco 1400 S. Harrison Olathe, KS 66061
Aeroshell grease 7
Shell Oil Co.
Royco 27A
Royal Lubricants Co., Inc.
Supermil grease No. A72832
Amoco Oil Co. 200 East Randolph Dr. Chicago, IL 60601
Braycote 6275
Burmah-Castrol, Inc. Bray Products Div. 16815 Von Karman Ave. Irving, CA 92714
Castrolease A1
Burmah-Castrol, Inc.
TG-11900 low temp grease EP
Southwest Petro-Chem,Inc.
Brayco 885
Brumah-Castrol, Inc.
Royco 363
Royal Lubricants Co., Inc.
Petrotect 7870A
Penreco 106 South Main Street Butler, PA 16001
OG
MIL-PRF-7870
Oil, general purpose
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MODEL 208 MAINTENANCE MANUAL Table 1. Recommended Lubricants (continued) SYMBOL
PROCUREMENT SPECIFICATION
DESCRIPTION AND USE
PRODUCT PART NUMBER
VENDOR
Windsor lube L-1018
Anderson Oil & Chemical Co., Inc. Portland, CT 06480
Octoil 70
Octagon Process, Inc. 596 River Road Edgewater, NJ 07020
PL
VV-P-236
Petrolatum technical
Commercially Available
PG
SS-G-659
Powdered graphite
Commercially Available
GL
MIL-G-21164
High and low temperature grease
Everlube 211-G Moly Grease
E/M Corporation Box 2200 Highway 52 N.W. West Lafayette, IN 47906
Royco 64
Royal Lubricants Co., Inc.
GP
None
Number 10 weight, non-detergent oil
Commercially Available
GW
MIL-G-24139
Grease, multi-purpose, water-resistant
OL
VV-L-800
Light Oil
Commercially Available
SG
None
Special DC33
Commercially Available
Aeroshell grease 6
Shell Oil Company One Shell Plaza Houston, TX 77001
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MODEL 208 MAINTENANCE MANUAL BATTERY RECEPTACLE - SERVICING 1.
General A.
2.
It is recommended the airplane be secured in an area free of contamination from sand, dust or other environmental conditions that may contribute to improper lubrication practices.
Servicing A.
Servicing Procedures (Refer to Figure 301). (1) Open right engine cowling to gain access to battery and receptacle. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices. (2) Lubricate fasteners. Refer to Figure 301. (3) Close right engine cowling. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices.
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Battery Receptacle Lubrication Figure 301 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FLIGHT CONTROLS - SERVICING 1.
General A.
If possible, do the airplane servicing in an area free of contamination from sand, dust or other environmental conditions that can contribute to incorrect lubrication procedures.
B.
Lubricate the flight controls at specified time intervals. Refer to Chapter 5, Inspection Time Limits.
C.
Refer to Table 301 and 302 to find the lubrication data.
Table 301. Lubrication Specifics ITEM DESCRIPTION
LUBRICATION TYPE
APPLICATION
FIGURE NUMBER
EFFECTIVITY
Rod Ends
GL
Hand
301, Sheet 1
Drive Sprocket
OG
Oil Can
302, Sheet 1
Drive Shaft
OG
Oil Can
302, Sheet 1
Drive Shaft Bearings
OG
Oil Can
302, Sheet 1
Forward Elevator Control Bell Crank (without Grease Fitting)
GR
Hand
303, Sheet 1
Airplanes 20800001 thru 20800156 and 208B0001 thru 208B0188 Except Airplanes Incorporating SNL89-17.
Forward Elevator Control Bell Crank
N/A
N/A
303, Sheet 1
Airplanes 20800157 and On and 208B0189 and On and Airplanes 20800001 thru 20800156 and 208B0001 thru 208B0188 Incorporating SNL89-17.
Elevator Bell Crank (without Grease Fitting)
GR
Hand
303, Sheet 1
Airplanes 20800001 thru 20800156 and 208B0001 thru 208B0188 Except Airplanes Incorporating SNL89-17.
Elevator Bell Crank (with one Grease Fitting)
GR
Gun
303, Sheet 1
Airplanes 20800199 and On and 208B0228 and On.
Elevator Bell Crank Pushrod (without Grease Fitting)
GR
Hand
303, Sheet 1
Airplanes 20800001 thru 20800156 and 208B0001 thru 208B0188 Except Airplanes Incorporating SNL89-17.
Elevator Bell Crank Pushrod (with one Grease Fitting)
GR
Gun
303, Sheet 1
Airplanes 20800199 and On and 208B0228 and On.
Aft Elevator Control Bell Crank End Bearings
GR
Hand
303, Sheet 1
Airplanes 20800001 thru 20800156 and 208B0001 thru 208B0188 Except Airplanes Incorporating SNL89-17.
Aft Elevator Control Bell Crank End Bearings
GR
Hand
303, Sheet 1
Airplanes 20800157 and On and 208B0189 and On and Airplanes 20800001 thru 20800156 and 208B0001 thru 208B0188 Incorporating SNL89-17.
Aft Elevator Control Bell Crank Center Bearing
GH
Hand
303, Sheet 1
Airplanes 20800001 thru 20800156 and 208B0001 thru 208B01888 Except Airplanes Incorporating SNL89-17.
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MODEL 208 MAINTENANCE MANUAL Table 301. Lubrication Specifics (continued) APPLICATION
ITEM DESCRIPTION
LUBRICATION TYPE
Aft Elevator Control Bell Crank Center Bearing
N/A
N/A
303, Sheet 1
Airplanes 20800157 and On and 208B0189 and On and Airplanes 20800001 thru 20800156 and 208B0001 thru 208B0188 Incorporating SNL89-17.
Pivot Bearing
GR
Hand
304, Sheet 1
Airplanes 20800001 thru 20800207 and 208B0001 thru 208B0277 not Incorporating SK208-83.
Pivot Bearing
N/A
N/A
304, Sheet 1
Airplanes 20800208 and On and 208B0278 and On and Airplanes 20800001 thru 20800207 and 208B0001 thru 208B0277 Incorporating SK208-83 and all spares.
Control Wheel Column Bearings
GR
Gun
304, Sheet 1
Bearing
OG
Oil Can
305, Sheet 1
Bearing
GR
Hand
306, Sheet 1
Airplanes 20800001 thru 20800142 and 208B0001 thru 208B0135 Except Airplanes Incorporating SK208-148.
Bearing
N/A
N/A
306, Sheet 1
Airplanes 20800143 and On and 208B0136 and On and Airplanes 20800001 thru 20800142 and 208B0001 thru 208B0135 Incorporating SK208-148 and all spares.
Rod End Bearing
OG
Oil Can
Figure 306, Sheet 1 and 2
Flap Actuator Screw Threads
GP
Oil Can
Figure 306, Sheet 1 and 2
FIGURE NUMBER
EFFECTIVITY
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MODEL 208 MAINTENANCE MANUAL Table 301. Lubrication Specifics (continued) ITEM DESCRIPTION
LUBRICATION TYPE
APPLICATION
FIGURE NUMBER
EFFECTIVITY
Rod End Bearings
OG
Oil Can
Figure 307, Airplanes 20800001 thru 20800142 Sheet 1 and 208B0001 thru 208B0135 Except Airplanes Incorporating SK208-148.
Bell Crank Bearings
GH
Hand
Figure 307, Sheet 1
Bell Crank Bearings
N/A
N/A
Figure 307, Airplanes 20800143 and On and Sheet 1 208B0136 and On and Airplanes 20800001 thru 20800142 and 208B0001 thru 208B0135 Incorporating SK208-148 and all spares.
Airplanes 20800001 thru 20800142 and 208B0001 thru 208B0135 not Incorporating SK208-148.
Table 302. Recommended Lubricants LUBRICATION TYPE
PROCUREMENT SPECIFICATION
LUBRICANT DESCRIPTION
GH
MIL-PRF-23827
Grease, aircraft and instrument, gear and actuator screw.
GL
MIL-G-21164
Grease, molybdenum disulfide, for low and high temperatures.
GP
None
Number 10 weight, non-detergent oil
GR
MIL-PRF-81322
Grease, aircraft, general purpose, wide temperature range.
OG
MIL-PRF-7870
Lubricating oil, general purpose, low temperature.
2.
Aileron Slot Lip Spoiler Pushrod Lubrication A.
Lubricate the aileron slot lip spoiler pushrods. Refer to Figure 301.
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MODEL 208 MAINTENANCE MANUAL
3.
4.
Aileron Trim Control Assembly Lubrication A.
Lubricate the aileron trim control assembly. Refer to Figure 302.
B.
Do the Aileron Trim System Lubrication. Refer to Aileron Trim System - Inspection/Check.
Lubrication of Elevator Bell Cranks A.
5.
Lubrication of Control Tube and Control Column Bearings A.
6.
Lubricate the control tube and control column bearings. Refer to Figure 304.
Rudder Bar Lubrication A.
7.
Lubricate the elevator bell crank. Refer to Figure 303.
Lubricate the rudder bar. Refer to Figure 305.
Wing Flap Actuator Lubrication A.
Lubricate the wing flap actuator system. Refer to Figure 306.
B.
Lubricate the actuator jack screw. Refer to the steps that follow. NOTE: (1) (2) (3) (4)
8.
This procedure, used in conjunction with Figure 306, lets you lubricate the screw jack while it is in position on the airplane.
Move the flaps to the full up position. Clean the jack screw threads with a rag soaked in solvent. Dry the jack screw with compressed air. Apply a light coat of oil to the jack screw threads.
Flap Bell Crank Lubrication A.
Lubricate the flap bell cranks, interconnect rods, and pushrods. Refer to Figure 307.
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Aileron Slot Lip Spoiler Pushrods Lubrication Figure 301 (Sheet 1)
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Aileron Trim Control Assembly Lubrication Figure 302 (Sheet 1)
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Lubrication of Elevator Bell Cranks Figure 303 (Sheet 1)
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Lubrication of Control Tube and Control Column Bearings Figure 304 (Sheet 1)
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Rudder Bar Lubrication Figure 305 (Sheet 1)
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Wing Flap Actuator Lubrication Figure 306 (Sheet 1)
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Wing Flap Actuator Lubrication Figure 306 (Sheet 2)
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Flap Bell Crank, Interconnect Rods and Pushrods Lubrication Figure 307 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL LANDING GEAR - SERVICING 1.
General A.
This section has the necessary servicing procedures necessary to keep the landing gear system in a serviceable condition.
Task 12-21-03-640 2.
Landing Gear Lubrication A.
If possible do the airplane servicing in an area free of contamination from sand, dust or other environmental conditions that can contribute to improper lubrication procedures.
B.
We recommend the equipment include a grease gun and other tools necessary to do the lubrication procedure.
WARNING: When you clean the wheel bearings, use low pressure shop air to dry the bearings. Do not spin the bearing cones with compressed air. Dry bearings without lubrication can explode at high rpm. CAUTION: Make sure you can put grease into the zerk fitting. If you cannot put grease into the zerk fitting, find the cause and repair it. This will help prevent damage to the equipment. C.
When the lubrication task is completed, clean the unwanted grease from the zerk fitting and from around the bearings where the old and new grease has come out.
D.
Refer to Figure 301 for the lubrication requirements on the nose landing gear. Refer to Figure 302 for the lubrication requirements on the main landing gear. (1) When the wheel is disassembled to lubricate the bearing, or for any other purpose, do the special corrosion protection procedures described in Chapter 32 or the bearing life will be decreased.
E.
Refer to Table 301 and 302 to find the lubrication data.
Table 301. Lubrication Specifics ITEM DESCRIPTION
LUBRICATION TYPE
APPLICATION
FIGURE NUMBER
EFFECTIVITY
Torque Links (with five Grease Fittings)
GL
Gun
301 Sheet 1
Airplanes 20800134 and On and 208B0099 and On and Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Incorporating SK208-51.
Shock Strut
GL
Gun
301 Sheet 1
Airplanes 20800134 and On and 208B0099 and On and Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Incorporating SK208-51.
Shimmy Damper Pivots
OG
Oil Can
301 Sheet 1
Airplanes 20800134 and On and 208B0099 and On and Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Incorporating SK208-51.
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MODEL 208 MAINTENANCE MANUAL Table 301. Lubrication Specifics (continued) ITEM DESCRIPTION
Wheel Bearings
LUBRICATION TYPE GWB
APPLICATION
FIGURE NUMBER
EFFECTIVITY
Hand
301 Sheet 1
Airplanes 20800134 and On and 208B0099 and On and Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Incorporating SK208-51.
Spring Yoke Bearings
OG
Oil Can
301 Sheet 1
Airplanes 20800134 and On and 208B0099 and On and Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Incorporating SK208-51.
Torque Links (with five Grease Fittings)
GL
Gun
301 Sheet 2
Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 not Incorporating SK208-51.
Shock Strut
GL
Gun
301 Sheet 2
Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 not Incorporating SK208-51.
Shimmy Damper Pivots
OG
Oil Can
301 Sheet 2
Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 not Incorporating SK208-51.
Hand
301 Sheet 2
Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 not Incorporating SK208-51.
Oil Can
301 Sheet 2
Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 not Incorporating SK208-51.
Hand
302 Sheet 1
Wheel Bearings
GWB
Spring Yoke Bearings
Wheel Bearings
OG
GWB
Table 302. Recommended Lubricants LUBRICATION TYPE
PROCUREMENT SPECIFICATION
LUBRICANT DESCRIPTION
ALTERNATE
GL
MIL-G-21164
Grease, molybdenum disulfide, for low and high temperatures.
AMS/Oil GHD
OG
MIL-PRF-7870
Lubricating oil, general purpose, low temperature.
GWB
None
Mobil Aviation Grease, SHC 100.
MIL-PRF-81322 Grease, aircraft, general purpose, wide temperature range or AMS/Oil GHD .
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Nose Gear Lubrication Figure 301 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Nose Gear Lubrication Figure 301 (Sheet 2)
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Main Gear Lubrication
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MODEL 208 MAINTENANCE MANUAL Figure 302 (Sheet 1) End of task
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MODEL 208 MAINTENANCE MANUAL PROPELLER (HARTZELL) - SERVICING 1.
General A.
This section has the necessary servicing procedures necessary to keep the Hartzell propeller in a serviceable condition.
Task 12-21-04-640 2.
Hartzell Propeller Lubrication A.
General (1) This task gives the procedures to do the lubrication of the Hartzell propeller.
B.
Access (1) Remove the propeller spinner to get access to the propeller grease fittings. Refer to Chapter 61, Propeller (Hartzell) - Maintenance Practices.
C.
Special Tools (1) Grease Gun
D.
Do the Hartzell Propeller Lubrication (refer to Figure 301).
CAUTION: One zerk fitting must be removed from the clamp to do the lubrication. If it is not removed, too much pressure from the grease gun can rupture the blade clamp seal. (1) (2) (3)
Remove one grease zerk from each clamp. Turn the propeller until one of the open zerk ports is as its highest possible point. Count and record the number of grease gun stokes when you do the lubrication at the initial blade. NOTE:
Servicing the same quantity of grease at each clamp will help keep the static and dynamic balance of the propeller.
(a)
(4) (5) (6) (7)
Pump grease into the zerk fitting until clean grease and all moisture is removed from the open zerk fitting port. Turn the propeller until the next open zerk port is as its highest possible point. (a) Use the recorded number of grease gun strokes from the initial blade to lubricate the blade. Turn the propeller until the open zerk port that remains is at its highest possible point. (a) Use the recorded number of grease gun strokes from the initial blade to lubricate the blade. Use a clean cloth to remove all excess grease from the propeller. Install the zerks in the open ports.
A.
Restore Access (1) Install the propeller spinner. Refer to Chapter 61, Propeller (Hartzell) - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL
Propeller (Hartzell) Lubrication Figure 301 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ENGINE CONTROL ROD ENDS - SERVICING 1.
General A.
It is recommended that the airplane be secured in an area free of contamination from sand, dust or other environmental conditions that may contribute to improper lubrication practices.
B.
Ensure no lubrication enters oilite bushings at either end of the power lever cross shaft.
C.
Lubricate all engine control rod ends after external engine wash or washing of the engine compartment.
D.
Refer to Figure 301 for lubrication requirements of the engine control rod ends.
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MODEL 208 MAINTENANCE MANUAL
Engine Control Rod Ends Lubrication Figure 301 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL EXTERNAL - CLEANING/PAINTING 1.
2.
General A.
The airplane should be washed frequently in order to maintain its appearance and minimize corrosion. The painted area of the airplane should be polished at periodic intervals to remove chalking paint and restore its gloss.
B.
Water/detergent cleaning is the preferred method to clean the exterior surface of the airplane.
Precautions A.
Read and adhere to all manufacturers instructions, warnings and cautions on the cleaning/solvent compounds used. NOTE:
Do not use silicone based car wax.
B.
Do not park or store airplane where it might be subjected to direct contact with fluid or vapors from methanol, denatured alcohol, gasoline, benzene, xylene, methyl n-propyl ketone, acetone, carbon tetrachloride, lacquer thinners, commercial or household window cleaning sprays, paint strippers or other types of solvents.
C.
Do not use solar screens or shields installed on inside of airplane or leave sun visors up against windshield. The reflected heat from these items causes elevated temperatures which accelerate crazing and may cause formation of bubbles in the inner ply of multiple ply windshields.
D.
Do not use a power drill motor or other powered device to clean, polish, or wax surfaces.
E.
Cover static ports prior to wash.
WARNING: Ensure static ports are uncovered after wash. Failure to do so could result in erroneous airspeed and altitude indications. 3.
Preventive Maintenance A.
Keep all surfaces of windshields and windows clean.
B.
If desired, wax acrylic surfaces.
C.
Carefully cover all surfaces during any painting, powerplant cleaning or other procedure that calls for use of any type of solvent or chemical. Table 701 lists approved coatings for use in protecting surfaces from solvent attack.
Table 701. Approved Protective Coatings NAME
NUMBER
MANUFACTURER
USE
Spray
MIL-C-6799, Type 1, Class II
Available Commercially
To protect surfaces from solvents.
Masking Paper
WPL-3
Champion Intl. Corp. Forest Product Division 7785 Bay Meadows Way Jacksonville, FL 32256
To protect surfaces from solvents.
Poly-Spot stick
SXN
Champion Intl. Corp.
To protect surfaces from solvents.
Mask Off Company 345 Marie Avenue Monrovia, CA
To protect surfaces from solvents.
Protex 40
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4.
Approved Windshield/Window Products
CAUTION: Do not use gasoline, alcohol, benzene, acetone, carbon tetrachloride, fire extinguisher fluid, deicer fluid, lacquer thinner or glass window cleaning spray. These solvents will soften and craze the plastic. A.
Table 702 lists Cessna approved materials required for cleaning and polishing windshield and windows used on the airplane.
Table 702. Approved Windshield and Window Cleaners/Polishers NAME
NUMBER
MANUFACTURER
USE
Commercially Available
To clean windshields and windows.
Commercially Available
To remove deposits which cannot be removed with mild soap solution on acrylic windshields and windows.
Turtle Wax (paste)
Commercially Available
For waxing acrylic windshields and windows.
Great Reflections Paste Wax
DuPont 1251 Brandywine Blvd. Wilmington, DE 19898
For waxing acrylic windshields and windows.
Slip-Stream Wax (Paste)
Classic Chemical Grand Prairie, TX 75050
For waxing acrylic windshields and windows.
Mild soap or detergent (hand dishwashing type without abrasives) Aliphatic Naphtha Type II
Federal Specification TTN-95
Permatex Plastic Cleaner No. 403D
Federal Specification P-P-560
Permatex Company, Inc. Kansas City, KS 66115
To clean windshields and windows.
REPCON (Refer to Note)
Federal Specification MIL-W-6862
UNELKO Corp. scottsdale, AZ 85260
Rain shedding on acrylic windshields.
Commercially Available
For applying and removing wax and polish.
Soft cloth (cotton flannel or cotton terry cloth) NOTE:
REPCON is the only windshield rain repellent approved by Cessna Aircraft Company for use on Cessna Model 208 series airplanes.
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5.
Cleaning Windshield and Windows
CAUTION: Windshields and windows are easily damaged by improper handling and cleaning techniques. CAUTION: Do not use any of the following for cleaning windshields and windows: methanol, denatured alcohol, gasoline, benzene, xylene, methyl n-propyl ketone, acetone, carbontetrachloride, lacquer thinners, commercial or household window cleaning sprays.
6.
A.
Refer to Table 702 for approved cleaning materials.
B.
Place airplane inside hanger or in shaded area and allow to cool from heat of suns direct rays.
C.
Using clean (preferably running) water, flood surface. Use bare hands with no jewelry to feel and dislodge any dirt or abrasive materials.
D.
Using a mild soap or detergent (such as dishwashing liquid) in water, wash surface. Again use only bare hands to provide rubbing force. (A clean cloth may be used to transfer soap solution to surface, but extreme care must be exercised to prevent scratching surface.)
E.
On acrylic windshields and windows only, if soils that cannot be removed by a mild detergent remain, Type II aliphatic naphtha applied with a soft clean cloth may be used as a cleaning solvent. Be sure to frequently refold cloth to avoid redepositing soil and/or scratching windshield with any abrasive particles.
F.
Rinse surface thoroughly with clean fresh water and dry with a clean cloth.
Waxing and Polishing Windshield and Windows NOTE:
7.
A.
Refer to Table 702 for approved polishing materials.
B.
Hand polishing wax should be applied to acrylic surfaces. (The wax has an index of refraction nearly the same as transparent acrylic and tends to mask any scratches on windshield surface).
C.
Acrylic surfaces may be polished using a polish meeting Federal Specification P- P-560 applied per manufacturers instructions.
Windshield Rain Repellent A.
8.
Refer to Chapter 56, Windshields and Windows - Maintenance Practices.
Aluminum Surfaces A.
9.
When applying and removing wax and polish, use a clean soft cloth.
Aluminum surfaces require a minimum of care, but should never be neglected. The airplane may be washed with clean water to remove dirt and may be washed with non-alkaline grease solvents to remove oil and/or grease. Household type detergent soap powders are effective cleaners, but should be used cautiously, since some of them are strongly alkaline. Many good aluminum cleaners, polishes and waxes are available from commercial suppliers of airplane products.
Painted External Surfaces A.
Approximately ten days are required for new paint to cure completely. In most cases, the curing period will have been completed prior to delivery of the airplane. If polishing or buffing is required within the ten day curing period, the work should be performed by people experienced in handling uncured paint. Cessna dealers can perform this work.
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10.
B.
Generally, the painted surfaces can be kept bright by washing with water and mild soap, followed by a rinse with water and drying with cloths or a chamois. Harsh or abrasive soaps or detergents which could cause corrosion or scratches should never be used. Remove stubborn oil and grease with a cloth moistened with Stoddard solvent.
C.
To seal any minor surface chips or scratches and protect against corrosion, the airplane should be waxed regularly with a good automotive wax applied in accordance with the manufacturers instructions. If the airplane is operated in a seacoast area or other salt water environment, it must be washed and waxed more frequently to assure adequate protection. Special care should be taken to seal around rivet heads and skin laps, which are the areas susceptible to corrosion. A heavier coating of wax on the leading edges of the wings and tail and on the cowl nose cap and propeller spinner will help reduce the abrasion encountered in these areas. Reapplication of wax will generally be necessary after cleaning with soap solutions or after chemical deicing operations.
Engine Compressor Wash A.
11.
The compressor section of the engine requires a desalination wash routinely. Operating environment determines washing frequency. Refer to Chapter 71, Compressor/Turbine Blade Wash - Maintenance Practices.
Engine Compartment A.
A wash down of engine and accessories should be performed routinely to remove oil, grease, salt corrosion, and other residue. Periodic cleaning can be an aid to discovering defects during inspection.
B.
Precautions should be taken when working with cleaning agents, such as, wearing of rubber gloves, an apron or coveralls and a face shield or goggles. Use the least toxic of available cleaning agents that will satisfactorily accomplish the work. These cleaning agents include; (1) Stoddard Solvent (Specification P-D-680 type II), (2) a water alkaline detergent cleaner (MIL-C-25769) mixed, 1 part cleaner, 2 to 3 parts water and 8 to 12 parts Stoddard solvent or (3) a solvent base emulsion cleaner (MIL-C-4361) mixed 1 part cleaner and 3 parts Stoddard solvent.
CAUTION: Cover propeller to prevent contact with alkaline detergent during wash. CAUTION: Do not use gasoline or other highly flammable substances for wash down. CAUTION: Do not attempt to wash an engine which is still hot or running. Allow the engine to cool before cleaning. C.
12.
Perform all cleaning operations in well ventilated work areas and ensure that adequate firefighting and safety equipment is available. Do not smoke or expose a flame, within 100 feet of the cleaning area. Compressed air, used for cleaning agent, application or drying, should be regulated to the lowest practical pressure. Use of a stiff bristle brush rather than a steel brush is recommended if cleaning agents do not remove excess grease and grime during spraying.
Propeller A.
Clean propeller regularly with water and a mild detergent to remove grass and bug stains.
CAUTION: Alkaline detergents not to be used on propellers. 13.
Wheels A.
The wheels should be washed periodically and examined for condition, chipped paint and cracks or dents. Sand smooth, prime and repaint minor defects; however, cracked wheel halves should be replaced.
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14.
Deice Boots A.
15.
Wing, wing strut, stabilizer, propeller boots, and cargo pod boot (if installed) should be washed and serviced routinely. Keep boots clean and free from oil, grease and other solvents which cause rubber to swell and deteriorate. For cleaning and servicing procedures of the deice boots, refer to Deicing Servicing.
TKS Panels A.
To keep a good appearance of the TKS panels and help make sure TKS system operates correctly, refer to TKS Anti-Ice System - Servicing.
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MODEL 208 MAINTENANCE MANUAL INTERIOR - CLEANING/PAINTING 1.
General A.
2.
This section is designed to assist the operator and recommend different types of cleaning materials and cleaning procedures for the interior of the airplane.
Airplane Interior Cleaning Materials
WARNING: Cleaning operations using solvent should be performed in a well vented atmosphere. Exercise normal safety precautions during use. NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Yosemite
Y-999
Yosemite Chemical Co. 1248 Wholesale St. Los Angeles, CA 90021
For cleaning vinyl-coated fabrics, mylar, Scotchcal murals, polyplastex, leathers, vinyl flooring, Formica, linoleum, finished flexwood or painted surfaces.
Aliphatic Naphtha
TT-N-95ir
Commercially Available
Cleaning interior decorative material and furnishings.
Host Dry Cleaning Compound
Host of California 2935 Coleridge Ave. Pasadena, CA 91107
For cleaning drapes, curtains, upholstery, fabrics and carpet.
Wet Rug Shampoo
Commercially Available
Carpeting.
Perchloroethylene
Commercially Available
Spot clean carpet.
Stoddard Solvent
Commercially Available
Cleaning nylon safety belts.
Mild Soap Detergent
Commercially Available
Cleaning nylon safety belts. Cleaning plastic.
A.
Cleaning Interior Decorative Materials. (1) Clean with Yosemite Y-999 (or equivalent) as follows: (a) Spray or wipe on the soiled surface. (b) Wipe off with a clean cloth dampened in water. (2) Clean with Aliphatic Naphtha as follows: (a) Wipe with a clean cloth dampened with naphtha and wipe dry with a clean cloth. (b) When removing tar, asphalt or chewing gum, remove as much as possible with a knife. Apply naphtha to the residue and then wipe dry with a clean cloth; this has a buffing effect that eliminates the possibility of stain from the solution.
B.
Cleaning Rugs, Drapes, Curtains and Upholstery Fabrics. (1) Dry clean commercially. (2) Host dry cleaning compound. (a) Sprinkle the compound liberally on the soiled area. (b) Rub the compound into the soiled area. (c) Remove the compound with a vacuum cleaner. NOTE:
This compound is nonflammable and may be used on fueled airplanes.
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Wet (a) (b) (c)
shampoo. Remove carpet or upholstery from airplane. If at all possible, use the spot cleaning method. Vacuum the carpet and upholstery, removing as much dirt and dust as possible. Place a tablespoon of shampoo in a pail and direct a jet of water into the shampoo to produce abundant foam. (d) Apply the foam uniformly over the surface to be cleaned. (e) Remove the suds by wiping with a brush or clean cotton cloth. Since there is very little moisture in the foam, wetting of the fabric or retention of moisture will not occur.
CAUTION: Use of mechanical shampooing equipment may distort the carpet. (4)
Spot cleaning. (a) Spot-clean tufted carpet in the airplane, if at all possible, rather than completely removing the carpet for shampooing. (b) Saturate a clean white cloth with perchloroethylene solution.
CAUTION: Do not pour perchloroethylene solution directly on the carpet. (c)
Hand-rub the perchloroethylene- saturated cloth in a circular motion on the soiled spot.
CAUTION: Do not use mechanical shampooing equipment; it may distort the carpet. Upholstery hand shampooing equipment may be utilized on areas which are difficult to clean. For cleaning acrylic plastic, refer to External - Cleaning. (d)
(5)
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MODEL 208 MAINTENANCE MANUAL UNSCHEDULED SERVICING - DESCRIPTION AND OPERATION 1.
General A.
This section outlines procedures and recommendations necessary to carry out servicing that is normally unscheduled. Instructions for ice and snow removal from the parked airplane are provided.
B.
Deicing procedures are included in the section to assist personnel in removing ice from the airplane. NOTE:
C. 2.
Ensure that all chemical supplier’s instructions including bulletins, warnings and cautions are adhered to.
This section also includes procedures used to clean and protect the deice boots on the airplane.
Tools, Equipment and Material NOTE:
The following table provides manufacturer names and addresses for materials used in unscheduled servicing.
NAME
NUMBER
MANUFACTURER
USE
MISCO
7084VP
American Optical Corp. 3401 Virginia Road Cleveland, OH 44122
To test refractive index of deice fluids.
ICEX
BFGoodrich 1555 Corporate Woods Parkway P.O. Box 1277 Uniontown, OH 44685
To promote ice shedding from deice boots.
AGE MASTER NO. 1
Chem-Pro Manufacturing P.O. Box 213 Buffalo, NY 14221
To protect deice boots against deterioration.
ShineMaster
BFGoodrich
To protect deice boots which have been resurfaced using Estane coating.
ShineMaster Prep
BFGoodrich
To remove ShineMaster coating from deice boots.
3.
Extreme Weather Maintenance A.
4.
Seacoast and Humid areas. (1) In salt water areas special care should be taken to keep engine, accessories and airframe clean to help prevent oxidation. (2) In humid areas, fuel and oil should be checked frequently and drained of condensation to prevent corrosion.
Ground Power Receptacle A.
Connect to 24-volt DC, negative-ground power unit with a maximum output of 28.8 volts, for cold weather starting, and lengthy ground maintenance of the airplane electrical equipment, with exception of electronic equipment. Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual for engine starting instructions with auxiliary power. NOTE:
The ground power receptacle circuit incorporates a polarity reversal protection. Power from the external power source will flow only if the ground service plug is connected correctly to the airplane.
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MODEL 208 MAINTENANCE MANUAL DEICING/ANTI-ICING - DESCRIPTION AND OPERATION 1.
General A.
FAA regulations require that all critical components (wings, control surfaces and engine inlets as an example) be free of snow, ice or frost before takeoff. The deicing process is intended to restore the airplane to a clean conÞguration so that neither aerodynamic characteristics nor mechanical interference from contaminants will occur.
B.
Deicing and anti-icing ßuids are aqueous solutions which work by lowering the freezing point of water in either the liquid or crystal phase which delays the onset of freezing. For this reason, they are referred to as a Freezing Point Depressant (FPD) ßuids. Deicing ßuid is classiÞed as Type I. Anti-icing ßuid is classiÞed as Type II, Type III, or Type IV. The one-step method of airplane deicing utilizes only Type I ßuid. The two-step approach to airplane deicing utilizes Type I ßuid to deice the airplane, which is followed rapidly by application of Type II, Type III, or Type IV ßuid to delay the onset of refreezing. (1) Type I, Type II, Type III, and Type IV ßuids have time limitations before refreezing begins. This time limitation is referred to as “holdover time,” Type II, Type III, and Type IV anti-icing ßuids have a much longer holdover time than Type I deicing ßuids. Because holdover time is highly dependent on a number of factors, charts can provide only approximate estimates. Refer to speciÞc manufacturers data sheets for holdover times. It remains the responsibility of the ßight crew to determine the effectiveness of any deicing procedure.
CAUTION: You cannot mix Type I, Type II, Type III, and Type IV ßuids because they are not compatible. Also, most manufacturer's prohibit mixing of brands within a type. C.
Deicing (1) Deicing may be accomplished using the ambient temperature available from a heated hangar or by mechanical means using a glycol-based Freezing Point Depressant (FPD) Type I ßuid. (a) Care must be exercised, however, to ensure that all melted precipitation is removed from the airplane to prevent refreezing once the airplane is moved from the hangar to the ßight line. (b) Type I deicing ßuids are applied in a temperature range from 160°F to 180°F (71°C to 82°C) using a moderate to high-pressure washer. Heated solutions of FPD are more effective than unheated solutions because thermal energy is used to melt the ice, snow or frost formations. Type I deicing ßuids are used in the diluted state, with speciÞc ratios of ßuid-to-water dependent on ambient temperature. Type I deicing ßuids have a very limited holdover time.
D.
Anti-icing (1) Anti-icing is accomplished by using Type II, Type III, or Type IV ßuids, and their purpose is to delay the reforming of ice, snow or frost on the airplane. This is accomplished by using chemically thickened formulas with pseudo-plastic properties. This feature enables the ßuid to form a protective Þlm on treated surfaces of the airplane, and is designed to ßow off airplane surfaces at high speeds.
CAUTION: Type II, Type III, and Type IV ßuids are designed for use on airplanes with a VR speed of 85 knots or greater. Type II, Type III, and Type IV ßuid is used undiluted and is typically applied to the airplane unheated. Holdover times for Type II, Type III, and Type IV ßuid can vary widely based on atmospheric conditions. Consult speciÞc manufacturers charts for holdover time. E.
Deicing ßuids are not intended for use in removing snow deposits. Snow is best removed by mechanically sweeping or brushing it from the airplane structure.
F.
Deicing procedures must be closely coordinated with the ßight crew and carried out in a timely manner. Ultimate responsibility for safety of ßight rests with the ßight crew, and any decisions to deice an airplane must be accomplished under their direct supervision.
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2.
Deicing/Anti-Icing Fluids NOTE: A.
Equivalent substitutes may be used for the following tables:
Currently available Type I Deicing Fluids are listed in Table 1:
Table 1. SAE Type I and ISO Type I Deicing Fluids NAME
MANUFACTURER
COLOR
CHEMICAL BASE
UCAR ADF Concentrate
Union Carbide 10235 West Little York Rd., Suite 300 Houston, TX 77040
Orange
Ethylene-glycol
UCAR ADF 50/50
Union Carbide
Orange
Ethylene-glycol
ARCOPLUS Dilute
ARCO Chemical Company 3801 West Chester Pike Newtown Square, PA 19073
Orange
Propylene-glycol
ARCOPLUS
ARCO Chemical Company
Orange
Propylene-glycol
B.
Currently available Type II Anti-Icing Fluids are listed in Table 2:
Table 2. SAE Type II Anti-Icing Fluids NAME
MANUFACTURER
COLOR
CHEMICAL BASE
KILFROST ABC-3
ARCO Chemical Company 3801 West Chester Pike Newtown Square, PA 19073
Pale Amber
Propylene-glycol
UCAR UC5-1
Union Carbide 10235 West Little York Rd., Suite 300 Houston, TX 77040
Pale Yellow
Ethylene-glycol
UCAR AAF ULTRA
Union Carbide
Emerald Green
Ethylene-glycol
C.
Currently available Type III Anti-Icing Fluids are listed in Table 3:
Table 3. QualiÞed Type III Deicing/Anti-Icing Fluids NAME
MANUFACTURER
COLOR
CHEMICAL BASE
Safewing MP III 2031 ECO
Clariant Corporation 60050 McHenry, IL
Bright yellow
50% Propylene Glycol
D.
Currently available Type IV Anti-Icing Fluids are listed in Table 4:
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Table 4. SAE Type IV Deicing/Anti-Icing Fluids NAME
MANUFACTURER
COLOR
CHEMICAL BASE
UCAR ADF/AAF ULTRA+
Union Carbide 10235 West Little York Rd., Suite 300 Houston, TX 77040
Emerald Green
Ethylene-glycol
E.
Currently available TKS Anti-Icing Fluids are listed in Table 5. Additional TKS Anti-Icing Fluid data is found in Replenishing - Description and Operation.
Table 5. TKS Anti-Icing Fluids NAME
MANUFACTURER
SPECIFICATION
CHEMICAL BASE
AVL-TKS
Aviation Laboratories 5401 Mitchelldale #B6 Houston, TX 77092
DTD406B
Monoethylene glycol/isopropyl alcohol/deionized
TKS-Fluid
DW Davies 3200 PHILLIPS AVENUE RACINE, WI 53403
DTD406B
Monoethylene glycol/isopropyl alcohol/deionized
AeroShell Compound 07
AeroShell Shell Oil Company One Shell Plaza Houston, TX 77001
DTD406B
Monoethylene glycol/isopropyl alcohol/deionized
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MODEL 208 MAINTENANCE MANUAL DEICING/ANTI-ICING - SERVICING 1.
2.
General A.
This servicing section is to supply maintenance personnel with information necessary to remove the snow, ice or frost from the airplane when conditions exist (or are possible). deicing and anti-icing procedures are to be completed in conjunction with the ßight crew. The pilot-in-command makes the Þnal decision if an airplane's components are free of frozen contaminants.
B.
The effectiveness of any Freezing Point Depressant (FPD) deicing or anti-icing treatment can only be estimated because of the many variables that inßuence holdover time. Those variables are: (1) Ambient temperature (2) Airplane surface temperature (3) Freezing Point Depressant ßuid application procedure (4) Freezing Point Depressant solution strength (5) Freezing Point Depressant Þlm thickness (6) Freezing Point Depressant ßuid temperature (7) Freezing Point Depressant ßuid type (8) Operation in close proximity to other airplanes, equipment and structures (9) Operation on snow, slush, wet ramps, taxiways and runways (10) Precipitation type and rate (11) Residual moisture on airplane surface (12) Relative humidity (13) Solar radiation (14) Wind speed and direction
Approved Products A.
3.
For a list of Type I deicing ßuids, refer to Deicing - Description and Operation.
Deicing/Anti-Icing Precautions A.
Before Type I deicing procedures begin, maintenance personnel must familiarize themselves with areas to be sprayed and areas to avoid a direct spray of ßuid. Refer to Figure 301 for areas to be sprayed. Refer to Figure 302 for areas to apply anti-ice ßuid. Refer to Figure 303 for critical areas to avoid spraying directly. Refer to Figure 304 for application sequence.
B.
Type I deicing ßuids must never be used at full strength (undiluted). Undiluted glycol ßuid is quite viscous below 14°F (-10°C) and can actually produce lift restrictions of about 20 percent. Undiluted glycol has a higher freezing point than glycol/water mixture.
C.
Deicing procedures must be completed with the engine off.
D.
Before Type II, Type III or Type IV anti- ice procedures begin, maintenance personnel must familiarize themselves with the areas to be sprayed and areas not to be sprayed. Anti-icing is applied primarily to protect the wings, control surfaces and fuselage. Refer to Figure 302 for areas that receive the anti-ice application, Figure 303 for areas that do not receive the anti-ice aplication, and Figure 304 for the sequence of the application.
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Essential Areas to be Deiced Figure 301 (Sheet 1)
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Areas to Apply Anti-Ice Fluid Figure 302 (Sheet 1)
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Deicing Fluid Minimum Direct Spray Areas Figure 303 (Sheet 1)
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Deicing Application Figure 304 (Sheet 1)
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CAUTION: Although irritation from freezing point depressant fumes is classiÞed as negligible, maintenance personnel must wear protective clothing during deicing/anti-icing procedures. Pure glycol, if swallowed in amounts of three ounces or more, can be fatal. Maintenance personnel must familiarize themselves with the Manufacturers Material Safety Data Sheet (MSDS) before deicing/anti-icing procedures begin. 4.
Type I Deicing Preparations A.
Before the deicing procedures begin, maintenance personnel must know the lowest possible outside air temperature (OAT). Based on this information, the glycol/water mixture must then be adjusted to lower the freezing point of the Type I solution to at least 18°F (10°C) below this OAT. The difference between the possible OAT and the freezing point of the solution is known as the “buffer.” (1) Each manufacturer has speciÞc instructions for mixing glycol/water and the freezing point that any given mixture will provide. Refer to these instructions when preparing Type I solutions. (2) Most manufacturers provide a refractive index of their products. This index is required to ascertain the freezing point of any given solution. Refer to Unscheduled Servicing - Description and Operation for a list of manufacturers offering refractive index testing kits.
WARNING: It is the responsibility of the pilot and deicing personnel to know the freezing point of any solution they apply. A refractive index coupled with speciÞc manufacturers data is the only positive method for identifying the freezing point of a previously mixed Type I solution whose glycol/water ratio is unknown. CAUTION: Type I deicing ßuid must not be intermixed between brands. Manufacturers add speciÞc dyes to their product for visual evidence of contamination. A ßuid that does not meet the color criteria made by the manufacturer must be considered to be contaminated and must not be used. B. 5.
Make sure that Type I deicing ßuid is between 160°F and 180°F (71°C and 82°C) before application begins.
Type II, Type III and Type IV Anti-Ice Preparations A.
Type II, Type III and Type IV anti-icing ßuids undiluted and at ambient temperature unless otherwise speciÞed by the manufacturer. (1) Type I, Type II, Type III, and Type IV ßuids have time limitations before refreezing begins. This time limitation is referred to as “holdover time,” Type II, Type III, and Type IV anti-icing ßuids have a much longer holdover time than Type I deicing ßuids. Because holdover time is highly dependent on a number of factors, charts can provide only approximate estimates. Refer to speciÞc manufacturers data sheets for holdover times. It remains the responsibility of the ßight crew to determine the effectiveness of any deicing procedure. NOTE:
Type II, Type III, and Type IV anti-ice ßuid has thickening agents, which are designed to remain on the wings of an airplane during ground operations or short term storage to provide some anti-ice protection. This ßuid is also made to ßow off readily during takeoff at speeds of approximately 85 knots. Type II, Type III, and Type IV anti-ice procedures supply longer holdover times than Type I deicing procedures.
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CAUTION: You must apply Type II, Type III, and Type IV anti-ice ßuid, must never be mixed with Type I deicing ßuid. Anti-ice ßuids, require dedicated equipment and must not be dispersed with equipment used for Type I deicing ßuid. Type II, Type III and Type IV anti-ice ßuids must not be mixed between brands. WARNING: Refer to the manufacturers instructions for low temperature limits. If a Type II, Type III, or Type IV anti-ice ßuid is applied at temperatures lower than those approved by the manufacturer, the ßuid may remain on the airplane and severely inhibit lift characteristics. B.
6.
Make sure to set the dedicated Type II, Type III, or Type IV equipment to apply low-to-moderate pressure ßuid. Because you apply Type II, Type III, or Type IV anti-ice ßuids immediately after the Type I deicing procedure, you must fully service the Type II, Type III, or Type IV equipment before Type I deicing begins.
Deicing Procedures A.
Remove heavy quantities of snow with brooms or other equivalent methods. Carefully brush around antennas, windows, ßight controls, TKS anti-ice panels, probes, vanes and other airframe equipment.
B.
Refer to Figure 301 for areas to spray. Refer to Figure 303 for areas to avoid spraying directly. Refer to Figure 304 for instructions of application.
C.
If the deicing is to be followed by anti-ice, the anti-ice must begin immediately after completion of the deicing procedure. NOTE:
It is the heat of the deicing ßuid that melts ice and snow. The only function of glycol in the deicing solution is to lower the freezing point of the ßuid which remains on the airplane.
D.
Spraying Hints For Type I Fluid (1) The ßuid must be sprayed on the airplane in a manner that decreases heat loss of ßuid to the air. The ßuid must be sprayed in a solid cone pattern of large coarse droplets. (2) The ßuid must be sprayed as close as possible to the airplane surfaces, but not closer than approximately 10 feet if a high pressure nozzle is used. (3) If a thick layer of frozen snow or ice is on the airplane surface, it is better to concentrate a directed spray of heated ßuid on one area until that section of the airplane is cleaned. The hot ßuid will heat the airplane surface, and the heated surface will help loosen the frozen bond of ice and snow around the cleaned area. (4) When spraying the wing and tail areas, spray from the tip inboard and from the leading to the trailing edge. This procedure takes advantage of dihedral to aid in ßuid dispersion. (5) Make sure the upper fuselage is clear. (6) Windshields and windows must not be sprayed directly. (7) Pitot tubes and static ports must not be sprayed directly.
E.
Deicing the airplane Refer to Figure 304 for application sequence. NOTE:
(1) (2) (3) (4) (5) (6)
Record the time the deicing procedures begin. The length of time that deicing ßuids remain effective is known as “holdover time” and is highly dependent on a number of variables. Refer to Pilot’s Operating Handbook and FAA approved Airplane Flight Manual for Type I deicing ßuid approximate holdover times.
Deicing pilot side fuselage from spinner to wing strut area. Deicing upper fuselage above cockpit area (pilot side). Deicing left wing. Deicing left fuselage from wing strut area to tail. Deicing tail section - left side. Deicing tail section - right side.
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9) (10) (11) (12)
7.
Deicing right fuselage from tail to wing strut area. Deicing right wing. Deicing upper fuselage above cockpit area (copilot side). Deicing copilot side fuselage from wing strut area to spinner. Complete post-deicing checks. Refer to Post- Application Checks. Convey deicing information to ßight crew with the following statement: “This airplane has been deiced using Type I deicing ßuid with a freezing point of _______°F. Holdover time began at _________.”
Do an Anti-Ice Procedure on the Airplane A.
Anti-Ice the airplane.
WARNING: You must never apply Type II, Type III, or Type IV anti-ice ßuids to pitot heads, control surface cavities, cockpit windows, the windshield, the fuselage nose, static ports, air inlets, or engine. NOTE:
Record the time that the anti-ice procedures begin. The length of time an anti-ice ßuid remains effective is known as “holdover time” and is dependent on a number of variables. Refer to the appropriate manufacturer's information or the latest FAA "ßight Standard Information Bulletin for Air Transportation (FAST)" for the approximate holdover time of anti-icing ßuid, in undiluted form.
NOTE:
Type II, Type III, or Type IV anti-ice ßuid must be applied in three minutes after deicing is complete due to the limited holdover time of Type I deicing ßuid. If Type II, Type III, or Type IV anti-ice ßuid has been applied and the airplane has not been dispatched before new ice has formed, the airplane must be completely deiced again and a second anti-ice treatment applied immediately.
NOTE:
Anti-ice ßuid is applied to the airplane at low pressure to form a thin Þlm on its surfaces. Type II, Type III or Type IV anti-ice ßuids must cover the airplane surfaces without runoff. Type II, Type III, or Type IV anti-ice ßuids are applied only from the wing section aft.
(1) (2) (3) (4) (5) (6) (7)
8.
Refer to Figure 302 for the areas where you must to apply anti-ice ßuid. Refer to Figure 303 for the speciÞc areas where you must not apply anti-ice ßuid. Refer to Figure 304 for the application sequence. Apply anti-ice ßuid to the left wing. Apply anti-ice ßuid to the left tail section and empennage. Apply anti-ice ßuid to the right tail section and empennage. Apply anti-ice ßuid to the right wing. Complete a post-application check. Refer to Post-Application Checks. Convey anti-ice information to the ßight crew with the following statement: “ This airplane has been anti-iced using Type II, Type III, or Type IV anti-ice ßuid. Holdover time began at _________”.
Post-Application Checks A.
After the airplane has been deiced, maintenance personnel must perform a post - application check to make sure that all the critical areas are free of ice, snow or slush. These critical areas are as follows: (1) Wing leading edges, upper surfaces and lower surfaces. (2) Horizontal and vertical stabilizers. (3) All control surfaces and control surface gaps. (4) Spoilers. (5) Windshields for clear visibility. (6) Engine inlets. (7) Antennas. (8) All pitot and static probes/ports. (9) Fuel tank and fuel cap vents.
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MODEL 208 MAINTENANCE MANUAL (10) Air inlet scoops. (11) Landing gear. 9.
Post-Flight Clean Up A.
10.
It is highly recommended that airplanes which have undergone deicing or anti-icing procedures be thoroughly cleaned after ßight operations are completed. Refer to External - Cleaning procedures.
Wheel Brake Deicing Procedure A.
Wheel Brake Deicing (1) If the brakes freeze from ice forming after the airplane has been parked on the ramp, and full deicing procedures are not required, the following must be completed to remove the ice from the brake area.
CAUTION: Exercise care when you use a ground heater to deice the brakes if the airplane is setting on ice or is in close proximity to other parked airplanes. (a) (b) (c) (d) 11.
Use a ground heater if available. Spray or pour isopropyl alcohol on the brakes. Cycle the brakes asymmetrically while you apply engine power. In slush conditions, spraying alcohol on the brakes before taxi and takeoff will help prevent the brakes from freezing in ßight.
Deicing Boot Cleaning A.
Clean the Deicing Boots. NOTE:
Boots on the wings, struts, stabilizers, propeller and cargo pod (if installed) must be washed and serviced routinely.
CAUTION: Do not clean with petroleum based liquids such as methyl-n-propl ketone, unleaded gasoline, etc. You must be careful with the deicing boots to prevent damage. The deicing boots have an electrical coating to bleed off the static charges and could make holes in the tail deicing boots. CAUTION: The temperature of the water must not be more than 140°f (60°C). (1)
Clean the deicing boots with mild soap and water, then rinse thoroughly with clean water. NOTE:
(2) 12.
Isopropyl alcohol can be used to remove grime which cannot be removed using soap. If isopropyl alcohol is used for cleaning, wash area with mild soap and water, then rinse thoroughly with clean water.
Keep boots clean and free from oil, grease and other solvents which will cause the rubber to swell and deteriorate.
Deicing Boot Preservation A.
You can get a longer service life out of the deicing boots and reduce the adhesion of ice to them if you apply AGE MASTER No. 1 and ICEX II or SHINEMASTER PREP, SHINEMASTER, and ICEX II to the them.
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NAME
MANUFACTURER
USE
ICEX II
Aviall Distribution Dallas, TX Phone: 800-284-2551 Email: www.aviall.com
To increase ice shedding from deicing boots.
API Memphis, TN Phone: 800-450-6777 Email: www.apiparts.com AAR Wood Dale, IL Phone: 877-227-6900 Email: www.aarcorp.com AGE MASTER No. 1
Any of the manufactures listed above
To protect the deicing boots against deterioration
RESURFACING KIT 74-451-L.
Any of the manufactures listed above
To resurface the deicing boots with a tough layer of estane
SHINEMASTER PREP
Any of the manufactures listed above
To clean the deicing boots before you apply SHINEMASTER
SHINEMASTER
Any of the manufactures listed above
To improve the cosmetic high gloss look of the deicing boots
(1) (2)
B.
Apply AGE MASTER No. 1 and ICEX II to the deicing boots if you have not resurfaced them with Goodrich Resurfacing Kit 74-451-L. Apply SHINEMASTER and ICEX II to the deicing boots if you have resurfaced them with Goodrich Resurfacing Kit 74-451-L. This kit applies a tough Þlm of estane to the deicing boots.
NOTE:
AGE MASTER No. 1 is used to protect the rubber of the deicing boots against deterioration from ozone, sunlight, weathering, oxidation, and pollution.
NOTE:
ICEX II is used to help retard ice adhesion and to extend the appearance of new deicing boots.
NOTE:
Goodrich Resurfacing Kit 74-451-L is used to put a new surface made of a tough layer of estane on the deicing boots.
NOTE:
SHINEMASTER PREP and SHINEMASTER restore a cosmetic, high-gloss look to the deicing boots and adds a layer of estane to the deicing boot surface.
You must do the application of both AGE MASTER No. 1 and ICEX II in accordance with the instructions that follow:
CAUTION: Protect adjacent areas and clothing, and use plastic or rubber gloves during applications. AGE MASTER No. 1 stains and ICEX II contains silicone which makes paint touch-up almost impossible. CAUTION: Make sure that you follow the manufacturers warnings and cautions when you apply AGE MASTER No. 1 and ICEX II. (1)
AGE MASTER No. 1 Application Instructions: NOTE:
Apply AGE MASTER No. 1 every 6 months or 150 ßight hours, whichever comes Þrst.
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CAUTION: You must not use AGE MASTER No. 1 on deicing boots if they have been treated with Goodrich Resurfacing Kit 74-451-L. The resurfacing kit applies a layer of estane to the deicing boots, which is not compatible with AGE MASTER No. 1 treatments. You must use SHINEMASTER in place of AGE MASTER No 1 on deicing boots that were treated with a layer of estane. CAUTION: Protect adjacent areas and clothing, and use plastic or rubber gloves during applications. AGE MASTER No. 1 stains and ICEX II contains silicone which makes paint touch-up almost impossible. (a) (b)
Clean all oil, grease, and wax from the deicing boot surfaces before you apply AGE MASTER No. 1. Put masking tape on the adjacent areas to protect the painted surfaces. For best results and appearance, apply AGE MASTER No. 1 in a single, ful,l and even continuous motion (spanwise) with a soft clean cloth. Let AGE MASTER No. 1 penetrate into the rubber.
CAUTION: You must not apply AGE MASTER No. 1 as a spray. It can cause damage to related components, and it is a Þre hazard. 1
You must apply three layers of AGE MASTER No. 1 for a complete treatment. NOTE:
Additional applications of AGE MASTER No. 1 can be necessary for older deicing boots.
NOTE:
The total amount of AGE MASTER No. 1 applied in all (three coats) must not exceed 0.3 to 0.4 ounces per square foot. One quart of AGE MASTER No. 1 will treat 80 to 106 square feet in a three application process.
Allow 5 to 10 minutes for AGE MASTER No.1 to dry before you apply another layer. The time it takes to dry can vary with weather conditions and/or temperature. (c) Use waterless hand cleaner to clean your hands and equipment and to remove stains from your clothing after you Þnish your work with AGE MASTER No.1. ICEX II Application Instructions: 2
(2)
NOTE: (a) (b) (c)
ICEX II must be applied every 50 ßight hours to airframe deicing boots and every 15 ßight hours to propeller deicing boots.
Clean the deicing boot surface with a mild soap and water solution before you apply ICEX II. After you clean the deicing boots, rinse them fully with clean water and let them dry fully. Use isopropyl alcohol to remove substances that you cannot remove with soap and water. After you clean the deicing boots with isopropyl alcohol, clean the surface again with mild soap and water, rinse fully with clean water, and let them dry fully. NOTE:
(d)
Do not apply ICEX II on surfaces that are treated with AGE MASTER No. 1 until they have dried a minimum of 25 hours.
Apply ICEX II with a soft clean cloth in a single continuous motion (spanwise) without further cleaning. Make sure you fully cover the deicing boot with ICEX II.
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CAUTION: You must not apply ICEX II as a spray. It can cause damage to related components, and it is a Þre hazard. NOTE:
(3)
If you use too much ICEX II in an application, the result will be a sticky surface that will collect dust and dirt. This will reduce the efÞciency of the ICEX II. One quart of ICEX II covers approximately 500 square feet.
SHINEMASTER Application Instructions. (a) Refer to the speciÞc manufacturer's instructions on how to apply SHINEMASTER on estane coated boots.
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE SYSTEM - SERVICING 1.
2.
General A.
The TKS anti-ice system fluid tank is attached to the bottom of the fuselage. Airplanes can have the TKS installed in the cargo pod or in the optional fairing. On airplanes with the cargo pod, the filler opening is on the left side of the pod. On airplanes with the fairing, the filler opening is on the left side of the fuselage. The fluid tank is vented overboard by vent lines in the top of the tank. A sight glass is installed in the left side of the fluid tank to give a visual indication of the fluid level.
B.
The TKS anti-ice system fluid tank is serviced with approved anti-icing fluids only. For a list of approved TKS anti-icing fluids, refer to Deicing/Anti-Icing - Description and Operation.
Servicing
WARNING: Keep anti-icing fluid away from heat, sparks, and open flame. Anti-icing fluid is classified as a combustible liquid. It is also harmful or fatal if swallowed. Avoid prolonged or repeated breathing of vapors. Avoid prolonged skin contact. Do not store in open or unlabeled containers. If swallowed, induce vomiting and get medical attention immediately. If fluid comes in contact with eyes, flush with large amounts of water and get medical attention immediately. A.
Cleanliness Precautions (1) The TKS system has filtration to protect the components against damage and blockage from particulate matter, but they are not always effective for liquid contaminants. To extend the life of the filter element and strainer, the following precautions are recommended: (a) Where possible minimize the number of containers that you use to store and fill the airplane anti-ice fluid tank. Purchase anti-ice fluid in small (2.5 gallon) containers that you can use to pour it directly in the airplanes fluid tank. (b) Always clean the top of the containers before you remove the cap to pour from the container. (c) Always replace the cap on the fluid containers that you use for the transfer of the fluid. (d) If you transfer fluid from the original container to other containers for storage, make sure that these containers are clean and of correct materials. NOTE: (e) (f) (g)
B.
Tin plated steel or similar containers that can rust are not the correct type.
Keep a set of containers and implements solely for use with anti-ice fluid. Keep the area around the airplane TKS filler clean. Keep the filler cap clean. Always replace the filler cap on the airplane immediately after you fill the tank.
Service the TKS Fluid Tank (Refer to Figure 301 for fairing installation and, Figure 302 for cargo pod installation as applicable). (1) Supply external electrical power to the airplane. (2) Monitor the fluid level indications in the cockpit to make sure the indications agree with the sight gage level. (3) On airplanes with the fairing open the fluid filler cap. (4) On airplanes with the cargo pod, open the cargo pod forward center door to get access to the fluid tank. (5) Remove the fluid filler cap. (6) Fill the fluid tank and continue to monitor fluid level indications as follows: (a) On airplanes with the fairing, monitor the fluid tank sight glass at the TKS fluid window. (b) On airplanes with the cargo pod, monitor the fluid tank sight glass at the cargo pod forward center door. (7) After you service the fluid tank, install the fluid filler cap and close the cargo pod door. (8) Check the fluid indication in the cockpit.
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TKS Anti-Ice Fluid Tank Servicing (Fairing Installation) Figure 301 (Sheet 1)
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TKS Anti-Ice Fluid Tank Servicing (Cargo Pod Installation) Figure 302 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (9) 3.
Remove external electrical power from the airplane.
TKS Deicing Porous Panel Care A.
General (1) The TKS Anti-Ice system is a durable system that requires little attention for day to day care. The basic care is cosmetic and preventive in nature, and will maintain the appearance and promote proper system operation.
B.
Cleanliness Precautions (1) The TKS system has filtration to protect the components against damage and blockage from particulate matter, but they are not always effective for liquid contaminants. To extend the life of the filter element and strainer, the following precautions are recommended: (a) Where possible minimize the number of containers that you use to store and fill the airplane anti-ice fluid tank. Purchase anti-ice fluid in small (2.5 gallon) (9.4607 liter) containers that you can use to pour it directly in the airplanes fluid tank. (b) Always clean the top of the containers before you remove the cap to pour from the container. (c) Always replace the cap on the fluid containers that you use for the transfer of the fluid. (d) If you transfer fluid from the original container to other containers for storage, make sure that these containers are clean and of correct materials. NOTE: (e) (f) (g)
C.
Tin plated steel or similar containers that can rust are not the correct type.
Keep a set of containers and implements solely for use with anti-ice fluid. Keep the area around the airplane TKS filler clean. Keep the filler cap clean. Always replace the filler cap on the airplane immediately after you fill the tank.
Porous Panel Cleaning
CAUTION: Porous panels contain a plastic membrane that can be damaged by certain solvents, especially methyl ethyl ketone (MEK), acetone, paint thinners, paint stripper, and other types of thinners and solvents. Do not use these materials to clean the panels. Put tape on the panels with non-porous solvent resistant material if you use solvents of this type on adjacent parts of the airplane. CAUTION: Do not paint the outer surface of the porous panels. CAUTION: Do not polish the surface of the porous panels when you polish painted surfaces. Wax or silicone polishes decrease the wetting qualities of the de-icing fluids and can degrade the ice protection efficiency. Repeated or intensive porous panel polishing can also block some of the pores in the panels. (1)
Clean dirt and insect debris from the panels when you clean the exterior of the airplane. NOTE:
(2) (3)
From the functional aspect porous panels are self cleaned by the back flushing action of the de-icing fluid when the system is operated. However additional cleaning is recommended when you clean the airplane.
Operate the ice protection system for a sufficient period of time to wet the leading edge TKS panels with fluid before you clean them. Approved ice protection fluid has a softening effect on insect debris. Spread the fluid over the insect encrusted area with a cloth or sponge while the ice protection system operates. Stop the system and let the fluid stand for about ten minutes before you clean the porous panels.
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WARNING: Take appropriate precautions to prevent fire. CAUTION: Only use the solvents that are listed below. • • • • • • • (4)
Clean deposits of oils, greases, adhesives, paint, etc. NOTE:
D.
Water (Soaps and detergents are permitted) DTD 406B Ice Protection Fluid Gasoline or Avgas Kerosene or Jet Fuel Isopropyl Alcohol Ethyl Alcohol Industrial Methylated Spirit
The removal of the deposit can be assisted by the use of "Scotch-Brite"™ and/or careful scraping.
Polishing of Porous Panels
CAUTION: Do not use liquid or wax polish on the TKS porous panels. (1)
Use "Scotch-Brite"™ to restore the panels to their initial condition. Refer to Chapter 30, Ice and Rain Protection-General. NOTE: (a) (b) (c)
If you use "Scotch-Brite"™ and polish the panels again and again it can cause a blockage of some of the pores in the panel.
Put low-adhesive tape on the airplane skin adjacent to the panel, where necessary, to prevent damage to the paint. Polish the porous panel with "Scotch-Brite"™ in a chordwise direction to get an initial surface texture. Clean and polish the porous panels with a Very Fine grade of "Scotch-Brite"™. Then clean and polish the panels with a Ultra Fine Grade of "Scotch-Brite"™.
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20 CHAPTER
STANDARD PRACTICES AIRFRAME
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
20-00-00
Pages 1-2
Mar 1/2000
20-00-01
Pages 1-4
Mar 1/2000
20-10-10
Pages 201-208
Mar 1/2000
20-10-20
Pages 201-214
Jun 1/2011
20-10-30
Pages 201-202
Aug 1/1995
20-10-40
Pages 401-407
Apr 1/2010
20-21-02
Pages 201-231
Apr 1/2010
20-21-03
Pages 601-604
Jun 1/2011
20-30-00
Page 1
Aug 1/1995
20-30-01
Pages 201-204
Mar 1/2008
20-30-03
Pages 201-227
Dec 1/2006
20-30-04
Pages 201-210
Sep 4/2001
20-30-05
Pages 201-203
Mar 1/2000
20-30-06
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Mar 1/2008
20-31-00
Pages 701-705
Mar 1/2000
20-40-00
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20-Title 20-List of Effective Pages 20-Record of Temporary Revisions 20-Table of Contents
20 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS STANDARD PRACTICES AIRFRAME - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20-00-00 Page 1 20-00-00 Page 1 20-00-00 Page 2
MATERIAL AND TOOL CAUTIONS - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . Titanium. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mercury . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Asbestos . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cadmium Plated Fasteners . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Usage Solvents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . National Emissions Standards for Hazardous Air Pollutants . . . . . . . . . . . . . . . . . . . . . Facilities and Equipment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20-00-01 20-00-01 20-00-01 20-00-01 20-00-01 20-00-01 20-00-01 20-00-01 20-00-01
Page 1 Page 1 Page 1 Page 1 Page 2 Page 2 Page 2 Page 3 Page 4
TORQUE DATA - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Requirements for Bolts, Screws and Nuts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Requirements for Hi-Lok Fasteners . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Requirements for Electrical Current Carrying And Airframe Ground Fasteners. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Requirements for Straight Threaded Fittings . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Requirements for Tubes and Hoses. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Requirements for V-Band Clamps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20-10-10 Page 201 20-10-10 Page 201 20-10-10 Page 204 20-10-10 Page 205
SAFETYING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety Wire. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lockwire Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cotter Pin Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Locking Clip Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety Cable Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20-10-20 Page 201 20-10-20 Page 201 20-10-20 Page 201 20-10-20 Page 202 20-10-20 Page 206 20-10-20 Page 207 20-10-20 Page 207
CONTROL CABLE WIRE BREAKAGE AND CORROSION LIMITATIONS MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Examination of Control Cables.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20-10-30 Page 201 20-10-30 Page 201
BEARINGS - REMOVAL/INSTALLATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bearings Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Application of Fastener Retaining Compounds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bearing/Bushing Retention. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sealant and Retaining Compounds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20-10-40 Page 401 20-10-40 Page 401 20-10-40 Page 406 20-10-40 Page 406 20-10-40 Page 407
ELECTRICAL BONDING - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hardware and Material Usage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bonding Surface Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Protective Coating Sealing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Bonding Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Bond Type (Class) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bonding Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bonding Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bonding Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
20-21-02 Page 201 20-21-02 Page 201 20-21-02 Page 201 20-21-02 Page 202 20-21-02 Page 203 20-21-02 Page 205 20-21-02 Page 207 20-21-02 Page 207 20-21-02 Page 213 20-21-02 Page 215 20-21-02 Page 217
20-10-10 Page 206 20-10-10 Page 206 20-10-10 Page 207 20-10-10 Page 208
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MODEL 208 MAINTENANCE MANUAL HIGH INTENSITY RADIATED FIELDS (HIRF) - INSPECTION/CHECK . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Access Panels and Doors Bonding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Bonding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wire Bundle Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Visual Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Wire Bundle Assembly Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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SOLVENTS, SEALANTS AND ADHESIVES - DESCRIPTION AND OPERATION . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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GENERAL SOLVENTS/CLEANERS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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FUEL, WEATHER AND HIGH-TEMPERATURE SEALING - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition of Sealing Terms. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sealant Curing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mixing of Sealants . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sealing Application . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sealant Repair. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Integral Fuel Tank Sealing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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ADHESIVE AND SOLVENT BONDING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Clear Polyurethane Topcoat. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Material Classification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Requirements for Bonding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Manual Cleaning and Deoxidizing of Aluminum Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . Liquid Solvent Cleaning. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adhesive Mixing, Application and Curing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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ANAEROBIC ADHESIVES - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Requirements for Bonding/Sealing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Primer Application. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adhesive Application . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adhesive Cure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wire Tacking. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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ADHESIVES, CEMENTS AND SEALANTS SHELF LIFE AND STORAGE DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Storage Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Testing Criteria. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL INTERIOR AND EXTERIOR FINISH - CLEANING/PAINTING . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Interior and Exterior Finishes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Paint Facility. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sanding Surfacer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Paint Stripping . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hand Solvent Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance of the Interior and Exterior Primary Coatings and Topcoat . . . . . . . . . .
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CONVERSION DATA - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conversion Factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL STANDARD PRACTICES AIRFRAME - GENERAL 1.
Scope A.
This Chapter describes the standard maintenance practices for maintaining and repairing items of the airframe, and systems that are typical to more than one area or system. Maintenance practices which are unique to a particular system, or subject are described in the appropriate chapter and section in the maintenance manual. NOTE: NAME
Equivalent substitutes may be used for the following items: NUMBER
MANUFACTURER
USE
Overall Base Paint
810 Series (Jet-Glo)
Sherman Williams 630 E. 13th St. Andover, KS 67002
Used as topcoat overall color.
Hisolids Corrosion Primer
483-928
Sherman Williams
Used as overall intermediate primer.
Hisolids Corrosion Primer
R4001-K14 MAX COR
U.S. Paint Corp. 831 S. 21st St. St. Louis, MO 63103
Epoxy Primer Base
Acid Resistant Enamel
Jet-Glo Urethane 571-567
Sherman Williams
Used as topcoat for battery box.
Polyester
PCA-1575-GP1
Polymer Corp 2120 Fairmont Ave. Reading, PA 19612-4235.
Used as topcoat for landing gear.
Heat Resistant Enamel (Gray)
521-520
Sherman Williams
Used for engine mount and acccompanying hardware in engine compartment.
Chromated Etch Primer
U2655 Wash Primer and 142265 Reducer
Sterling Laquer 3150 Brannon Ave. St. Louis, MO 63139
To prime bonded or spot welded assemblies.
Epoxy Primer
454-4-1 Base and CA-109 Converter 513J102 Base and 91J138 Converter
DuPont 251 Brandywine Bldg. Wilmington, DE 19803
To prime interior surfaces of wing fuel bays and all detail parts. To prime interior surfaces of wing fuel bays.
Intermediate Epoxy Primer
825-8136 Base and VG8392 Activator T3871 Thinner
DuPont
To prime aluminum airplane exterior. To thin intermediate epoxy primer.
Polishing Compound
808 Polishing Compound
DuPont
To rub out overspray.
Cloth
Hex Wiping Cloth
Commercially Available
With solvent to clean airplane exterior.
Tape
Masking Tape, Y-231
3M Co. 3M Center Minneapolis, MN 55144
To hold masking paper in place.
Cleaning Solvent
Aliphatic Naptha
Commercially Available
To clean acrylic.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Paint Remover
Oakite EPA
Oakite Products Inc. 50 Valley Rd. Berkeley Heights, NJ 07922
To remove difficult epoxies and polyurethane.
Paint Remover
1717 SC
B & B Chemical Co. Inc. 875 W 28th St Hialeah, FL 33010
To remove epoxies and urethane paint.
Paint Remover
5075 NP
B & B Chemical Co. Inc.
To remove epoxies and urethane paint.
Paint Remover
PR-4028
Eldorado Chemical Co. 14358 Lookout Road San Antonio, TX 78265-4837
To remove epoxies and urethane paint.
Paint Remover
Turco T-6776 LO
ELF Atochem North America, Inc. Turco Div. P.O. Box 5780 Winston-Salem, NC 27113
To remove lacquers, alkyds, epoxies and urethane paint.
Fireproof Coating (Intumescent Paint)
173
Flame Control Coatings P.O. Box 786 4120 Hyde Park Blvd. Niagra Falls,, NY 14302
Fireproof coating of electrical power box.
Urethane-Acrylic Paint (Gray)
WP593
Sherman Williams
Top coat for electrical power box.
Modified Urethane Paint (Seal Gray)
WP46
Sherman Williams
Finish coat for electrical powerbox.
2.
Definition A.
This chapter is divided into subjects and sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief definition of the subjects and sections incorporated in this chapter is as follows. (1) The subject on Material and Tool Cautions describes general cautions and warnings applicable to maintenance on or around the airplane. (2) The subject on Torque Data describes maintenance practices for torquing tools, torquing requirements, formulas and torque limits for various type fasteners. (3) The subject on Safetying describes the proper methods and use of safety wire/lockwire, cotter pins and lock clip installations. (4) The subject on Control Cables and Pulleys describes the construction, examination and storage of cable assemblies and pulleys. (5) The section on Solvents, Sealants and Adhesives provides the description and uses for solvents and cleaners; fuel, weather, pressure and high-temperature sealing; and the application of adhesives and solvent bonding. (6) The subject on Conversion Data contains information converting the more commonly used measuring units found in the Maintenance Manual.
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MODEL 208 MAINTENANCE MANUAL MATERIAL AND TOOL CAUTIONS - DESCRIPTION AND OPERATION 1.
Titanium
CAUTION: Cadmium plated tools must not be used on titanium parts, particularly if parts are mounted where they may be subjected to temperatures above 250°F. Small cadmium deposits, which may be left on such parts, will react with titanium when heated, resulting in brittleness and possibly cracks. CAUTION: Cadmium plated fasteners must not be used in contact with titanium parts. 2.
Mercury
CAUTION: Mercury-containing thermometers and other test equipment must not be used on the airplane. A.
3.
Mercury, by the amalgamation process, can penetrate any break in the finish, paint or sealing coating of a metal structural element. An oxide coating on a dry metallic surface will tend to inhibit an immediate action while a bright, polished, shining or scratched surface will hasten the process. Moisture will also promote the amalgamation process. Soils, greases or other inert contaminants, present on the metal surfaces, will prevent the start of the action. The corrosion and embrittlement which results from an initial penetration, can be extremely rapid in structural members under load. Once it has begun, there is no known method of stopping it. Complete destruction of the load carrying capacity of the metal will result.
Asbestos
WARNING: Asbestos fibers are harmful when ingested into the body. The following steps must be adhered to when working with parts containing asbestos. A.
Avoid inhalation of dust with either the following methods. (1) Use engineering controls, such as working with properly filtered exhaust chamber, or use wet methods to maintain exposure below OSHA personnel exposure limits. (2) If methods in step (1) above cannot be used, use respiratory protection, including high efficiency filters. Other protection must include coveralls, gloves and eye protection.
B.
Dispose of all asbestos containing material in accordance with local, state and federal regulations.
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MODEL 208 MAINTENANCE MANUAL
4.
Cadmium Plated Fasteners
CAUTION: Cadmium plated fasteners, utilized in areas where contact with jet fuel exists, must be completely covered with fuel tank sealant to prevent contact with fuel. All avenues that fuel can travel to reach head or shank or fastener must be sealed. Cadmium fasteners when continuously in contact with jet fuel may disperse cadmium from the fastener into the fuel system, which will be detrimental to the engines. 5.
Maintenance Precautions
WARNING: During maintenance, repair and servicing of the airplane, many substances and environments encountered may be injurious if proper precautions are not observed.
6.
A.
Carefully read and follow all instructions, and especially adhere to all cautions and warnings provided by the manufacturer of the product being used. Use appropriate safety equipment as required including goggles, face shields, breathing apparatus, respirators, protective clothing and gloves. Fuel, engine oil, solvents, volatile chemicals, adhesives, paints and strong cleaning agents may be injurious when contacting the skin or eyes, or when vapors are breathed. When sanding composites or metals or otherwise working in an area where dust particles may be produced, the area should be ventilated and the appropriate respirator must be used.
B.
In case of a spill of hazardous material, consult the latest version of the Emergency Response Guidebook DOT P 5800.5 for guidance in dealing with the hazard. As soon as possible, notify CHEMTREC at 1-800-424-9300 (in the United States, including Alaska and Hawaii; and in Canada) for more detailed information in dealing with the hazard.
General Usage Solvents A.
During the course of daily work routines, many will have the occasion to work with solvents. Webster’s dictionary defines a solvent as, “a substance, usually liquid, that dissolves or can dissolve another substance.” An example would be the reaction that is obtained when salt is added to water. The salt is dissolved by the water. General usage solvents include the following: Methyl n-Propyl Ketone Isopropyl Alcohol Naptha
B.
These chemicals/solvents are generally colorless, evaporate quicker than water, and tend to give off vapors in higher quantities as their temperature increases. The vapors are generally heavier than air, which causes them to collect in low lying areas or push normal oxygen and air out of a confined area. This situation can lead to oxygen deficient atmospheres. Many general usage solvents are also flammable.
C.
Solvents are hazardous to work with because of their flammability, rate of evaporation and reaction to oxidizers. Solvents can also be an irritant to the skin and eyes.
D.
Solvent flammability can be induced by a single spark, a smoldering cigarette, or even atmospheric conditions can ignite gasoline vapors. The vapors may also flash back to the original source which can explode. The same reaction can take place with Toluene or Isopropyl Alcohol. The lower the flash point of the chemical, the more likely it is to become flammable. Generally, flash points of less than 100°F (37.8°C) are considered flammables. Examples of solvent flash points are shown below:
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MODEL 208 MAINTENANCE MANUAL
SOLVENT
7.
FLASH POINT
Methyl n-Propyl Ketone
45°F (7.2°C)
Isopropyl Alcohol
53.6°F (12°C)
E.
The rate of evaporation is closely tied to flammability because normally the vapors must be present to ignite the liquid. Vaporization also allows the solvents, even those that are not flammable, to get into the air and into the body’s blood stream through the lungs.
F.
Solvents can also react explosively with oxidizers (chemicals which release oxygen). A very violent and uncontrollable reaction takes place which generates heat rapidly. For this reason, it is very important for each person to be aware of specific chemicals in use in the work area, and to comprehend the labeling of containers. Chemical manufacturers are required to label each container with a diamond shaped symbol: red for flammable and yellow for oxidizers.
G.
Solvents can also damage the hands and skin. Solvents dry out skin and dissolve the natural oils. The condition can develop into an irritation or, if left untreated with continuous exposure it may progress to a dermatitis. Damaged skin allows other contaminants to worsen the condition, because the contaminants have easier access to the deeper levels of the skin. In serious cases, blood poisoning is also possible.
H.
The best defense against skin irritation is not to be exposed, but many times exposure to solvents/ chemicals is unavoidable. Fortunately, the body is capable of filtering small amounts of solvents out of the body. This filtration function takes place in the liver. The liver receives blood which may be contaminated with solvents from both the lungs and the skin. If the quantities are low enough and infrequent enough, the liver can filter out the contaminants. This is one of the scientific facts on which OSHA based its Permissible Exposure Limits. However, when exposures are constantly above these levels over an extended period of many years, the filter (liver) becomes clogged and the solvents can then affect other parts/portions of the body.
National Emissions Standards for Hazardous Air Pollutants A.
National Emissions Standards for Hazardous Air Pollutants (NESHAP). (1) The NESHAP standards have restricted the use of certain chemicals and solvents. (2) For complete details of the regulatory standards, see Federal Register, 40 CFR Part 63, [AdFRL-5636-1], RIN 2060-AG65. (3) The currently acceptable replacements for chemicals that have been restricted or prohibited by the standards are listed in Exterior Finish - Cleaning/Painting, and these supersede materials which may be specified elsewhere in this manual.
B.
NESHAP Requirements. (1) Hand-Wipe Cleaning. (a) All hazardous air pollutants or volatile organic compounds that are used as hand wipe cleaning solvents must meet a composition requirement and have a vapor pressure less than or equal to 1.75Hg at 69° (45 mm Hg at 20°C.) (b) The requirements specified may be met by an alternative compliance plan administered by the permitting authority and approved under Section 112(1) of the Clean Air Act. (2) Primer Application. (a) The organic hazardous air pollutant content is limited to 350 g/l (2.9 pounds per gallon), less water, as applied. (b) The volatile organic compound limit is 350 g/l (2.9 pounds per gallon), less water, as applied. (c) Achieve the content limits by using coatings below the content limit or use monthly volumeweighted averaging to meet content limits. (3) Topcoat Application. (a) The base coat organic hazardous air pollutant content must be less than 420 g/l (3.5 pounds per gallon), less water, as applied. (b) The volatile organic compound limit is 420 g/l (3.5 pounds per gallon), less water, as applied. (c) The topcoats must meet the requirements of MIL-C-85285.
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MODEL 208 MAINTENANCE MANUAL (d)
(4)
8.
Stripe paint requirements are the same as the base coat requirements. If the recommended supplier cannot be used, then use base coat materials. to paint stripes. NOTE:
All paints and primers must have specific application techniques. If alternative is sought, use only the materials that are less than or equal in emissions, to less than HVLP or electrostatic spray application techniques.
NOTE:
Operate all application equipment according to the manufacturer’s specifications, company procedures or locally specified operating procedures.
Depainting (a) Depainting operations apply to the outer surface of the airplane and do not apply to parts or units normally removed. Fuselage, wings and stabilizers are covered. Radomes and parts normally removed are exempt from the following requirements: No organic hazardous air pollutants are to be emitted from chemical strippers or 1 softeners. 2 Inorganic hazardous air pollutant emissions must be kept to a minimum during periods of nonchemical based equipment malfunctions. The use of organic hazardous air pollutant material for spot stripping and decal 3 removal is limited to 190 pounds per airplane per year. (b) Operating requirements for depainting operations generating airborne inorganic hazardous air pollutants include control with particulate filters or water wash systems. (c) Mechanical and hand sanding are exempt from these requirements.
Facilities and Equipment A.
Facilities (1) A system must be provided to collect processing waters to treat for chromium and pH or to be hauled away. (2) Facilities must have proper safety equipment.
B.
Equipment (1) Spray application of cleaning solvents, paint removers or color chemical film treatment solutions is prohibited unless all requirements of NESHAP are met. (2) Spraying equipment to wash the airplane with alkaline cleaner may be used. This equipment should be adequate to spray deoxidizer, chemical film solutions and rinse water. (3) A high pressure washer is recommended, with or without hot water. (4) Respirators and/or dust masks should be used.
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MODEL 208 MAINTENANCE MANUAL TORQUE DATA - MAINTENANCE PRACTICES 1.
General A.
To ensure security of installation and prevent overstressing of components during installation, the torque values outlined in this section and other applicable chapters of this manual should be used during installation and repair of components.
B.
The torque value tables, listed in this section, are standard torque values for the nut and bolt combinations shown. Components which require special torque values will have those values listed in the applicable maintenance practices section.
C.
Torque is typically applied and measured using a torque wrench. Different adapters, used in conjunction with the torque wrench, may produce an actual torque to the nut or bolt which is different from the torque reading. Figure 201 is provided to help calculate actual torque in relation to specific adaptors used with the torque wrench.
D.
Free Running Torque Value. (1) Free running torque value is the torque value required to rotate a nut on a threaded shaft, without tightening. Free running torque value does not represent the torque values listed in the tables of this section. Torque values listed in the tables represent the torque values above free running torque. EXAMPLE:
(2)
E.
If final torque required is to be 150 inch-pounds and the free running torque is 25 inch-pounds, then the free running torque must be added to the required torque to achieve final torque of 150 +25 = 175 inch-pounds.
Breakaway torque value is the value of torque required to start a nut rotating on a thread shaft, and does not represent free running torque value. It should be noted that on some installations the breakaway torque value cannot be measured.
General Torquing Notes. (1) These requirements do not apply to threaded parts used for adjustment, such as turnbuckles and rod ends. (2) Torque values shown are for clean nonlubricated parts. Threads should be free of dust, metal filings, etc. Lubricants, other than that on the nut as purchased, should not be used on any bolt installation unless specified. (3) Assembly of threaded fasteners, such as bolts, screws and nuts, should conform to torque values shown in Table 201. (4) When necessary to tighten from the bolt head, increase maximum torque value by an amount equal to shank friction. Measure shank friction with a torque wrench. (5) Sheet metal screws should be tightened firmly, but not to a specific torque value. (6) Straight threaded connections using O-rings or gaskets for seal, such as AN924 or AN6298 nuts, and fittings conforming to MS33656, Style E, need not be tightened to a specific torque value, but shall be installed per AN10064. (7) Countersunk washers used with close tolerance bolts must be installed correctly to ensure proper torquing (refer to Figure 202). (8) For Hi-Lok Fasteners used with MS21042 self-locking nuts. Fastener and nut should be lubricated prior to tightening. (9) Tighten accessible nuts to torque values per Table 201. Screws attached to nutplates, or screws with threads not listed in Table 201 should be tightened firmly, but not to a specific torque value. Screws used with dimpled washers should not be drawn tight enough to eliminate the washer crown. (10) Table 201 is not applicable to bolts, nuts and screws used in control systems or installations where the required torque would cause binding, or would interfere with proper operation of parts. On these installations, the assembly should be firm but not binding.
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Washer Installation close Tolerance Bolts Figure 201 (Sheet 1)
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Torque Wrench and Adapter Formulas Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (11) Castellated Nuts. (a) Self-locking and non self-locking castellated nuts, except MS17826, require cotter pins and should be tightened to the minimum torque value shown in Table 201. The torque may be increased to install the cotter pin, but this increase must not exceed the alternate torque values. (b) MS17826 self-locking, castellated nuts shall be torqued per Table 201. (c) The end of the bolt or screw should extend through the nut at least two full threads including the chamfer. (12) Joints containing wood, plastics, rubber or rubberlike materials should be torqued to values approximately 80 percent of the torque at which crushing is observed, or to the requirements of Table 201, whichever is lower, or as specified. 2.
Torque Requirements for Bolts, Screws and Nuts A.
Use Table 201 to determine torque requirements for bolts, screws and nuts.
Table 201. Torque Values Nuts, Bolts and Screws (Steel) (Inch-Pounds) Size of Bolt, Nut or Screw
• •
Fine Threaded Series (Tension Type Nuts)
Fine Threaded Series (Shear Type Nuts Except MS17826)
Standard
Alternate
Standard
Alternate
8-32
12-15
- -
7-9
- -
10-32
20-25
20-28
12-15
1/4-28
50-70
50-75
5/16-24
100-140
3/8-24
MS17826 Nuts
Standard
Alternate
- -
- -
12-19
12-15
12-20
30-40
30-48
30-40
30-45
100-150
60-85
60-100
60-80
60-90
160-190
160-260
95-110
95-170
95-110
95-125
7/16-20
450-500
450-560
270-300
270-390
180-210
180-225
1/2-20
480-690
480-730
290-410
290-500
240-280
240-300
9/16-18
800-1000
800-1070
480-600
480-750
320-370
320-400
5/8-18
1100-1300
1100-1600
660-780
660-1060
480-550
480-600
3/4-16
2300-2500
2300-3350
1300-1500
1300-2200
880-1010
880-1100
7/8- 14
2500-3000
2500-4650
1500-1800
1500-2900
1500-1750
1500-1900
1-14
3700-4500
3700-6650
2200-3300
2200-4400
2200-2700
2200-3000
1-1/8-12
5000-7000
5000-10000
3000-4200
3000-6300
3200-4200
3200-5000
1-1/4-12
9000-11000
9000-16700
5400-6600
5400-10000
5900-6400
5900-7000
Fine Thread Tension application Nuts include: AN310, AN315, AN345, MS17825, MS20365, MS21044 through MS21048, MS21078, NAS679, NAS1291 Fine Thread Shear application Nuts include: AN316, AN320, MS21025, MS21042, MS21043, MS21083, MS21245, NAS1022, S1117
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Table 202. Torque Values Nuts, Bolts and Screws (Steel) (Newton meters) Fine Threaded Series (Tension Type Nuts)
Size of Bolt, Nut or Screw
Fine Threaded Series (Shear Type Nuts Except MS17826)
MS17826 Nuts
Standard
Alternate
Standard
Alternate
Standard
Alternate
8-32
1.4-1.7
- -
0.8-1.0
- -
- -
- -
10-32
2.3-2.8
2.3-3.2
1.4-1.7
1.4-2.2
1.4-1.7
1.4-2.3
1/4-28
5.6-7.9
5.6-8.5
3.4-4.5
3.4-5.4
3.4-4.5
3.4-5.0
5/16-24
11.3-15.8
11.3-17.0
6.8-9.6
6.8-11.3
6.8-9.0
6.8-10.1
3/8- 24
18.0-21.4
18.0-29.4
10.7-12.4
10.7-19.2
10.7-12.4
10.7-14.1
7/16-20
50.8-56.5
50.8-63.2
30.5-33.8
30.5-44.0
20.3-23.7
20.3-25.4
1/2-20
54.2-77.9
54.2-82.4
32.7-46.3
32.7-56.4
27.1-31.6
27.1-33.8
9/16-18
90.3-112.9
90.3-120.8
54.2-67.8
54.2-84.7
36.1-41.8
36.1-45.1
5/8-18
124.2-146.8
124.2-180.7
74.5-88.1
74.5-19.7
54.2-62.1
54.2-67.7
3/4-16
259.8-282.4
259.8-378.5
46.8-169.4
46.8-248.5
99.4-114.1
99.4-124.2
7/8-14
282.4-338.9
282.4-545.3
169.4-203.3
169.4-327.6
169.4-197.7
169.4-214.6
1-14
418.0-508.4
418.0-751.3
248.5-372.8
248.5-497.1
248.5-305.0
248.5-338.9
1-1/8-12
564.9-790.8
564.9-1129.8
338.9- 474.5
338.9-711.8
361.5-474.5
361.5-564.9
1-1/4-12
1016.8-1242.8
1016.81886.8
610.1-745.7
610.1-1129.8
666.6-723.1
666.6-790.8
• • 3.
Fine Thread Tension application Nuts include: AN310, AN315, AN345, MS17825, MS20365, MS21044 through MS21048, MS21078, NAS679, NAS1291 Fine Thread Shear application Nuts include: AN316, AN320, MS21025, MS21042, MS21043, MS21083, MS21245, NAS1022, S1117 Torque Requirements for Hi-Lok Fasteners A.
Use Table 203 to determine torque requirements for Hi-Lok fasteners. NOTE:
This table is used in conjunction with MS21042 Self-Locking nuts.
Table 203. Torque Values For Hi-Lok Fasteners (Alloy Steel, 180 to 200 KSI) NOMINAL FASTENER DIAMETER
TORQUE VALUE (INCH-POUNDS)
6-32
8 to 10
8-32
12 to 15
10-32
20 to 25
1/4-28
50 to 70
5/16-24
100 to 140
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MODEL 208 MAINTENANCE MANUAL Table 203. Torque Values For Hi-Lok Fasteners (Alloy Steel, 180 to 200 KSI) (continued)
4.
NOMINAL FASTENER DIAMETER
TORQUE VALUE (INCH-POUNDS)
3/8-24
160 to 190
7/16-20
450 to 500
1/2-20
480 to 690
Torque Requirements for Electrical Current Carrying And Airframe Ground Fasteners A.
Use Table 204 to determine torque requirements for threaded electrical current carrying fasteners. (1) Torque values shown are clean nonlubricated parts. Threads shall be free of dust and metal filings. Lubricants, other than on the nut as purchased, shall not be used on any bolt installations unless specified in the applicable chapters of this manual. (2) All threaded electrical current carrying fasteners for relay terminals, shunt terminals, fuse limiter mount block terminals and bus bar attaching hardware shall be torqued per Table 204. NOTE:
B.
There is no satisfactory method of determining the torque previously applied to a threaded fastener. When retorquing, always back off approximately 1/4 turn or more before reapplying torque.
Use Table 205 to determine torque requirements for threaded fasteners used as airframe electrical ground terminals.
Table 204. Torque Values For Electrical Current Carrying Fasteners FASTENER DIAMETER
TORQUE VALUE (INCH-POUNDS)
6-32
8 to 12
8-32
13 to 17
10-32
20 to 30
3/16
20 to 30
1/4
40 to 60
5/16
80 to 100
3/8
105 to 125
1/2
130 to 150
Table 205. Torque Values For Airframe Electrical Ground Terminals
5.
FASTENER DIAMETER
TORQUE VALUE (INCH-POUNDS)
5/16
130 to 150
3/8
160 to 190
Torque Requirements for Straight Threaded Fittings A.
Use Table 206 to determine torque requirements for straight threaded fittings.
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Table 206. Torque Values For Straight Threaded Fittings (Inch-Pounds) Tube Outside Diameter
Steel Tubing
6061-T Aluminum Tubing (Steel Sleeve)
6061-0 Aluminum 5052-0 Aluminum Tubing or Aluminum Hose Insert
Minimum Torque
Maximum Torque
Minimum Torque
Maximum Torque
1/8
45
55
20
30
3/16
90
100
30
1/4
135
150
5/16
180
3/8
Tube Wall **(Inches)
Minimum Torque
Maximum Torque
---
---
---
40
0.028
45
55
40
65
0.022 0.018 0.035 0.049
80 80 80 90
105 105 105 115
200
60
80
0.028 0.035 0.042
80 80 125
105 105 175
270
300
75
125
0.028 0.035 0.049
125 125 125
175 175 175
1/2
450
500
150
250
0.028 0.015 0.049 0.058 0.065
135 200 400 400 400
180 300 500 500 500
5/8
700
800
200
450
All
500
600
3/4
1100
1150
300
500
All
600
700
1
1200
1400
500
700
All
1000
1300
1 1/4
1300
1450
600
900
All
1300
1500
1 1/2
1350
1500
600
900
All
1400
1700
2
1500
1700
600
900
---
---
---
** Tube wall thickness is applicable to 6061-T aluminum tubing only. 6.
Torque Requirements for Tubes and Hoses A.
Use Table 207 to determine torque requirements for tubes and hoses.
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Table 207. Torque Values for Tubing and Hoses (Inch-Pounds) Hose Tubing Size O.D
Aluminum Tubing (Flared)
Steel Tubing (Flared)
Aluminum Tubing (Flareless)
Steel Tubing (Flareless)
Min Torque
Max Torque
Min Torque
Max Torque
Min Torque
Max Torque
Min Torque
Max Torque
-3
3/16
---
---
90
100
75
90
90
100
-4
1/4
40
65
135
150
80
100
135
150
-5
5/16
60
80
180
200
100
130
180
200
-6
3/8
75
125
270
300
100
130
270
300
-8
1/2
150
250
450
500
200
240
450
500
-10
5/8
200
350
700
800
360
400
700
800
-12
3/4
300
500
1100
1150
390
430
1100
1150
-16
1
500
700
1200
1400
600
900
1200
1400
-20
1 1/4
600
900
1300
1450
600
900
1300
1450
-24
1 1/2
600
900
1350
1500
600
900
1350
1500
Hose Size
7.
Tubing O.D.
Aluminum Fittings Oxygen Lines Only
Steel Hose End (Flared)
Steel Hose End (Flareless)
Min
Max
Min
Max
Min
Max
-3
3/16
- - -
- - -
70
100
95
105
-4
1/4
- - -
- - -
70
120
135
145
-5
5/16
100
125
85
180
175 dry
195 dry
-6
3/8
- - -
- - -
100
250
215
245
-8
1/2
- - -
- - -
210
420
470
510
-10
5/8
- - -
- - -
300
480
620
680
-12
3/4
- - -
- - -
500
850
855
945
-16
1
- - -
- - -
700
1150
1140
1260
-20
1 1/4
- - -
- - -
- - -
- - -
- - -
- - -
-24
1 1/2
- - -
- - -
- - -
- - -
- - -
- - -
Torque Requirements for V-Band Clamps A.
V-band clamps are used on engine bleed air lines and on the starter/generator. Clamp torque is dependent on V-band size and manufacturer. Clamps should be torqued according to torque value stamped on each individual clamp.
CAUTION: Do not exceed torque value stamped on clamp.
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MODEL 208 MAINTENANCE MANUAL SAFETYING - MAINTENANCE PRACTICES 1.
General A.
Lockwire. (1) Inconel (Uncoated), Monel (Uncoated). (a) Used for general lock wiring purposes. Lock wiring is the application of wire to prevent relative movement of structural or other critical components subjected to vibration, tension, torque, etc. Monel to be used at temperatures up to 700°F and inconel to be used at temperatures up to 1500°F. Identified by the color of the finish, monel and inconel color is natural wire color. (2) Copper, Cadmium Plated and Dyed Yellow in Accordance with FED-STD 595. (a) This will be used for shear and seal wiring applications. Shear applications are those where it is necessary to purposely break or shear the wire to permit operation or actuation of emergency devices. Seal applications are those where the wire is used with a lead seal to prevent tampering or use of a device without indication. Identified by the color of the finish, copper is dyed yellow. (3) Aluminum Alloy (Alclad 5056), Anodized and Dyed Blue in Accordance with FED-STD 595. (a) This wire will be used exclusively for safety wiring magnesium parts. NOTE:
2.
Surface treatment which obscures visual identification of safety wire is prohibited.
B.
Safety Cable. (1) Used as an alternative to corrosion-resistant steel lockwire.
C.
Cotter Pin. (1) The selection of material should be in accordance with temperature, atmosphere and service limitations.
D.
Locking Clips. (1) Used to safety turnbuckles.
Safety Wire NOTE: A.
You can use safety cable as an alternative to safety wire. Refer to Safety Cable Installation, in this section.
Wire Size. (1) The size of the wire should be in accordance with the requirements of Table 201.
Table 201. Safety Wire Material
Number (MS20995-XXX)
Ni-Cu Alloy (Monel)
NC20 NC32 NC40
NC51 NC91
Ni-Cr-Fe Alloy (Inconel)
N20
N32
N51
Carbon Steel, Zinc-Coated
F20
F32
N40 F41
F47
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MODEL 208 MAINTENANCE MANUAL Table 201. Safety Wire (continued) Material Corrosion-Resistant Steel
Number (MS20995-XXX) C15
Aluminum Alloy (Blue) Copper (Yellow)
CY15
C20
C32
C41
C47
C91
AB20
AB32
AB41
AB47
AB91
CY20
Example of Part Numbers: MS20995 CY20 = Copper, Shear or Seal Wire, 0.020 inch Diameter MS20995 AB32 = Aluminum Alloy, Anodized, 0.032 inch Diameter NOTE 1: The dash numbers indicate wire material and diameter in thousandths of an inch. (a)
(b) (c) 3.
0.032 inch minimum diameter for general purpose lock wiring except that 0.020 inch diameter wire may be used on parts having a nominal hole diameter of less than 0.045 inch; on parts having a nominal hole diameter between 0.045 inch and 0.062 inch with spacing between parts of less than two inches; or on closely spaced screws and bolts of 0.25 inch diameter and smaller. 0.020 inch diameter copper wire should be used for shear and seal wire applications. When employing the single wire method of locking the largest nominal size wire for the applicable material or part which the hole will accommodate should be used.
Lockwire Installation A.
Method (Refer to Figure 201). (1) The double-twist method of lock wiring should be used as the common method of lock wiring. The single wire method of lock wiring may be used in a closely spaced, closed geometrical pattern (triangle, square, circle, etc.), on parts in electrical systems, and in places that would make the single wire method more advisable. Closely spaced should be considered a maximum of two inches between centers.
CAUTION: Screws in closely spaced geometric patterns which secure hydraulic or air seals, hold hydraulic pressure, or used in critical areas, should use the double-twist method of lock wiring. (2)
Use single copper wire method for shear and seal wiring application. Make sure that the wire is so installed that it can easily be broken when required in an emergency situation. For securing emergency devices where it is necessary to break the wire quickly, use copper wire only.
B.
Spacing. (1) When lock wiring widely spaced multiple groups by the double-twist method, three units should be the maximum number in a series. (2) When lock wiring closely spaced multiple groups, the number of units that can be lockwired by a twenty-four inch length of wire should be the maximum number in a series. (3) Widely spaced multiple groups should mean those in which the fastenings are from four to six inches apart. Lockwiring should not be used to secure fasteners or fittings which are spaced more than six inches apart, unless tie points are provided on adjacent parts to shorten the span of the lockwire to less than six inches.
C.
Tension. (1) Parts should be lock wired to put tension on lock wires when the parts tend to loosen. The lockwire should always be installed and twisted so the loop around the head stays down and does not tend to come up over the bolt head and leave a slack loop. NOTE: (2)
This does not necessarily apply to castellated nuts when the slot is close to the top of the nut; the wire will be more secure if it is made to pass along the side of the stud.
Care should be exercised when installing lockwire, to ensure it is tight but not overstressed.
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Lockwire Safetying Figure 201 (Sheet 1)
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Lockwire Safetying Figure 201 (Sheet 2)
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Lockwire Safetying Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL D.
4.
Usage. (1) A pigtail of 0.25 to 0.50 inch (three to six twists) should be made at the end of the wiring. This pigtail should be bent back or under to prevent it from becoming a snag. (2) Safety wire (lockwire) should be new upon each application. (3) When castellated nuts are to be secured with lockwire, tighten the nut to the low side of the selected torque range unless otherwise specified, and, if necessary, continue tightening until a slot aligns with the hole. (4) In blind tapped hole applications of bolts or castellated nuts on studs, lock wiring should be as described in these instructions. (5) Hollow head bolts are safetied in the manner prescribed for regular bolts. (6) Drain plugs and cocks may be safetied to a bolt, nut or other part having a free lock hole in accordance with the instructions described in this text. (7) External snaprings may be locked if necessary in accordance with the general locking principles as described and illustrated. Internal snaprings should not be lock wired. (8) When locking is required on electrical connectors which use threaded coupling rings, or on plugs which employ screws or rings to fasten the individual parts of the plug together, they should be lock wired with 0.020 inch diameter wire in accordance with the locking principles as described and illustrated. It is preferable to lock wire all electrical connectors individually. Do not lock wire one connector to another unless it is necessary to do so. (9) Drilled head bolts and screws need not be lock wired if installed into self-locking nuts or installed with lockwashers. Castellated nuts with cotter pins or lockwire are preferred on bolts or studs with drilled shanks, but self-locking nuts are permissible within the limitations of MS33588. (10) Lockwire shall not be used to secure or be dependent on fracture as the basis for operation of emergency devices such as handles, switches, guards covering handles, etc., that operate emergency mechanisms such as emergency exits, fire extinguishers, emergency cabin pressure release, emergency landing gear release and the like. However, where existing structural equipment or safety-of-flight emergency devices require shear wire to secure equipment while not in use, but which are dependent on shearing or breaking of the lockwire for successful emergency operation of equipment, particular care should be exercised to assure that lock wiring under these circumstances will not prevent emergency operations of these devices.
Cotter Pin Installation A.
General instruction for the selection and application of cotter pins (refer to Figure 202). (1) Select cotter pin material in accordance with temperature, atmosphere and service limitations. Refer to Table 202.
Table 202. Cotter Pin Material Material
Temperature
Service
MS24665 Cotter Pins Carbon Steel
Ambient Temperature up to 460°F
Normal atmosphere cotter pins contacting cadmium plated bolts or nuts.
MS24665 Cotter Pins Corrosion Resistant Steel
Ambient Temperature up to 800°F
Non magnetic requirements cotter pins contacting corrosion resistant steel bolts or nuts in a corrosive atmosphere.
(2) (3)
(4)
Cotter pins should be new upon each application. When nuts are to be secured to the fastener with cotter pins, tighten the nut to the low side (minimum) of the applicable specified or selected torque range, unless otherwise specified, and if necessary, continue tightening until the slot aligns with the hole. In no case should the high side (maximum) torque range be exceeded. Castellated nuts mounted on bolts may be safetied with cotter pins or lockwire. The preferred method is with the cotter pin bent parallel to the axis of the bolt. The alternate method, where the cotter pin is mounted normal to the axis of the bolt, may be used when the cotter pin in the preferred method is apt to become a snag.
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7) (8) (9) 5.
Locking Clip Installation A.
6.
In the event when more than 50 percent of the cotter pin diameter is above the nut castellation, a washer should be used under the nut or a shorter fastener should be used. A maximum of two washers may be permitted under a nut. The largest nominal diameter cotter pin (listed in MS24665) which the hole and slots will accommodate should be used; but in no application to a nut, bolt or screw should the pin size be less than the sizes described in Figure 202. Install the cotter pin with the head firmly in the slot of the nut, with the axis of the eye at right angles to the bolt shank. Bend prongs so the head and upper prong are firmly seated against the bolt. In the pin applications, install the cotter pin with the axis of the eye parallel to the shank of the clevis pin or rod end. Bend the prongs around the shank of the pin or rod end. Cadmium plated cotter pins should not be used in applications bringing them in contact with fuel, hydraulic fluid or synthetic lubricants.
Safetying Turnbuckles (Refer to Figure 203). (1) Prior to safetying, both threaded terminals should be screwed an equal distance into the turnbuckle barrel, and should be screwed in, at a minimum, so no more than three threads of any terminal are exposed outside the body. (2) After the turnbuckle has been adjusted to its locking position, with the groove on terminals and slot indicator notch on barrel aligned, insert the end of the locking clip into the terminal and barrel until the "U" curved end of the locking clip is over the hole in the center of the barrel. (a) Press the locking clip into the hole to its full extent. (b) The curved end of the locking clip will latch in the hole in the barrel. (c) To check proper seating of locking clip, attempt to remove pressed "U" end from barrel hole with fingers only. Do not use a tool as the locking clip could be distorted. (3) Locking clips are for one time use only and should not be reused. (4) Both locking clips may be inserted in the same hole of the turnbuckle barrel or in opposite holes of the turnbuckle barrel.
Safety Cable Installation A.
Tools and Equipment. Name
Number
Manufacturer
Use
Ferrule, Safety Cable
SAE AS4536
Commercially available
To use with the safety cable.
Safety Cable
SAE AS4536
Commercially available
To prevent the movement of structural or other critical components that have had vibration, tension, or torque applied to them.
Safety Cable Application Tool
SCT Series
Daniels Manufacturing Corporation 526 Thorpe Rd. Orlando, FL 32824-8133
To install the Daniels safety cable.
Safety Cable Terminator Tool
BM Series
Bergen Cable Technology, LLC 343 Kaplan Drive Fairfield, NJ 07004
To install the Bergen safety cable.
B.
Procedure (Refer to Figure 204). (1) Make sure that you obey the precautions for the safety cable as follows: (a) Wear eye protection when you cut the safety cable.
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Cotter Pin Safetying Figure 202 (Sheet 1)
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Safetying Turnbuckle Assemblies Figure 203 (Sheet 1)
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Safetying Turnbuckle Assemblies Figure 203 (Sheet 2)
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Safety Cable Installation Figure 204 (Sheet 1)
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Safety Cable Installation Figure 204 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (b) (c) (d)
Do not use a safety cable or a ferrule more than one time. Always discard the safety cable that you remove. Make sure that you use the correct type and dimension of safety cable for the applicable procedure. NOTE:
(e) (f) (g) (h) (i)
Safety cable that is not of the correct type, length, and dimension can break. This can occur when there is more than the specified tension limit for that type and dimension of safety cable.
Examine the safety cable for kinks, nicks, frayed edges, or other damage. If you find damage on the safety cable, you must discard the cable. Replace it with a new safety cable. The maximum span of the safety cable between two fasteners is 6 inches (15.24 cm). You must install the safety cable through the holes supplied for safetying. It is not permitted to install the safety cable in other locations not for safetying. Do not torque the bolt (or other fastener) to less or more than the specified value to align the holes. This is not permitted. Install the safety cable in the two-bolt pattern or the three-bolt pattern. NOTE:
The two-bolt pattern is the recommended procedure when there is an even number of fasteners.
(j) (k) (l)
Crimp the ferrule to the safety cable with one of the correct mechanical procedures. After installation, you must cut the unwanted safety cable from the ferrule that you crimped. The maximum permitted length of the safety cable that can extend from the ferrule is 0.031 inch (0.79 mm). (m) Safety the cable to the maximum extension limits. Refer to Table 203 and Figure 204. 1 Refer to Figure 204 to find the middle of the span between the two bolts. 2 Apply a light force of approximately 2 pounds (8.90 N) to the safety cable at the middle of the span. 3 Make sure that the safety cable does not stretch more than the maximum extension limits.
Table 203. Maximum Extension Limits A
B
C
0.5 inch (12.70 mm)
0.152 inch (3.17 mm)
0.062 inch (1.57 mm)
1.0 inch (25.40 mm)
0.250 inch (6.35 mm)
0.125 inch (3.17 mm)
2.0 inches (50.80 mm)
0.375 inch (9.52 mm)
0.188 inch (4.77 mm)
3.0 inches (76.20 mm)
0.375 inch (9.52 mm)
0.188 inch (4.77 mm)
4.0 inches (101.60 mm)
0.500 inch (12.70 mm)
0.250 inch (6.35 mm)
5.0 inches (127.00 mm)
0.500 inch (12.70 mm)
0.250 inch (6.35 mm)
6.0 inches (152.40 mm)
0.625 inch (15.87 mm)
0.312 inch (7.92 mm)
(2)
A fastener will stay tight if you install the safety cable correctly. While movement or tension on the fastener causes it to loosen, the cable tension increases. This will hold the fastener in its position. Refer to Figure 204 for examples of safety cable installation.
CAUTION: Do not use the safety cable or the ferrule again after you remove it. It can break if you apply too much force to it and cause damage to the equipment. (3) (4)
Install the safety cable through the holes in the fasteners. Put a loose ferrule on the safety cable.
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7) (8)
Put the end of the safety cable through the safety cable tool. Apply tension to the preset load with the safety cable tool. Crimp the ferrule with the safety cable tool. Cut the unwanted cable from the crimped ferrule. (a) Make sure that the maximum length of the cable that extends from the ferrule is not more than 0.031 inch (0.79 mm).
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MODEL 208 MAINTENANCE MANUAL CONTROL CABLE WIRE BREAKAGE AND CORROSION LIMITATIONS - MAINTENANCE PRACTICES 1.
Examination of Control Cables. A.
Control cable assemblies are subject to a variety of environmental conditions and forms of deterioration. Some deterioration, such as wire or strand breakage, is easy to recognize. Other deterioration, such as internal corrosion or cable distortion, is harder to identify. The following information will aid in detecting these cable conditions.
B.
Broken Wire Examination (Refer to Figure 201). (1) Examine cables for broken wires by passing a cloth along length of cable. This will detect broken wires, if cloth snags on cable. Critical areas for wire breakage are those sections of cable which pass through fairleads, across rub blocks, and around pulleys. If no snags are found, then no further inspection is required. If snags are found or broken wires are suspected, then a more detailed inspection is necessary which requires that the cable be bent in a loop to con firm broken wires. Loosen or remove cable to allow it to be bent in a loop as shown. While rotating cable, inspect bent area for broken wires. (2) Wire breakage criteria for cables in flap, aileron, rudder, and elevator systems are as follows: (a) Individual broken wires at random locations are acceptable in primary and secondary control cables when there are no more than six broken wires in any given ten-inch cable length.
C.
Corrosion. (1) Carefully examine any cable for corrosion that has a broken wire in a section not in contact with wear-producing airframe components, such as pulleys, fairleads, rub blocks, etc. It may be necessary to remove and bend cable to properly inspect it for internal strand corrosion, as this condition is usually not evident on outer surface of cable. Replace cable if internal corrosion is found. If a cable has been wiped clean of its corrosion-preventive lubricant and metal-brightened, the cable shall be examined closely for corrosion. For description of control cable corrosion, refer to Chapter 51, Corrosion and Corrosion Control - Maintenance Practices.
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Cable Broken Wire Examination Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL BEARINGS - REMOVAL/INSTALLATION 1.
Bearings Removal/Installation A.
Remove the Bearing (Refer to Figure 401). (1) Remove the bearing with its supporting bracket or component housing from the airplane, refer to the applicable chapter. (2) Push the worn bearing from its housing or supporting bracket. (3) After you remove the bearing, examine the component housing or bracket for structural damage (cracks, warpage or bends). (4) Examine the hole in the housing for damage, cracks or other abnormal conditions of the material and hole diameter. NOTE:
B.
The gap between the bearing outside diameter and the hole inside diameter must be within 0.0010 to 0.0035 inch.
Install the Bearing (Refer to Figure 401, Figure 402, Figure 403, and Figure 404). NOTE:
The new bearings must stay in their packages until the time of the actual installation.
CAUTION: Do not let the cleaner penetrate into the bearing. This will remove the lubrication from the bearing. (1)
Clean the outer surfaces of the bearing and hole in the component housing with a clean cloth and remove all traces of oil or grease. NOTE:
(2)
The cloth must be moist with an approved cleaning solvent.
Use a clean cloth to dry the bearing and hole.
CAUTION: Make sure that the retaining compound does not go into the bearing. (3) (4) (5)
(6) C.
Apply retaining compound to the outer surface of the bearing and the mating surface of the hole in the housing. Refer to the Application of Fastener Retaining Compounds and Table 401. Push the bearing into position. Use a staking tool to stake the bearing in place. (a) Stake between the current stake marks around the hole (refer to Figure 404). (b) If a new component housing or bracket is necessary, stake the bearing in the same pattern as the original installation. (c) If the bearing is not kept in position on the opposite side of the stake (refer to Figure 402), use a support during the staking operation (refer to Figure 403). 1 Stake the bearing housing on both sides only if the bearing is not kept in position on the opposite side of the stake. Install the bearing component or bracket assembly on the airplane, refer to the applicable chapter.
Riveted-On Bearing Brackets or Housings: (1) The replacement bearing brackets, housings or bearing and bracket assemblies can be supplied without holes. The riveted installation must be put into position and drilled. NOTE:
If you replace a bearing that is attached with rivets, the alignment of the removed bracket must be marked. The new bracket must be installed in the original alignment.
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Bearing Removal and Installation Figure 401 (Sheet 1)
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Staking Dimension Figure 402 (Sheet 1)
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Support During Staking Figure 403 (Sheet 1)
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Staking Tool - Typical Figure 404 (Sheet 1)
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2.
Application of Fastener Retaining Compounds A.
General (1) This procedure gives general methods to apply materials used to seal, lock and keep metal parts in position. The retaining compounds cure anaerobically, that is, they will become hard only when put between properly prepared mating surfaces where no air can touch the retaining compound. Refer to Table 401 for the retaining compounds and surface primers included in this procedure.
CAUTION: Make sure that the primer and retaining compounds do not touch the synthetic rubber. CAUTION: Make sure that the primer and retaining compounds do not go into the bushings or bearings. NOTE:
It is not necessary to apply primer to surfaces other than cadmium, zinc, anodized or corrosion-resistant steel.
NOTE:
For optimum strength properties, the gap between the bushing (bearing) outside diameter and the housing hole inside diameter must be within 0.0010 to 0.0035 inch.
NOTE:
Primer and retaining compounds must be kept in an enclosed building that will give the containers protection from direct sunlight, wind and rain.
Table 401. Primer and Retaining Compounds LOCQUIC SURFACE PRIMER MIL-S-22473
GRADE
FORM
COLOR
Primer (Catalog Number 747-56)
T
R
Yellow
APPLICATION (Ready to use) Can be used with Loctite 609 or 680 and Permabond HL038 or HH020.
RETAINING COMPOUND MIL-R-46082
MANUFACTURER
APPLICATION
Loctite 609
Loctite Corp. Newington, CT 06111
Used as high strength retaining compound that cures quickly for studs, bearings and bushings.
Loctite 680
Loctite Corp.
Used as extra-high strength retaining compound that cures quickly for press fits on cylindrical parts.
Permabond HL038
Permabond International Corp. 480 S. Dean Street Englewood, NJ 07631
Used as high strength retaining compound that cures quickly for studs, bearings and bushings.
Permabond HH020
Permabond International Corp.
Used as extra-high strength retaining compound that cures quickly for press fits on cylindrical parts.
3.
Bearing/Bushing Retention A.
Preparation (1) Prepare the parts to be kept as follows: (a) Before you apply retaining compound to a surface, clean it with an approved cleaning solvent. Use a clean cloth and remove all traces of grease or oil.
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Keep contamination from the cleaned surfaces, especially if they are not assembled immediately after they are cleaned. Do not touch the clean parts with your bare hands. Use a clean cloth or clean white cotton gloves when you assemble the mating parts.
CAUTION: Primer has a chemical effect on materials such as thermoplastics or titanium. NOTE: B.
Locquic Primer NOTE: (1) (2)
C.
Cadmium, zinc, anodized, corrosion resistant steel and plastic surfaces must be primed with Locquic primer, Grade T, Form R (yellow).
Apply MIL-S-22473 Locquic primer, Grade T, Form R (yellow), to all surfaces that touch. Do not apply the primer to the oil grooves or ports of the bearings. Let the surfaces air dry at room temperature for a minimum of 30 minutes.
Installation NOTE:
(1)
Bearings or bushings can be installed dry and the retaining compound applied as stated in the following step. Or a thin coat of retaining compound specified for repair can be applied to the primed surfaces that are to be joined and assembled wet.
After the installation (wet or dry), apply MIL-R-46082 retaining compound. (a) Touch the application nozzle of the retaining compound container to the mating joint between the bearing outside diameter and the housing. NOTE:
4.
Some materials that are affected by softening or crazing include vinyl, cellulosic, styrene and methacrylate plastics. Thermosetting plastics are not affected.
The compound will go into the joint by capillary action. When a ring of compound stays just outside the joint, the capillary penetration is complete.
Sealant and Retaining Compounds A.
Cure Methods (1) Two methods to cure sealant and retaining compounds are: (a) Method 1 - The parts must stay undisturbed for 24 hours at room temperature to get to full strength. NOTE:
(b) (c)
Method 2 - The parts must cure at 275°F, +10° or -10°F for 15 minutes after the part gets to the necessary temperature. Examine the bearing for damage before it is put into position. NOTE:
(d)
If the bearing or bushing moves out of position before the full cure of the retaining compound is complete, the parts must be cleaned, primed and assembled again.
Any damaged area of the bearing must be repaired before it is put into position.
Lubricate the bearing or bushing after the retaining compound has cured.
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL BONDING - MAINTENANCE PRACTICES 1.
General A.
This section describes airplane electrical bonding requirements and procedures. The following procedures and specification MIL-STD-464 - Electromagnetic Environmental Effects Requirements for Systems, govern installation and testing of electrical bonds and ground returns.
B.
Some electrical bonding requirements for static discharge wicks are shown in Figure 214 in this section. Refer to Chapter 23, Static Discharging - Maintenance Practices for more required static discharging procedures.
C.
Poor electrical bonding can cause or contribute to a variety of operational problems in electrical, avionics, and communication systems, such as complete failure, reduced performance, or electromagnetic interference (EMI), or radio frequency interference (RFI) with navigation and communication systems.
D.
Maintenance personnel must follow recommended practices for establishing, remaking, testing, and protecting electrical bonds, particularly during routine maintenance activities. NOTE: (1) (2) (3)
2.
If a component is moved or the bond of an installed component is otherwise broken, the resistance of the connection must be verified again after installation.
Removal and installation of avionics and electrical equipment and mounting trays. Assembly and reassembly of supporting structure for avionics or electrical equipment. Reinstallation of control surfaces and removable fairings (including radome and stinger).
Definitions A.
In this section, the applicable definitions are as follows: (1) Method A (nonconductive) bonding is the usual method to seal the surfaces. This is done by cleaning a surface area that is larger than the connector that will be installed. A fillet seal of nonconductive sealant is then applied on and around the connector and the metal bonding surface. To make sure that the finish is nonconductive, it is necessary that you use a microohmmeter to do an electrical bonding test. (a) The cleaned area must be between 0.063 inch (1.59 mm) and 0.250 inch (6.35 mm) more than the connector. (b) The sealant must not be applied on screws, rivets, or other mounting hardware. (2) Method B (conductive) bonding is a different method to seal the surfaces. This is done by cleaning a surface area that is smaller than the connector that will be installed. A fay seal of conductive sealant is then applied between the connector and the metal bonding surface. (a) The cleaned area must not be more than, and not less than, the diameter of the contact area. (b) The sealant must be applied between the connector and the bonding surface. (c) The sealant on and around the bonding must then be removed (cleaned). (3) Bond (Electrical) - A fixed union existing between two objects that provides good electrical conductivity due to a low-resistance path between the objects. (4) Bonding (Electrical) - The process of making the necessary connections to provide good electrical conductivity between units or between unit and airplane ground. (5) Ground - The common connection of the electrical circuits of a system or subsystem to a conductive medium (airplane primary structure) that becomes a common reference plane for all voltage potentials in the airplane. (6) Grounding - The process of making the necessary connection(s) to provide a ground for an electrical, electronic, or radio frequency circuit. (7) Primary structure - Load carrying members of the airframe, such as bulkheads, ribs, webs, and stringers, extending through two or more bulkheads or ribs. All primary structure is considered to be airplane ground. (8) Secondary structure - Sheet metal or extruded metal parts attached to primary structure (or to secondary structure that is ultimately attached to primary structure) in at least two places by structural fasteners or by three or more rivets on each location.
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3.
Hardware and Material Usage A.
The hardware and materials that are recommended to be used and those that are not to be used for electrical bonding are shown in the tables that follow:
Table 201. Cleaning Material Aluminum Wool Sandpaper Stainless Steel Wool Aluminum Oxide Cloth, High Purity Bonding Rotary Brush P-D-680 Solvent Table 202. Bolts, Nuts, and Screws Cadmium Plated Steel
Recommended for all areas other than engine compartment and fuel system.
Corrosion Resistant Steel
Recommended for engine compartment and fuel system.
Aluminum
Recommended for all areas other than engine compartment.
Self-Tapping Screws
Not to be used for bonding application.
Zinc Plated Screws
Not to be used for bonding application.
Spring, Self-Locking, Clip-in Instrument Mounting Nut
Not to be used for bonding application.
Wing Nuts
Not to be used for bonding application.
Table 203. Washers Anodized
Not to be used for bonding application.
Zinc Plated
Not to be used for bonding application.
Unplated
Not to be used for bonding application.
NAS1149, MS35337, MS35339
Recommended for bonding in all areas.
Tooth Lock Washer
Not to be used for bonding application.
Table 204. Bonding Jumpers MS25083
Recommended for bonding in all areas, other than fuel tanks.
On Aluminum Alloys
Use aluminum or MS25083 tinned copper jumpers.
On Steel Alloys
Use MS25083 copper, brass or bronze tinned coated jumpers only.
In Fuel System
Use S2876 aluminum only.
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Table 205. Clamps AN735
Recommended for bonding in all areas.
Cushion Clamps
Not to be used for bonding applications.
AN742
Not to be used for bonding applications.
Table 206. Nutplates Cadmium and Silver Plated (Floating or Nonfloating)
Recommended for direct bonding applications.
Nonmetallic Insert, Dry Film Lube-Type
Not to be used for direct bonding applications.
Table 207. Low-Resistance Test Set (Bonding Meter) Keithley Model 580 Micro-Ohmmeter (or Equivalent)
Keithley Instruments, Inc. Instrument Division 28775 Aurora Road Cleveland, OH 44139
Megohmmeter Model 2850 (or Equivalent)
Associated Research, Inc. 3773 West Belmont Ave. Chicago, IL 60618
Table 208. Permanent Ground Studs VNS1924CA3-1-8
Recommended for permanent ground bonding in all areas.
Table 209. Permanent Ground Stud Pulling Head Adapter VST1116-10 VST1116-8
4.
Voi-Shan 8463 Higuera Street Culver City, CA 90232-0512
Bonding Surface Cleaning NOTE:
If you can bond through the fasteners, it is recommended that you use Method B to install static wicks. If you cannot bond through the fasteners, it is necessary that you use Method B to install static wicks.
NOTE:
Bonds between two metal surfaces and bonding jumper attachment points must be free of insulating material, such as paint, primer, grease, oil, and materials that prevent corrosion. A clean area is needed to make sure there is an adequate bond and to help get resistance measurements. The cleaned surface must be a clean and smooth finish that has not had too much material removed under the protective finish.
A.
Method A (nonconductive) Surface Cleaning (1) Clean the bonding surface area that is larger than the connector to be installed. (a) Make sure that the cleaned area is between 0.063 inch (1.59 mm) and 0.250 inch (6.35 mm) more than the connector.
B.
Method B (Conductive) Surface Cleaning (1) Clean the bonding surface area that is smaller than the connector to be installed.
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MODEL 208 MAINTENANCE MANUAL (2) C.
A fay seal of conductive Type XV, Class B sealant is then applied between the metal surfaces.
Method A Steel and Aluminum Surface Cleaning (1) Use medium Roloc surface condition disc pads (Scotchbrite) or medium EXL wheel 6A (Scotchbrite) that is no more than 0.50 inch (12.7 mm) larger than the diameter of the bonding surface of the connector.
(2) (3)
NOTE:
You are permitted to use 400 through 600 grit emery paper or cloth, or an equivalent fine sandpaper and/or aluminum oxide paper or cloth, stainless steel wool, or a stainless steel or monel bonding brush.
NOTE:
You can use aluminum wool only on aluminum.
Use IPA, MPK, or equivalent to clean the bonding surfaces. To clean aluminum that is not for a bonding jumper, make sure that the bonding surface is between 0.50 inch (12.7 mm) and 0.250 inch (6.35 mm) more than the connector. NOTE:
(4) (5)
To attach a bonding jumper, clean the bonding surface area to 150 percent of the diameter of the bonding jumper terminal. For bare aluminum, before an electrical bond is made, apply a chemical film treatment to the bonding surface. NOTE:
D.
This will give the bond electrical and some corrosion protection.
Method B Steel and Aluminum Surface Cleaning (1) Use medium Roloc surface condition disc pads (Scotchbrite) or medium EXL wheel 6A (Scotchbrite) that is no larger than the diameter of the bonding surface of the connector.
(2) (3)
NOTE:
You are permitted to use 400 through 600 grit emery paper or cloth, or an equivalent fine sandpaper and/or aluminum oxide paper or cloth, stainless steel wool, or a stainless steel or monel bonding brush.
NOTE:
You can use aluminum wool only on aluminum.
Use IPA, MPK, or equivalent to clean the bonding surfaces. To clean aluminum that is not for a bonding jumper, make sure that the bonding surface is no larger than the diameter, but no smaller than 50 percent of the bonding surface of the connector. NOTE:
(4)
If the cleaned area is larger than the diameter of the bonding surface, the surface must be primed, dried, then cleaned again.
For bare aluminum, before an electrical bond is made, apply a chemical film treatment to the bonding surface. NOTE:
E.
If the cleaned area is more than 0.50 inch (12.7 mm) more than the connector, you must apply primer, let it dry, then clean it again.
This will give the bond electrical and some corrosion protection.
Method A Magnesium Surface Cleaning
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CAUTION: Do not use steel wool, stainless steel wool, or aluminum wool to clean magnesium alloys. (1)
Use medium Roloc surface condition disc pads (Scotchbrite) or medium EXL wheel 6A (Scotchbrite) that is no larger and no smaller than the diameter of the bonding surface of the connector. NOTE:
(2) F.
You are permitted to use 400 through 600 grit emery paper or cloth, or an equivalent fine sandpaper.
Use IPA, MPK, or equivalent to clean the bonding surfaces.
Method B Magnesium Surface Cleaning
CAUTION: Do not use steel wool, stainless steel wool, or aluminum wool to clean magnesium alloys. (1)
Use medium Roloc surface condition disc pads (Scotchbrite) or medium EXL wheel 6A (Scotchbrite) that is no larger and no smaller than the diameter of the bonding surface of the connector. NOTE:
(2) 5.
You are permitted to use 400 through 600 grit emery paper or cloth, or an equivalent fine sandpaper.
Use IPA, MPK, or equivalent to clean the bonding surfaces.
Protective Coating Sealing NOTE:
Although you must apply a corrosion inhibitor to seal the perimeter of some electrical bonds after they are assembled, some electrical bonds must have a finish applied before you can apply the corrosion inhibitor.
NOTE:
Method A (nonconductive) Sealing is done after the electrical bond is assembled. To make sure that the finish is nonconductive, it is necessary that you use a micro-ohmmeter to do an electrical bonding test.
NOTE:
Method B (conductive) Sealing is done during the electrical bond assembly.
A.
All bonded surfaces requiring protective coating must be refinished per the original finish or color chemical film treated within as short a time as possible. Refinishing within a 24-hour period is highly recommended.
B.
Sealant as Protective Coatings (1) Do not use sealants as a protective coating on bulkhead electrical connectors in the pressure vessel (unless required for pressure seal). (2) Do not apply sealants to equipment racks and equipment mounting surfaces. (3) Do not apply sealants to stud-type ground blocks. (4) Do not use sealants as a protective coating to feed-thru plates (unless required for the pressure seal). NOTE:
In areas where the surface already has chemical film applied, such as feed-thru plates on pressure bulkheads, it is not required to remove this finish and reapply chemical film to achieve bonding unless the bonding requirements cannot be met.
NOTE:
A cleaned area must not be refinished until the electrical bond connection has been inspected and approved.
NOTE:
Bonding jumpers do not need painting.
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Method A (Nonconductive) Sealing
WARNING: Do not use conductive sealant in fuel tanks. Use only nonconductive sealants that are approved for use in fuel tanks. The metal material in the conductive sealant can create a spark (arcing) in the fuel tank. NOTE:
For Method A sealing, it is recommended that you use corrosion-inhibitive polysulfidebased sealant.
NOTE:
If you can bond through the fasteners, it is recommended that you use Method B to install static wicks. If you cannot bond through the fasteners, it is necessary that you use Method B to install static wicks.
(1)
(2)
(3)
D.
Method A sealing is applicable to use in the areas that follow: • Feed-thru connectors out of the pressure vessels • All electrical bonding in the fuel tanks • Surface-mounted ground blocks where the fayed surface must be removed for electrical bonding • Ground studs out of the pressurized area • Bonding jumpers and straps out of the pressurized area For Method A sealing, you can use one of the sealants that follow: • Type X, Class B sealant (for all but in fuel tank areas) • Type I, Class B sealant (Polysulfide–based fuel, weather, and pressure sealant) • Type V, Class A sealant (RTV silicone sealant) • Type V, Class B sealant (RTV silicone sealant) Apply a fillet seal on and around the connector and the metal bonding surface. (a) To make sure that the finish is nonconductive, it is necessary that you use a micro-ohmmeter to do an electrical bonding test. (b) Do not apply sealant on screws, rivets, or other mounting hardware. (c) Make sure that the sealant is applied between 0.063 and 0.125 inch (1.6 and 3.2 mm) larger than the cleaned area.
Method B (Conductive) Sealing
WARNING: Do not use conductive sealant in fuel tanks. Use only nonconductive sealants that are approved for use in fuel tanks. The metal material in the conductive sealant can create a spark (arcing) in the fuel tank. NOTE:
For Method B sealing, you must use corrosion-inhibitive polysulfide-based sealant.
NOTE:
If you can bond through the fasteners, it is recommended that you use Method B to install static wicks. If you cannot bond through the fasteners, it is necessary that you use Method B to install static wicks.
(1)
(2)
Method B sealing is applicable to use in the areas that follow: • Feed-thru connectors out of the pressure vessels • Surface-mounted ground blocks where the fayed surface must be removed for electrical bonding • Ground studs out of the pressurized area • Bonding jumpers and straps out of the pressurized area Use Type XV, Class B conductive sealant to apply a fillet seal on and around the bonding surface and the connector. (a) Do not use this sealant in fuel tank areas. (b) Make sure that the area of the sealant is no larger than the bonding surface of the connector.
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MODEL 208 MAINTENANCE MANUAL (c) 6.
Clean all sealant off of the area that is more than permitted.
Electrical Bonding Test A.
Electrical Bonding Criteria (1) Bonding to primary structure (grounding) - Do this type of bonding only after the resistance between the bonded object and primary structure has been measured with a low-resistance test set (bonding meter) and found to be no more than the maximum allowable resistance for that application. For typical resistance values, refer to Table 210. NOTE:
(2)
B.
Bonding one object to another - This type of bonding must be considered satisfactory only after the resistance between the objects has been measured with a bonding meter and found to be no more than the maximum allowable resistance for that application.
Using a Low-Resistance Test Set (Bonding Meter) (1) Follow manufacturers instructions included with test set for setup, operation, and reading of test set display. (2) Place or connect the probes on bare metal surfaces. (3) The probes should be placed as close as possible to the bonding area, preferably within six inches along surface of the object or structural member. (4) If it is necessary to remove paint or primer from a surface in order to provide good probe contact, apply the original (or equivalent) finish after an electrical bonding test. NOTE:
C.
7.
In many cases the object to be grounded is mounted on or attached to secondary structure, or otherwise separated from direct contact with primary structure. Grounding depends on both the satisfactory bonding of the mounting tray to secondary structure, and of the secondary structure to primary structure. The electrical bonding test is then done to measure resistance across each bond to identify the source(s) of poor grounding.
To make sure that the finish is nonconductive, it is necessary that you use a microohmmeter to do an electrical bonding test.
Electrical Bonding Test of Composite Panels (1) Leave out or remove one screw per 4 lineal feet of panel edge. (2) Make sure that countersink is free of paint or other insulating material. (3) Do the electrical bonding test between countersink(s) and primary structure.
Electrical Bond Type (Class) A.
Electrical Bond Classes NOTE: (1) (2)
(3) (4)
The classes of electrical bonds are given in accordance to the type of material that is bonded together and the method used to bond the materials.
Type I usually applies to metallic components bonded together with direct metal to metal contact. Some examples are riveted skin bonds, equipment racks, and bulkhead connectors. These components will have a requirement of less than 0.0025 ohm maximum. Type II usually applies to aluminum or steel components (i.e., landing gear, doors, and/or airplane structure) bonded together electrically by bonding jumpers, but can also be used when applied to bonds with multiple metal to metal contacts, such as, radome diverter strips to airplane structure or antennas mounted on metal fairings to the airplane structure with a requirement of less than 0.005 ohms. Type III usually applies to adhesively bonded aluminum components bonded together electrically by bonding jumpers with a requirement of less than 0.005 ohms. Type IV usually applies to connections between expanded aluminum mesh (or similar material) and the airframe through bonding jumpers and clips with a requirement of less than 0.005 ohms.
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7) (8)
(9) (10)
(11)
(12) (13)
Type V usually applies to connections between conductive composite structure, such as, Carbon Fiber Composite (Graphite), and its metallic attachment hardware with a requirement of less than 0.005 ohms. Type VI usually applies to connections for P-static protection between composite materials and metallic airplane structure with a requirement of less than 100,000 ohms. Type VI-A usually applies to connections for P-static protection between P-static paint and a high-resistive, low-conductive gasket used in nonconductive fuel doors and the airplane structure with a requirement of less than 10,000,000 ohms Type VII usually applies to connections for P-static protection between nonconductive materials used in radomes or electrically-heated windshields and their metallic attachment hardware and/ or airplane structure with a requirement of 1,000,000 ohms, but not more than 100,000,000 ohms. Type VIII usually applies to connections for low conductive gaskets and metallic airplane structure with a requirement of 1,000,000 ohms or greater. Type IX usually applies to connections for hydraulic and fuel lines and tubes, metallic tubing, seat frames (nonelectrical components) and electrical switches, circuit breakers, and ducts bonded to the airplane structure by different means, such as, clamps or attachment screws. Because this type includes many different installations, the maximum permitted resistance value can be different from one installation to another. These differences are specified in Table 210. Type X usually applies to connections for different fuel system hardware, such as, fuel filler nozzle, fuel vents, and fuel gages. Because this type includes many different installations, the maximum permitted resistance value can be different from one installation to another. These differences are specified in Table 210. Type XI usually applies to connections for flaps, slats, piano hinged surfaces, and roller bearing surfaces. This type includes several different installations. The maximum permitted resistance will be different from one installation to another. These differences are specified in Table 210. Type XII usually applies to return-path grounds. This is for ground studs installed in metal airplane structures. Refer to Table 210.
Table 210. Typical Resistance Values OBJECT TO BE BONDED
METHOD OF BONDING
MAXIMUM ALLOWABLE RESISTANCE VALUE (Ohms)
BOND TYPE (CLASS)
All electrical and electronic equipment ground return bonds to basic structure
Direct metal case to structure
0.0025
I
0.005
II
Antenna base
Metal base to metal fuselage through fasteners
0.0025
I
Metal base to metal screen or expanded metal
0.05
IV
Metal base to composite fairing through fasteners
0.5
V
Battery box to basic structure
Direct metal case to structure
0.0025
I
Bulkhead feed-thru connectors
Metal to metal
0.0025
I
Cable bundle shields
Direct attachment to connector backshell
0.0025
I
Electric trim (actuator assembly)
Direct metal to metal
0.0025
I
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MODEL 208 MAINTENANCE MANUAL Table 210. Typical Resistance Values (continued) OBJECT TO BE BONDED
METHOD OF BONDING
MAXIMUM ALLOWABLE RESISTANCE VALUE (Ohms)
BOND TYPE (CLASS)
Electrical devices to enclosure
Direct attachment through attachment hardware or bonding jumper
0.0025
I
Electrical motors to adjacent structure
Direct metal to metal
0.0025
I
Engine to nacelle structure bond
Direct attachment by fasteners (metal nacelle)
0.0025
I
Honeycomb panel assemblies
Direct metal to metal attachment by fasteners
0.0025
I
Instrument panels to stationary panel
Direct metal to metal attachment by fasteners
0.0025
I
Instruments
Direct metal to metal
0.0025
I
Radio racks, shelves and brackets to adjacent primary structure
Direct metal to metal
0.0025
I
Radome external diverter strips
Attachment to radome frame
0.0025
I
RFI noise filters (across joint)
Direct metal to metal
0.0025
I
Rivet skin joints and breaks (across joint) or structural joints or breaks (across joint)
Direct metal to metal
0.0025
I
Servos amplifier, gaging equipment instruments, etc.
Direct metal to metal
0.0025
I
Side console and electrical equipment panels to basic structure
Direct metal to metal by fasteners
0.0025
I
Starters, generator and alternator grounds (case to engine frame)
Direct metal case to structure
0.0025
I
Static wicks (metal surface)
Direct metal to metal attachment by fasteners
0.0025
I
Stationary instrument panels to primary structure
Direct metal to metal by fasteners
0.0025
I
Structural joints or breaks (across joint)
Direct metal to metal
0.0025
I
Wing tie down and ground point
Attachment through fasteners
0.0025
I
Wing to fuselage
Direct metal to metal or attachment bolts
0.0025
I
Wire Bundle Shields
Direct attachment to backshell
0.0025
I
0.005
II
All electrical and electronic equipment ground return bonds to basic structure
Bonding jumper to structure
0.005
II
Baggage/Avionics compartment door
Bonding jumper across hinge
0.005
II
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MODEL 208 MAINTENANCE MANUAL Table 210. Typical Resistance Values (continued) OBJECT TO BE BONDED
METHOD OF BONDING
MAXIMUM ALLOWABLE RESISTANCE VALUE (Ohms)
BOND TYPE (CLASS)
Battery box cover
Fastener to battery box
0.005
II
Bearings (roller and ball) Piano hinged surfaces
Bonding jumper
0.005
II
Cable bundle shields
Bonding jumper
0.005
II
Composite aileron
Bonding jumper across hinge or graphite structure or embedded metal screen/mesh to airplane structure
0.005
II
0.5
V
0.05
IV
Bonding jumper across hinge or graphite structure or embedded metal screen/mesh to airplane structure
0.005
II
0.5
V
0.05
IV
Bonding jumper across actuator or roller/track (loaded configuration)
0.005
II
0.5
XI
0.5
V
Bonding jumper across hinge or graphite structure or embedded metal screen/mesh to airplane structure
0.005
II
0.5
V
0.05
IV
Doors and Inspection plates
Fastener to airplane structure
0.005
II
Electric trim (actuator assembly)
Direct metal to metal or bonding jumper
0.0025
I
0.005
II
Electrical devices to enclosure
Bonding jumper
0.005
II
Electrical motors to adjacent structure
Bonding jumper
0.0025
I
0.005
II
Composite elevator
Composite flap
Composite rudder
Engine
Bonding jumper across mount
0.005
II
Honeycomb panel assemblies
Direct metal to metal attachment by fasteners or by bonding agent
0.0025
I
0.005
II
Instruments
Bonding jumper
0.005
II
Landing gear
Bonding jumper
0.005
II
Landing gear doors
Bonding jumper across hinge
0.005
II
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MODEL 208 MAINTENANCE MANUAL Table 210. Typical Resistance Values (continued) OBJECT TO BE BONDED
METHOD OF BONDING
MAXIMUM ALLOWABLE RESISTANCE VALUE (Ohms)
BOND TYPE (CLASS)
Metal aileron
Bonding jumper across hinge
0.005
II
Metal cowls - removable
Fastener to engine frame
0.005
II
Metal elevator
Bonding jumper across hinge
0.005
II
Metal flap
Bonding jumper across actuator or roller/track (loaded configuration)
0.005
II
0.5
XI
Metal nacelle
Fastener to engine frame
0.005
II
Metal rudder
Bonding jumper across hinge
0.005
II
Nose wheel doors
Bonding jumper across hinge
0.005
II
Panel feed-thru plates
Direct metal to adhesively bonded honeycomb or bonding jumper
0.03
III
0.005
II
Radome external diverter strips
Attachment to radome frame
0.005
II
Radome
Diverter strips through radome frame to airplane by attachment hardware
0.010
II (NOTE: This measurement involves multiple bond paths. Therefore, the allowable resistance value is twice standard Type II)
Servos amplifier, gaging equipment instruments, etc.
Bonding jumper
0.005
II
Spoiler
Bonding jumper across hinge or directly through hinge
0.005
II
Trim tab
Bonding jumper across hinge
0.005
II
Wire Bundle Shields
Bonding jumper to structure
0.005
II
Panel feed-thru plates
Direct metal to adhesively bonded honeycomb
0.03
III
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MODEL 208 MAINTENANCE MANUAL Table 210. Typical Resistance Values (continued) METHOD OF BONDING
MAXIMUM ALLOWABLE RESISTANCE VALUE (Ohms)
BOND TYPE (CLASS)
Composite aileron
Embedded metal screen/mesh to airplane structure
0.05
IV
Composite elevator
Embedded metal screen/mesh to airplane structure
0.05
IV
Composite rudder
Embedded metal screen/mesh to airplane structure
0.05
IV
Antenna base
Metal base to composite fairing through fasteners
0.05
V
Composite aileron
Graphite composite to airplane structure
0.5
V
Composite cowls - removable
Fastener to engine frame
0.5
V
Composite elevator
Graphite composite to airplane structure
0.5
V
Composite flap
Graphite composite to airplane structure
0.5
V
Composite nacelle
Fastener to engine frame
0.5
V
Composite rudder
Graphite composite to airplane structure
0.5
V
Engine to nacelle structure bond
Direct attachment by fasteners (composite nacelle)
0.5
V
Static wicks (composite surface)
Direct metal to composite surface and attachment by fasteners
0.5
V
Control cables and rods to movable surfaces equipment
Bonding through attachment hardware
0.01
IX
Electrical switches, circuit breakers and potentiometer in circuits exceeding 50 volts
Direct attachment through hardware
0.10
IX
Hydraulic cylinders
Direct metal to metal or bonding jumper
0.01
IX
Metal ducts
Direct metal to metal and attachment anchors
0.005
IX
Metal ducts (nonelectrical: rigid and flexible)
Bonding through clamp and attachment hardware
1.0
IX
Metallic tubing
Cable runs which terminate in a direct metal to metal contact
1.0
IX
Metallic tubing
Direct metal to metal
0.0025
I
OBJECT TO BE BONDED
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MODEL 208 MAINTENANCE MANUAL Table 210. Typical Resistance Values (continued) METHOD OF BONDING
MAXIMUM ALLOWABLE RESISTANCE VALUE (Ohms)
BOND TYPE (CLASS)
Metallic tube wiggins fittings
Feed-thru hardware and attachment anchors
0.1
IX
Oil and Anti-Ice tank
Attachment through fasteners
0.1
IX
Oxygen cylinders
Direct metal to metal or bonding jumper
0.01
IX
Radiators and heat exchangers
Bonding through attachment hardware
0.01
IX
Seat frame
Attachment through fasteners
0.01
IX
Shock mounts
Direct metal to metal through attachment hardware
0.01
IX
Fuel filler nozzle
Attachment through fasteners
0.005
X
Fuel nozzle jumper ground receptacle
Attachment through fasteners
0.003
X
Fuel vents scoops
Attachment through fasteners
0.005
X
Bearings (roller and ball) Piano hinged surfaces
Metal to metal through bearing or hinge
0.01
XI
Composite flap
Roller/track (loaded configuration)
0.5
XI
Metal flap
Roller/track (loaded configuration)
0.5
XI
Slats
Roller/track (loaded configuration)
0.5
XI
Hoods and canopies
Not applicable
Not required
Not required
Ground Studs
Base of the ground stud to the aircraft structure
0.0005
XII
OBJECT TO BE BONDED
8.
Bonding Requirements A.
Current Path Return Bonds (1) Current return bonds are those required to complete the ground return path to the battery and/ or the power generator source for all electrical and avionics equipment. This type of bond is accomplished with a standard hook-up wire. The location of the ground bond connection should be to primary structure. In some cases where the equipment is internally case grounded, current return may be accomplished by direct bonding of the surfaces and through the mounting hardware. (2) If the bonding surface resistance permitted in Table 210 cannot be met by direct surface bonding, then the equipment or component in question must be bonded by a bolted bonding jumper.
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MODEL 208 MAINTENANCE MANUAL B.
Radio Frequency and Static Bonds (1) All electrical and electronic equipment and/or components should be installed in a manner to provide a continuous low-impedance path from the equipment enclosure to airplane primary structure. (2) All metallic pipes, tubes and hoses carrying fluids in motion should be bonded to basic structure. (3) Table 210 includes equipment and areas of radio frequency and static bonds. (4) All control surfaces should have a bonding jumper between the airframe and the control surface. Where necessary, additional jumpers should be used between the control surface and structure to achieve the resistance level. A piano-type hinge may be considered as self- bonded, provided the resistance across the hinge halves is satisfactory. (5) All conducting items, such as metal lines and/or tubing carrying fluids or air in motion having a linear dimension of 24 inches or more and installed within one foot of unshielded transmitting antenna lead-ins should have a bond to structure. Refer to bonding of pipes and tubing.
C.
Shock Hazard and Lightning Protection Bonds (1) If the requirements of current path return bonds and radio frequency and static bonds have been successfully accomplished, then shock hazard and lightning protection bonds have been partially fulfilled. (2) Shock hazard pertaining to exposed conducting frames or surfaces (such as elevators, flaps, trim tabs) or parts of electrical or electronic equipment must have a low- resistance bond to primary structure. (3) Lightning protection is the bonding of cover assemblies, such as fuel fillers, fuel vents, pitot tubes, radome, plastic and fiberglass surfaces and control surfaces. (4) Typical resistance values are shown in Table 210.
D.
Bonding in Hazard Areas (1) To eliminate any possible source of ignition in areas prone to explosion or fire hazards, do not add any new bonding installations in hazard areas. NOTE:
E.
Bonding Verification (1) For Type I through Type V, a small area of the protective finish must be removed or omitted to allow each probe to make electrical contact with the metallic or conductive composite surface. These test areas should be in close proximity to each other on adjacent sections of the airplane. On large panels it is recommended to do several tests approximately 4 feet (1.22 m) apart. (2) Electrical bonding test for composite panels is to be performed at clean screw countersink to metallic structure.
(3)
F.
Current return grounds must be avoided in fuel vapor areas.
NOTE:
To make sure that the finish is nonconductive, it is necessary that you use a microohmmeter to do an electrical bonding test.
NOTE:
Upon completion of the electrical bonding test, the metallic surface must be refinished to protect from corrosion.
For Type VI thru VIII the bond between antistatic paint and metallic attachment hardware is measured using a megohmmeter. A piece of metallic tape is placed directly on the antistatic paint approximately 1.0 inch (25.4 mm) away from attachment hardware. The measurement is accomplished by attaching one probe to the metallic tape and the other probe to the metallic hardware. The measurement should be made with the megohmmeter placed on the 500 volt setting.
Bonding Connection (1) Bonding connections must be installed so vibration, expansion, contraction or relative movement, incident to normal service use, will not break or loosen the connection to the extent that resistance will vary during the movement. (a) Bonding connections must be located in a protected area and whenever possible near an inspection door or an accessible location to permit inspection or replacement.
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MODEL 208 MAINTENANCE MANUAL (b)
Parts must be bonded directly to the primary structure, rather than through other bonded parts, such as plumbing, conduits, etc. (c) All parts must be bonded to the primary structure with as short a lead as possible. (d) Bonding jumpers must be installed so that movable components are not impeded in their operation by the jumper. (e) Bonding connections must not be made by compression fastened through nonmetallic materials. (f) All bonding surfaces must be cleaned prior to installation of bond joint. (g) All nuts used in bonding must be of the all metal self-locking type (no nonmetallic inserts). (h) Radio frequency current returns must not be made through magnesium alloys. (i) Solder joints alone must not be used for bonding parts that are subject to movement and/or vibration. (j) All electrical bonding must be accomplished without affecting the structural integrity of the airframe. (k) Nonmetallic inserts or dry film lube nutplates must not be used for bonding application, such as antenna installation. (l) All AC ground returns must be connected separate from DC ground returns. (m) Shielded wire grounds must be attached directly to the primary structure unless otherwise noted. (n) Where possible, multiple bonding jumpers or dual system grounds (left system and right system) must not be connected to the same ground point on the primary structure. (o) Electrical bonding procedure must not reduce corrosion protection of bonded objects. NOTE: 9.
Apply sealant or corrosion resistant primer to bare metal for corrosion protection.
Bonding Methods A.
The following bonding methods are provided to accomplish satisfactory bonds on the airplane. In most cases, a single method will satisfy the requirements, while in others it may be necessary to use more than one method. (1) Typical Bolted Bonding Jumper Installation (a) All bolted type jumpers should be as shown in Figure 201 and Figure 202. All jumper connections are made with number eight screws. Number six screws are used where edge distance will not permit the use of number eight screws. (2) Bonding by Riveted and/or Bolted Skin Construction (a) Close riveted and/or bolted skin construction is considered an adequate bond, provided the resistance value between bonding surfaces is 0.0025 ohm or less for current path return areas and 0.005 ohm for other areas. (b) When bonding by riveted and/or bolted skin construction only is not possible, bonding as shown in Figure 203 and Figure 204 should be done. (3) Bonding by Riveted and/or Bolted Bracket and Angle Construction (a) Close riveted and/or bolted bracket and angle construction is considered to be an adequate bond provided the resistance value across bonding surfaces is 0.005 ohm or less. (b) If the bracket and angle construction is used for current return path, then the resistance value across bonding surfaces and to primary structure must be 0.0025 ohm or less. (c) Areas not meeting the requirements noted above must be bonded in accordance with Figure 205 or by adding a bonding jumper across each joint to the primary structure. (4) Bonding of Pipes and Tubing (a) Metallic pipes and tubes supported with clean metal clamps to a metal structure member is thought to be a sufficient bond if the resistance value between the pipe or tube and primary structure is 0.1 ohm or less, except for inline Wiggins couplings, which are 1.0 ohm or less. Lines not meeting this requirement must be bonded as shown in Figure 206 and Figure 207. (b) Metallic tubing routed from a valve through the airplane structure is a sufficient bond if the resistance value between the tubing and the airplane structure is 1.0 ohm or less. The resistance value between the bulkhead fitting or the valve to the airplane structure must be 0.0025 ohm or less. Refer to Figure 213.
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MODEL 208 MAINTENANCE MANUAL (5)
(6)
Typical Access Panel or Door (a) Fastening hardware such as screws, latches, and hinges are considered sufficient bonding for access panels and doors if the resistance value to the structure is 0.005 ohm or less. Refer to Figure 208. Typical Antenna Bonding Installation (a) Fastening hardware, such as screws, nuts and nutplates, are considered adequate bonding for radio antennas, such as ADF loop, marker beacon, NAV, COM, etc., provided the fasteners have a direct contact with the metal base of the antenna and the fasteners are used in conjunction with nutplates or attachment hardware which are riveted into the structure of the airplane. (b) It is critical that all finishes which are nonconducting are removed from the interfacing contact area of the countersink and fastening hardware; however, it is usually not required to remove the finishes from the antenna mounting flange or bearing surface, providing the antenna meets the resistance value to structure as defined in Table 210. NOTE: (c) (d)
To make sure that the finish is nonconductive, it is necessary that you use a micro-ohmmeter to do an electrical bonding test.
If the bonding requirement cannot be met by the above procedure, it may be required to remove the finishes from both the antenna flange and the airplane mounting surface. If an antenna is mounted on a composite fairing, it may be required to remove the finishes from the antenna flange and clean the composite surface down to the outer layer. NOTE:
The outer layer will usually be the lightning protection material and care must be taken not to sand through the layer and conductive sealant used in conduction with the fasteners to achieve the bond.
(e)
(7)
On antennas not meeting the requirements noted above, refer to Figure 209 and, if necessary, do the bonding as follows: 1 The bearing surface between the mounting screw head and antenna metal insert must be clean, free of paint and all insulating material. This must be done on at least 25 percent of the total mounting screws used for the installation of the antenna. 2 Screw head, nut and/or nutplate structure bearing surface must be clean and bonded by riveted and/or bolted skin construction. If required, the antenna mounting doubler must be bonded to basic structure by the same method. Ground Studs NOTE:
If ground stud is positioned where it will be exposed to moisture, such as in wheel well, RTV 108 sealant must be used to seal ground stud.
(a)
(8)
It is recommended to attach single ground wires to permanently installed ground studs rather than attach them to primary structure with removable screws. The stud must be correctly bonded to the primary structure and sealed for a permanent installation. The ground wires can then be attached to and removed from the stud with no effect to the bond. Refer to Figure 210, Figure 211, and Figure 212. Static Wicks NOTE:
(a)
If you can bond through the fasteners, it is recommended that you use Method B to install static wicks. If you cannot bond through the fasteners, it is necessary that you use Method B to install static wicks.
Some electrical bonding requirements for bonding static wicks are shown in Figure 214. Refer to Chapter 23, Static Discharging - Maintenance Practices for more required static discharging procedures.
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MODEL 208 MAINTENANCE MANUAL
10.
Bonding Protection A.
Bonding protection preserves the integrity of the electrical bond by preventing the entry of water, dirt, grease, oil, and corrosion between the bonded surfaces. It also prevents corrosion damage to structure.
B.
Finish (1) All bonded surfaces requiring protective coating must be brushed chem film and refinished per the original finish within as short a time as possible. (2) Refinishing within a 24-hour period is highly recommended. The surface should be brought back to original condition. (3) Sealants are not recommended in the areas that follow: (a) Bulkhead connectors within the pressure vessel (unless required for pressure seal). (b) Equipment racks and equipment mounting surfaces. (c) Stud-type grounding blocks. (d) Feed-thru plates (unless required for pressure seal). (4) In areas where the surface already has chem film applied (i.e., feed-thru plates on aft and mid pressure bulkheads, etc.), it is not required to remove this finish and reapply chem film to achieve the bonding unless the bonding requirement cannot be met. (5) The cleaned area must not be refinished until the electrical bond connection has been inspected and approved. (6) Bonding jumpers do not need to be painted.
C.
Sealing (1) Fillet seal the perimeter of all electrical bonding areas with Type I, or Type V, Class C sealant. (Refer to Fuel, Weather, Pressure, and High Temperature Sealing - Maintenance Practices.) (2) Some areas, but not necessarily all, where it is applicable to seal, are as follows: (a) Feed-thru connectors which are exposed to wide temperature changes, such as temperature changes outside the pressure vessel. (b) Any bonding surface required within the fuel cells. (c) Grounding blocks which require a large amount of the surface to be removed to achieve the electrical bond. (3) Apply the same type (or equivalent) of protective coating that was initially applied to bare areas remaining around the bonding area.
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Electrical Bonding - Typical Bonding Jumper Installation for Aluminum and Magnesium Alloys Figure 201 (Sheet 1)
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Electrical Bonding - Typical Bonding Jumper Installation for Steel and Titanium Alloys Figure 202 (Sheet 1)
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Electrical Bonding - Typical Bonding Joint Installation for Riveted Sheet Metal Construction Figure 203 (Sheet 1)
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Electrical Bonding - Typical Bonding Joint Installation Using Sheet Metal Screws Figure 204 (Sheet 1)
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Electrical Bonding - Typical Bonding Installation on Bracket and/or Angle Construction Figure 205 (Sheet 1)
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Electrical Bonding - Typical Bonding Jumper Installation of Plumbing to Structure Figure 206 (Sheet 1)
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Electrical Bonding - Typical Bonding Jumper Installation for Continuity of Plumbing Figure 207 (Sheet 1)
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Electrical Bonding - Typical Inspection Plate Bonding Figure 208 (Sheet 1)
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Electrical Bonding - Typical Antenna Bonding Figure 209 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Electrical Bonding - Typical Permanent Ground Stud Installation Figure 210 (Sheet 1)
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Electrical Bonding - Cockpit Voice Recorder Removable Ground Stud Installation Figure 211 (Sheet 1)
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Electrical Bonding - Electrical and Avionic Ground Stud Installation Figure 212 (Sheet 1)
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Electrical Bonding - Typical Metallic Tubing Route with a Bulkhead Fitting and a Valve Figure 213 (Sheet 1)
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Electrical Bonding - Typical Static Wick Installation Figure 214 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL HIGH INTENSITY RADIATED FIELDS (HIRF) - INSPECTION/CHECK 1.
2.
3.
4.
General A.
This section includes the procedures to do the high intensity radiated fields (HIRF) inspection/checks. The inspection/checks are necessary to find if the protection is not sufficient, and to restore the it to its original condition.
B.
This inspection/check can be performed any time you think that the system protection is not sufficient.
Access Panels and Doors Bonding A.
Get the maximum allowable resistance value in ohms for the applicable access panel or door. For resistance listings refer to, Electrical Bonding - Maintenance Practices, Table 210.
B.
Use an approved bonding meter to make sure that the access panel or door has the maximum allowable resistance value.
Equipment Bonding A.
Get the maximum allowable resistance value in ohms for the applicable equipment. For resistance listings refer to, Electrical Bonding - Maintenance Practices, Table 210.
B.
Use an approved bonding meter to make sure that the equipment has the maximum allowable resistance value.
Wire Bundle Protection NOTE:
5.
Before you do a wire bundle shield bonding check, examine the shield connections where the testing is to be performed and the points from which the testing is to be performed. Refer to the Model 208 Wiring Diagram Manual.
A.
Get the maximum allowable resistance value in ohms for the applicable wire bundle shield. For resistance listings refer to, Electrical Bonding - Maintenance Practices, Table 210.
B.
Use an approved bonding meter to make sure that the wire bundle shielding has the maximum allowable resistance value.
Visual Check A.
General. (1) In this section, a visual check of an affected system or area will be performed. The checks look for general condition, corrosion, environmental concerns, correct mounting, and wire routing.
B.
Visual Check. (1) This inspection is used following maintenance that disturbs the wiring. The extent of maintenance determines what needs to be inspected. This includes moving, removing or replacing wire bundles, wire ties, clamps, brackets, and wiring feed-thru parts. Mounting trays, mount holders, mounting brackets, and shock mounted components are also inspected.
C.
Wire Bundles Visual Inspection (1) In the area where maintenance was performed, examine the installation of the wiring for corrosion, environmental damage, and general condition. (2) Examine all wire bundles in the general area for mounting security. (3) Examine the wire routing to make sure that the routing has not changed from its initial routing. (4) Make sure that no unnecessary strains are placed on wire segments between the clamps and the wire holders. (5) Examine the external shielding and coverings as necessary to make sure that the shielding is not interrupted and the coverings are in good condition.
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MODEL 208 MAINTENANCE MANUAL (6)
Examine the adjacent access panels and doors for necessary bonding jumpers and/or local grounding studs. (a) Perform bonding and grounding checks as necessary. NOTE:
6.
After maintenance is complete, it is necessary to do an electrical check after the grounding studs are installed.
D.
Data Bus Wiring Visual Inspection. (1) Data bus wiring can be identified by markings or referring to the Model 208 Wiring Diagram Manual to find reference designators of connectors. Data bus wiring is critical for input/output signals to operate correctly. (2) Examine the end terminations where the shielded and twisted shielded wiring has been disconnected because of maintenance requirements. (a) Examine the pins for corrosion. (3) Examine the connector shells for corrosion and to make sure that the connector threads are serviceable. (4) Examine the backshell for corrosion. (a) If the backshell has a ground terminal, examine it for a looseness and condition of the wire to terminal. (5) After the maintenance is complete, make sure that the installation of the data bus wire is the same as before the maintenance. (6) Examine the exterior of the wiring for obvious damage and shielding continuation. Breaks in shielding must be repaired. (7) Examine the clamps, standoffs, and cable guides for correct positioning, correct hardware, and security of installation.
E.
Rack Mounted Components Visual Inspection. (1) If an electrical component that is mounted on a tray or rack is removed for maintenance, examine the component to the rack interface to make sure that the connections are correct. (2) Examine the rack or mounting tray for correct grounding. (a) If a screw is used, make sure that it is tight and that the bond integrity is good. (3) If a bond jumper is installed, make sure that the jumper is in good condition and mounted correctly.
Electrical Wire Bundle Assembly Inspection
CAUTION: Coaxial cable assemblies identified as impedance matching units must be of a specified length. Do not shorten the cable to remove too much slack. NOTE:
Electrical wire bundle assemblies are examined when the individual component or system is examined. When you work in a zone on the airplane and the access panels or floorboards are removed, the wire bundle assemblies in that area must be examined at the same time.
NOTE:
If a component is disturbed in such a way as to disturb the electrical bond to the primary structure, you must do the electrical bonding check again. Refer to Electrical Bonding - Maintenance Practices.
A.
Examine the Wire Bundle Assemblies. (1) Examine the wire bundle assemblies for correct routing and clamp installation. (2) Examine for any signs of chafing or other damage. (3) Make sure that the electrical connectors and the wiring do not show signs of overheating or arcing. (4) Examine the Connector Backshells. (a) If connector backshells are installed, examine them for security and signs of corrosion. (b) Examine the condition of the silicone sealant (RTV-157). (c) If sealant is missing, do the steps that follow: 1 Make sure that the backshell is tight. 2 Clean the outside of the backshell.
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MODEL 208 MAINTENANCE MANUAL 3
Apply a bead of RTV-157 sealant approximately 1/4 inch long on outside of connector, between the connector and backshell. Refer to Electrical Bonding - Maintenance Practices. a Make sure that the sealant does not interfere with disconnecting or connecting the connector. (5) Examine for any crossovers, twists, sharp bends, or kinks. (6) Make sure that the wiring is not attached to or supported by the plumbing lines that contain flammable liquids or oxygen. (a) Wiring that is less than six inches from lines that contain flammable liquids or oxygen must be firmly supported and where possible, routed above the lines. (7) Cables installed in locations where fluid can be trapped must be protected against contamination by the fluid. (8) Where possible, wires must be kept away from high temperature equipment. (9) Make sure that a nylon grommet is installed where wiring passes through cutouts or holes in the airplane structure. (10) Make sure that wire bundles have enough slack to do the following: (a) Allow staking of terminals. (b) Allow ease of maintenance. (c) Allow free movement of shock mounted equipment. (d) Prevent tension and strain on wires and supports. (e) Give sufficient drip and service loops. 1 Make sure that drip loops are arranged so that a drip from the loop does not fall on the equipment. NOTE:
On moisture resistant connectors, drip loops can be omitted where space is limited.
(11) Examine the wires for any sharp bends. (a) The minimum bend radius for wires is 10 times the outside diameter of the wire. (b) At terminal strips where the wire is sufficiently supported, the minimum bend radius is 3 times the outside diameter of the wire. (12) Make sure that the wire groups or bundles are tied at intervals not more than 12 inches. (a) Make sure that the ties are not too tight that it can cut or penetrate the insulation. (b) Make sure that the tying cord is not used to support a wire or wire bundle. (c) Make sure that when the wires or bundles cross, they are tied together at the point of contact. (d) In areas where the operating temperature is 200°F (93°C) or less, the ties can be made with waxed twine. (e) In areas where the operating temperature is above 200°F (93°C), the ties must be made with fiberglass cord. (f) Wire bundle assemblies installed outside of the engine compartment that have QUIK-TY connector clamps can be secured by ties and lacing cords or with plastic tie wraps. (13) To keep the wires from shorting to ground, especially in the engine compartment, make sure that the wiring is not tied to the plumbing, and that standoff clamps are used to correctly secure the wires. (14) Visually examine the cable assembly that routes through the cabin floor for the following: (a) The cable assembly is secured with tie wraps, anchors, and clamps. (b) The cable assembly must be supported and clamped to make sure that it does not chafe the structure or contact the bleed air ducts. (c) Make sure that grommets are installed where the cables routes through the bulkhead. (d) Do a visual examination of the shield terminations for damage or corrosion.
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MODEL 208 MAINTENANCE MANUAL B.
Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields (1) General (a) In this section, a visual check of an affected system or area will be performed. The checks look for signs of a lightning strike, general condition, corrosion, environmental concerns, correct mounting, and wire routing. NOTE:
(b) (c) (d) (e) (f) (g) (h) (i) (j) (k) (l) C.
This includes wire bundles, wire ties, back shells, bonding straps, clamps, brackets and wiring feed-thru parts. Mounting trays, mount holders, mounting brackets and shock mounted components are also examined.
Examine for corrosion, environmental damage, and general condition. Examine for any signs of chafing or other damage. Examine all wire bundles in the general area for mounting security. Examine the external shielding and coverings as necessary to make sure that the shielding is not interrupted and the coverings are in good condition. Examine the clamps, standoffs, and cable guides for correct positioning, correct hardware, and security of installation. If a bond jumper is installed, make sure that the jumper is in good condition and is mounted correctly. Examine the adjacent access panels and doors for necessary bonding jumpers and/or local grounding studs. Examine the backshell for corrosion. 1 If the backshell has a ground terminal, examine it for a looseness and condition of the wire to terminal. Examine the rack or mounting tray for correct grounding. 1 If a screw is used, make sure that it is tight and that the bond integrity is good. Examine the wire ties for security. Examine for signs of lightning strike, look for burnt wires, discolored structure, and exit holes in the structure and/or skin.
External Zonal Visual Inspection of Lightning and High Intensity Radiated Fields (1) General (a) In this section, a visual check of an affected system or area will be performed. The checks look for signs of a lightning strike, general condition, corrosion, environmental concerns, correct mounting and wire routing. NOTE: (b) (c) (d) (e) (f) (g) (h) (i) (j)
If applicable this includes wire bundles, wire ties, back shells, bonding straps, clamps, brackets, and wiring feed-thru parts.
Examine for corrosion, environmental damage, and general condition. Examine all wire bundles in the general area for mounting security. Examine the external shielding and coverings to make sure that the shielding is not interrupted and the coverings are in good condition. Examine the clamps, standoffs, and cable guides for correct positioning, correct hardware, and security of installation. If a bond jumper is installed, make sure that the jumper is in good condition and is mounted correctly. Examine the adjacent access panels and doors for necessary bonding jumpers and/or local grounding studs. Examine the backshell for corrosion. 1 If the backshell has a ground terminal, examine it for a looseness and condition of the wire to terminal. Examine the wire ties for security. Examine for evidence of a lightning strike, burnt wires, discolored structure, and exit holes in the structure and/or skin.
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MODEL 208 MAINTENANCE MANUAL SOLVENTS, SEALANTS AND ADHESIVES - DESCRIPTION AND OPERATION 1.
2.
General A.
Solvents, sealants and adhesives are composed of a group of chemicals that often prove toxic. Anyone engaged in maintenance, repair and operation of airplane and airplane accessories may be exposed to these chemicals.
B.
To help avoid the effects of these toxic substances, work only in a clean, well-lit and well-ventilated area. Rubber gloves and protective clothing should be worn. Avoid breathing spray vapors as they are highly toxic.
C.
When working with toxic substances, always be alert for symptoms of poisoning. If symptoms are observed, immediate removal of the victim from the contaminated area is most important.
Description A.
For clarification, the description of solvents, sealants and adhesives are presented in individual paragraphs. (1) Solvents. (a) Solvents are composed of chemicals which are capable of dissolving other materials and are primarily used as a cleaning agent. Solvent cleaning should be used when it is not practical to clean parts by vapor degreasing or immersion in chemical cleaners. (2) Sealants. (a) Sealants are composed of chemical compounds which are primarily used as a seal against the passage of air and liquids. Classification of sealants are categorized by type according to their application. (3) Adhesives. (a) Adhesives are composed of a mixture of chemicals which make an adherent that is primarily used for bonding like or unlike materials, and are classified according to their application.
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MODEL 208 MAINTENANCE MANUAL GENERAL SOLVENTS/CLEANERS - MAINTENANCE PRACTICES 1.
General A.
2.
Solvents are used in a wide range of cleaning activities and selected solvents can be used in the removal of oil, grease and dirt from objects without harm to metal, plastics or elastomeric parts. Refer to Tools, Equipment and Materials.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items.
NAME
NUMBER
Detergent
MANUFACTURER
USE
Commercially available
General cleaning.
ScotchBrite Pads
Type A
3M Co. 3M Center St. Paul, MN 55101
Light abrasion of metal surfaces.
Sandpaper
320 Grit
Commercially available
Light abrasion of metal surfaces.
Rymple Cloth
Commercially available
Wiping and applying cleaning agents.
Wiping cloth white, oil free, absorbent
Commercially available
Wiping and applying cleaning agents.
3.
Description A.
Solvents exhibit a selective solvent action which permits its use in the removal of oil, grease or dirt. For selection of proper solvent, refer to Table 201. For the cleaning of metal, plastics or rubber, proceed as follows: (1) Metal. NOTE:
Prior to bonding or priming, lightly abrade surface with either a ScotchBrite pad or sandpaper prior to cleaning.
(a) (b)
(2)
Wipe off all excess oil, grease or dirt from surface. Apply solvent to a clean cloth by pouring solvent on the cloth from a safety can or other approved container. The cloth should be well saturated but not to the point of dripping. (c) Wipe the surface with the moistened cloth as required to dissolve or loosen soil. Work on small enough area so the surface being cleaned remains wet. (d) With a clean dry cloth, immediately wipe the surface while the solvent is still wet. Do not allow the surface to evaporate dry. (e) Repeat steps (b) through (d) until there is no discoloration on the drying cloth. Plastic or Rubber. NOTE: (a) (b) (c) (d) (e)
If cleaning a bonding surface, lightly abrade the bonding surface with sandpaper prior to cleaning.
Remove heavy soil from surface by washing with a water detergent solution. Apply solvent to a clean cloth by pouring solvent onto cloth from a safety can or other approved container. The cloth should be well saturated but not to the point where dripping. Wipe the surface with the moistened cloth as required to dissolve or loosen soil. Work on a small enough area so that the surface being clean remains wet. Using a clean dry cloth, immediately wipe the surface while the surface is still wet. Do not allow the surface to evaporate dry. Repeat steps (b) through (d) until there is no discoloration on the drying cloth.
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Table 201. General Solvents CLEANER/ SOLVENT
FEDERAL SPECIFICATION
TYPE CLASSIFICATION
USE/ DESCRIPTION FUNCTION
CAUTION/ WARNING
Dry
MILPRF-680
Type I - 100°F Type II 140°F
General cleaning solvent. Dry cleaning of textile materials. Grease removal.
FLAMMABLE.
Type I Regular Type II - with dauber Type III Aerosol
Spot removing from fabrics. General cleaning solvent. Cleaning of assembled equipment.
USE WITH ADEQUATE VENTILATION. AVOID PROLONGED BREATHING OF VAPOR. AVOID PROLONGED CONTACT WITH SKIN.
1-1-1 ASTM Inhibited D4126 Technical Trichloroethane
Turco Seal Solvent Turco Products
Cleaning/Degreasing metal parts.
Penwalt 2331
Preparing metal plate for painting.
ACID ACTIVATED SOLVENT, DO NOT USE ON PLASTICS. REMOVES PAINT. AVOID CONTACT WITH SKIN.
Carbon Removing Compound
P-C111A
Use in soak tank to facilitate removal of carbon, gum, oil and other surface contaminants except rust or corrosion from engine and other metal parts.
Cleaning Compound
P-C-535
Heavy duty electro cleaner used for removal of soils from ferrous metal surfaces prior to electroplating or other treatments.
Cleaning Compound, UnÞnished Aluminum
MIL-C5410
Type I Viscous Emulsion Type II - Clear Liquid
Used full strength for overhaul of unÞnished aluminum surfaces. Use full strength or diluted with mineral spirits and water for maintenance of airplane unÞnished aluminum surfaces.
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MODEL 208 MAINTENANCE MANUAL Table 201. General Solvents (continued) CLEANER/ SOLVENT
FEDERAL SPECIFICATION
TYPE CLASSIFICATION
Trichloroethylene
ASTM D4080
Type I - Cleaning of metal parts. Regular Degreasing of metal parts. Type II Special purpose solvent. Vapor Degreasing
REMOVES PAINT AND DAMAGES PLASTICS. USE ONLY WITH ADEQUATE VENTILATION. HIGH CONCENTRATIONS OF VAPOR ARE ANESTHETIC AND DANGEROUS TO LIFE. VERY TOXIC.
Polish, Metal Aluminum
MIL-P6888C
Type I - Liquid Type II Paste
Metal polish for use on airplane aluminum surfaces.
FLAMMABLE.
Naphtha, Aliphatic
TT-N958
Type I Type II
For use with organic coatings only. Cleaner for acrylic plastics and may be used in place of Type I General cleaning agent.
DO NOT USE WITH ACRYLIC PLASTICS. FLAMMABLE. VAPOR HARMFUL. AVOID PROLONGED OR REPEATED BREATHING OR CONTACT WITH SKIN.
DeSoclean 110, 020K019
X547506
Paint and adhesive thinner, cleaning agent.
FLAMMABLE.
Isopropyl Alcohol
TT-I-735
For use with organic coatings and as an anti-icing ßuid. General Solvent for synthetic rubbers.
USE DISCRIMINATELY WITH ACRYLIC PLASTICS.
Wax, Airplane, Waterproof Solvent Type
MIL-W18723C
A waterproof wax that can be dissolved or dispersed with an organic solvent.
DO NOT USE SOLVENTS THAT MAY DAMAGE PAINT OR FINISH FOR REMOVAL OF WAX.
Grade B -0.4% water
USE/ DESCRIPTION FUNCTION
CAUTION/ WARNING
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MODEL 208 MAINTENANCE MANUAL Table 201. General Solvents (continued) CLEANER/ SOLVENT
FEDERAL SPECIFICATION
TYPE CLASSIFICATION
USE/ DESCRIPTION FUNCTION
CAUTION/ WARNING
Cleaning Compound, Aluminum
MIL-C5410B
Type I Viscous Emulsion Type II - Clear Liquid
Use full strength for maintenance of unÞnished aluminum surfaces. Use full strength or diluted with mineral spirits and water for maintenance of unÞnished aluminum surfaces.
RUBBER OR SYNTHETIC RUBBER GLOVES AND EYE PROTECTION SHOULD BE USED WHEN HANDLING THE COMPOUND. WASH FROM SKIN IMMEDIATELY WITH WATER OR A SOLUTION OF SODIUM BICARBONATE AND APPLY GLYCERIN OR PETROLEUM JELLY. WASH FROM EYES AS PER MANUFACTURER’S INSTRUCTIONS AND REPORT TO NEAREST MEDICAL FACILITY.
Type II
Cleaner
FLAMMABLE. EYE PROTECTION SHOULD BE USED WHEN HANDLING. USE ONLY WITH ADEQUATE VENTILATION. VAPOR CONCENTRATIONS MAY CAUSE DROWSINESS AND IRRITATION OF EYES OR RESPIRATORY TRACT.
Use as a solvent or thinner for organic coatings, various resins, and chlorinated rubber. Also used to dilute cellulose lacquers and dopes.
FLAMMABLE VAPOR. VAPOR HARMFUL.
Methyl nPropyl Ketone
Toluene
4.
A-A59107D
Safety Precautions A.
Caution should be exercised during cleaning operations. Solvents should be considered ßammable and should not be exposed to ßame or spark under any circumstances. Fresh air masks and/or adequate ventilation are required for all closed areas.
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MODEL 208 MAINTENANCE MANUAL FUEL, WEATHER AND HIGH-TEMPERATURE SEALING - MAINTENANCE PRACTICES 1.
2.
General A.
This section provides instructions and details for various sealing applications. This section is generic to many Cessna products and may cover applications which are not used on the Model 208. Refer to speciÞc maintenance practices to determine sealing applicability.
B.
Sealing is intended to prevent the leakage of liquids, vapors or air pressure through airframe structure. Sealing is required for protection of personnel and equipment.
Tools and Equipment NOTE:
SpeciÞed sealants, cleaning solvents, parting agents, adhesion inhibitors and equipment are listed for use. Suitable substitutes may be used for sealing equipment only.
Table 201. Sealants Type I, Class A-1/2, or A-2 - AMS-S-8802 NAME Sealants
NUMBER
MANUFACTURER
P/S 890 Class A-2
PRC-DeSoto International 5426 San Fernando Rd. Glendale, CA 91209
PR-1440 Class A-1/2 Class A-2
PRC-DeSoto International
AC-236
Advanced Chemistry Technology Garden Grove, CA 92641
USE Fuel, pressure and weather sealant brush application.
Table 202. Sealants Type I, Quick Repair - MIL-S-83318 NAME Sealants
NUMBER PS 860 Class B-1/6
MANUFACTURER PRC-DeSoto International
USE Fuel, pressure and weather sealant. For limited repairs requiring rapid curing sealant.
Table 203. Sealants Type I, Class B-1/2, B-2 or B-4 - AMS-S-8802 NAME Sealants
NUMBER
MANUFACTURER
Pro-Seal 890 Class B-1/2 Class B-2 Class B-4
PRC-DeSoto International
PR-1440 Class B-2 Class B-4
PRC-DeSoto International
AC-236 Class B-2 Class B-4
Advanced Chemistry Technology
USE Fuel, pressure and weather sealant spatula, faying seals application.
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Table 204. Sealants Type I, Class C-20, C-48 or C-80 NAME Sealant
NUMBER
MANUFACTURER
Pro-Seal 890 Class C-20 Class C-48 Class C-80
PRC-DeSoto International
PR-1440 Class C-20
PRC-DeSoto International
USE Fuel, pressure and weather sealant. Suitable for faying surface sealing.
Table 205. Sealants Type II NAME
NUMBER
Sealant
PR1448 Class B-2
MANUFACTURER PRC-DeSoto International
USE Void/hole Þlling compound.
Table 206. Sealant Type III NAME Sealant
NUMBER
MANUFACTURER
PR-810
PRC-DeSoto International
USE High temperature sealing.
Table 207. Sealants Type IV NAME Sealants
NUMBER
MANUFACTURER
Dapco 2100
D. Aircraft Inc. Anaheim, CA 92807
USE Firewall sealing.
Table 208. Sealants Type V NAME
NUMBER
MANUFACTURER
USE
Sealant
RTV106
General Electric Co. Silicone Products Dept. Waterford, NY 12301
Extreme high temperature sealing.
Sealant
RTV162 Class E
General Electric Co. Silicone Products Business Dept.
High temperature and very strong bond and sealant.
Table 209. Sealants Type VI NAME Sealant (Acrylic Latex)
NUMBER
MANUFACTURER
USE
FA-0606 125
H. B. Fuller Company St. Paul, MN 55116
Water and weathertight sealing.
SM8500
Schnee-Moorehead Irving, TX 75017
Water and weathertight sealing.
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Table 210. Sealants Type VII NAME Sealant
NUMBER
MANUFACTURER
Pro-Seal 895
PRC-DeSoto International
USE Aerodynamic smoothing compound.
Table 211. Sealant Type VIII NAME Sealant
NUMBER
MANUFACTURER
PR-1428 Class B-1/2 Class B-2
PRC-DeSoto International
FR-1081 Class B-1/2 Class B-2
Fiber Resin Corporation Burbank, CA 91502
USE Used in areas for access.
Table 212. Sealants Type IX NAME Sealant
NUMBER
MANUFACTURER
Fluorosilicone RTV 730
Dow Corning Corp. Midland, MI 48686
USE Used in areas exposed to fuel.
Table 213. Sealants Type X NAME Sealant
NUMBER
MANUFACTURER
USE
Pro-Seal 870 Class A Type I Class B Type II Class C Type IV
PRC-DeSoto International
Corrosion-inhibitive sealant.
AC-635 Class B Type II Class C Type IV
Advanced Chemistry Technology
Corrosion-inhibitive sealant.
Table 214. Sealants Type XI NAME Sealant Tape
NUMBER
MANUFACTURER
EP-7191 T-0877 (0.062 in. x 0.5 in).
Fiber Resin Corp.
USE Weather-tight window sealant tape.
Table 215. Sealants Type XII NAME Sealant
NUMBER
MANUFACTURER
PR-1829
PRC-DeSoto International
PR-1425
PRC-DeSoto International
USE Windshield and window sealant requires PR-142 adhesion promoter.
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Table 216. Sealants Type XIII NAME Sealant
NUMBER
MANUFACTURER
PR-1776 Class B-1/2 Class B-2
PRC-DeSoto International
USE Low density fuel tank sealant.
Table 217. Cleaning Solvents NAME
NUMBER
MANUFACTURER
Methyl n-Propyl Ketone Naphtha Type III
MIL-PRF-680
Desoclean 110 Isopropyl alcohol
TT-I-735
USE
Commercially Available
Cleaning organic coating.
Commercially Available
Presealing cleaning.
PRC-DeSoto International
Presealing cleaning.
Commercially Available
Cleaning plastic transparencies.
Table 218. Parting Agents NAME
NUMBER
MANUFACTURER
USE
Silicone compound
AS 8660
Commercially available
Prevent sealant sticking.
Petrolatum technical
Federal SpeciÞcation VV-P-236
Commercially available
Prevent sealant sticking.
Table 219. Equipment NAME
NUMBER
Pneumatic sealing gun
Semco No. 250 with accessories (or equivalent)
PRC-DeSoto International
Injection sealing.
Hand-operated sealing gun
Semco No. 850
PRC-DeSoto International
Injection sealing.
PRC-DeSoto International
Application of sealant.
Commercially available
Application of sealant.
MANUFACTURER
Nozzles, Round 1/16 oriÞce
Semco No. 420
Round 1/8 oriÞce
Semco No. 440
Duckbill
Semco No. 8615
Duckbill
Semco No. 8648
Comb
Semco No. 8646
Polyethylene cartridges with plungers and caps for sealant gun
USE
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MODEL 208 MAINTENANCE MANUAL Table 219. Equipment (continued) NAME
NUMBER
MANUFACTURER
USE
Metal spatulas with either stainless steel or glass plates
Commercially available
Mixing sealant.
Plastic lined cups, wax-free with caps
Commercially available
Mixing sealant.
Sealant fairing tools
Commercially available
To fair-in sealant.
Cheesecloth, lint-free
Commercially available
Cleaning.
Plastic scraper, 45-degree cutting edge
Commercially available
Removing old sealant.
Rex Gauge Company, Inc. 3230 West Lake Avenue P.O. Box 46 Glenview, IL 60025
Testing cure of sealant.
Gloves, lightweight lint-free white cotton
Commercially available
Removing old sealant.
Nylon bristle brushes
Commercially available
Removing old sealant.
Pipe cleaners
Commercially available
Cleaning.
Funnel brushes
Commercially available
Cleaning.
Durometer
3.
Rex Model 1500 (or equivalent)
DeÞnition of Sealing Terms A.
The following deÞnitions are included to provide a basic concept of the special terms used in sealing. This list is not all inclusive but the more common terms are listed. (1) Absolute Sealing - There can be no leakage allowed. All openings of any nature through the seal plane are positively sealed. This is the Þrst level of sealing. (All holes, slots, joggles, fasteners and seams must be sealed.) (2) Accelerator (Activator) - Curing agent for sealants. (3) Application Time - The length of time sealant remains workable or suitable for application to structure by brush, extrusion gun, spatula or roller. (4) Base Compound - The major component of a two-part sealing compound which is mixed with the accelerator prior to application to produce a fuel, temperature, pressure, weather and/or Þrewall sealing material. (5) Brush Coat - Apply an over coating or continuous Þlm of appropriate sealing compound by use of a brush. (6) Electrical Seal Fitting - A device used for sealing electrical wires which pass through bulkheads, etc. Not to be used through the integral fuel tank wall. (7) Fay Seal or Faying Surface Seal - A seal barrier created by the sandwiching of sealant between mating surfaces of structure. Special attention must be taken to avoid metal chips or dirt at the faying surface. (8) Fillet Seal - Applying a bead of sealant to a seam, joint or fastener after the assembly has all permanent fasteners installed. (9) Hole - An opening that has no appreciable depth, such as a tool hole. Holes that penetrate the seal plane must be metal Þlled with a fastener, gusset or patch. (10) Injection Seal - Filling of channels by forcing sealant into a void or cavity after assembly. (11) Integral Tank - Composition of structure and sealant material which forms a tank that is capable of containing fuel without a bladder.
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MODEL 208 MAINTENANCE MANUAL (12) Intermediate Seal - The second level of sealing. All holes, slots, joggles and seams in the seal plane must be sealed. A minor amount of leakage is tolerable and permanent fasteners are not required to be sealed. (13) Post-assembly Seal - A seal that is applied after the structure is assembled. (Fillet and injection seals.) (14) Pre-assembly Seal - Sealant material that must be applied during or prior to the assembly of the structure. (Faying surface and prepack seals.) (15) Prepack Seal - A pre-assembly seal used to Þll voids and cavities; can be a primary seal used to provide seal continuity when used in conjunction with a Þllet seal. It can be used as a backup seal to support a Þllet across a void. Fill the entire cavity to be pre-packed. Usage as a primary seal should be kept to a minimum. (16) Primary Seal - Sealant material that prevents leakage and forms a continuous seal plane. This seal is in direct contact with the fuel, vapor, air and acid. With few exceptions, it is in the form of a Þllet seal. (17) Sealant - A compound applied to form a seal barrier. (18) Seal Plane - A surface composed of structure, sealant and fasteners on which the continuity of seal is established. (19) Shank Sealing - Sealant compound shall be applied to the hole or to both the shank and the under head area of the fastener in sufÞcient quantity that the entire shank is coated and a small continuous bead of sealant is extruded out around the complete periphery of each end of the fastener when installed. The fastener shall be installed within the application time of the sealing compound used. (20) Squeeze-Out Life - Length of time sealant remains suitable for structure assembly in faying surface seal application. (21) Tack-Free Time - Tack-free time is a stage, during the cure of the sealant compound, after which the sealant compound is no longer tacky. When the sealant compound is pressed Þrmly with the knuckles, but no longer adheres to the knuckles, the sealant compound is tack-free. 4.
Materials A.
Type of Sealants - Sealants are categorized by type of usage. Type I sealants are separated by class to differentiate the material to use by method of application. Dash numbers following the class designation indicate the minimum application time (in hours) for Class A and Class B and minimum work life (in hours) for Class C. Reference Table 220 for application time and curing rate for Type I sealants. (1) Type I - Fuel, pressure and weather sealant. (a) Class A - Sealant which is suitable for brush application. (b) Class B - Sealant which is suitable for application by extrusion gun, spatula, etc. (c) Class C - Sealant which is suitable in faying surface applications. (d) Quick Repair Sealant - This material is for use only in making repairs when an extremely rapid curing sealant is required. A possible application includes sealing a leaking fuel tank on an airplane which must be dispatched within a few hours.
Table 220. Curing Properties of Type I Sealant TACK-FREE TIME (HOURS, MAXIMUM)
CURING RATE (HOURS, MAXIMUM)
1/2
10
40
A-2
2
40
72
B-1/2
1/2
4
6
B-2
2
40
72
B-4
4
48
90
CLASS
APPLICATION TIME (HOURS, MINIMUM)
A-1/2
WORK LIFE (HOURS, MINIMUM)
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MODEL 208 MAINTENANCE MANUAL Table 220. Curing Properties of Type I Sealant (continued) CLASS
APPLICATION TIME (HOURS, MINIMUM)
WORK LIFE (HOURS, MINIMUM)
TACK-FREE TIME (HOURS, MAXIMUM)
CURING RATE (HOURS, MAXIMUM)
C-24
8
24
96
168 (7 days)
C-48
8
48
110
336 (14 days)
C-80
8
80
120
504 (21 days)
NOTE 1: Time periods are based on a temperature of 77°F and 50 percent relative humidity. Any increase in either temperature or relative humidity may shorten these time periods and accelerate the sealant cure. (2)
Type II - Hole Þlling compound. This material is for holes and slots that cannot be Þlled with one application of Type I; Class B sealant. Type II sealant shall not be used for the sealing of an integral fuel tank. (3) Type III - High-temperature sealant. This material is for use where exposure to fuel is moderate and for intermittent exposures up to 450°F, but is not suitable for pressure sealing. (4) Type IV - Firewall sealant. This material is for use when exposure to fuel is minimal and for intermittent temperature exposures up to 500°F, but is not suitable for pressure sealing. (5) Type V - Extreme high-temperature sealant. This material is for use where exposure to fuel is minimal and for intermittent exposures up to 600°F, and is also suitable for pressure sealing. (6) Type VI - Watertight and weather tight sealant. This material is for use where there is no exposure to fuel, high temperature or pressure. (7) Type VII - Aerodynamic Smoothing Compound. This material is used for Þlling skin gaps to obtain a smooth aerodynamic surface. (8) Type VIII - Low Adhesion Access Sealant. This Class B material is designed for sealing faying surfaces where easy separation of the joined surfaces is required. The sealant has low adhesion and forms a gasket that molds itself to Þll all irregularities between two surfaces. It is exceptionally resistant to fuels, greases, water, most solvents and oils, including red hydraulic oil. (9) Type IX Fluorosilicone RTV Sealant. This sealant is a room temperature vulcanizing sealant that will withstand fuel. (10) Type X - Corrosion Inhibitive Sealant. These materials are 2-part, room temperature curing, synthetic rubber compounds used in the sealing and coating of metal components for protection against corrosion. NOTE:
Type X may be used in all applications where Type I is used except that it shall not be used for fuel tank sealing.
(11) Type XI - Sealant Tape. These materials are permanently pliable and can be used to set windshields before sealing or to seal covers. (12) Type XII - Windshield and Window Sealant. These materials are 2-part, room temperature curing synthetic rubber compounds used to seal glass, polycarbonate, or acrylic transparencies. (13) Type XIII - Low-Density (1.35 sp gr max) Fuel Tank Sealant. This material is manganese dioxide cured, for applications at service temperatures of -65° to 250°F (-54 to 121°C).
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CAUTION: Quick repair sealant must be applied within its working life of 15 minutes. Attempts to work quick repair sealant beyond working life will result in incomplete wetting of surface and will result in a failed seal. 5.
General Requirements A.
When working with sealants, observe the following requirements. (1) Unmixed sealants shall not be more than two months old when received. These sealants shall not be more than six months old when used. (2) Unmixed sealants stored at temperatures exceeding 80°F shall be used within Þve weeks. (3) Sealants which have been pre-mixed, degassed and ßash frozen shall be maintained at -40°F or lower and shall not be received more than two weeks beyond the date of mixing. These sealants shall not be used more than six weeks after the date of mixing. (4) Frozen sealant shall be thawed before being used. If sealant were applied at a temperature below 60°F, it would not be sufÞciently pliable for proper application and adhesion could be critically reduced by condensation of moisture. On the other hand, although sealant must extrude freely for proper application, it would be subject to excessive slumping if applied at a temperature above 80°F. Frozen sealant may be thawed by any suitable means which does not cause contamination or overheating of the sealant and does not shorten the application time of the sealant to an impractical period. Examples: Thawing by exposure to ambient air temperature, accelerated thawing by exposure in a constant temperature bath (using clean, hot water), accelerated thawing in a microwave oven. In any case, thawing temperature and time shall be adjusted to give a thawed sealant temperature between 60°F and 80°F at the time the sealant is applied. (5) Mixed, frozen sealants which have thawed shall not be refrozen. (6) Complete pre-assembly operations, such as Þtting, Þling, drilling, countersinking, dimpling and deburring, prior to cleaning and sealant application. (7) Surfaces must be clean and dry, free from dust, lint, grease, chips, oil, condensation or other moisture, and all other contaminating substances prior to the application of sealant. (8) Naphtha Type II or Isopropyl Alcohol (TT-I-735) are the only cleaners which may be used on plastic transparencies. (9) Sealant materials may be applied to unprimed or primed surfaces. Nonchromated or epoxy primers shall have good adhesion to the substrate material and shall have aged at least 48 hours prior to sealant application. (10) Sealants shall not be applied when the temperature of either the sealant or the structure is below 60°F. (11) The sealants Pro-Seal 890 B-1/2, B-2 or B-4 are the only sealants which may be used on plastic transparencies. (12) Sealant applied by the Þllet or brush coat methods shall always be applied to the pressure side of a joint if possible. (13) After application, sealants shall be free of entrapped air bubbles and shall not exhibit poor adhesion. All Þllets shall be smoothed down and pressed into the seam or joint with a Þlleting tool before sealant application time has expired. (14) Where fasteners have been sunk or under-head sealed, extruded sealant shall be evident around the complete periphery of the fastener to indicate adequate sealing. Sealant extruded through a hole by a rivet shall be wiped from the end of the rivet before bucking. Threaded fasteners which have been shank or under-head sealed shall not be retorqued after expiration of the application time of sealant. In torquing, turn the nut rather than the bolt if possible. (15) Pressure testing shall not be accomplished until the sealant is cured. (16) Sealant shall not be applied over ink, pencil or wax pencil marks. If these materials extend into the sealing area, they must be removed. (17) If sealing is to be accomplished over primer and the primer is removed during the cleaning process, it is permissible to seal directly over the cleaned area and then touch up the exposed areas after the sealant has been applied and is tack free. (18) Sealed structure shall not be handled or moved until sealant is tack free (sealant may be dislodged or have the adhesion damaged). Excessive vibration of structure, such as riveting and engine run up, is not permitted.
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MODEL 208 MAINTENANCE MANUAL (19) Drilling holes and installing fasteners through a fay sealed area shall be performed during the working life of faying sealant, or the entire shank and area under fastener head shall be fay sealed. 6.
7.
Sealant Curing A.
Room Temperature. (1) Room temperature curing properties are based on a temperature of 75°F, +5° or -5°F, and a relative humidity of 50 percent. Curing times of two-part sealants will shorten with increased temperature and/or relative humidities. (2) Room temperature curing properties of Type I sealants are given in Table 201. (3) Room temperature curing properties of Type II sealant are: Application Time 2 Hours (Minimum); Tack- Free Time 20 Hours (Maximum); Curing Rate 40 Hours (Maximum). (4) Room temperature curing properties of Type III sealant are dependent on solvent release. Type III sealant should cure for a minimum of 17 days at room temperature before being subjected to temperatures as high as 400°F. (5) Room temperature curing properties of Type IV sealant are: Application Time 1-1/2 Hours (Minimum); Tack-Free Time 24 Hours (Maximum); Curing Rate 48 Hours (Maximum). Type IV sealant should cure for a minimum of 72 hours at room temperature before being subjected to temperatures as high as 400°F. (6) Room temperature curing properties of Type V sealant are: Tack-Free Time 1/2 Hour (Maximum); Curing Rate 24 Hours (Maximum). Type V sealant should cure for a minimum of 48 hours at room temperature before being subjected to temperatures as high as 400°F. (7) Room temperature curing properties of Type VI sealant are: Tack-Free Time 2 Hours (Maximum); Curing Rate 16 Hours (Maximum). (8) Room temperature curing properties of Type VII sealant are: Class B-1/2 Application Time 1/2 Hour; Tack-Free Time 10 Hours; Cure Time 24/35R Hours/Hardness. Class B-2 Application Time 2 Hours; Tack-Free Time 24 Hours; Cure Time 48/35R Hours/Hardness. (9) Curing properties of Type VIII, Class B sealants are the same as for Type I, Class B. Adhesion to aluminum should be (peel) less than 2 pounds/inch width.
B.
Accelerated Curing. (1) Accelerated curing of sealant can be accomplished in several ways. The procedure to be used is dependent on the type of sealant and other factors. (2) The cure of Type I and Type II sealants can be accelerated by an increase in temperature and/ or relative humidity. Warm circulating air at a temperature not to exceed 120°F may be used to accelerate cure. Heat lamps may be used if the surface temperature of the sealant does not exceed 140°F. At temperatures above 120°F, the relative humidity will normally be so low (below 40 percent) that sealant curing will be retarded. If necessary, the relative humidity may be increased by the use of water containing less than 100 parts per million total solids and less than 10 parts per million chlorides. (3) The cure of Type III sealants can be accelerated after Þrst curing for a minimum of 72 hours at room temperature by heating for 8 hours with warm circulating air or heat lamps in such a manner that the surface temperature of the sealant does not exceed 120°F. (Lowered relative humidity is helpful.) Curing should be completed before the sealant is subjected to temperatures as high as 400°F. (4) The cure of Type IV sealants can be accelerated by reducing the relative humidity. However, the sealants should be cured for a minimum of 72 hours at room temperature before being subjected to temperatures as high as 400°F. (5) The cure of Type V and Type VI sealants can be accelerated by the same procedures given for Type I or Type II sealants.
Mixing of Sealants A.
Requirements. (1) Sealants shall be mixed or thinned in accordance with the manufacturer's recommendations and thoroughly blended prior to application. All mixed sealant shall be as void free as possible.
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MODEL 208 MAINTENANCE MANUAL (2)
Prior to mixing, the sealing compound base and its curing agent, both in their respective original unopened containers, shall be brought to a temperature between 75°F and 90°F. All required mixing equipment should also be brought to a temperature between 75°F and 90°F.
B.
Hand Mixing of Sealant. (1) The correct amount of base and curing agent, per manufacturer’s instructions, shall be weighed in a clean, wax-free container immediately prior to mixing. An alternate method is to mix the sealant on a ßat plate with a spatula. The scales and weighing process must be controlled within +2 or -2 percent to ensure good quality. (2) Do not allow the accelerator to come in contact with the sides of the container. (3) Materials shall be accurately weighed on scales that are calibrated and maintained for required accuracy. (4) Mix the components until the color is uniform taking care not to trap air in the sealant. (5) Transfer the sealant to another clean container and complete the mix.
C.
Mixing Two-Part Sealant Cartridges. Refer to Figure 201.
WARNING: The cartridge should be held Þrmly, but must not be squeezed, as the dasher blades may penetrate the cartridge and injure the hand. (1) (2)
Pull dasher rod to the FULL OUT position, so the dasher is at the nozzle end of cartridge. Insert ramrod in the center of dasher rod against the piston and push the piston in approximately 1 inch. NOTE:
(3)
Move the dasher rod in approximately 1 inch, then push piston in another inch. Repeat this action until accelerator is distributed along the entire length of the cartridge. NOTE:
(4)
The accelerator has been fully injected into the cartridge when the ramrod is fully inserted into the dasher rod.
Remove and properly discard the ramrod. NOTE:
(5)
Extra force will be needed on the ramrod at the beginning of accelerator injection into the base material.
Mixing the accelerator and base material can be accomplished manually, or as an alternate method, with the use of a drill motor.
Manual Mixing. (a) Begin mixing operation by rotating the dasher rod in a clockwise direction while slowly moving it to the FULL OUT position. NOTE: (b)
Do not rotate the dasher rod counterclockwise; the four-blade dasher inside the cartridge will unscrew and separate from the dasher rod.
Continue clockwise rotation and slowly move the dasher rod to the FULL IN position. 1 A minimum of Þve full clockwise revolutions must be made for each full-out stroke and for each full-in stroke of the dasher rod. Approximately sixty strokes are necessary for a complete mix. NOTE:
(c) (d)
If streaks are present in the sealant (viewing through the side of the cartridge), the sealant is not completely mixed.
End mixing operation with the four-blade dasher at the bottom of the cartridge. Hold cartridge upright; unscrew dasher rod from the four-blade dasher by gripping the cartridge at the four-blade dasher and turn the dasher rod counterclockwise. Remove dasher rod.
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Two-Part Sealant Cartridge Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (e)
(6)
Screw appropriate nozzle into the cartridge. If sealant gun is to be used, install cartridge in gun. Drill Motor Mixing. NOTE: (a)
A tapered rotary Þle or a 25/64 inch drill bit may be used with a drill motor to turn the dasher rod.
Insert the rotary Þle/drill bit into dasher rod approximately 0.5 inch.
WARNING: The cartridge should be held Þrmly, but not squeezed, as the dasher blades may penetrate the cartridge and injure the hand. (b) (c)
Verify the drill motor will rotate the dasher rod clockwise (looking toward the nozzle end of the cartridge). With the cartridge held Þrmly in one hand and the drill motor in the other, rotate the dasher rod at approximately 50 revolutions-per-minute while moving the dasher rod to FULL IN and FULL OUT positions. 1 Mix sealant for at least 50 strokes (a stroke is one complete full-in and full-out stroke of the dasher rod). NOTE:
(d) (e) (f) 8.
If streaks are present in the sealant (viewing through the side of the cartridge), the sealant is not completely mixed.
End mixing operation with the four-blade dasher at the bottom of the cartridge. Hold cartridge upright; remove drill motor and rotary Þle/drill bit from the dasher rod; unscrew dasher rod from the four-blade dasher by gripping the cartridge at the four-blade dasher and turn the dasher rod counterclockwise. Remove dasher rod. Screw appropriate nozzle into the cartridge. If sealant gun is to be used, install cartridge in gun.
Cleaning A.
All surfaces to which sealant is to be applied shall be clean and dry.
B.
Remove all dust, lint, chips and shavings with a vacuum cleaner where necessary.
C.
Cleaning shall be accomplished by scrubbing the surface with clean cheesecloth moistened with solvent. The cloth shall not be saturated to the point where dripping will occur. For channels and joggles, pipe cleaners and/or funnel brushes may be used instead of cheesecloth. (1) Scotch Brite pads should be used to clean all nutplates (except domed nutplates) and all exposed bonding primer on all bonded assemblies. (2) The solvents to be used on all surfaces to be sealed, except the integral fuel tank and on plastic transparencies, shall be A-A-59281, cleaning compound, ASTM D4126, 1, 1, 1 - Trichloroethane, Technical, Inhibited. (3) The solvents to be used for the cleaning in the integral fuel tank are A-A-59281 for the Þrst or preliminary cleaning. For the Þnal cleaning, ASTM D4126 only must be used. (4) The only solvent to be used on plastic transparencies shall be TT-I-735, isopropyl alcohol.
D.
The cleaning solvent should never be poured or sprayed on the structure.
E.
The cleaning solvent shall be wiped from the surfaces before evaporation using a piece of clean, dry cheesecloth so oils, grease, wax etc. will not be redeposited.
F.
It is essential that only clean cheesecloth and clean solvent be used in the cleaning operations. Solvents shall be kept in safety containers and shall be poured on the cheesecloth. The cheesecloth shall not be dipped in the solvent containers and contaminated solvents shall not be returned to the clean solvent containers.
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9.
G.
Final cleaning shall be accomplished immediately prior to sealant application by the person who is going to apply the sealant. (1) The area which is to be sealed shall be thoroughly cleaned. A small clean paint brush may be needed to clean corners, gaps, etc. Always clean an area larger than the area where the sealant is to be applied. Never clean an area larger than 30 inches (0.76 meter) in length when practical. When the area is being scrubbed with a moistened cloth in one hand, another clean dry cloth shall be held in the other hand and shall be used to dry the structure. The solvent must be wiped from the surfaces before it evaporates. (2) The above procedure shall be repeated until there is no discoloration on the clean drying cloth. Marks resulting from wax or grease pencils must be removed from parts prior to sealing.
H.
Allow all cleaned surfaces to dry a minimum of 5 minutes before application of sealant materials.
I.
Sealant shall be applied as soon as possible after cleaning and drying the surfaces to be sealed. Do not handle the parts between the cleaning and sealing operations. Sealant application personnel handling cleaned surfaces shall wear clean white gloves to prevent surface contamination. In the event contamination does occur, the surfaces shall be recleaned.
J.
Safety precautions should be observed during the cleaning and sealing operation. Cleaning solvents are toxic and ßammable in most cases. Fresh air masks and/or adequate ventilation are required for all closed areas. The structure shall be electrically grounded before starting any cleaning or sealing operation.
Sealing Application A.
General. (1) All new sealing shall be accomplished using the type of sealing material required for the area being sealed. All sealant repairs shall be accomplished using the same type of sealing material as that being repaired. (2) Application time of the sealing compound shall be strictly observed. Material which becomes too stiff and difÞcult to work or which does not wet the surface properly shall be discarded even though the application time has not expired. (3) Prior to sealant application, all surfaces to be sealed shall be cleaned per Cleaning.
B.
Faying Surface Sealing - The application of a faying surface seal shall be made only when new structure is being added to the airplane and requires a faying surface seal or when the structure and/or parts have been disassembled for reasons other than a faulty seal. (1) Immediately prior to Þnal closure of the joint, sealant shall be applied to one mating surface of the joint with a sealant gun, spatula, roller or other suitable tool. SufÞcient sealant shall be applied so the space between the assembled faying surfaces is completely Þlled with sealant and a small excess is squeezed out in a continuous bead around the periphery of the joint when the joint is secured. Refer to Figure 202. (2) Place parts in assembly position and install the fasteners within the application time of the faying surface sealant. When assembly with permanent type fasteners is not feasible, temporary fasteners (clecos or bolts) may be used, but when the temporary fasteners are used, they must be replaced by permanent type fasteners prior to the expiration of the work life of the faying surface sealant. Removal of each individual temporary fastener shall be followed immediately by the installation of the permanent fastener. (3) When a Þllet seal is required around the periphery of a fay sealed joint, it is not necessary to remove the sealant squeeze-out where the Þllet is to be applied, provided the material which was squeezed out has been shaped into a small Þllet conÞguration prior to expiration of the application time. When the squeeze-out has been shaped, a Þnal or full bodied seal can be applied over the shaped squeeze-out without waiting for the squeeze-out to cure. If the squeezed out material was not shaped before expiration of its application time, it shall be cured to a tack-free condition and then removed, by use of a plastic tool, from locations where a Þllet is to be applied. (4) Immediately after assembly is completed and all permanent type fasteners have been installed, remove uncured sealant, which extrudes on the exterior of the airplane, using clean rags moistened with A-A-59107, Toluene.
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Fay Sealing Figure 202 (Sheet 1)
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Fay Sealing Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL C.
Injection Sealing (1) Sealant shall be injected in the channel, joggle, void or cavity from one point only, using a sealant gun. No air shall be entrapped, the channel, joggle, void or cavity shall be completely Þlled, and sealant shall emerge from the prescribed opening. Refer to Figure 203. If multiple exits or channels exist, block each channel exit after it is Þlled, without stopping the injection, so that sealant extrudes into all necessary channels. (2) Remove excess sealant before expiration of its application time, and, using a suitable tool, smooth ßush with the surface.
D.
Fillet Sealing. (1) Fastener considerations: (a) Do not Þllet seal any parts until they are held completely together by permanent fasteners. (b) Prior to Þlleting the periphery of bolted structure and Þttings, it is necessary that all bolts, accomplishing the attachment, be properly torqued. (2) The sealant shall be applied using a sealant gun or spatula. (3) When using a sealant gun for Þllet sealing, the nozzle tip shall be pointed in the seam or joint and shall be maintained nearly perpendicular to the line of travel. A continuous bead of sealant shall precede the tip and the tip size, shape and rate of travel shall be such that sufÞcient sealant is applied to produce the required Þllet. (4) Fillets shall be shaped or formed to meet the size and shape requirements as shown in applicable Þgures using the nozzle tip and/or fairing tools to press against the sealant while moving parallel to the bead. Exercise caution to prevent folds and entrapment of air during application and shaping of the Þllet and work out any visible air bubbles. The Þllet shall be formed so the highest portion of the Þllet is centered over the edge of the structure or Þtting. Lubrication in any form shall not be used for smoothing purposes. In all cases, Þllet size shall be kept as near minimum as practical. (5) Where it is more convenient or Þllet slumping is encountered, the Þllet may be applied in two stages. A small Þrst Þllet shall be applied and allowed to cure to a tack-free state, and then followed by a second application of sealant sufÞcient to form the Þnal Þllet conforming to the speciÞed dimensions for a Þllet seal. If the Þrst Þllet has cured, it must be cleaned before the second application of sealant is made. If the Þllet has only cured to a tack-free state, it shall be wiped lightly with a gauze pad or cheesecloth pad dampened with cleaning solvent. (6) Allow the sealant to cure to a tack-free condition prior to the airplane being moved, handled and/or worked on. (7) In cases where a Þllet seal connects to an injection seal, the full bodied Þllet shall extend past the end of the injection and then taper out. (8) Seal lap joints and seam Þllets. Refer to Figure 204. (9) Seal butt joint Þllets. Refer to Figure 205. (10) Fillet seal bolts. Refer to Figure 206. The area for sealing shall consist of the area of the structure surrounding the base of the fastener end, plus the entire exposed area of the fastener. An optional method of sealing threaded fasteners is to apply a brush coat of Type I, Class A sealant. Where brush coating is used as the method of sealing threaded fasteners, the sealant must be worked around each fastener with a stiff brush and considerable care for effective sealing. A simple pass of the brush with the sealant is not sufÞcient to produce an effective seal. (11) Fillet seal dome-type nutplates. Refer to Figure 207. The area for sealing shall consist of the area of the structure surrounding the base of the fastener and from there up over the rivets to the dome. (12) Rivetless, self-sealing nutplates requiring sealing for lightning protection should be brush coated over the entire surface and mating structure. (13) Fill holes and Þllet seal slots. Refer to Figure 208. NOTE: (a)
A hole or slot through the wall of an integral fuel tank must not be sealed by this method.
Holes and slots that are too large to be Þlled with one application of Type I, Class B sealant shall be Þlled with Type II sealant. Large holes or slots may be backed with masking tape to prevent excessive extrusion of sealant through the holes or slots, but the masking tape shall be removed after the sealant has cured to a tack-free condition.
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Injection Sealing Figure 203 (Sheet 1)
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Lap Joint and Seam Fillets Figure 204 (Sheet 1)
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Butt Joint Fillets Figure 205 (Sheet 1)
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Bolt Head, Nut and Thread Sealing Figure 206 (Sheet 1)
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Dome Type Fillets Figure 207 (Sheet 1)
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Slot, Hole and Mismatch Sealing Figure 208 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (b) (c) E.
In all locations where Type II sealant has been applied, after the Type II sealant has cured to a tack-free condition it shall be brush coated with Type I, Class A sealant. The brush coat shall overlap the edge of the Type II sealant sufÞciently to ensure complete coverage. Tooling holes shall be plugged with a shank sealed soft rivet and then brush coated with Type I, Class A sealant. Refer to Figure 209.
Firewall Sealing - The engine Þrewall shall be sealed to an intermediate level of sealing using Type IV sealant. (1) Clean the areas to be sealed per Cleaning. (2) Mix, by weight, 1 part of curing agent with 100 parts of Type IV (Coast Pro-Seal #700) sealant. NOTE: (3) (4)
10.
Sealant should be mixed by weight. It is important that accelerator be completely and uniformly dispersed throughout the base compound.
Using a spatula and fairing tool, apply a Þllet of sealer along all cracks, seams, joints and also over all fasteners in the Þrewall. Type IV sealant shall be cured for a minimum of 72 hours at room temperature before being subjected to temperatures of 400°F.
Sealant Repair A.
B.
Materials - Repairs, in general, shall be accomplished with the same type of material as that being repaired. NOTE:
Type I, Class B-1/2 is recommended for use during cold weather to obtain an accelerated cure.
NOTE:
Type I, Quick Repair sealant may be used as a repair for sealant in pressure vessels and fuel tanks if desired for fast cure and rapid dispatch.
Temperature Requirements. (1) The structure shall be above 60°F before the sealant is applied and shall remain above 60°F until the sealant is tack- free. NOTE:
(2) C.
For outside operations only, the temperature of the structure may be allowed to drop below 60°F but not below 58°F after application and for a period of time not to exceed 48 hours; however, the structure must be subsequently heated to above 60°F and the sealant allowed to become tack-free before the tanks are refueled.
The maximum air temperature allowed to come in contact with the curing sealant is 120°F.
Fillet and Fastener Sealing Repairs. (1) Repair of damaged or faulty sealant applications shall be accomplished as follows: (a) Remove all damaged or faulty sealant to ensure solid residual material. (b) Sealant shall be cut so as to produce a smooth continuous scarfed face. Refer to Figure 210. The sealant shall be completely removed in the affected areas. The cutting tools should only be made from nonmetallic materials that are softer than aluminum. (c) Inspect repair areas for clean and smooth cuts. Loose chunks or ßaps of sealant on the cut areas shall be removed. (d) Clean the area to be sealed, including the scarfed face of the old seal, per Cleaning. (e) Apply new Þllet seals per Sealing Application, Fillet Sealing. Slight overlapping of the fresh material over the existing Þllet is permissible. A large buildup of sealant shall not be allowed. Type VI sealant may be used over Type I, II and III sealant except in the integral fuel tank sealing. Type VI will cure more rapidly for weather and pressurization repairs. (f) Rework of a Þllet which has been over sprayed or brushed with primer shall be accomplished by a scarfed joint and removal of the Þllet having primer on it, in the area of the repair. The primer shall not be sandwiched in between the old and new sealants.
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Shank Sealing Figure 209 (Sheet 1)
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Cutaway View of Sealing Bead Figure 210 (Sheet 1)
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Cutaway View of Sealing Bead Figure 210 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (g)
11.
If the primer is removed during the cleaning operation, it is permissible to apply the new Þllet seal directly over the clean bare metal and then touch up all exposed areas of bare metal with the proper primer after the sealant has been applied.
D.
Faying Surface Sealing Repair - After determining the area which contains the faulty and/or leaking faying surface seal, the repair shall be accomplished by applying a Þllet seal along the edge of the part adjacent to the faying surface seal long enough to fully cover the area of the faulty and/or leaking seal.
E.
Brush Coat Sealing Repair - Repair of damaged or leaking brush coat seals shall be accomplished by removing the discrepant brush coat. Clean the area of sealant removal and the surrounding structure and sealant per Cleaning. Apply a new brush coat of sealant.
Integral Fuel Tank Sealing NOTE: A.
For complete fuel tank sealing procedures, refer to Chapter 28, Fuel Tank Sealing - Maintenance Practices.
Integral wing fuel tank sealing is a reÞnement of fuel sealing process. With an integral fuel tank, the fuel is conÞned in a sealed cavity in the wing structure. (1) All damaged or leak areas must be completely and carefully repaired. (2) Cleaning shall be performed with a clean cheesecloth dampened with solvent. Brush or pipe cleaners may be used to clean corners, gaps, joggles and channels. (3) After application, the sealant must be free of entrapped air bubbles. (4) All Þllets are to be smoothed down and pressed into the seam or joint with a Þlleting tool. (5) The sealant shall be tack-free and additional 50 percent of normal cure time shall be allowed prior to refueling. (6) Before pressure testing, the sealant must be cured.
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MODEL 208 MAINTENANCE MANUAL ADHESIVE AND SOLVENT BONDING - MAINTENANCE PRACTICES 1.
2.
3.
General A.
This section describes the application of adhesives and solvent bonding.
B.
A protective coating is applied to particular areas like exterior placards to protect the placard from hydraulic fluid and weather elements.
Safety A.
Cleaning and bonding operations should be performed in well ventilated areas away from sparks or flames.
B.
Cleaning solvents should be dispensed from approved containers. Solvent wetted cheesecloth should be disposed of in special safety containers provided solely for this purpose.
C.
Rubber gloves should be worn when practical, and hands should be washed prior to eating or smoking after handling solvents and adhesives.
Clear Polyurethane Topcoat A.
Mix the clear polyurethane C63C with the AA-92-C-39 catalyst according to manufacturer's directions. Apply the clear polyurethane coating in three uniform, 50 percent overlay spray coats to an approximate thickness of 2 1/2 to 3 mils dry film thickness. Air dry 4 to 6 hours or force dry at approximately 135°F for 1 hour. NOTE:
4.
All equipment should be cleaned immediately after use with T732A thinner.
Material Classification A.
Type I, Epoxy Base Adhesive. (1) Used for bonding metal to metal, fiberglass, wood and thermoplastics.
B.
Type II, Oil Resistant, Synthetic Rubber Base Adhesive. (1) Used for bonding fabric, leather, rubber, insulation batting, metals and ABS thermoplastics.
C.
Type III, Fuel Resistant, Synthetic Rubber Base Adhesive. (1) Used for bonding cork, leather and rubber gaskets to metals where there may be some exposure to fuel; also for rubber, wood, glass, vinyls and some plastics.
D.
Type IV, Synthetic Resin Base Adhesives. (1) Used for bonding vinyl materials to themselves or metals, glass, plastics and wood.
E.
Type V, Silicone Rubber Base Adhesives. (1) Used for bonding metals, plastics, glass, ceramic and rubber insulation.
F.
Type VI, Solvent Cementing. (1) Used for cementing thermoplastics to themselves. Solvents should be either C.P., U.S.P. or Reagent Grade. Heat and pressure may be used as an alternate method.
G.
Type VII, Cyanoacrylate Base Adhesive. (1) Quick setting adhesive used for plastics, metals and rubber (not waterproof).
H.
Type VIII, Pressure Sensitive Adhesive. (1) Used for quick mounting of small parts of metal, plastic, glass, wood or fabric.
I.
Type IX, Polyurethane Base Adhesive. (1) Used to bond plastics to themselves or other plastics.
J.
Type X, Acrylic Plastic Base Adhesive. (1) Used for bonding acrylic plastics to themselves, other plastics or metals.
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5.
Material
Table 201. Adhesives NAME
NUMBER
MANUFACTURER
USE
Type/Class IA
EA-9304.1 EA-9346.5
The Dexter Corporation Aerospace Material Division Ft. Lauderdale, FL
Adhesive
Type/Class IB
EA-9309.3NA EA-9314NA EA-9330.3 EA-9339
The Dexter Corporation Aerospace Material Division
Adhesive
EPK-9340
The Dexter Corp. Engineering Adhesives Division Seabrook, NH
Adhesive
Type/Class IC
EA-9394 EA-9396
The Dexter Corporation Aerospace Material Division
Adhesive
Type/Class ID
A-1186-B
SIA Adhesives Inc. 123 West Bartges St. Akron, OH 44311
Adhesive
Type/Class IE
Fastweld 10
Ciba-Geigy Corp. Furane Aerospace Products Los Angeles, CA
Adhesive
608 Epoxy-Patch
The Dexter Corp. Engineering Adhesives Division
Adhesive
Polystrate 5-Minute Epoxy
ITW Devcon 30 Endicott St. Danvers, MA 01923
Adhesive
EA-960F
The Dexter Corporation Aerospace Material Division
Adhesive
Aluminum Putty F
ITW Devcon
Adhesive
16307
Dayton-Granger Inc. 3299 SW 9th Ave. P.O. Box 350550 Ft. Lauderdale, FL 33335
`Adhesive
KE4238/HD3475
The Dexter Corp. Electronic Materials Divsion Olean, NY
Adhesive
Type/Class IH
15348
Dayton-Granger Inc.
Adhesive
Type/Class IIA
SC-1589
H. B. Fuller Company 1200 Willow Lake Blvd. P.O. Box 64683 St. Paul, MN 55164
Adhesive
Vangrip 14-30
Mid-West Industrial Chemical Co. St. Louis, MO
Adhesive
Type/Class IF
Type/Class IG
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MODEL 208 MAINTENANCE MANUAL Table 201. Adhesives (continued) NAME
NUMBER
MANUFACTURER
1300-L
3M Adhesives, Coatings Division St. Paul, MN
USE Adhesive
& Sealers
Type/Class IIB
30-NF2000-NF/ Spray
3M Co.
Adhesive
Type/Class IIIA
Scotch-Grip 847
3M Co.
Adhesive
Type/Class IIIB
EC-776
3M Co.
Adhesive
EC-776SR
3M Co.
Adhesive
CS-3600
Flamemaster Corp. Chem Seal Division Sun Valley, CA
Adhesive
EC2262
3M Co.
Adhesive
4693
3M Co.
Adhesive
RTV102 RTV103 RTV108
General Electric Company Silicone Products Dept. Mechanicville Rd. Waterford, NY12188
Adhesive
RTV732 RTV734
Dow Corning Corp.
Adhesive
Type/Class VB
RTV106
General Electric Company
Adhesive
Type/Class VC
RTV157 RTV159
General Electric Company
Adhesive
Type/Class VD
93-076 RTV
Dow Corning Corp.
Adhesive
PSA529/ SRC18
General Electric Company
Adhesive
Type/Class VF
Silastic 730
Dow Corning Corp.
Adhesive
Type/Class VIIA
Loctite 49550
Loctite Corp. Newington, CT 06111
Adhesive
Type/Class VIIB
Blak Max 38050 38061
Loctite Corp.
Adhesive
Type VIII
950 Transfer Tape 4930 VHB Tape 4945 VHB Tape Scotch-Mount 4962
3M Co. Industrial Tape Division St. Paul, MN
FasTape 1191 UHA Transfer Tape FasTape 3099 UHA Transfer Tape
Avery Dennison Specialty Tape Division Painesville, OH
Type IV
Type/Class VA
Adhesive and
Specialties
Adhesive
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MODEL 208 MAINTENANCE MANUAL Table 201. Adhesives (continued) NAME Type/Class IXA
Type/Class IXB
NUMBER
MANUFACTURER
Thixon 405
Morton Intl. Inc. Specialty Chemicals Adhesives West Alexandria, OH
USE Adhesive Group/
HE17017
Hartel Enterprises, Inc.
Adhesive
Uralane 8089A/B Uralane 5774A/B
Ciba-Geigy Corp. Reliable Division Fountain Valley, CA 92708
Adhesive
32555 707
Loctite Corp.
Adhesive
PS-18 PS-30
Caseway Industrial Products 6624 Prospect St. P.O. Box 249 Caseville, MI 48725
Adhesive
Type XI
Hot Melt Adhesive 6363
Bostik Inc. Middleton, MA
Adhesive
Type/Class XIIA
Duco Cement
ITW Devcon
Adhesive
Type/Class XIIB
Velcro #40
Velcro USA Inc. Manchester, NH
Adhesive
Type X
Table 202. Solvent Cements NAME
NUMBER
(All solvents should be either C.P., U.S.P. or Reagent Grade)
Class VI Methyl n-Propyl Ketone MIBK Acetone Cyclohexanone Tetrahydrofuran Methylene Dichloride Ethylene Chloride EC4801 6.
MANUFACTURER
USE Cleaning
3M Co.
Requirements for Bonding A.
Surfaces to be bonded must be clean and dry, free from dust, lint, grease, oil, condensation, other moisture and all other contaminating substances.
B.
Jelled or over aged adhesives should not be used until they are tested.
C.
Bonds should be free of wrinkles and entrapped air bubbles. They should not be loose at the edges or exhibit poor adhesion.
D.
Containers for adhesives should be kept tightly closed when the adhesives are not being used, unless otherwise specified.
E.
Adhesives should not be applied when the temperature of either the adhesive or the surfaces to be bonded is below 65°F.
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MODEL 208 MAINTENANCE MANUAL F.
Two component adhesives require weighing and mixing unless pre-weighed kits are used. Weighing equipment should be kept clean and personnel doing the weighing should use good personal hygiene.
G.
Metals should be chemically cleaned. In general, cleaning of metal surfaces for bonding consists of removing oils and greases by solvent or soap solutions. This is followed by mechanical abrading of the surface. Loose abraded particles should be removed before bonding.
H.
In general, plastics, rubber, leather, cork, wood, etc., should be cleaned of oil and grease by use of solvent. Bonding of plastics and rubber will be improved by abrading the surface after degreasing. Loose abraded particles should be removed before bonding. NOTE:
(1) (2)
7.
All surfaces prepared for adhesive bonding should be free of grease, fingerprints, paint, heat scale, corrosion, smut, powder, etc.; slight water stains are permissible providing the surface passes the following:
Examine metal surfaces while they are still wet from the rinsing operations for continuity of water film. Formation of water droplets or discontinuity of the water film (water break) indicates the presence of oily or greasy residues and parts should be reprocessed. If doubt exists to quality of rinse, the following test may be applied to parts while they are still wet from the rinsing operation. Select a representative area of the bonding surface and test this area with pH indicating paper. A pH of less than 5.0 or greater than 9.5 requires rerinsing and retesting of surface.
Manual Cleaning and Deoxidizing of Aluminum Alloys A.
Procedure. NOTE: (1) (2) (3) (4) (5)
(6)
Exercise care to prevent trapping solutions at the edge of joints.
Remove oil, grease, ink, etc. by solvent cleaning. Mask off dissimilar metals or surface not to be deoxidized. Spray, brush or swab alkaline cleaner (Turco 4215) on surface. Keep area to be cleaned wet for at least 5 minutes. Spray rinse thoroughly with room temperature water for a minimum of 3 minutes. With an acid brush, apply paste cleaner 0.06 inch to 0.12 inch thick on surface to be bonded. Allow the paste cleaner to remain on the surface for 45 to 60 minutes. (a) Paste cleaner (all measure by weight). Sulfuric Acid (Concentrated, Technical Grade): 38 percent, +2 or -2 percent; Sodium 1 Dichromate Dihydrate: 7 percent, +1 or -1 percent; Cab-O-Sil: 7 percent, +1 or -2 percent. Balance: distilled water. 2 Remove the paste with dry cheesecloth. Wash the area with a clean cheesecloth saturated with high purity water. Parts should be water-break free. If not, repeat procedure above beginning with Step 3. Dry the area for 15 minutes minimum with heat lamps before bonding. NOTE:
Parts processed should be handled so as to prevent recontamination by dirt, grease, fingerprints, etc. Personnel handling prepared surfaces for adhesive bonding should wear clean, white, low-lint gloves. Change gloves frequently to avoid contamination. Gloves may be contaminated easily by contact of working surface with body oils or hair. These soils must not be transferred to bonding surfaces. Prepared surfaces of items that will require transportation or short time storage should be wrapped with clean kraft paper.
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8.
Liquid Solvent Cleaning
WARNING: Solvents should be considered flammable and should not be exposed to flames or sparks under any circumstances. Fresh air masks and/or adequate ventilation is required for all closed areas. A.
Requirements. (1) Plastic or rubber materials should not be immersion cleaned or vapor degreased. (2) Solvents should never be poured or sprayed on surface to be cleaned. (3) It is essential that clean cloths and clean solvents are used during the final cleaning operation. (4) Bonding or subsequent priming operations should be accomplished as soon as possible after cleaning and drying of surfaces. (5) Solvent cleaned surfaces should be dry and free of all visible soils. Iridescent surfaces are evidence of improper cleaning.
B.
Procedures. (1) Liquid solvent cleaning should generally be used when it is not practical to clean parts of assemblies by vapor degreasing or immersion in chemical cleaners. However, some finishing codes require solvent cleaning. One or more steps may be eliminated if the surfaces to be cleaned are not soiled enough to warrant the inclusion of all steps. (2) Wipe off excess oil, grease or dirt from surface. (3) Apply solvent to a clean, oil-free cloth, preferably by pouring solvent on the cloth from a safety can or other approved container. The cloth should be well saturated but not to the point where it is dripping. (4) Wipe the surface with the moistened cloth as required to dissolve or loosen soil. Work on a small enough area so that the surface being cleaned remains wet. (5) With a clean dry cloth, immediately wipe the surface while the solvent is still wet. Do not allow the surface to evaporate dry. (6) Repeat Steps (3) through (5) until there is no discoloration on the drying cloth.
C.
Additions or Exceptions. (1) Metals. (a) Prior to bonding or priming, lightly abrade surface with ScotchBrite brand pads, Clean N’ Finish material, Type A fine or aluminum oxide 320 grit sandpaper. Remove loose abraded particles and follow by solvent cleaning.
CAUTION: Abrasives containing silicone carbide are not suitable for this purpose and should not be used. (b)
Metal surfaces should be cleaned with a solvent chosen from reference Table 203.
Table 203. Solvent Metal Cleaners METAL
SOLVENT
All
Methyl n-Propyl Ketone MIL-PRF-680 Solvent, Dry Cleaning, Type III TT-I-735 Isopropyl Alcohol
Table 204. Solvent Cleaners for Plastic Materials PLASTIC TYPE
SOLVENT
ABS (Acrylonitrite-Butadiene-Styrene)
TT-I-735 Isopropyl Alcohol
Cellulose Acetate
MIL-PRF-680 Solvent, Dry Cleaning, Type III
CAB (Cellulose- Acetate-Butyrate)
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MODEL 208 MAINTENANCE MANUAL Table 204. Solvent Cleaners for Plastic Materials (continued) PLASTIC TYPE
SOLVENT
PPO (Polyphenylene Oxide) Polystyrenes Polyurethanes Silicones Vinyls Acrylics
TT-I-735 Isopropyl Alcohol
Polycarbonates Epoxies Melamines Nylons Phenolics Polyesters
Methyl-n-Propyl Ketone
Polyethylenes
Detergent, Liquid Dishwashing
Polypropylenes Polyimides
TT-I-735 Isopropyl Alcohol
Fluoroplastics (TFE, FEP, KEL-F)
MIL-PRF-680 Solvent, dry cleaning, Type III Methyl n-Propyl Ketone TT- I-735 Isopropyl Alcohol
(2)
Plastic or rubber. (a) Removal of heavy soil from surfaces may be accomplished by washing the surface with a mild water and liquid dishwashing detergent solution prior to solvent cleaning.
CAUTION: Abrasives containing silicone carbide are not suitable for this purpose and should not be used. (b) (c)
Prior to bonding, lightly abrade surface with aluminum oxide 180 grit sandpaper. Remove loose abraded particles and follow by solvent cleaning. Surfaces should be cleaned with a solvent chosen from Table 204 for plastic or Table 205 for rubber materials.
Table 205. Solvent Cleaners for Rubber Materials RUBBER TYPE Buna S
SOLVENT TT-I-735 Isopropyl Alcohol
Buna N Neoprene Thiokol Butyl
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MODEL 208 MAINTENANCE MANUAL Table 205. Solvent Cleaners for Rubber Materials (continued) RUBBER TYPE
SOLVENT
Natural TT-I-735 Isopropyl Alcohol
Silicones Ethylene Propylene
NOTE 1: When cleaning rubber, use fluid sparingly and dry dampened area thoroughly. 9.
Adhesive Mixing, Application and Curing A.
Type I Epoxy. (1) These adhesives are two component materials and require weighing to obtain the proper amount of each component. Thorough mixing of the weighed components is required for the adhesive to perform properly. Do not mix large batches of the adhesives at one time as this reduces the pot life of the adhesive. Four hundred grams of the adhesive will generally give a pot life of 30 minutes. Small batches and shallow containers will lengthen pot life. Apply a coat of adhesive 0.020 to 0.030 inch (0.5 to 0.8 mm) thick to the surface to be bonded. Press bonded surfaces together to extrude excess adhesive and air so that the resultant bondline is 0.005 inch to 0.010 inch (0.13 mm to 0.25 mm) thick. Pressure may be applied by clamps or weights until cured. Cure for 24 hours at 77°F (25°C) or 2 hours at 180°F (82°C). (a) Type IA (EA9309.3NA). Combine 100 parts by weight of component A with 23 parts by weight of component B. Mix thoroughly. Weight and mix per instructions on containers. (b) Type IB (EA907). Combine 100 parts by weight of component A with 80 parts by weight of component B. Mix thoroughly and follow instructions on container. (Devcon F). Combine 1 part by weight Devcon F hardener with 9 parts by weight of Devcon F. Mix thoroughly to a lump free mixture. Devcon F will cure in 2 hours at 77°F (25°C). Both of these materials are primarily fillers for hole repair. Apply and shape to the desired thickness or contour, allow to cure, then sand to the desired shape or size. (c) Type IC (EA9394NA and 380/6). These high temperature adhesives are mixed by combining 100 parts by weight of base material to 33 parts of hardener. Observe mixing instructions on containers. (d) Type ID (A1186-B). Combine 1 part by weight of A1186-B catalyst A with 8 parts by weight of A1186-B, then mix thoroughly. The pot life of the mixed material is approximately 8 hours at 77°F (25°C). Apply a coat of adhesive on the surfaces to be bonded and allow them to air dry until the solvent odor is gone (approximately 3 to 4 minutes at 77°F (25°C)). Press the faying surfaces firmly together, preferably using a hard rubber or plastic roller and allow them to remain together for 16 hours at 77°F (25°C) before handling. Cure for 24 hours at 77°F (25°C) before applying stress to the bond. Maximum strength develops in 5 days at 77°F (25°C). Pressure may be applied by clamps or weights during part or all of the 5-day period as desired. (e) Type IE (EC2216). Combine 100 parts by weight of component B with 140 parts by weight of component A. Mix until the components blend to a uniform medium gray color. (f) Type IF (number 10). Combine equal weights or volume of both components. Mix together until material is one color. Apply to joint. Work life is only 5 minutes and material sets in 10 minutes. Apply pressure to the joint or component being bonded. Adhesive should carry a load within 1 hour.
B.
Type II, III and IV Synthetic Rubber and Resin Adhesives. NOTE: (1)
These adhesives are single component solvent blends of rubber or resin. They have high initial tack and will bond a wide variety of different materials.
Apply a coat of adhesive on the surfaces to be bonded and allow them to air dry until most of the solvent has evaporated and the adhesive exhibits an aggressive tack. This condition can be determined by touching the adhesive lightly using the back of the knuckle instead of the fingertips in order to minimize contamination. When the adhesive is quite tacky but no longer
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(2)
(3) (4) C.
transfers to the back of the knuckle, the surfaces are ready for bonding. This normally requires from 3 minutes to 30 minutes depending on film thickness, nature of the surfaces, temperature and humidity. (a) Very porous surfaces may require two coats. If two coats are applied, let the first coat dry completely from 30 minutes to 60 minutes before applying the second coat and testing for tack as described above. (b) When bonding two nonporous surfaces, the coat of adhesive to both surfaces may be allowed to dry completely and then one surface reactivated with a very light coat of adhesive and tested for tack as described above. This latter procedure will reduce the amount of solvent trapped in the bond and is especially useful in the case of bonding nonporous surfaces since trapped solvents can greatly prolong the time required for the bond to reach full strength. (c) Press the faying surfaces firmly together, preferably using a hard rubber or plastic roller and apply any needed clamps or weight. Cure for at least 24 hours at 77°F (25°C) before applying any stress to the bond. Type II. (a) EC880. Bond according to procedures in step (1). (b) EC847. Bond according to procedures in step (1). (c) EC1300L. Bond according to procedures in step (1). (d) 5452 Contact Adhesive. Bond according to procedures in step (1). (e) 5431 Tuf-Grip Cement. Bond according to procedures in step (1). (f) 1636. Bond according to procedures in step (1). Type III. (a) EC847. Bond according to procedures in step (1). Type IV. (a) Type IV (EC2262). Bond according to procedures in step (1).
Type V Silicone Rubber Adhesives. NOTE:
(1) (2) (3) (4)
(5) (6) D.
These adhesives are one part silicone rubber material which will bond a wide variety of different materials. Cure of these adhesives is initiated by moisture in the air. Nonporous surfaces being bonded with these adhesives will cure very slowly or not at all on wide bond lines.
Apply a coat of adhesive to the surfaces to be bonded and press them firmly together within 10 minutes. Apply pressure by clamps or weights for at least 24 hours at 77°F (25°C) before handling. Type VA (RTV157). Bond according to procedures in step (1). Type VB (RTV159). Bond according to procedures in step (1). Type VC. (a) RTV732. Bond according to procedures in step (1). (b) RTV734. Bond according to procedures in step (1). (c) RTV738. Bond according to procedures in step (1). (d) RTV102. Bond according to procedures in step (1). (e) RTV103. Bond according to procedures in step (1). (f) RTV108. Bond according to procedures in step (1). (g) RTV109. Bond according to procedures in step (1). (h) RTV162. Bond according to procedures in step (1). Type VD (RTV106). Bond according to procedures in step (1). Type VE (RTV94-034). Bond according to procedures in step (1).
Type VI Solvent Bonding. (1) This type bonding depends on the solvent softening the plastic surfaces to be bonded. The softened surfaces are pressed together and held until the solvent evaporates and the plastic hardens. The appropriate solvent may be applied to plastic surfaces by brushing, spraying, dipping or by use of a felt pad. (2) Allow the solvent to remain on the plastic until both surfaces soften then immediately join the surfaces while wet.
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Apply clamps, weight or a holding fixture to keep the bonding surfaces in solid contact until the bond is firmly set. Allow the bond to set for 24 hours at 77°F (25°C) before applying any stress to the bond.
E.
Type VII Cyanoacrylate. (1) Apply adhesive to surface to be bonded. Do not apply excess adhesive. Mating parts should fit well so the bond line will be 0.005 inch (0.13 mm) or as specified on label. Apply clamping pressure.
F.
Type VIII Pressure Sensitive. (1) Clean surfaces to be bonded. Apply adhesive to bond surface and assemble with pressure. Adhesive gains strength with time under pressure although most parts may be handled within 5 to 10 minutes after application of pressure.
G.
Type IX Polyurethane. (1) This adhesive is a tough, flexible material which bonds a variety of plastic materials as well as aluminum. The adhesive offers excellent low temperature performance plus good peel. Weigh and mix adhesive in accordance with directions on label. Apply adhesive to area to be bonded. Clamp parts together so the resultant bond line will be within 0.005 inch to 0.020 inch (0.13 mm to 0.51 mm) and there is no entrapped air. Parts may be lightly handled after 6 hours but 24 hours of cure is preferred.
H.
Type X Acrylic. (1) PS-18. (a) The mixing of PS-18 cement is based on 4 fluid ounces (118 ml) of cement. The 4 fluid ounces have a useful life of 30 minutes. Mix at a temperature of 65°F to 80°F (18°C to 27°C). Batches larger than 4 fluid ounces (118 ml) should not be mixed at one time. Do not mix more cement than can be used in 30 minutes. Unused cement should be discarded after 30 minutes. (b) Add one capsule (2.4 g) of catalyst mixture (component B) to 4 fluid ounces (118 ml) of base cement (component A). Dissolve by stirring. The base cement with catalyst added may be stored in a refrigerator at 40°F (4°C) or below for 24 hours.
WARNING: Do not mix catalyst (Component B) directly with promoter (Component C). A violent reaction will take place when these two materials are directly mixed together. If promoter (Component C) is accidentally spilled on skin, remove immediately by washing with soap and water. (c) (d)
Just before using the catalyzed cement, add 5 ml of promoter (component C) to the mixture. Stir thoroughly. Do not add component C to more cement than can be used in 30 minutes. Surfaces of acrylic to be bonded should have these areas sanded and cleaned with aliphatic naphtha Type II (TT-N-95) before application of cement. Apply cement to bond area in sufficient quantity so the bond line, after pressure application, will be 0.005 inches to 0.015 inch (0.13 mm to 0.38 mm). Apply clamps or pressure by other means and hold for at least 3 hours at 75°F (24°C). After this time, parts may be lightly handled. Allow 24 hours at 75°F (24°C) for more complete cure.
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MODEL 208 MAINTENANCE MANUAL ANAEROBIC ADHESIVES - MAINTENANCE PRACTICES 1.
General A.
This maintenance practice provides procedures for application of anaerobic adhesives. Anaerobic adhesives are retaining compounds which will harden between properly prepared mating surfaces where air is excluded.
WARNING: Some of these materials contain Aromatic Amines and/or Cyano-Acrylate fluids and are mildly poisonous. Avoid prolonged or repeated contact with the liquid or breathing of the vapors. Use with adequate ventilation. Cyano-Acrylate adhesives will instantly bond skin and can cause severe eye injury. Apply only to surface to be bonded. In case of skin contact, flush with water. In case of eye or internal contact, get medical attention. 2.
Materials A.
For anaerobic adhesive materials and application, refer to Table 201., Adhesives and Applications. NOTE: NAME
Equivalent substitutes may be used for the following items. NUMBER
MANUFACTURER
USE
Methyl n-Propyl Ketone
Commercially available
Cleaning solvent.
Sotoclean 110
PRC-DeSoto International 5340 San Fernando Rd. Glendale, CA 94710
Cleaning solvent.
DS108
Dynamold, Inc. 2905 Shamrock Ave. Fort Worth, TX 76107
Cleaning solvent.
Locquic Primer N (Ready to use)
Catalog Number 764-56
Loctite Corp. 705 North Mountain Road Newington, CT 06111
May be used with Loctite products 515, 569 and 592.
Locquic Primer T
Catalog Number 747-56
Loctite Corp.
May be used with Loctite products 222, 242, 271, 277, 290, 416, 601, 620 and 680.
Loctite Corp.
Tacking wires or wire bundles in place.
Wire Tacking Kit (Loctite) 3.
General Requirements for Bonding/Sealing A.
All surfaces to be bonded and/or sealed must be free of paints and corrosion preventive organic coatings.
B.
Surfaces must be clean and dry prior to application of adhesive. (1) Surface must be free from dust, lint, grease, chips, oil, condensation or other moisture and all other contaminating substances.
C.
Primers and/or adhesives must not be applied when the temperature of the primer, adhesive or parts is below 60°F.
D.
Excess adhesive must squeeze out of the joint when it is secured.
E.
Correct primer and/or adhesive must be used.
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Correct cleaning solvent and procedure must be used.
Table 201. Adhesives and Applications LOCTITE PRODUCT
COLOR
APPLICATION
222
Purple
Low strength locking and sealing or threaded fasteners (removable).
242
Blue
Medium strength locking and sealing of threaded fasteners and key assemblies (removable) (one-fourth inch diameter or larger).
271
Red
High strength locking and sealing of threaded fasteners (removable) (three-eighth inch diameter or larger).
277
Red
High strength locking and sealing welds and locking preassembled threaded fasteners.
290
Green
Penetrating action for sealing welds and locking preassembled threaded fasteners.
515
Purple
Form-in-place gaskets and dressing cut gaskets.
569
Brown
Seals hydraulic fluids, including fire resistant synthetics, to the working pressure of the hydraulic line.
592
White
Low strength sealing of pipe threads and threaded fittings.
601
Green
High strength retaining for studs, bearings, and bushings.
620
Green
High strength retaining at temperatures up to 450°F.
680
Red
Extra-high strength retaining for strengthening press fits and bonding cylindrical parts.
416
Clear
Wire tacking adhesive.
4.
Cleaning A.
All surfaces to which adhesive is to be applied must be clean and dry.
WARNING: Caution must be observed during cleaning. Most cleaning solvents are toxic and/or flammable. (1) (2) (3) 5.
Parts will be vapor degreased, cleaned with solvents, or, when a primer is to be used, the primer may also be used for cleaning. Allow all cleaned surfaces to dry a minimum of five minutes prior to application of primer and/or adhesive materials. In the event that contamination occurs, the surfaces must be recleaned.
Primer Application A.
When primer is used, it may be applied by dipping, brushing or spraying the surfaces to be bonded and/or sealed. The parts should be allowed to drain and air dry prior to application of the adhesive materials. (1) Apply only a thin uniform coating, avoiding any excess.
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6.
Adhesive Application NOTE: A.
7.
Application of Adhesive. (1) Apply the adhesive to both mating surfaces of the joint when possible. (2) Sufficient adhesive must be applied so the space between the assembled surfaces is completely filled with adhesive and a small excess is squeezed out around the periphery of the joint when the joint is secured. (3) After application of adhesive over primed surfaces, the joint should be secured immediately. (4) After the assembly is secured, remove uncured adhesive which extrudes onto the exterior of the joint. Use a wiper dampened with water or solvent as applicable.
Adhesive Cure A.
8.
All applications must be made using the proper adhesive material.
In most cases, adhesives applied to active surfaces will cure in 24 to 26 hours at 77°F. (1) Factors that influence cure are: (a) Activity of the surfaces. (b) Clearance in joint fit. (c) Specific adhesive material used. (d) Specific primer used, if any. (e) Temperature.
Wire Tacking A.
Wires and/or wire bundles may be tacked in place with adhesive Loctite 416 or using a Wire Tacking Kit. (1) When utilizing a Wire Tacking Kit, follow the instructions provided with the kit. (2) When utilizing adhesive Loctite 416 not in a kit, perform the following: (a) Hand solvent clean the surface. Refer to Cleaning. (b) Apply a thin stripe of "Speed" activator (primer) with the felt applicator in the bottle. (c) Position the wire and hold it against the surface. (d) Apply one or two drops of wire tacking adhesive over the wire. (e) Hold for 15 to 30 seconds.
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MODEL 208 MAINTENANCE MANUAL ADHESIVES, CEMENTS AND SEALANTS SHELF LIFE AND STORAGE - DESCRIPTION AND OPERATION 1.
2.
General A.
This section provides information which deÞnes the proper storage and usable life (shelf life) of adhesives, cements and sealants which are used for maintenance and/or repair of the airplane. Also included in this section is the criteria used for testing these materials after the normal shelf life has expired, to determine if an extension to the shelf life is possible.
B.
Shelf life refers to a speciÞed period of time usually from the date of manufacture (normally stamped or printed on the product container) to the expiration date (which should be determined using limits speciÞed in Table 1) or if applicable, the manufacturer's expiration date printed or stamped on the product container) The speciÞed shelf life is dependent on proper storage in accordance with the limits speciÞed in this section and/or the manufacturer's instructions.
Storage Criteria A.
Storage of Adhesives and Cements. (1) All adhesives and cements shall be stored under controlled temperature conditions. If open shop storage becomes necessary, these products shall in no case be stored in an area which will subject them to temperatures in excess of 95°F. Containers shall be tightly closed prior to placing them in the proper storage environment. For deÞnition of the proper storage environment, refer to Table 1 and the following paragraphs. For identiÞcation of adhesive and cement classiÞcations, refer to Adhesive and Solvent Bonding - Maintenance Practices. (a) Class I - These adhesives are epoxy base materials and have one-year storage at room temperature. 0°F storage will extend the storage life. Refer to the product container instructions for storage temperature and life. (b) Class II, III and IV - These adhesives are rubber and resin base and are good for six months at room temperature storage. 40°F storage will extend the storage life. Refer to the product container instructions for limits of each adhesive. (c) Class V - These are silicone rubber adhesives. If stored in their original containers at a temperature below 80°F, the shelf life is one year or as indicated on the storage container. (d) Class VI - These are solvent bonding solvents. They should be stored in tightly closed, original containers at 40°F. (e) Class VII - Cyanoacrylate base materials must be stored in the original containers at 40°F or as speciÞed on the container instructions. (f) Class VIII - These are pressure sensitive materials. The shelf life is two years when stored at 75°F and 50 percent relative humidity. (g) Class IX - These are polyurethane products. Store in original container, between 70° and 100°F. Urethanes are moisture sensitive and precautions should be taken to ensure complete protection from moisture contamination. Container must be tightly closed at all times. (h) Class X - These are acrylic base materials. They require storage at 40°F or per instructions on product container.
B.
Storage of Sealants. (1) All sealants shall be stored under controlled temperature conditions. If open shop storage becomes necessary, these products shall in no case be stored in an area which will subject them to temperatures in excess of 95°F or below 40°F. Containers shall be tightly closed prior to placing them in the proper storage environment. For proper storage environment, refer to Table 1 and the following paragraphs. For identiÞcation of sealant classiÞcation, refer to Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (a) Premixed and frozen sealants shall be stored at -40°F or colder and shall not be used more than six weeks after the date of mixing, even if all storage is at -40°F or colder. If storage temperatures rise above -40°F, but are not warmer than -30°F, the material may be stored for a maximum of two weeks warmer than -40°F plus time at -40°F or colder for a combined total not to exceed Þve weeks beyond the date of mixing. If storage temperatures rise
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(b)
C.
3.
above -40°F but are not warmer than - 20°F, the materials may be stored for a maximum of one week above -30°F plus time at -40°F or colder for a combined total not to exceed four weeks beyond the date of mixing. Unmixed sealants shall be stored at a controlled temperature of between 40°F and 80°F and have a shelf life of approximately six months when stored within this temperature range. Unmixed sealants stored at temperatures exceeding 80°F shall be used within Þve weeks.
All materials should be used on a "Þrst in, Þrst out" basis. The adhesives, cements and sealants should be rotated so this requirement can be accomplished. All material containers should be clearly marked with a "use by" date, consisting of the year and month. All materials not used by this date must be tested prior to use. Refer to Testing Criteria and Table 1.
Testing Criteria A.
Any material (adhesive, cement or sealant) not used within its shelf life will be tested and the results reviewed to determine if the material is usable. If there is doubt about the material being usable, it must be properly disposed of. Material that has exceeded its original shelf life may be retested to determine if the material meets its requirements. Materials meeting their requirements will have their shelf life extended as speciÞed in Table 1. Materials with shelf life extensions must be retested after a speciÞed period of time. Refer to Table 1.
B.
Testing of Overage Adhesives and Cements. NOTE: (1)
(2)
C.
Overage adhesives and cements are those that have exceeded their original shelf life and must be tested prior to use and/or given extended shelf life.
For identiÞcation of adhesive and cement classiÞcation, refer to Adhesive and Solvent Bonding - Maintenance Practices. (a) Class I Epoxy Adhesive - Examine both components to ensure that they are still workable. Check for gelling and/or contamination. Stir components and mix a small amount of adhesive. Verify that adhesive sets up and hardens. (b) Class II, III and IV Rubber and Resin Base Adhesives - Open containers and check for gelling and/or contamination. Check for spreading and drying. (c) Class V Silicone Rubber Adhesives - Examine adhesive for hardness. If adhesive is still soft and can be spread, it is acceptable. Verify that adhesive will harden. (d) Class VI Solvent Bonding Solvents - Check for signs of apparent contamination. Solvents should be clean and clear with no signs of cloudiness. (e) Class VII Cyanoacrylic Base Adhesives - Verify that product is still liquid with no visible signs of contamination. (f) Class VIII Pressure Sensitive Materials - Open containers and inspect for hardening, gelling and contamination. Stir components and mix a small amount of adhesive. Verify that adhesive sets up properly. (g) Class X Acrylic Adhesives - Inspect base material to ensure that it is still liquid. Mix a small amount of the components and verify that it sets up properly. In general, if these materials exhibit normal physical properties, with no signs of hardening, gelling or contamination and set up and/or harden properly as applicable, the shelf life may be extended as speciÞed in Table 1.
Testing of Overage Sealants. NOTE: (1) (2)
Overage sealants are those that have exceeded their original shelf life and must be tested prior to use and/or given extended shelf life.
For identiÞcation of sealant classiÞcation, refer to Fuel, Weather and High-Temperature Sealing - Maintenance Practices. Overage sealants to be tested for possible shelf life extension shall be properly mixed using the correct materials, procedures and equipment.
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(4)
(5) (6) (7) (8) (9) (10) (11) (12) (13) (14)
Overage premixed frozen sealants, along with unmixed sealants should be visually inspected. Sealants which show conclusive evidence of separation, discoloration and/or gelling prior to the addition of a thinner or curing agent shall be discarded. When in doubt of the sealant quality, the overage sealant should be compared with the same type of sealant, under six months old, which is known to be satisfactory. The mixed sealants may be tested by placing a small amount of sealant (sample buttons) on a sheet of paper. After the sample buttons have cured, they should be cut in half and examined. The sealant should show no signs of spots or streaks of unmixed base compound or curing agent. However, sample buttons containing spots, streaks, discoloration and/or variations in uniformity of color are acceptable if these spots, streaks, etc., are tack free upon inspection. All mixed sealant should be as void free as possible. Contaminated sealant and premixed sealant that have been thawed and refrozen shall be discarded. Type I, Class A sealants should be checked for appearance, application time, tack-free time, cure time and adhesion. Type I, Class B sealants should be checked for appearance, application time, cure time, tackfree time and adhesion. In addition, Class B-2 and B-4 sealants should be checked for initial ßow. Type I, Class C sealants should be checked for appearance, application time, cure time and adhesion. In addition, Class C sealants should be tested to determine that they are not at a tack-free condition at the end of their rated work life (squeeze out life). Type II sealants should be checked for appearance, application time, tack-free time and cure time. Type III sealants should be easily thinned with methyl n-propyl ketone. When difÞculty is encountered in thinning this sealant, it should be discarded. Type IV sealants should be checked for appearance, application time, tack-free time and cure time. Type V and VI sealants should be checked for appearance, tack-free time and cure time. Type VII sealants should be checked for appearance, application time, tack-free time and cure time. Type VIII sealants should be checked for appearance, application time, tack-free time, cure time and adhesion. Adhesion to aluminum should be (peel) less than two pounds per inch of width. NOTE:
For application time, tack-free time and cure time for the above listed sealant types and classes, refer to Sealant Curing in Fuel, Weather and High- Temperature Sealing - Maintenance Practices. Also refer to Table 1,. Storage and Shelf Life of Adhesives, Cements and Sealants.
Table 1. Storage and Shelf Life of Adhesives, Cements and Sealants. PRODUCT
STORAGE CONDITION (TEMPERATURE)
SHELF LIFE
EXTEND SHELF LIFE
RETEST IN
ADHESIVES AND CEMENTS EA9309.3NA
40 to 80°F
12 Months
6 Months
6 Months
EA9339
40 to 80°F
12 Months
6 Months
6 Months
EA9314
40 to 80°F
12 Months
6 Months
6 Months
EA9330
40 to 80°F
12 Months
6 Months
6 Months
EA907
40 to 80°F
12 Months
6 Months
6 Months
Devcon F
40 to 80°F
12 Months
6 Months
6 Months
EA934NA
40 to 80°F
12 Months
6 Months
6 Months
380/6
40 to 80°F
12 Months
6 Months
6 Months
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MODEL 208 MAINTENANCE MANUAL Table 1. Storage and Shelf Life of Adhesives, Cements and Sealants. (continued) STORAGE CONDITION (TEMPERATURE)
SHELF LIFE
EXTEND SHELF LIFE
RETEST IN
A1186-B
40 to 80°F
12 Months
6 Months
6 Months
EC2216
40 to 80°F
12 Months
6 Months
6 Months
#10 Fastset
40 to 80°F
12 Months
6 Months
6 Months
608 Quickset
40 to 80°F
12 Months
6 Months
6 Months
EC880
40 to 80°F
8 Months
3 Months
3 Months
EC847
40 to 80°F
8 Months
3 Months
3 Months
EC1300L
40 to 80°F
* 6 Months
* 3 Months
* 3 Months
5452
40 to 80°F
12 Months
6 Months
6 Months
5431
40 to 80°F
12 Months
6 Months
6 Months
1636
40 to 80°F
12 Months
6 Months
6 Months
EC2262
40 to 80°F
12 Months
6 Months
6 Months
RTV-157
40 to 80°F
12 Months
6 Months
6 Months
RTV-158
40 to 80°F
12 Months
6 Months
6 Months
RTV-159
40 to 80°F
12 Months
6 Months
6 Months
RTV-732
40 to 80°F
12 Months
6 Months
6 Months
RTV-102
40 to 80°F
12 Months
6 Months
6 Months
RTV-103
40 to 80°F
12 Months
6 Months
6 Months
RTV-106
40 to 80°F
12 Months
6 Months
6 Months
RTV-108
40 to 80°F
12 Months
6 Months
6 Months
RTV-109
40 to 80°F
12 Months
6 Months
6 Months
RTV-94-034
40 to 80°F
12 Months
6 Months
6 Months
Loctite 222
40 to 80°F
12 Months
6 Months
6 Months
Loctite 242
40 to 80°F
12 Months
6 Months
6 Months
Loctite 271
40 to 80°F
12 Months
6 Months
6 Months
Loctite 277
40 to 80°F
12 Months
6 Months
6 Months
Loctite 290
40 to 80°F
12 Months
6 Months
6 Months
Loctite 416
40 to 80°F
12 Months
6 Months
6 Months
Loctite 495
40 to 80°F
12 Months
6 Months
6 Months
Loctite 515
40 to 80°F
12 Months
6 Months
6 Months
Loctite 569
40 to 80°F
12 Months
6 Months
6 Months
Loctite 592
40 to 80°F
12 Months
6 Months
6 Months
Loctite 595
40 to 80°F
12 Months
6 Months
6 Months
Loctite 609
40 to 80°F
12 Months
6 Months
6 Months
Loctite 620
40 to 80°F
12 Months
6 Months
6 Months
PRODUCT
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MODEL 208 MAINTENANCE MANUAL Table 1. Storage and Shelf Life of Adhesives, Cements and Sealants. (continued) PRODUCT
STORAGE CONDITION (TEMPERATURE)
SHELF LIFE
EXTEND SHELF LIFE
RETEST IN
Loctite 680
40 to 80°F
12 Months
6 Months
6 Months
Loctite 12829
40 to 80°F
12 Months
6 Months
6 Months
Loctite 12839
40 to 80°F
12 Months
6 Months
6 Months
DA-552-1
40 to 80°F
12 Months
6 Months
6 Months
PS-18
40 to 80°F
12 Months
6 Months
6 Months
PS-30
40 to 80°F
12 Months
6 Months
6 Months
XA-3678
40 to 80°F
12 Months
6 Months
6 Months
XF-3585
40 to 80°F
12 Months
6 Months
6 Months
LR-100-226
40 to 80°F
12 Months
6 Months
6 Months
EC776
40 to 80°F
* 8 Months
* 3 Months
* 3 Months
SB and P-2
40 to 80°F
12 Months
6 Months
6 Months
Pro-Seal 890
40 to 80°F
6 Months
2 Months
2 Months
GC-408
40 to 80°F
6 Months
2 Months
2 Months
PR-1422
40 to 80°F
6 Months
2 Months
2 Months
PR-1440
40 to 80°F
6 Months
2 Months
2 Months
GC-435
40 to 80°F
6 Months
2 Months
2 Months
Pro-Seal 567
40 to 80°F
6 Months
2 Months
2 Months
PR-810
40 to 80°F
6 Months
2 Months
2 Months
Pro-Seal 700
40 to 80°F
6 Months
2 Months
2 Months
GC-1900
40 to 80°F
6 Months
2 Months
2 Months
PR-366
40 to 80°F
6 Months
2 Months
2 Months
Pro-Seal 706B
40 to 80°F
6 Months
2 Months
2 Months
Pro-Seal 735
40 to 80°F
6 Months
2 Months
2 Months
Pro-Seal 870
40° to 80°F
9 Months
4.5 Months
4.5 Months
Pro-Seal 895
40 to 80°F
6 Months
2 Months
2 Months
PR-1321
40 to 80°F
6 Months
2 Months
2 Months
GC-200
40 to 80°F
6 Months
2 Months
2 Months
RTV-730
40 to 80°F
6 Months
2 Months
2 Months
Pro-Seal 815
40 to 80°F
6 Months
2 Months
2 Months
GC-402
40 to 80°F
6 Months
2 Months
2 Months
SEALANTS
NOTE 1: * Do not use after three months of storage in the 81°F to 90°F range. Do not use after Þve days of storage above 90°F.
20-30-06 © Cessna Aircraft Company
Page 5 Mar 1/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL INTERIOR AND EXTERIOR FINISH - CLEANING/PAINTING 1.
General A.
2.
Interior and exterior finish cleaning/painting consists of general information and instructions for applying chemical film treatments, primer and topcoats to the airplane.
Interior and Exterior Finishes A.
Detail aluminum parts are chemically pretreated and epoxy primed prior to assembly. The chem-film pretreatment and the epoxy primer are primary coatings and must be maintained and preserved for corrosion control. Exterior assemblies that are to be topcoated receive ScotchBrite, hand solvent cleaning and another overall application of epoxy primer. The airplane exterior then receives an overall topcoat of polyurethane paint including stripes.
CAUTION: All plastic and fiberglass parts, except bushings, bearings, grommets and certain purchased antenna covers which are not colored or painted, shall be colored or painted to match adjacent surface. The head of the pitot tube must be open and free from paint and other foreign objects. The surface adjacent to static port must be smooth and free from all paint imperfection. Do not paint pitot tube, fuel caps, trim tab pushrods where they operate in an actuator, oleo strut sliding surfaces, standard polished spinners, exhausts stall warning vanes, chromed items (handles, locks, ect.) or the tie-down lugs (located on struts) or light lens. Paint the landing gear barrels and torque links to match the overall color. 3.
Paint Facility A.
4.
Painting facilities must include the ability to maintain environmental control of temperature at a minimum of 65°F (18°C). All paint equipment must be clean. Accurate measuring containers should be available for mixing protective coatings. Use of approved respirators while painting is a must for personal safety. All solvent containers should be grounded to prevent static buildup. Catalyst materials are toxic, therefore, breathing fumes or allowing contact with skin can cause serious irritation. Material stock should be rotated to allow use of older materials first, because its useful life is limited. All supplies should be stored in an area where temperature is higher than 50°F (10°C), but lower than 90°F (32°C). Storage at 90°F (32°C) is allowable for no more than sixty days, providing it is returned to room temperature for mixing and use. (1) Areas in which cleaning or painting are done shall have adequate ventilation and shall be protected from uncontrolled spray, dust, or fumes. (2) Areas for prolonged storage of cleaned parts and assemblies awaiting painting shall be free from uncontrolled spray, dust, or fumes, or else positive means of protecting part cleanliness such as enclosed bins or wrapping in kraft paper shall be provided. (3) Areas in which cleaning or painting are done shall be periodically cleaned and dusted. (4) Compressed air used for dusting and paint spraying shall be free from oil, water and particulate matter.
Sanding Surfacer A.
Purpose and Requirements. (1) Surfacer is applied over fiberglass and Kevlar assemblies to provide aerodynamic contour, smoothness and to seal porous surfaces. Application of surfacer also provides a good surface for a polyurethane finish. (2) The objective of a surfacer is to fill local depressions, pits, pin holes and other small surface defects so a smooth surface is obtained for paint. The total surfacer thickness shall not be greater than 15 mils (0.38 mm). Only enough surfacer shall be applied to obtain a smooth surface for paint. If less thickness will provide a smooth surface, this is better. A thick layer of surfacer is less flexible and may crack in service.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (3) (4)
(5)
B.
To complete the airplane's polyurethane finish over surfacer, begin by applying the intermediate coat. Apply topcoat (polyurethane enamel) using same procedure. Should a repair be required (cracked or chipped paint) to areas where surfacer is applied, sanding surfacer should be removed to expose fiberglass or Kevlar. It may be necessary to remove all sanding surfacer on that individual assembly and/or component to obtain a satisfactory finish. For additional information, refer to Cleaning. Sanding surfacer methods. (a) Do not intermix vendor material or substitute material. Also, do not substitute instructions. Select and use one vendor's material and use the corresponding instructions.
Cleaning.
CAUTION: Do not use chemical strippers on fiberglass, kevlar and graphite composite assemblies. Paint stripper solvent will damage these assemblies. Exterior composite assemblies include: inside of nose compartment doors, inside of nose landing gear doors, wing tips, aileron tips, inside tailcone access door, tailcone stinger cap, pylon ram air scoop and vertical stabilizer bullet CAUTION: Sanding of paint and/or sanding surfacer must be very carefully accomplished. Do not sand into the fabric layers of composite assemblies as this will result in loss of strength. (1) (2) (3) (4) 5.
Remove paint covering sanding surfacer by sanding. Paint should be removed well beyond damaged area. For best results, it is recommended to remove all paint covering sanding surfacer of that individual composite component. Remove sanding surfacer by sanding from individual component to expose fabric. Scuff sand area to be refinished with 320 grit paper. Do not over expose fabric. Clean surface with Methyl n-Propyl Ketone. Follow manufacturer's instructions for final cleaning prior to sanding surfacer application.
Paint Stripping A.
Mechanical Stripping (1) Mechanical methods of stripping include power sanding with a disc or jitterbug type sander, grinder, hand sanding, and wire brushing. (a) Ensure mechanical methods do not damage surfaces being stripped. Damage may include, but is not limited to, cutting fibers of composite structures or scratches in the surface of metallic surfaces.
CAUTION: Do not use low carbon steel brushes on aluminum, magnesium, copper, stainless steel or titanium surfaces. Steel particles may become embedded in the surfaces, and later rust or cause galvanic corrosion of the metal surfaces. (2) (3) (4) B.
Mechanical stripping must be used for stripping composite or plastic surfaces. Mechanical stripping is recommended for surfaces which might entrap chemical strippers and result in corrosion. Mechanical stripping is required for painted surfaces masked during chemical stripping.
Chemical Stripping.
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MODEL 208 MAINTENANCE MANUAL
WARNING: All paint strippers are harmful to eyes and skin. All operators should wear goggle-type eyeglasses, rubber gloves, aprons and boots. In case of contact with skin, flush with water. In case of contact with eyes, flush eyes thoroughly with water and consult physician immediately. Paint stripping should be done in a well ventilated area. CAUTION: Use of a heater with an open flame in an area in which stripping with a methylene chloride type stripper is used produces hydrochloric acid fumes. If acid is deposited on airplane it will corrode all surfaces. (1) (2)
(3)
Thoroughly clean airplane surfaces to remove all grease and other dirt which might keep stripping agent from attacking paint. All seams and joints must be protected by applying a tape, resistant to strippers, to every joint to prevent stripping chemicals from entering the skin joints. Chemicals used for stripping polyurethane paint are very difficult to remove from joints, and may promote corrosion or deteriorate bonding agents used in assembly of airplane. Mask following surfaces using plastic sheeting or waxed paper and plastic tape so as to make a safety margin of at least one-half inch (13 mm) between protected surface and surface to be stripped. NOTE: (a)
Do not use masking tape.
Mask all windows and transparencies.
CAUTION: Acrylic windows may be softened or otherwise damaged by paint stripper, solvent or paint. Use water and grease-proof barrier material and polyethylene coated tape to protect windows. Place barrier material over window and seal around periphery with polyethylene backed masking tape. Cut second sheet of barrier material an inch (26 mm) or more larger than window. 2 Place second sheet of barrier material over window and seal with polyethylene tape. 3 Mask all rubber and other non metals. Composites if possible, shall be removed from airplane prior to stripping. Mask all honeycomb panels and all fasteners which penetrate honeycomb panels. Mask all pivots, bearings and landing gear. Titanium, if used on airplane, must be protected from strippers. Mask all skin laps, inspection holes, drain holes, or any opening that would allow stripper to enter airplane structure.
1
(b) (c) (d) (e) (f) (g)
CAUTION: Do not allow paint stripper to contact high heat treated steel pins, such as pins attaching landing gear components. Paint strippers may induce hydrogen embrittlement in high heat treated steel. (4)
Apply approved stripper by spray or brush method.
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MODEL 208 MAINTENANCE MANUAL
WARNING: Use normal safety precautions when using flammable materials during cleaning and painting procedures. WARNING: Paint stripper solution is harmful to eyes and skin. Wear goggles, rubber gloves, apron and boots when working with paint stripper. Also wear appropriate respirator when applying “spray-on” strippers. The chemical supplier bulletins and instructions should be closely followed for proper mixing of solution, application methods and safety precautions. (a)
(5) (6) (7) (8) (9) (10) (11) (12) (13) C.
If using spray method, apply a mist coat to area to be stripped, then when paint begins to lift, apply a second heavy coat. (b) If applying with brush, brush across the surface only once, in one direction. Allow stripper coating to lay on the surface until paint lifts. After paint begins to lift, use a propylene bristle brush to agitate stripper to allow deeper penetration of stripper. Remove lifted paint with a plastic squeegee. Dispose of residue in accordance with local regulations. Inspect all surfaces for incomplete paint removal. (a) Repeat previous procedural steps as necessary until all paint is removed. After stripping airplane, thoroughly rinse to remove any stripping residue. Remove tape applied to protect joints and other masked areas. Carefully remove remaining paint at skin joints and masked areas by sanding with a hand or jitterbug type sander. If necessary to remove paint from inside skin joints, refer to Cleanout of Skin Joints. If corrosion is encountered, refer to Structural Repair Manual, Chapter 51, Corrosion/Repair, for corrosion treatment.
Cleanout of Skin Joints. (1) Install a surface conditioning disc on a pneumatic drill. (2) Taper edge of disc to an edge which will allow edge to fit into skin joint seam. (a) Run disc against a piece of coarse abrasive paper or a mill file until edge is tapered.
CAUTION: Excessive pressure or dwell time will cause scratches or grooves in metal. Ensure doubler at bottom of joint is not damaged or gouged in any way by this process. (3) (4)
Using tapered surface conditioning disc, remove paint and other material from joint seams. Carefully, and using as low speed as possible, remove paint and all other material from joint. NOTE:
6.
Surface conditioning disc will wear rapidly, it will be necessary to resharpen (retaper) disc frequently.
Hand Solvent Cleaning
WARNING: Work in a well ventilated area free from sources of ignition. Use only approved solvents and materials. CAUTION: Airplane shall be grounded during solvent wipe. A.
Surface Cleaning. (1) Apply solvent to a clean wiping cloth by pouring from a safety can or other approved container. The cloth should be well saturated with solvent. Avoid dipping wipers into open solvent containers as this contaminates the solvent. (2) Wipe the surface with the wet cloth as required to dissolve or loosen soils. Work on a small enough area so that the area being cleaned remains wet with solvent.
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Page 704 Mar 1/2000
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MODEL 208 MAINTENANCE MANUAL (3) (4) 7.
With a clean dry cloth, immediately wipe dry the area being cleaned. Do not allow the surface to evaporate dry. Repeat steps (1) through (3) as required and change cloths often.
Maintenance of the Interior and Exterior Primary Coatings and Topcoat A.
Rework and repair primary coatings on airplane interior and exterior surfaces for protection and corrosion control. (1) Minor scratches or defects, which do not penetrate the epoxy primer or which penetrate the primer and expose bare metal, with the total area of exposed bare metal less than the size of a dime, touch up as follows: (a) Hand solvent clean and sand with 320 grit or finer sandpaper. (b) Clean with compressed air, hand solvent clean again, then wipe with a tack rag. (c) Mix and reapply epoxy primer (MIL P-23377 or equivalent) as directed by the primer manufacturer or supplier. (d) On a properly prepared surface, mix and apply polyurethane topcoat as directed by the paint manufacturer or supplier. (2) Major defects which expose bare metal to an area larger than the size of a dime, touch up as follows: (a) Hand solvent clean and sand with 320 grit or finer sandpaper. (b) Clean with compressed air, hand solvent clean again, then wipe with a tack rag. (c) Apply a spray wash primer or (preferred method) brush chem film primer. Mask the area to minimize the amount of primer from spreading over the existing epoxy primer. Let cure according to the product manufacturers recommendations. (d) Mix and apply epoxy primer (MIL P-23377 or equivalent) to the affected area within four hours. (e) If an exterior painted surface, mix and apply polyurethane topcoat as directed by the paint manufacturer or supplier.
20-31-00 © Cessna Aircraft Company
Page 705 Mar 1/2000
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL CONVERSION DATA - DESCRIPTION AND OPERATION 1.
2.
General A.
This section contains information for converting the more commonly used measuring units found in this manual from the common United States system to the International System of Units (metric system).
B.
Other conversion factors may be found in manuals such as Standard for Use of the International System of Units (SI): The Modern Metric System, prepared by ASTM, 100 Bar Harbor Drive, West Conshohocken, PA 19428-2959 USA.
Conversion Factors A.
Distance and Length (1) Multiply inches by 25.4 to obtain mm (millimeters). (2) Multiply feet by 0.3048 to obtain m (meters).
B.
Mass (1) Multiply ounces by 28.35 to obtain g (grams). (2) Multiply pounds by 0.436 to obtain kg (kilograms).
C.
Temperature (1) Subtract 32 from degrees Fahrenheit and multiply by 5/9 to obtain degrees Celsius.
D.
Torque (1) Multiply inch-pounds by 0.11298 to obtain Newton-meters. (2) Multiply foot pounds by 1.3588 to obtain Newton-meters.
E.
Force (1) Multiply pounds of force by 4.4482 to obtain N (Newtons).
F.
Pressure (1) Multiply pressure (psi) by 6.8948 to obtain kPa (kiloPascals).
G.
Mass flow (1) Multiply pounds-per-hour by 1.26 X 10 -4 to obtain kg/sec.
20-40-00 © Cessna Aircraft Company
Page 1 Mar 1/2000
21 CHAPTER
AIR CONDITIONING
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
21-00-00
Pages 1-2
Mar 3/1997
21-20-00
Page 1
Aug 1/1995
21-21-00
Pages 201-216
Apr 1/2010
21-21-00
Page 501
Jun 1/2011
21-22-00
Page 1
Aug 1/1995
21-22-00
Pages 101-102
Aug 1/1995
21-22-00
Pages 201-209
Aug 1/1995
21-24-00
Page 1
Apr 1/2010
21-24-00
Page 601
Jun 1/2011
21-24-01
Pages 401-402
Oct 15/1999
21-24-02
Pages 401-402
Apr 1/2010
21-41-00
Pages 1-5
Aug 1/1995
21-41-00
Pages 101-109
Aug 1/1995
21-41-00
Pages 201-217
Nov 3/2003
21-50-00
Page 1
Aug 1/1995
21-50-00
Page 601
Jun 1/2011
21-51-00
Pages 1-4
Apr 1/2010
21-51-00
Pages 101-110
Aug 1/1995
21-51-00
Pages 201-220
Apr 1/2010
21-52-00
Pages 1-4
Oct 15/1999
21-52-00
Pages 101-110
Oct 15/1999
21-52-00
Pages 201-221
Apr 1/2010
21-60-00
Pages 1-2
Aug 1/1995
21-Title 21-List of Effective Pages 21-Record of Temporary Revisions 21-Table of Contents 21-List of Tasks
21 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS AIR CONDITIONING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-00-00 21-00-00 21-00-00 21-00-00
DISTRIBUTION - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-20-00 Page 1 21-20-00 Page 1
FRESH AIR DISTRIBUTION - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inlet Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cabin Air Ventilation System Valves and Controls Removal/Installation. . . . . . . . . . . Cabin Air Outlet Valve Removal/Installation (Model 208 only) . . . . . . . . . . . . . . . . . . . Cabin Air Outlet Valves Removal/Installation (Model 208B Passenger) . . . . . . . . . . . Cabin Ventilation Fans and Switches Removal/Installation (Model 208 and 208B Passenger) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-21-00 Page 201 21-21-00 Page 201 21-21-00 Page 201 21-21-00 Page 201 21-21-00 Page 210 21-21-00 Page 211 21-21-00 Page 211
FRESH AIR DISTRIBUTION - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ventilation Outlet and Controls Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-21-00 Page 501 21-21-00 Page 501 21-21-00 Page 501
HEATING AND DEFROSTING AIR DISTRIBUTION - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-22-00 Page 1 21-22-00 Page 1 21-22-00 Page 1
HEATING AND DEFROSTING AIR DISTRIBUTION - TROUBLESHOOTING . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-22-00 Page 101 21-22-00 Page 101
HEATING AND DEFROSTING AIR DISTRIBUTION - MAINTENANCE PRACTICES . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Heater Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Defroster Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Defroster Nozzle Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Return Air Duct Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-22-00 Page 201 21-22-00 Page 201 21-22-00 Page 201 21-22-00 Page 201 21-22-00 Page 208 21-22-00 Page 208
AVIONICS COOLING - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-24-00 Page 1 21-24-00 Page 1
AVIONICS COOLING - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Avionics Cooling Fan Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-24-00 Page 601 21-24-00 Page 601 21-24-00 Page 601
CENTER CONSOLE AVIONICS COOLING - REMOVAL/INSTALLATION . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Center Console Avionics Cooling Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . .
21-24-01 Page 401 21-24-01 Page 401 21-24-01 Page 401
GARMIN DISPLAY UNIT (GDU) COOLING FAN - REMOVAL/INSTALLATION . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin Display Unit (GDU) Fan Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin Display Unit (GDU) Fan Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-24-02 Page 401 21-24-02 Page 401 21-24-02 Page 401 21-24-02 Page 401
COMPRESSOR BLEED AIR HEATER - DESCRIPTION AND OPERATION . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation (Airplanes 20800001 Thru 20800179 and 208B0001 Thru 208B0209) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation (Airplanes 20800180 and On, 208B0210 and On, and Airplanes Incorporating CAB90-9). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-41-00 Page 1 21-41-00 Page 1
21-21-00 Page 211
21-41-00 Page 1 21-41-00 Page 4
21 - CONTENTS © Cessna Aircraft Company
Page 1 Page 1 Page 1 Page 1
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MODEL 208 MAINTENANCE MANUAL COMPRESSOR BLEED AIR HEATER - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-41-00 Page 101 21-41-00 Page 101
COMPRESSOR BLEED AIR HEATER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Bleed Air Heater Components Removal/Installation . . . . . . . . . . . . . . . . Individual Component Disassembly/Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Component Cleaning/Servicing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cabin Heat Functional Test (Airplanes 20800001 Thru 20800179 and 208B0001 Thru 208B0209 Except Airplanes Incorporating CAB90-9) . . . . . . . . . . . . . . . . . . . . Cabin Heat Functional Test (Airplanes 20800180 and On and 208B0210 and On and Airplanes 20800001 Thru 20800179 and 208B0001 Thru 208B0209 Incorporating CAB90-9) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Heater Output Check (Airplanes 20800180 and On and 208B0210 and On) . . . . . .
21-41-00 Page 201 21-41-00 Page 201 21-41-00 Page 201 21-41-00 Page 208 21-41-00 Page 212
21-41-00 Page 215 21-41-00 Page 216
COOLING - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-50-00 Page 1 21-50-00 Page 1
COOLING - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Drive Belt Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-50-00 Page 601 21-50-00 Page 601 21-50-00 Page 601
FREON AIR CONDITIONING - DESCRIPTION AND OPERATION (Airplanes 20800112 and On, and 208B0214 and On) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-51-00 Page 21-51-00 Page 21-51-00 Page 21-51-00 Page
FREON AIR CONDITIONING - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21-51-00 Page 101 21-51-00 Page 101
FREON AIR CONDITIONING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . General Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Drive Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Drive Belt Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Condenser Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Receiver/Dryer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Conditioning Plumbing Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Mounted Evaporator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Mounted Return Air Check Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . Forward Evaporator Return Air Grill. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tailcone Mounted Evaporator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Evaporator Distribution and Return Air System Removal/Installation . . . . . . . . . . System Operational Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Switch Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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R134A AIR CONDITIONING SYSTEM - DESCRIPTION AND OPERATION (Airplanes 20800274 And On, and 208B0655 And On) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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R134A AIR CONDITIONING SYSTEM - TROUBLESHOOTING (Airplanes 20800274 and On, and 208B0655 and On) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Troubleshooting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL R134A AIR CONDITIONING SYSTEM - MAINTENANCE PRACTICES (Airplanes 20800274 And On, and 208B0655 And On) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Safety Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Drive Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Drive Unit Disassembly/Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Drive Belt Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drive Belt Tension Adjustment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Condenser Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Receiver-Dryer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Conditioning Plumbing Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing-Mounted Evaporator Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Mounted Return Air Check Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . Forward Evaporator Return Air Grill. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tailcone Mounted Evaporator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Evaporator Distribution and Return Air System Removal/Installation . . . . . . . . . . System Operational Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Switch Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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TEMPERATURE CONTROL - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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LIST OF TASKS 21-24-00-710
Avionics Cooling Fan Operational Check
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Compressor Drive Belt Functional Check
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MODEL 208 MAINTENANCE MANUAL AIR CONDITIONING - GENERAL 1.
Scope A.
2.
Definition A.
3.
This chapter describes those units and components which furnish a means of heating, cooling and ventilating the cockpit and cabin/cargo areas of the airplane.
This chapter is divided into sections to aid maintenance technicians in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief description of the sections follows: (1) The section on distribution describes that portion of the system used to distribute fresh and heated air throughout the cockpit and cabin area. (2) The section on heating describes those components used to generate (but not distribute) heat for the airplane. (3) The section on cooling describes the freon air conditioning system used to generate and distribute cool air throughout the cockpit and cabin. (4) The section on temperature control describes components used to control heat in the cabin area.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following listed items:
NAME
NUMBER
MANUFACTURER
USE
Primer
S4155
General Electric Silicone Products Dept. Hudson River Rd. Waterford, N.Y. 12188
Priming mating surface of instrument panel and inlet.
Sealant
RTV-102
General Electric
Sealant for air ducts.
Sealant
RTV-103
General Electric
Sealant used on compressor bleed air installation.
Clear Adhesive
A1186B
BFGoodrich 250 N. Cleveland Massillion Rd. P. O. Box 5501 Akron, OH 44318-0501
Resealing component parts in cabin air valves.
Sealant
GC-1900
Goal Chemical Sealants Corp. 3137 E. 26th St. Los Angeles, CA 90023
Sealant for heater and defroster valve flanges.
Stoddard Solvent
P- D-680
Commercially available
Cleaning.
Commerically available
Cleaning.
Methyl n-Propyl Ketone Isopropyl Alcohol
Federal Specification TT-l-735
Commercially available
Cleaning.
Silicone Lubricant
113A10010
Parker Hannifin Airborne Air & Fuel Products 711 Taylor Street Elyria, OH 44035
Lubricant for Bal-seals.
Sealant
Silastic E
Dow Corning
Used on diverter valves.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Primer
DC-1200
Dow Corning
Used in diverter valve reassembly.
Anti-Seize Compound
26316503
Parker Hannifin Airborne Air & Fuel Products
Used on bleed air flow control valve.
Sealant
RTV-157
General Electric
Used on diverter valves.
Refrigerant Oil (500 viscosity minimum)
Capella WF 100
Texaco Oil Company Box 1601 White Plains, NY 10650
Lubricate compressor, fittings and O-rings.
Refrigerant Oil (500 viscosity minimum)
Suniso 5GS
Sun Oil Company 1801 Market Street Philadelphia, PA 19103
Lubricate compressor, fittings and O-rings.
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MODEL 208 MAINTENANCE MANUAL DISTRIBUTION - DESCRIPTION AND OPERATION 1.
Description and Operation A.
This section is concerned with those systems which induct and distribute air for both environmental and equipment needs. It does not cover those devices which produce heated or cooled air, nor does it cover those devices used to change warm air temperature. (1) Environmental distribution system can be broken down to sub systems which provide a means of distributing fresh air, heated/defrosted air, and (optional) cool air within the airplane. Each system typically utilizes its own plenums and ducts to distribute air to respective outlets. (a) For a description of how fresh air is inducted and distributed throughout ventilation sub system, refer to Fresh Air Distribution - Maintenance Practices. (b) For a description of how hot air is distributed throughout heating/defrosting sub system, refer to Heating and Defrosting Air Distribution - Maintenance Practices. For a description of how hot air is produced, refer to Compressor Bleed Air Heater - Description and Operation. For a description of how warm air temperature is altered, refer to Temperature Control - Description and Operation. (c) For a description of how optional freon cooled air is produced and distributed, refer to Freon Air Conditioning - Maintenance Practices. (2) Equipment distribution system is limited to coverage of avionics cooling system. Refer to Avionics Cooling - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL FRESH AIR DISTRIBUTION - MAINTENANCE PRACTICES 1.
2.
General A.
Fresh air ventilation is provided for airplane via a series of hoses and ducts. Primary ventilation system picks up ram air at upper end of each wing strut and distributes it to a central plenum located in the cockpit headliner area. From plenum, various ducts are used to distribute fresh air throughout cockpit and cabin/cargo area. Control knobs, located in overhead console, provide a means to modulate fresh air flow from each ram air source before it reaches the plenum. (1) Model 208 and 208B passenger ventilation systems include wemac valves and associated duct work for pilot, copilot and each passenger seat position. (2) Model 208B, 208 Cargomaster and 208B Super Cargomaster ventilation systems include wemac valves and associated duct work for pilot and copilot seat positions only. (3) Model 208 and 208B Passenger may be equipped with optional blowers upstream of the plenum. These blowers can be used in ground and/or flight operations to draw additional fresh air into the plenum.
B.
Fresh air may also be drawn into cockpit area by small inlet doors located on the left and right side of forward fuselage. These doors are cable-actuated and open or close to allow a variable amount of ram air to flow into cockpit. Ducts connect doors to adjustable outlets (left and right) on the instrument panel, which further directs flow of fresh air in cockpit area.
Tools, Equipment and Materials A.
3.
For a list of required tools, equipment and materials, refer to Air Conditioning - General.
Inlet Door Removal/Installation A.
Remove Inlet Doors And Associated Components (Refer to Figure 201). (1) Cut straps securing duct to door inlet and instrument panel inlet. (2) Remove duct from airplane. (3) Remove screen from between duct and door inlet. (4) Remove nut, bolt and washer securing end of control cable to door lever. (5) Remove screw and clamp securing control cable to door inlet. (6) Drill out rivets which attach door inlet to fuselage. Remove door inlet from airplane and discard seal. (7) Place door inlet on work bench. (8) Remove cotter pin, washer and pin. Disconnect door lever from door. (9) Remove adjustable outlet from instrument panel. (10) Remove instrument panel inlet from instrument panel. Clean sealant from mating surfaces of instrument panel inlet and instrument panel.
B.
Install Inlet Doors and Associated Components (Refer to Figure 201). (1) Attach door to door inlet. Install pins, washers and cotter pins. (2) Attach door lever to door. Install pin, washer and cotter pin. (3) Locate and rivet door inlet (with new seal) to forward fuselage. (4) Attach control cable to door lever; secure with bolt, washer and nut. (5) Attach control cable to door inlet using clamp and hardware. (6) Prime mating surfaces of instrument panel and instrument panel inlet with SS4155 and apply thin bead (maximum thickness shall not exceed one- quarter inch) of sealant, RTV-102, to mating surfaces. (7) Attach instrument panel inlet to instrument panel using screws, washers and nuts. (8) Clean screen. (9) Install cleaned screen between duct and door inlet. (10) Secure duct at both ends using sta-straps.
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Cabin Air Ventilation System Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Cabin Air Ventilation System Figure 201 (Sheet 2)
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Cabin Air Ventilation System Figure 201 (Sheet 3)
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Cabin Air Ventilation System Figure 201 (Sheet 4)
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Cabin Air Ventilation System Figure 201 (Sheet 5)
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Cabin Air Ventilation System Figure 201 (Sheet 6)
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Cabin Air Ventilation System Figure 201 (Sheet 7)
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MODEL 208 MAINTENANCE MANUAL
Cabin Air Ventilation System Figure 201 (Sheet 8)
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4.
Cabin Air Ventilation System Valves and Controls Removal/Installation A.
Remove Valves (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7) (8) (9)
Wing-mounted valves control volume of air allowed to pass into cabin ventilation system.
Remove lower wing access panel 501AB/601AB to gain access to wing-mounted components. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Disconnect control cable at cabin air control valve elbow by loosening lock nut. Remove screws connecting outboard duct to elbow. Remove nuts, bolts and washers securing elbow to inboard duct. Remove nut, bolt and clamp securing control cable to elbow. Remove elbow from airplane. Remove nuts and screws securing butterfly to shaft. Remove pin which secures control arm to shaft. Disassemble and remove shaft from elbow. Note position of friction washers and nylon washers for later reassembly.
B.
Install Valves (Refer to Figure 201). (1) Install shaft to cabin air control valve elbow. Ensure friction washers and nylon washers are in correct position. (2) Attach butterfly valve to shaft using screws and nuts. (3) Attach control arm to shaft using pin. (4) Install cabin air control valve elbow to inboard duct using nuts, bolts, washers and shims. (5) Install outboard duct to elbow using screws. (6) Attach control cable to elbow using clamp, nut and bolt. (7) Attach end of control cable to arm using washer and new lock nut. (8) Check for freedom of movement and full travel by rotating cockpit control knob to extreme positions. Valve should fully open and close with control knob movement. (9) Install removed access panel 501AB/601AB . Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
C.
Remove Controls (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7)
D.
Controls are mounted overhead and are manually connected to wing-mounted fresh air ventilation valves using cables.
Unzip headliner and remove overhead console to gain access to controls. Loosen setscrew on control knob and remove knob from converter. Remove screws and nuts securing control cable to converter. Disconnect control cable from converter. Remove screws securing converter to cabin top structure. Unscrew wemac valve from wemac retainer. Remove hardware which secures wemac retainer to duct.
Install Controls (Refer to Figure 201). (1) On early models, position wemac retainer in channel and install duct using spacers, washers, nuts and screws. (2) On later models, position retainer in channel and install duct using spacers, washers, nuts and screws. (3) Screw wemac valve into wemac retainer. (4) Attach converter to cabin top structure using spacers, washers and nuts. (5) Attach control cable to converter using nuts, screws and brackets. (6) Attach control knob to converter and tighten set screw. (7) Rotate control knob and verify full travel of converter. (8) Replace overhead console and close headliner.
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MODEL 208 MAINTENANCE MANUAL
5.
Cabin Air Outlet Valve Removal/Installation (Model 208 only) A.
Remove Valve (Refer to Figure 201). (1) Unzip headliner to gain access to outlet valves. (2) Remove control knob from shaft by loosening set screw. (3) Remove screws securing cover and plate to air outlet plenum. Remove cover and plate. (4) Remove pin from index and disconnect index from shaft. (5) Disconnect transition from duct. (6) Disconnect coupling from plenum. NOTE: (7) (8)
B.
Drill rivet from butterfly, and disconnect shaft from plenum and butterfly. Remove butterfly from plenum.
Install Valve (Refer to Figure 201). (1) Install butterfly in plenum. (2) Install shaft in plenum and butterfly. Place new rivet in butterfly and shaft. (3) Connect transition and coupling to duct and plenum. (4) Clean sealant from mating surfaces of plenum, duct, transition and coupling with a cloth moistened in aliphatic naphtha. Cloth should be folded each time surfaces are wiped in order to present a clean area and avoid smearing the adhesive being removed. Wipe cleaned surfaces with a clean dry cloth before the naphtha evaporates. NOTE:
(5) (6) (7) (8) (9) 6.
7.
It is not necessary to disconnect transition from coupling.
A1186B clear adhesive may be used for resealing component parts. Mix eight (8) parts A1186B with one (1) part catalyst by volume. Shelf Life: One year, below 80°F, Work Life: Eight (8) hours at 75°F. Maximum Cure Time: 24 hours at 75°F. Accelerated Cure Time: 20 minutes at 200°F.
Apply A1186B sealant to mating surfaces of plenum, duct, transition, and coupling with caulking gun in one- quarter inch bead overlapping edges of mating parts. Connect index to shaft and install pin. Connect plate and cover to plenum using screws. Connect control knob to transition shaft and tighten setscrew. Close headliner.
Cabin Air Outlet Valves Removal/Installation (Model 208B Passenger) A.
Remove Cabin Air Outlet Valves (Refer to Figure 201). (1) Unscrew wemac valve from outlet assembly by turning counterclockwise.
B.
Install Cabin Air Outlet Valves (Refer to Figure 201). (1) Screw wemac valve into outlet assembly by turning clockwise.
Cabin Ventilation Fans and Switches Removal/Installation (Model 208 and 208B Passenger) A.
Remove Cabin Ventilation Fans and Switches (Refer to Figure 202 ). (1) Remove lower wing access panels 501AB and 601AB to gain access to wing mounted blower. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Disconnect electrical plug from fan. (3) Remove hardware which secures fan to inboard and outboard ducts. (4) Remove fan from wing area. (5) Unzip headliner and remove overhead console. (6) Disconnect electrical leads from switch and remove switch from overhead assembly.
B.
Install Cabin Ventilation Fans and Switches (Refer to Figure 202). (1) Attach fan to inboard and outboard ducts using screws, nut and washers. (2) Connect electrical connector to fan. (3) Reinstall lower wing access panels 501AB and 601AB . Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
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MODEL 208 MAINTENANCE MANUAL
Cabin Air Ventilation Fans Installation Figure 202 (Sheet 1)
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Cabin Air Ventilation Fans Installation Figure 202 (Sheet 2)
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Cabin Air Ventilation Fans Installation Figure 202 (Sheet 3)
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Cabin Air Ventilation Fans Installation Figure 202 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (4) (5)
Install switch to overhead mount using screws. Connect electrical leads to switch and ensure switch operates properly. NOTE:
(6)
Switch is actuated by a lug on converter. Lug contacts the switch when wing- mounted butterfly valve is approximately three-quarters open. When lug contacts switch, fan should begin operating.
Install overhead console and zip headliner.
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MODEL 208 MAINTENANCE MANUAL FRESH AIR DISTRIBUTION - ADJUSTMENT/TEST 1.
General A.
2.
This section has the inspections and checks necessary to keep the fresh air distribution system in a serviceable condition.
Ventilation Outlet and Controls Operational Check A.
General (1) This section gives the information needed to complete the operational check of the ventilation outlet and controls.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Ventilation Outlet and Controls Operational Check. (1) Examine the control and cable attachment for condition and security of installation. (2) Examine the wiring at the switch for evidence of damage and security of the connections. (3) Examine the WEMAC outlet for condition, security, and correct operation. (a) Make sure that the WEMAC does not stick in the closed position. (4) Examine the clamps, hoses, valves, inlet screens, and ventilation system for condition, evidence of moisture intrusion, and security of the components. (5) Examine the internal hose wire for condition. (6) Examine all vent outlets for condition, security, and correct operation of the ON/OFF control and outlet vanes. (7) Visually examine the accessible air ducts for condition and security.
E.
Restore Access (1) None
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MODEL 208 MAINTENANCE MANUAL HEATING AND DEFROSTING AIR DISTRIBUTION - DESCRIPTION AND OPERATION 1.
2.
General A.
This section is concerned with those components which distribute heated air to the heating and defrosting outlets. It does not include those components and sub systems which are used to produce or control temperature of heated air.
B.
For a description of how heated air is produced, refer to Compressor Bleed Air Heater - Description and Operation.
C.
For a description of how heated air is temperature controlled, refer to Temperature Control Description and Operation.
Description and Operation A.
Heating/defrost system consists of heater valve, defroster valve, heater valve control, defroster valve control, defroster nozzles in the glareshield, plenums on left and right sidewalls near floor level, forward cabin heater ducts on aft side of firewall ducts, valves and clamps as required to connect system components.
B.
System is controlled by two push-pull knobs on the cabin heat control panel. These knobs control volume of air allowed to pass into and through various heating and defrosting ducts located throughout airplane.
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MODEL 208 MAINTENANCE MANUAL HEATING AND DEFROSTING AIR DISTRIBUTION - TROUBLESHOOTING 1.
General A.
Troubleshooting of heating/defrost air distribution system should be performed anytime output flow falls below normal parameters.
B.
A troubleshooting chart has been prepared to aid maintenance technician in system troubleshooting. Refer to Figure 101.
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Heater and Defroster Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL HEATING AND DEFROSTING AIR DISTRIBUTION - MAINTENANCE PRACTICES 1.
General A.
2.
This section deals with maintenance to heating and defrosting air distribution system. Maintenance is typically limited to removal/installation of components.
Heater Valve Removal/Installation A.
Remove Heater Valve (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7) (8) (9)
Heater valve is located on aft side of firewall between pilot and copilot rudder pedals.
Remove clamps securing left and right ducts to wye. Remove clamp securing upper duct to heater valve. Remove ducts from heater valve and wye. Remove hardware securing heater control valve cable to arm of flapper shaft. Remove brackets securing heater control valve cable to heater valve and to firewall. From forward side of firewall, remove screws securing heater valve to firewall. Remove heater valve from airplane and place on bench. On bench, remove pin from flapper shaft arm. Carefully remove arm from flapper shaft. Remove nuts, washers and screws securing flapper to flapper shaft. NOTE:
B.
Check flapper and grommets for condition. Replace seal and grommets if worn or damaged.
Install Heater Valve (Refer to Figure 201). (1) Install flapper to shaft using nuts, washers and screws. (2) Install flapper shaft arm on flapper shaft and secure using pin. (3) Clean existing sealer from heater valve flanges and mating surfaces. (4) Apply GC-1900 sealant (or equivalent) to heater valve flanges. NOTE: (5) (6) (7)
Cure time at 77°F and 50 percent humidity is 72 hours. Accelerated cure time at 150°F is 4 hours. Accelerated cure time at 300°F is 15 minutes.
Attach heater valve to firewall using screws. Attach control cable to heater valve and firewall using applicable brackets. Pull cockpit heater valve control knob open approximately one-eighth inch from closed position while applying sufficient pressure on flapper shaft arm to secure flapper closed.
WARNING: Minimum installed bend radii for wire supported heater ducts in plane of bend, measured from wall of duct, shall be one-third diameter of maximum duct dimension. NOTE:
(8) 3.
When cutting heating system ducts to length, support wire should be cut back far enough to bend back (minimum bend radius one-eighth inch) under clamp and protrude one-quarter inch. Do not break the bond between wire and fabric. Before tightening clamps, make sure there is not twist or torque on the duct.
Connect all ducts to heater valve and heater valve body. Secure all ducts using clamps.
Defroster Valve Removal/Installation A.
Remove Defroster Valve (Refer to Figure 201). NOTE: (1) (2)
Defroster valve is located on aft side of firewall above heater valve.
Loosen clamps at wye and at defroster nozzles. Remove ducting between wye and defroster nozzles.
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Heating and Defrosting System Installation Figure 201 (Sheet 1)
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Heating and Defrosting System Installation Figure 201 (Sheet 2)
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Heating and Defrosting System Installation Figure 201 (Sheet 3)
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Heating and Defrosting System Installation Figure 201 (Sheet 4)
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Heating and Defrosting System Installation Figure 201 (Sheet 5)
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Heating and Defrosting System Installation Figure 201 (Sheet 6)
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6)
Remove screws securing left and right ducts to defroster valve. Remove left and right ducts. Loosen clamps and remove heater to defroster duct. Remove hardware and brackets securing control cable to primary arm. Remove nut and washer on secondary arm. Detach control cable from clamp and remove clamp from secondary arm. (7) Remove screws securing defroster valve to firewall and place defroster valve on bench. (8) Loosen and remove bar connected to primary and secondary arms. (9) Remove pins from primary and secondary arms. (10) Detach arms from respective shafts. (11) Remove nuts, washers and screws securing primary flapper to primary shaft. Remove primary flapper from defroster valve body. (12) Remove nuts, washers and screws securing secondary flapper to secondary shaft. Remove secondary flapper from defroster valve body. NOTE: B.
Install Defroster Valve (Refer to Figure 201). (1) Attach primary flapper and secondary flapper to primary shaft and secondary shaft. Align holes and secure flappers to shafts using screws, washers and nuts. (2) Attach primary arm and secondary arm to respective shafts. Secure to shafts using pins. (3) Attach bar between primary and secondary arms using washers and nuts. NOTE: (4) (5)
(6) (7) (8) (9) (10) (11)
5.
Let bar extend through clamp approximately three-eighths inch, and tighten nut and washers on secondary arm. Leave nut on primary arm loose.
Secure primary flapper and secondary flapper at 45 degree ends. Tighten nut on primary arm and release flappers. Apply GC-1900 sealant (or equivalent) to heater valve flanges. NOTE:
4.
Check flappers and grommets for condition. Replace seals and grommets if worn or damaged.
Cure time at 77°F and 50 percent humidity is 72 hours. Accelerated cure time at 150°F is 4 hours. Accelerated cure time at 300°F is 15 minutes.
Locate heater valve on firewall and secure to firewall using screws. Attach ducts between wye and defroster nozzles. Attach defroster to heater duct. Secure all ducts using clamps. Attach control cable to defroster valve and firewall using clamps and hardware. Attach end of control cable to primary arm and finger tighten nut. Pull defroster valve control knob open approximately one-eighth inch from closed position. Apply sufficient force on primary arm to keep primary and secondary flappers closed. Tighten nut on primary arm.
Defroster Nozzle Removal/Installation A.
Remove Defroster Nozzle (Refer to Figure 201). (1) Loosen clamps and remove duct between defroster nozzles and wye. (2) Remove bolts and spacers from insulation blanket. Detach nozzle from glareshield.
B.
Install Defroster Nozzle (Refer to Figure 201). (1) Attach defroster nozzle to glareshield using bolts and spacers. (2) Install duct between nozzle and wye. Secure duct using clamps.
Return Air Duct Removal/Installation A.
Remove Return Air Duct (Refer to Figure 201). (1) Remove bolts and washers to allow return air mount to be removed from airplane. (2) Remove screws and spacers securing flange and cap to return air mount.
B.
Install Return Air Duct (Refer to Figure 201). (1) Attach flange and spacers to cap using screws, spacers and nuts.
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Attach cap and flange to return air mount using screws. Locate and align return air mount to firewall. Secure using screws.
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MODEL 208 MAINTENANCE MANUAL AVIONICS COOLING - DESCRIPTION AND OPERATION 1.
General A.
On airplanes that do not have the Garmin G1000, avionics cooling is provided by a blower motor mounted behind the instrument panel. Flexible ducts are mounted to the outlet end of the blower motor and provide dedicated cooling lines for various avionics components. For removal/installation procedures, refer to the Model 208 Avionic Installations Service/Parts Manual
B.
Garmin Display Units (GDU) each have a cooling fan that blows air at the aft side of the display. For GDU cooling fan removal and installation, refer to Garmin Display Unit (GDU) Cooling Fan - Removal/ Installation.
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MODEL 208 MAINTENANCE MANUAL AVIONICS COOLING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fresh avionics cooling system in a serviceable condition.
Task 21-24-00-710 2.
Avionics Cooling Fan Operational Check A.
General (1) This section gives the information needed to complete the operational check of the avionics cooling system. NOTE:
B.
Special Tools (1) None
C.
Access (1) For (a) (b) (2) For (a)
The operational check for airplanes with and without the Garmin G1000 are included in this task.
airplanes without the G1000. Remove the screws that attach the center console to the floor. Tilt the console over toward the copilot side to get access to the blower motor. airplanes with the G1000. Remove the GDU to get access to the GDU fan. Refer to Chapter 34, Garmin Display Unit (GDU) - Maintenance Practices.
D.
Do the Avionics Cooling Operational Check for Airplanes Without the Garmin G1000. (1) Examine the blower for security, condition, and the connection of the duct hoses at the blower and at the radio racks. (2) Examine the wiring at the blower motor for condition and security. (3) Set the MASTER switch to the ON position. (4) Put the AVIONICS 1 and the AVIONICS 2 switch to the ON position (5) Make sure that the blower motor operates correctly. (6) Put the AVIONICS 1 and the AVIONICS 2 switch to the OFF position (7) Set the MASTER switch to the OFF position.
E.
Do the Avionics Cooling Operational Check for Airplanes With the Garmin G1000. NOTE: (1) (2) (3) (4) (5) (6) (7)
The operational check for the different GDU fans is typical.
Examine the GDU fan for security and condition. Examine the wiring at the GDU fan for condition and security. Set the MASTER switch to the ON position. Put the AVIONICS 1 and the AVIONICS 2 switch to the ON position Make sure that the GDU fan operates in the correct direction. Put the AVIONICS 1 and the AVIONICS 2 switch to the OFF position Set the MASTER switch to the OFF position.
F.
Restore Access (1) For airplanes without the G1000. (a) Put the center console in its position and on the floor. (b) Install the screws that attach the center console to the floor. (2) For airplanes with the G1000. (a) Install the GDU. Refer to Chapter 34, Garmin Display Unit (GDU) - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL CENTER CONSOLE AVIONICS COOLING - REMOVAL/INSTALLATION 1.
General A.
2.
A blower motor with flexible ducts is located in the lower portion of the center console. This motor provides ducted air to the transponder and autopilot computer.
Center Console Avionics Cooling Removal/Installation A.
Remove Blower Motor (Refer to Figure 401). (1) Disengage circuit breaker PED AVN FAN on left circuit breaker panel. (2) Remove crew seats from airplane. Refer to Chapter 25, Flight Compartment Maintenance Practices. (3) Remove and retain screws securing center console to floor. (4) Tilt console over toward copilot side to gain access to blower motor. (5) Disconnect electrical connector from blower motor. (6) Loosen clamps and remove duct hose from blower motor. (7) Remove screws securing blower motor. (8) Remove blower motor from airplane. (9) Tilt center console back to its original position.
B.
Install Blower Motor (Refer to Figure 401). (1) Ensure center console is tilted toward copilots side. (2) Position blower motor and secure with screws. (3) Attach duct hose to blower motor and secure with clamps. (4) Connect electrical connector to blower motor. (5) Position center console upright and secure to floor with retained screws. (6) Install crew seats. Refer to Section 25 - Flight Compartment Maintenance Practices. (7) Engage circuit breaker PED AVN FAN on left circuit breaker panel..
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Center Console Avionics Cooling Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GARMIN DISPLAY UNIT (GDU) COOLING FAN - REMOVAL/INSTALLATION 1.
General A.
2.
3.
The display cooling fans are found behind the Garmin Display Units (GDU). Maintenance on the system is only to remove and install the cooling fans.
Garmin Display Unit (GDU) Fan Removal/Installation A.
Remove the GDU Fan (Refer to Figure 401). (1) Make sure that the MASTER and AVIONICS switches are in the OFF position. (2) Remove the GDU. Refer to Chapter 34, Garmin Display Unit (GDU) - Maintenance Practices. (3) Record the fan airflow direction. (4) Remove the screws and nuts that attach the fan to the fan bracket. (5) Disconnect the electrical connector from the avionics fan. (6) Remove the fan from the airplane.
B.
Install the GDU Fan (Refer to Figure 401). (1) Connect the electrical connector to the GDU fan. (2) Make sure that the fan airflow is towards the GDU. (3) Install the screws and nuts that attach the fan to the fan bracket. (4) Do a fan operation check. Refer to Garmin Display Unit (GDU) Fan Operational Check (5) Set the MASTER switch and the AVIONICS switch in the OFF positions. (6) Install the GDU. Refer to Chapter 34, Garmin Display Unit (GDU) - Maintenance Practices.
Garmin Display Unit (GDU) Fan Operational Check A.
Garmin Display Unit (GDU) Fan Operational Check (Refer to Figure 401). (1) Remove the GDU. Refer to Garmin Display Unit (GDU) - Maintenance Practices. (2) Put the MASTER and AVIONICS switches in the ON position. (3) Observe and make sure that all GDU fans operate in the correct direction. (4) Install the GDU. Refer to Garmin Display Unit (GDU) - Maintenance Practices.
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Garmin Display Unit (GDU) Cooling Fan Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL COMPRESSOR BLEED AIR HEATER - DESCRIPTION AND OPERATION 1.
2.
General A.
The temperature and volume of air flow to the cabin is regulated by the cabin heating, ventilating and defrosting system. In the heating system, hot compressor outlet air and interstage compressor bleed air (P2.5 and P3 bleed air extracted from the engine) provide the source of heat. This heat is routed via a gate valve and a mixer/muffler to the cabin air distribution system. Controls are provided to direct the heated air to the forward or aft portions of the cabin for heating and to the windshield for defrosting.
B.
Two configurations of compressor bleed air heater may be used on the airplane. These configurations are described in Description and Operation below.
C.
Schematics and flow diagrams are provided to aid the maintenance technician in system understanding. These schematics and flow diagrams are applicable to both configurations of compressor bleed air heaters. Refer to Figure 1 for a compressor bleed air heater schematic. Refer to Figure 2 for a heating and defrosting flow diagram.
Description and Operation (Airplanes 20800001 Thru 20800179 and 208B0001 Thru 208B0209) A.
Component Descriptions are as follows: (1) Temperature Limiter Switch. (a) The temperature limiter switch is installed in the cabin heat firewall shutoff valve. The switch will open and de-energize the gate valve solenoid if bleed air temperature exceeds 210°F, +10 or -10°F. The switch will close when the temperature reduces to 196°F. (2) Flow Control Valve and Solenoid Valve Assembly. (a) The flow control valve acts as a variable (low) pressure regulator. It consists of a pressure operated poppet valve with a solenoid operated control pressure valve. Control pressure from the temperature control valve causes the spring-loaded poppet valve to open as control pressure increases. A diaphragm separates the control pressure cavity from P3 turbine bleed air which acts in conjunction with the spring tending to close the valve. (3) Regulator and Gate Valve Subassembly. (a) The regulator and gate valve subassembly consists of a pressure regulator, a gate valve assembly and an interconnecting control pressure line. (4) Pressure Regulator Assembly. (a) The function of the poppet valve regulator is to reduce P3 compressor outlet bleed air control pressure to 18.0 PSIG, +1 or -1 PSIG. A relief valve is provided to prevent excess downstream pressure in the event of regulator failure. The relief valve is set to open at 22.0 PSIG, +1 or -1 PSIG, and to reseat at 20.0 PSIG minimum. A small, screened opening below the inlet port allows the unpressurized side of the poppet valve rolling diaphragm to vent to atmosphere. A tapped port on the downstream side of the pressure regulator provides regulated air via the control pressure line to the gate solenoid valve. (5) Temperature Control Valve Assembly. (a) The needle control valve regulates the control pressure at the gate valve. The control valve has approximately 270 degree rotation. Clockwise rotation closes the valve, increasing control pressure up to 17.0 PSIG and increasing heat. Counterclockwise rotation opens the valve, decreasing pressure and decreasing heat. NOTE: (6)
(7)
The valve does not close completely. A bleed of 800 to 1000 CC/minute flow is allowed to vent to atmosphere to accommodate hysteresis of the gate valve.
Air Ejector Assembly (Mixer/Muffler). (a) The air ejector assembly consists of a muffler and bleed air ejector. The assembly combines regulated P3 air with either P2.5 air or cabin recirculation air and routes this flow to the cabin. The ejector configuration ensures complete mixing of the air sources, thereby reducing the P3 primary flow to a usable cabin heat source temperature. The cavity between the inner perforated tube and the outer shell is insulated with Nomex and fiberglass to attenuate noise and to act as a muffler. Valve Assembly - Air Diverter (Mixing Air Valve).
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Compressor Bleed Air Heater Schematic Figure 1 (Sheet 1)
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Heating and Defrosting Flow Diagram Figure 2 (Sheet 1)
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WARNING: This position is not to be used in flight. In the FLT PUSH position, P2.5 air is dumped to atmosphere. Cabin air is drawn into the muffler to mix with regulated P3 air. The FLT PUSH position may be used on the ground if P2.5 heat augmentation is not required. (a)
(8)
(9)
3.
The diverter valve (mixing air valve) has two doors mechanically linked together, with both doors operated by a single MIXING AIR control. In the GRD PULL position, P2.5 interstage compressor bleed air provides heat at power settings below 89 percent Ng and is used to augment the regulated P3 compressor outlet bleed air on the ground in cold temperatures. In the FLT PUSH position, P2.5 interstage compressor bleed air is dumped to atmosphere. Cabin air is drawn into the muffler to mix with regulated P3 compressor outlet bleed air. The FLT PUSH position may be used on the ground if interstage compressor bleed air heat augmentation is not required. Heater Valve, Firewall Shutoff. (a) The two firewall shutoff valves are operated by a single control located on the lower right side of the pilot’s control pedestal. With the control pushed IN, both valves are open. The lower valve controls bleed air supply from the mixer/muffler. The temperature limiter switch is installed on the right side of the valve and just above it is a cam operated microswitch in series with the temperature limiter switch. The microswitch is closed when the shutoff valve door is open. The upper valve provides the cabin air return to the mixing valve. Controls. (a) The two firewall shutoff valves are operated by a single push/pull control knob, CABIN HEAT FIREWALL SHUTOFF PULL OFF, located on a panel on the lower right side of the pilot’s control pedestal. When the knob is placed in the OFF position, it closes firewall doors and flow control valve. With the shutoff knob in this position, heated bleed air from the engine is shut out of the heating system, and heated air from the heating system is kept out of the cabin.
Description and Operation (Airplanes 20800180 and On, 208B0210 and On, and Airplanes Incorporating CAB90-9) A.
Components of the system are as follows: (1) Temperature Limiter Switch. (a) The temperature limiter switch is installed in the cabin heat firewall shutoff valve. The switch will open and de-energize the flow control valve solenoid if the bleed air temperature exceeds 210°F, +10° or - 10°F, preventing P3 air flow into the mixer/muffler and cabin. The switch will close when the temperature reduces to 196°F, +21° or - 21°F. (2) Flow Control Valve and Solenoid Valve Assembly. (a) The flow control valve acts as a variable (low) pressure regulator. It consists of a pressure operated poppet valve with a solenoid operated control pressure valve. Control pressure from the temperature control valve causes the spring-loaded poppet valve to open as control pressure increases. A diaphragm separates the control pressure cavity from P3 turbine bleed air which acts in conjunction with the spring tending to close the valve. (3) Pressure Regulator Assembly. (a) The function of the poppet valve regulator is to regulate P3 bleed air pressure between 17.0 to 20.0 PSIG for instrument vacuum and deice systems operation. It is functionally independent of the heater system. A relief valve is provided to prevent excess downstream pressure in the event of regulator failure to protect the deice boots from over inflation. The relief valve is set to open at 22.0 PSIG, +1.0 or -1.0 PSIG. (4) Temperature Control Valve Assembly. (a) The instrument panel mounted valve is an adjustable relief valve which varies flow control valve dome control pressure, thereby changing flow through the flow control valve and cabin air temperature. The control valve has approximately 270 degree rotation. Clockwise rotation closes the valve, increasing control pressure and increasing heat; counterclockwise rotation opens the valve, decreasing pressure and decreasing heat.
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MODEL 208 MAINTENANCE MANUAL (5)
(6)
Air Ejector Assembly (Mixer/Muffler). (a) The air ejector assembly consists of a muffler and bleed air ejector. The assembly combines regulated P3 air with either P2.5 air or cabin recirculation air, and routes this flow to the cabin. The ejector configuration ensures complete mixing of the air sources, thereby cooling the P3 primary flow to a usable cabin heat source temperature. The cavity between the inner perforated tube and the outer shell is insulated with Nomex and fiberglass to attenuate noise and to act as a muffler. Valve Assembly - Air Diverter (Mixing Air Valve).
WARNING: This position is not to be used in flight. In the FLT PUSH position, P2.5 air is dumped to atmosphere. Cabin air is drawn into the muffler to mix with regulated P3 air. The FLT PUSH position may be used on the ground if P2.5 heat augmentation is not required. (a)
(7)
(8)
(9)
The diverter valve (mixing air valve) has two doors mechanically linked together, which are both operated by a single MIXING AIR control. In the GRD PULL position, P2.5 air provides heat at power settings below 89 percent Ng and is used to augment the regulated P3 heat on the ground in cold temperatures. Heater Valve - Firewall Shutoff. (a) The two firewall shutoff valves are operated by a single control located on the lower right side of the pilot's control pedestal. With the control pushed IN, both valves are open. The lower valve controls bleed air supply from the mixer/muffler. The temperature limiter switch is installed on the right side of the valve; just above it is a cam-operated microswitch in series with the temperature limiter switch. The microswitch is closed when the shutoff valve door is OPEN. The upper valve provides the cabin air return to the mixing valve. Microswitch - P3 Flow Shutoff. (a) The microswitch is installed in the cabin heat firewall shutoff valve. The switch will open and de-energize the flow control valve solenoid when the firewall shutoff valves are closed, preventing P3 air flow to the mixer/muffler and cabin. Controls. (a) The cabin heat control panel is located at the lower edge of the instrument panel, to the right of airplane centerline. Individual controls are described from left to right. Temperature Control. Rotary control needle valve for temperature control. Rotate 1 clockwise to increase flow of heated air; rotate counterclockwise to decrease flow. Bleed Air Heat Switch. ON/OFF switch controls electrical power to the flow control 2 valve solenoid. (P3 air flow on/off control.) Mixing Air Push/Pull Control. Controls mixing air valve. (Mixing P2.5 air with P3 air 3 not to be used in flight.) Aft Cabin/Fwd Cabin Push/Pull Control. Controls cabin heat selector valve and 4 diverts heat to forward or aft cabin. Defrost/Fwd Cabin Push/Pull Control. Controls air selector valve to divert heat for 5 defrost or forward cabin. Cabin Heat Firewall Shutoff Push/Pull Control. Located on the lower right side of the 6 pilot's control pedestal. Pull to isolate all nacelle bleed air components and flow to and from the cabin. Bleed Air Heat Circuit Breaker. Located on the left side wall circuit breaker panel. 7 Source of electrical power for the flow control gate valve solenoid.
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MODEL 208 MAINTENANCE MANUAL COMPRESSOR BLEED AIR HEATER - TROUBLESHOOTING 1.
General A.
Troubleshooting of the compressor bleed air heater and/or flow control valve should be performed anytime output flow falls below normal parameters.
B.
Troubleshooting charts have been prepared to aid the maintenance technician in system troubleshooting. Refer to Figure 101 for compressor bleed air troubleshooting diagram. Refer to Figure 102 for flow control valve troubleshooting diagram. Refer to Figure 103 and Figure 104 for cabin heating and windshield defrosting troubleshooting diagrams.
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Compressor Bleed Air Heater Troubleshooting Diagram Figure 101 (Sheet 1)
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Flow Control Valve Troubleshooting Chart Figure 102 (Sheet 1)
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Cabin Heating and Windshield Defrosting Troubleshooting Diagram Figure 103 (Sheet 1)
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Cabin Heating and Windshield Defrosting Troubleshooting Diagram Figure 103 (Sheet 2)
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Cabin Heating and Windshield Defrosting Troubleshooting Diagram Figure 103 (Sheet 3)
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Cabin Heating and Windshield Defrosting Troubleshooting Diagram Figure 103 (Sheet 4)
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Cabin Heating and Windshield Defrosting Troubleshooting Diagram Figure 104 (Sheet 1)
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Cabin Heating and Windshield Defrosting Troubleshooting Diagram Figure 104 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL COMPRESSOR BLEED AIR HEATER - MAINTENANCE PRACTICES 1.
General A.
2.
This section covers maintenance practices and testing procedures for components of the compressor bleed air system.
Compressor Bleed Air Heater Components Removal/Installation A.
Remove Compressor Bleed Air Heater Components (Refer to Figure 201). NOTE: (1)
Components are removed in the following procedures. If individual components need to be disassembled, refer to Individual Component Disassembly/Assembly below.
Remove connector tube. (a) Cut safety wire and remove bolts securing bottom end of connector tube to mixer/muffler. Discard gasket. (b) Remove bolts and washers securing top end of connector tube to flow control valve. Discard gasket and remove connector tube from airplane. NOTE:
(2) (3)
(4)
(5)
(6)
On some airplanes the connector tube may be secured to engine mount with a clamp. This clamp must be removed before connector tube can come out.
Remove pressure regulator. (Refer to Chapter 36, Pneumatic Distribution - Maintenance Practices). Remove flow control valve. (a) Disconnect pneumatic line from flow control valve. (b) Cut safety wire and disconnect electrical connector from flow control valve. (c) Remove V-type clamp (or tube nut) from compressor duct. (d) Remove flow control valve with cast tee fitting from airplane. Remove compressor duct. (a) Loosen and remove clamp securing compressor duct to top of engine. (b) At top of engine, remove bolts attaching flange seal to compressor section cover. Discard flange seal. (c) With flange seal removed, cut safety wire and remove screws/washers securing compressor duct to compressor section cover. Remove compressor duct and gasket from compressor section cover. Discard gasket. Remove mixer/muffler. (a) Remove clamps securing flexible duct between mixer/muffler and cabin heating air valve. (b) Remove clamp between diverter valve and mixer/muffler. (c) Detach propeller control cable and clamp from mixer/muffler. (d) Remove nuts, bolts and washers securing mixer/muffler mounting bracket and clamp to engine mount. (e) Detach mixer/muffler with mounting bracket and clamp from airplane. Disconnect diverter valve. (a) Disconnect control cable at diverter valve lower arm. NOTE:
It is not necessary to disconnect turnbuckle between upper and lower diverter valve arms.
Loosen clamps securing flexible duct to diverter valve and remove flexible duct from diverter valve. (c) Remove hardware securing diverter valve to engine mount bracket. (d) Remove diverter valve from airplane. Remove cabin return air valve and cabin heating air valve from firewall. (a) Disconnect electrical connector from cabin heating air valve. (b) Remove interconnecting rod between cabin heating air valve and cabin return air valve. (c) Disconnect control cable from cabin heating air valve.
(b)
(7)
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Compressor Bleed Air Heater Installation Figure 201 (Sheet 1)
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Compressor Bleed Air Heater Installation Figure 201 (Sheet 2)
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Compressor Bleed Air Heater Installation Figure 201 (Sheet 3)
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Compressor Bleed Air Heater Installation Figure 201 (Sheet 4)
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Compressor Bleed Air Heater Installation Figure 201 (Sheet 5)
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Compressor Bleed Air Heater Installation Figure 201 (Sheet 6)
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MODEL 208 MAINTENANCE MANUAL (d) (e) B.
3.
Remove bracket, clamp and spacer securing control cable to lower corner of cabin heating air valve. Remove valves from firewall.
Install Compressor Bleed Air Heater Components (Refer to Figure 201). (1) Install cabin return air valve and cabin heating air valve to firewall. (a) Clean off existing sealant from the firewall and from valves. (b) Apply a 0.125 inch bead of RTV 103 (or equivalent) sealant to mating surfaces of firewall and valves. (c) Attach valves to firewall. (d) Reattach connecting rod between valves. (e) Attach control cable to corner of cabin heating air valve using bracket, clamp and spacer. (f) Rig control cable by pulling cabin heating air control knob open 0.125 inch from fully closed position. Hold cabin heating and return air doors fully closed, then tighten end of cable at heating valve using locknut. (2) Install diverter valve. (a) Install diverter valve to engine mount bracket. (b) Connect flexible duct between diverter valve and cabin return air valve. Secure duct using clamps. (c) Attach control cable to diverter valve. Rig. Pull diverter valve control knob open 0.125 inch from fully closed position. Hold 1 diverter door and recirculating door fully closed, then tighten end of control cable at diverter valve lower arm using locknut. (3) Install mixer/muffler. (a) Place mixer/muffler in engine compartment and connect mixer/muffler to diverter valve using clamp. (b) Secure mixer/muffler to engine mounting bracket using clamp and hardware. (c) Install duct between mixer/muffler and cabin heating air valve. Secure using clamp. (d) Reattach propeller control cable to mixer/muffler using clamp and hardware as required. (4) Install connector tube. (a) Using new gasket, attach bottom end of connector tube to mixer/muffler. (b) Safety wire bolts. (5) Install compressor duct. (a) Attach compressor duct to compressor section cover using new gasket and new flange seal. Safety wire bolts attaching compressor duct to compressor section cover. (b) Secure compressor duct to top of engine using clamp. (6) Install flow control valve. (a) Using new gasket, attach flow control valve to top of connector tube. (b) Safety wire bolts. (c) Connect flow control valve to compressor duct using V-type clamp or tube nut. Torque V-type clamp to value stamped on clamp. (d) Attach pneumatic line to side of flow control valve. (e) Connect electrical connector to flow control valve. (7) Install pressure regulator (Refer to Chapter 36, Pneumatic Distribution - Maintenance Practices).
Individual Component Disassembly/Assembly A.
Disassemble Diverter Valve (Refer to Figure 202). (1) Remove rivets from end cap. Detach end cap from valve body. (2) Remove hardware securing turnbuckle and connectors to upper and lower arms. (3) Remove hinge rivets from diverter door. Detach door from diverter wall. (4) Remove pins and detach arms and from shaft adapters. NOTE: (5) (6)
If diverter door seal or hinge is worn or damaged, replace diverter door and hinge.
Remove rivets from wall and detach wall from diverter valve body. Remove seals from shaft adapters. Discard seals.
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Diverter Valve Installation Figure 202 (Sheet 1)
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Remove rivets and detach spacer, hinge, hinge shaft, and recirculating door from diverter valve body. NOTE:
B.
If recirculating door or hinge is worn or damaged, it should be replaced.
Assemble Diverter Valve (Refer to Figure 202 ). (1) Rivet spacer and recirculating door hinge to diverter valve body. (2) Rivet diverter wall to diverter valve body. NOTE:
(3)
Rivet diverter door hinge to wall. NOTE:
(4)
Before assembling end cap on body, apply a 0.15 inch wide bead of Silastic E, approximately 0.40 inch from end of body. Let sealant cure four hours at 77°F before operating heating system.
Slide end cap into body, line up with diverter door and rivet to body. Install new seals on shaft adapters. NOTE:
(7) (8)
Check that diverter door is in line and fits flat on opening in wall. Ensure this area is free of sealant.
Clean existing sealant from mating surfaces of end cap and diverter valve body. NOTE:
(5) (6)
Before assembling diverter wall to body, clean all existing sealant from mating surfaces. Seal sides and ends of diverter wall with Silastic E. Clean excessive sealant from diverter wall in areas covered by diverter door. Let sealant cure four hours at 77°F before proceeding to the next step.
Before assembling seals, clean mating surfaces of seal and body. Prime with DC1200 and apply RTV-157 silicone sealant to seals.
Attach upper and lower arms to shaft adapters. Install and safety wire pins. Attach turnbuckle and connectors to upper and lower arms using hardware and new cotter pins.
C.
Disassemble Cabin Return Air and Cabin Heating Air Valves (Refer to Figure 203). (1) Remove pins from shaft adapters. Detach arms, seals and switch actuator. (2) Remove four rivets from base. Detach body from base. (3) Remove three rivets from base. Detach spacer, spring and hinge assembly from base.
D.
Assemble Cabin Return Air and Cabin Heating Air Valves (Refer to Figure 203). (1) Attach spacer, spring, and hinge assembly to base. Install three rivets in base. NOTE:
(2) (3)
Slip shaft adapter through hole in valve body. Attach body to base and install four rivets. Install new seals on shaft adapter. NOTE:
(4)
During installation, compress spring 90 degrees and slip over end of hinge shaft. Thread short end of spring through mounting hole in base, and bend short end to hold spring in place when door is actuated. Before installation, clean mating surfaces of valve bodies and base and apply a 0.125 inch bead of RTV-102, RTV-103, or equivalent to mating surfaces.
Before installing seals, clean mating surfaces of seal and body. Prime with DC-1200 and apply RTV-157 silicone sealant to seals.
Attach arms and switch actuator to shaft adapters. Install and safety wire pins.
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Cabin Return Air and Cabin Heating Air Valves Installation Figure 203 (Sheet 1)
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4.
Component Cleaning/Servicing A.
Flow Control Valve Cleaning Procedures (Refer to Figure 204). NOTE:
(1) (2) (3) (4) (5)
Perform this procedure when the following conditions exist: Flow control valve output is too high, too low, erratic, or sluggish; and solenoid operation/dome pressure variation with power changes and temperature control knob changes are normal.
Remove flow control valve from airplane. Remove screws holding lower housing assembly to housing. Remove poppet valve, spring retainer and spring. Remove and discard O-ring. Remove poppet screw holding bal-seal on poppet valve. Remove and discard bal-seal. Thoroughly clean poppet parts and interior sliding surface of lower housing assembly with Stoddard solvent (P-D-680 Type III), Methyl n-Propyl Ketone or isopropyl alcohol. Use a soft, nonmetallic bristle brush if necessary. Rinse with clean water and dry with a clean, lint free cloth and/or clean shop air. NOTE:
Use the minimum necessary amount of scrubbing or wiping to avoid removing dry film lubricant finish. Presoaking parts in solvent should aid in softening baked on deposits.
(6)
Apply 113A10010 Silicone Lubricant to new bal-seal. Ensure open edge of bal-seal is toward top of poppet valve and install bal-seal on poppet valve. Retain with poppet screw. Torque poppet screw to 85 inch-pounds, +5 or -5 inch-pounds. (7) Lubricate new O-ring with 113A10010 Silicone Lubricant and install on lower housing assembly. (8) Position spring, spring retainer and poppet assembly in lower housing. (9) Apply 26316503 Anti-Seize compound to screws and install lower housing assembly to regulator housing using screws. Torque to 32 inch-pounds, +2 or -2 inch-pounds. (10) Reinstall flow control valve on airplane. (11) Perform an engine run and check flow control valve for proper operation. B.
Flow Control Valve Solenoid Replacement (Refer to Figure 204). NOTE:
(1)
Use this procedure when solenoid is stuck open or closed, as shown by absence of an audible click when the cabin heat switch is operated, or is electrically defective, (open or short through pins combined with the absence of the click. Airplane circuitry provides 28.0 VDC with cabin heat switch ON, and 0.0 VDC with switch OFF.
Remove electrical connector from solenoid. NOTE:
(2) (3) (4) (5) (6) (7)
C.
It is not necessary to remove valve from airplane to perform the following procedure.
Remove screws attaching solenoid to housing. Remove and discard O-ring. Inspect mating surface of valve body. Clean as required with Stoddard solvent, Methyl n-Propyl Ketone or isopropyl alcohol. Dry with a clean, lint free cloth. Install new O-ring on housing. Position new solenoid on housing and attach with screws. Connect electrical connector and safety wire. Check solenoid operation by noting an audible click each time heater switch is turned on or off (or power is applied or removed using a jumper). Repeat approximately twenty times to check for intermittent sticking. If satisfactory, check flow control valve during engine run.
Temperature Control Valve Knob Replacement (Refer to Figure 205). (1) Remove knob by loosening two set screw attaching knob to shaft.
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Flow Control Valve Installation Figure 204 (Sheet 1)
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Temperature Control Valve Installation Figure 205 (Sheet 1)
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Install knob on shaft. Adjust knob so it is pointing downward when shaft is rotated fully counterclockwise (clocking may be varied based on operator preference) and tighten setscrews. NOTE:
5.
Friction may be varied by adjusting setscrews in valve body to provide desired amount of friction.
Cabin Heat Functional Test (Airplanes 20800001 Thru 20800179 and 208B0001 Thru 208B0209 Except Airplanes Incorporating CAB90-9) A.
Functional Test Procedures. (1) Set controls in the following positions: (a) Firewall Shutoff Valve - OPEN (b) Temperature Control Valve - FULL HOT (clockwise) (c) Bleed Air Heat Switch - ON (d) Mixing Air - FLT PUSH (e) Aft Cabin/Fwd Cabin - FWD CABIN PUSH (f) Defrost/Fwd Cabin - FWD CABIN PUSH (2) Set power lever for 60 percent Ng. (3) Check for unrestricted flow at both forward cabin heat outlets (above pilot's and copilot's rudder pedals) and for minimal or no flow through defroster outlets. (4) Set Defrost/Fwd Cabin Push/Pull Control in DEFROST position. Check for unrestricted flow through both defroster outlets and for minimal or no flow through forward cabin outlets. Return control to FWD cabin push position. (5) Set Aft Cabin/Fwd Cabin control in AFT CABIN PULL position and check for unrestricted flow from all cabin outlets. Return control to FWD CABIN PUSH position. NOTE:
On 208B and Cargomaster, check flow from floor mounted outlets, just aft of pilot/ copilot seats.
Set mixing air control GRD PULL position and check for a substantial increase in cabin heat flow. Set bleed air heat switch to OFF position. Increase power to above 90 percent Ng and return mixing air control to FLT PUSH position. Reset bleed air heat switch to ON position. (8) Rotate temperature control knob counterclockwise until cabin flow heat ceases and return it to FULL HOT position. Flow should gradually reduce to zero or a minimal amount in approximately 270 degrees of counterclockwise knob rotation from FULL HOT. Return temperature control knob to FULL clockwise position. (9) Power remaining at 60 percent Ng, pull firewall shutoff valve control to PULL OFF position and ensure cabin heat flow ceases, indicating that gate valve has closed. (10) Switch bleed air heat switch to OFF position, wait approximately 30 seconds to allow back pressure against shutoff valves to dissipate, push control to OPEN position and return switch to ON position. (11) Disengage BLEED AIR HEAT circuit breaker and ensure cabin heat flow ceases. (12) Engage BLEED AIR HEAT circuit breaker.
(6) (7)
6.
Cabin Heat Functional Test (Airplanes 20800180 and On and 208B0210 and On and Airplanes 20800001 Thru 20800179 and 208B0001 Thru 208B0209 Incorporating CAB90-9) A.
Functional Test Procedures. (1) Set controls in the following positions: (a) Firewall Shutoff Valves - OPEN (b) Temperature Control - FULL HOT (clockwise) (c) Bleed Air Heat Switch - ON (d) Mixing Air - FLT PUSH (e) Aft Cabin/Fwd Cabin - FWD CABIN PUSH (f) Defrost/Fwd Cabin - FWD CABIN PUSH (2) Set power lever for 60 percent Ng.
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Check for unrestricted flow at both forward cabin heat outlets, above the pilot's and copilot's rudder pedals, and for minimal or no flow through the defroster outlets. NOTE:
(4) (5)
If flow is restricted, refer to Compressor Bleed Air Heater - Troubleshooting.
Set the defrost/fwd cabin control in the DEFROST position. Check for unrestricted flow through both defroster outlets and for minimal or no flow through the forward cabin outlets. Return control to the FWD CABIN PUSH position. Set aft cabin/fwd cabin control to AFT CABIN PULL position and check for unrestricted flow from all cabin outlets. Return control to FWD CABIN PUSH position. NOTE:
On 208B and Cargomaster, check flow from floor mounted outlets, just aft of pilot/ copilot seats.
(6)
Set the mixing air control in the GRD PULL position and check for a substantial increase in cabin heat flow. (7) Set bleed air heat switch to OFF position. Increase power to above 90 percent N g and check for a reduction in cabin heat flow, indicating compressor bleed valve (P2.5 air supply) has closed. Reduce power to 60 percent Ng and return mixing air control to the FLT PUSH position. Reset bleed air heat switch to ON position. (8) Increase power to approximately 75 percent N g, rotate temperature control knob counterclockwise until cabin flow heat ceases and then return it to the FULL HOT position. Flow should gradually reduce to zero in approximately 270 degrees of counterclockwise knob rotation from FULL HOT. Return the temperature control knob to the FULL clockwise position. (9) Reduce power to 60 percent Ng, pull firewall shutoff valve control to the PULL OFF position and ensure cabin heat flow ceases, indicating microswitch has de-energized the flow control valve solenoid preventing P3 air flow. (10) Switch bleed air heat switch to OFF, wait approximately 30 seconds to allow back pressure against the shutoff valves to dissipate, push control to the OPEN position and return switch to the ON position. (11) Disengage BLEED AIR HEAT circuit breaker and ensure cabin heat flow ceases. Engage BLEED AIR HEAT circuit breaker. (12) Check for normal instrument vacuum and deice function. A quick check of pressure regulator output is that at low idle, the low vacuum annunciator will normally be off with no P3 bleed air heat and will come on with full heat. 7.
Heater Output Check (Airplanes 20800180 and On and 208B0210 and On) A.
Check heater output. (1) Connect hoses and gauges to measure heater output at the pressure tap between the flow control valve and the mixer/muffler, and to measure control pressure at the tee where the control line connects to the flow control valve. (2) Check that maximum heater output is 15.0 to 20.0 PSIG at 70 percent Ng and above, and that control pressure is approximately 0.0 to 5.0 PSIG above output pressure. Check that heater output is approximately 5.0 to 7.0 PSIG at low 52 percent Ng, and 10.0 to 13.0 PSIG at 65 percent Ng.
CAUTION: When operating with the control port capped, gradually advance power until an output pressure of 15.0 to 20.0 PSIG is reached. Do not increase power further, as an overpressure/overtemperature condition can occur. (3)
(4)
If output and control pressures are both low, recheck heater output with control line disconnected and control port capped at flow control valve tee. If not OK, replace flow control valve. If OK, check for leaks in control line from flow control valve to temperature control valve. If there are no leaks, replace temperature control valve. If output pressure is low and control pressure is normal, replace flow control valve.
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If both pressures are high, replace temperature control valve.
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MODEL 208 MAINTENANCE MANUAL COOLING - DESCRIPTION AND OPERATION 1.
General A.
This section describes those optional systems and components used to produce cool air. It does not include components used to distribute cool air, nor those components used to control temperature.
B.
For a description of how cool air is distributed, refer to Air Conditioning System - Description and Operation. For a description of how the temperature of cool air is controlled, refer to Temperature Control - Description and Operation.
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MODEL 208 MAINTENANCE MANUAL COOLING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the cooling system in a serviceable condition.
Task 21-50-00-720 2.
Compressor Drive Belt Functional Check NOTE:
The functional check of the compressor belt for the freon air conditioning system and the R134A air conditioning system is typical.
A.
General (1) This section gives the information needed to complete the functional check of the compressor drive belt.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Compressor Drive Belt Functional Check. (1) Examine the compressor drive belt for condition, wear and alignment. (2) Examine, and if necessary, adjust the compressor drive belt tension. Refer to Adjust Drive Belt , Freon Air Conditioning - Maintenance Practices. NOTE:
The adjustment is typical for the compressor drive belt for R134A air conditioning system.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL FREON AIR CONDITIONING - DESCRIPTION AND OPERATION (Airplanes 20800112 and On, and 208B0214 and On) 1.
2.
General A.
The air conditioning system is used to provide cool air for cockpit and cabin area. The system uses one compressor in conjunction with three evaporators to distribute freon cooled air through overhead ducts. (1) The compressor is mounted on the engine accessory section and driven by a V-belt from a drive unit assembly. (2) Two evaporator units with integral blowers are located in the wing root areas (left and right). The third evaporator unit is located in the tailcone behind the aft cabin bulkhead.
B.
An air conditioning schematic is provided to aid maintenance technicians in system understanding. Refer to Figure 1.
Description A.
Component Descriptions are as follows: (1) Compressor - The air conditioning compressor is a automotive type unit mounted on the aft left side of the engine and driven by a V-belt from a drive unit mounted on the engine accessory section. Service valves are located on the suction and discharge ports on the compressor. The compressor has a dual function. First, it is a means of moving refrigerant through the system, and second, it compresses the gaseous refrigerant, raising the pressure and temperature simultaneously. The temperature rise is the actual desired outcome and is accomplished with a corresponding rise in pressure. (2) Compressor Drive Unit - A compressor drive unit is installed on an accessory pad located on the aft left side of the engine. The drive unit is driven by the engine which in turn drives the air conditioning compressor by means of a pulley and V-belt. A drain hose is installed on the unit and routed from the underside of the drive unit to an outlet in the lower right cowl. The forward support assembly of the drive unit also provides for the attachment of the air conditioning compressor along with a clevis-turnbuckle arrangement which provides V-belt tension adjustment. (3) Compressor Drive Belt - The air conditioning compressor is driven by a V-belt from the drive unit pulley to the pulley on the compressor. (4) Condenser - The condenser is a flat tube fin coil located in the lower left section of the engine compartment. The condenser is interfaced with louvers in the lower left cowl by means of an inlet duct. The inlet duct extends from the condenser to the forward cowl opening, and a series of four seals are connected to the bottom of the condenser and extend downward to meet the aft opening in the cowl. The condenser and inlet duct are attached to the engine mount with clamps and hardware and to the firewall by means of a support bracket and attaching hardware. The condenser receives hot, high pressure gaseous refrigerant and converts it to a cooler, high pressure liquid. Ambient air, which is cooler than the super heated refrigerant, is blown across the condenser coil. Heat from the hot gas passes into the cooler air stream, and in the process, changes the state of the refrigerant back to a liquid. The liquid refrigerant is routed to the receiver/dryer for recycling. (5) Receiver/Dryer - The receiver/dryer is a canister type using a desiccant to remove moisture and a filter to remove larger particles of impurities and hold the desiccant in place. The unit also stores liquid refrigerant during the operation cycle. The receiver/dryer is installed in the lower right side of the engine compartment. (6) Pressure Switch - A high pressure safety is located in the lower right engine compartment just forward of the receiver/dryer. The switch disengages the compressor clutch and stops system operation in the event the system becomes overloaded. The system will cycle on again when the pressure reduces. (7) Air Conditioning Plumbing - Refrigerant lines in the engine compartment, under floorboards and fuselage side walls, interconnect the compressor, condenser, receiver/dryer and evaporators. (8) Wing Mounted Evaporator - Two evaporator units with integral blowers are located, one each in the left and right wing bays just outboard of the wing root rib. The evaporator units both contain an evaporator coil with an expansion valve, a shroud, and a scroll and blower assembly. Cabin
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MODEL 208 MAINTENANCE MANUAL air is drawn through each evaporator coil and conditioned air is routed into the distribution ducts and cabin area. The blower motors are dual speed and circulate conditioned air or ventilation air into the cabin. (9) Overhead Distribution Ducts - The two wing mounted evaporator units utilize existing fresh air distribution ducts for freon-cooled air. For a complete description of how fresh air is distributed through the cabin and cockpit area, refer to Fresh Air Distribution - Maintenance Practices. (10) Tailcone Mounted Evaporator - The tailcone mounted evaporator is mounted aft of the aft cabin bulkhead on the left side. The evaporator unit consists of an evaporator coil with an expansion valve, a shroud, a scroll and blower assembly. Cabin air is drawn through the evaporator coil and the conditioned air is routed through the distribution ducts into the aft cabin area. The blower motor is dual speed. (11) Aft Evaporator Distribution and Return Air System - The distribution and return air system consists of two return air grills mounted on the upper portion of the aft cabin bulkhead, two elbow assemblies and two ducts routed to the duct assembly mounted on the forward side of the evaporator. A duct connected to the evaporator blower assembly and routed to a Wye-duct, which is connected to two distribution ducts mounted in the aft cabin overhead and directs cooled air into the aft cabin area. Louver assemblies in the distribution ducts control the direction and amount of cooled air into the aft cabin area. (12) Check Valve - A check valve is installed in the fuselage root rib and ties into the ducting feeding into the wing mounted evaporators. The check valve allows air to exit the cabin for recirculation over the evaporator, but prevents air from entering the cabin through the return air duct and forces all air into the plenum distribution system. 3.
Operation A.
The evaporator units direct cooled air through the cabin air ventilation system to the cabin air outlets. The condenser, located in the lower left section of the engine compartment, is provided with an inlet and an outlet in the lower left side of the engine cowling to supply cooling airflow through the condenser. A receiver/dryer is installed in the lower right side of the engine compartment. A sight glass, used to determine when the system has been properly charged, is installed in the high pressure line near the Schrader valve. Two Schrader valves are installed, one in the high pressure line and one in the low pressure line, for servicing. The sight glass and service valves are located beneath the floorboard inspection covers between the pilot and copilot seats. Refrigerant lines run under the fuselage floorboards and interconnect system components with each other.
B.
Controls for the air conditioning system consist of a three-position toggle air conditioning switch and three two-position toggle fan switches. The controls are located at the lower edge of the instrument panel directly above the control pedestal, and two ventilation system controls are located in the overhead console. Placing the three-position switch, labeled OFF, VENTILATE, COOL, from the OFF position to the COOL position starts the compressor and evaporator fans. Placing the switch in the VENTILATE position activates only the evaporator fans, producing uncooled vent air to the cabin. The three two-position switches, all labeled AC FANS, provide separate HIGH or LOW speed control of each evaporator fan. System electrical protection is provided by four 15-ampere "pull-off" type circuit breakers, labeled LEFT VENT BLWR, RIGHT VENT BLWR, AFT VENT BLWR and AIR COND CONT. The circuit breakers are located on the left side wall circuit breaker panel.
C.
When the air conditioning system is operating, cooled air is supplied to the cabin through 16 overhead adjustable outlets (two each above the pilot and front passenger, one above each rear passenger seat and two directing air forward and one directing air downward from the aft cabin bulkhead area). The pilot's and front passenger outlets are the swivel type for optimum positioning. Airflow volume is controlled by rotating the outlet nozzle which controls an internal valve. The eight rear passenger seat outlets and three aft cabin outlets are directionally adjustable. Each rear passenger outlet has a separate rotary type control beside the outlet, with positions labeled AIR ON and AIR OFF, to control airflow volume through the outlet.
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Air Conditioning System Schematic Figure 1 (Sheet 1)
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Air Conditioning System Schematic Figure 1 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL FREON AIR CONDITIONING - TROUBLESHOOTING 1.
General A.
Troubleshooting charts are provided to aid maintenance technicians in system diagnosis of the freon air conditioning system. Refer to Figure 101.
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 1)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 2)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 3)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 4)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 5)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 6)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 7)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 8)
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Freon Cooling System Troubleshooting Chart Figure 101 (Sheet 9)
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MODEL 208 MAINTENANCE MANUAL FREON AIR CONDITIONING - MAINTENANCE PRACTICES 1.
General Precautions A.
Handling Freon. NOTE:
The effect of the Montreal Protocol and U. S. Environmental Protection Agency’s Clean Air Act of 1990 is to ban the unnecessary release of CFC-12 refrigerant (also known as R-12) into the atmosphere. In compliance with the preceding, Cessna Aircraft recommends the refrigerant be captured and recycled. For additional information, refer to Federal Clean Air Act, EPA 40 CFR Part 82.
WARNING: Liquid R-12 at normal atmospheric pressure and temperature will freeze anything it contacts. The eyes are especially susceptible to damage. Safety glasses are the absolute minimum protection and shall be worn at all times when servicing the Freon system. WARNING: Do not attempt to treat yourself, should any liquid refrigerant get into the eyes. Follow these instructions: Do not rub the eye. Splash large quantities of cool water into the eye to raise the temperature. Apply a few drops of mineral oil to the eye to wash it, followed by a weak solution of boric acid to flush out all of the oil. Seek the aid of a doctor immediately. (1) (2) B.
Observe safety precautions when handling refrigerant or servicing and performing maintenance on air conditioning system. Use of protective clothing, gloves and goggles will protect the skin and eyes.
General system notes. NOTE:
C.
Cleanliness is of the utmost importance to avoid system contamination and useless wear to the compressor and other equipment items. All plumbing and hoses shall be cleaned and capped after fabrication and shall remain capped during storage and installation until connected to their mating components. All ports shall also be capped with clean caps or plugs. During the time components are open, extreme care shall be exercised to assure that no contaminating matter enters the parts or system. The receiver/dryer is easily contaminated with moisture from the atmosphere. All care shall be exercised to prevent moisture from entering the receiver/dryer.
Removing hoses under pressure.
WARNING: Do not remove hoses under pressure. This procedure will result in the release of refrigerant into the atmosphere. Removing hoses under pressure may also result in personal injury if hose ends are not restrained. D.
Use of intense heat.
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WARNING: To avoid explosion, never weld, use a blow torch, steam clean, bake aircraft finish or use excess amounts of heat on or in the immediate area of any part of the air conditioning system or refrigerant supply tank, while they are closed to atmosphere, charged or not. Although R-12 gas, under normal conditions, is nonpoisonous, the discharge of refrigerant gas near a flame can produce a very poisonous gas (phosgene). This gas will also attack all bright metal surfaces. WARNING: Do not use a flame-type leak detector because of fire hazard on airplanes and production of minor amounts of phosgene gas. WARNING: Do not smoke in the vicinity of refrigerant discharge. Inhaling refrigerant through burning tobacco will produce a poisonous gas like an open flame. E.
Use of nitrogen. NOTE:
2.
All nitrogen pressure checks are to be made only with regulated nitrogen.
Compressor Removal/Installation A.
Remove Compressor (Refer to Figure 201). (1) Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (2) Disconnect electrical connector from compressor. (3) Unscrew discharge service valve from compressor. (4) Cap discharge service valve and compressor fitting. (5) Unscrew suction service valve from compressor. (6) Cap suction service valve and compressor fitting. (7) Release tension on compressor by loosening nut and bolt at bottom of support plate. (8) Remove clips from turnbuckle and loosen turnbuckle. (9) Remove turnbuckle from compressor. (10) Remove nut, bolt and washer from bottom of support plate. (11) Remove belt from compressor. (12) Remove compressor from airplane. NOTE:
For compressor refurbishing procedures, refer to vendor's component maintenance manual.
(13) If compressor is being replaced, perform the following steps: (a) Remove oil plug and drain compressor oil into measuring cup. Record amount of oil removed. B.
Install Compressor (Refer to Figure 201). (1) If new compressor is being installed, perform the following steps: (a) Drain oil from new compressor. NOTE:
Compressors are shipped from the factory with approximately 6.0 ounces of fluid. This fluid should be drained, discarded and replaced before compressor is attached to airplane.
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Compressor Installation Figure 201 (Sheet 1)
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CAUTION: Do not leave compressor oil containers uncapped. containers of refrigerant oil absorb moisture rapidly.
Open
CAUTION: Do not operate system without refrigerant oil in compressor. (b)
(2) (3) (4) (5) (6) (7) (8) 3.
Determine amount of oil removed from old compressor and add 1.0 ounce to this Add this amount of new, uncontaminated compressor oil to new measurement. compressor. Refer to Air Conditioning - General for a list of approved compressor oils. (c) Reinstall drain plug. Attach compressor to support assembly using nut, bolt and washer. Do not tighten. Lift up on compressor far enough to position belt around compressor pulley. Connect turnbuckle to adjuster plate using nut, washer and bolt. Adjust compressor belt tension. Refer to Compressor Drive Belt Removal/Adjustment. Remove protective caps from discharge and suction service valves and reconnect lines to compressor. Connect electrical connector to compressor. Charge system. Refer to Chapter 12, Freon Air Conditioning - Servicing.
Compressor Drive Unit Removal/Installation A.
Remove and Disassemble Compressor Drive Unit (Refer to Figure 202). (1) Remove compressor. Refer to Compressor Removal/Installation. (2) Remove drive belt. Refer to Compressor Drive Belt Removal/Adjustment. (3) Loosen and disconnect drain hose from elbow at bottom of support assembly. (4) Cut safety wire and remove bolts and washers securing support assembly to engine. Note position of shims for later reassembly. (5) Carefully pull entire support assembly aft to disengage drive shaft from engine. Discard gasket. (6) If disassembling drive unit, perform the following steps. (Refer to Figures 202 and 203. (a) Remove retaining ring, closure disc and shim at end cap. (b) Remove retaining rings which hold drive shaft to both ends of pulley. (c) Remove drive shaft from support assembly. NOTE: (d) (e) (f) (g)
B.
Drive shaft and retaining rings are removed to prevent damage when bearings are pulled.
Carefully pull end cap off of bearing. Remove pulley, bearings, splined coupling and retaining ring from support housing. Separate bearings and splined coupling from pulley. Remove oil seal from support housing.
Assemble and Install Compressor Drive Unit (Refer to Figure 202). (1) If drive unit was disassembled, reassemble in the following steps. (Refer to Figure 202 and Figure 203). (a) Press new oil seal into support housing. (b) Press bearing into support housing. (c) Press bearings on pulley. (d) Install retaining ring on splined coupling. (e) Install splined coupling to pulley and secure with retaining ring. (f) Install drive shaft into splined coupling and secure with retaining rings. (g) Press assembly shown in Detail B into assembly shown in Detail A, Figure 203. (h) Reinstall shim and closure disc to end cap using retaining ring. NOTE: (i)
Remove laminates from shim as required to install closure disc. Shim is required to keep closure disc in position.
Press end cap onto bearing. NOTE:
Do not insert spacer at this time. Spacer must be removed to install belt.
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Compressor Drive Unit Installation Figure 202 (Sheet 1)
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Compressor Drive Unit Cutaway View Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2)
Apply Plastilube (MIL W-G-632) lubricant to the forward splines of the Compressor Unit compressor drive shaft. NOTE:
(3) (4) (5) (6) (7) (8) 4.
Plastilube (MIL W-G-632) is not to be used on phenolic splines.
Using new gasket, align drive shaft on support assembly with accessory pad coupling. Secure support assembly on accessory pad using bolts, washers and shims as required. Safety wire bolts at accessory pad. Refer to Chapter 20, Safetying - Maintenance Practices. Connect drain hose to support assembly elbow. Install compressor. Refer to Compressor Removal/Installation. Install drive belt. Refer to Compressor Drive Belt Removal/Adjustment.
Compressor Drive Belt Removal/Installation. A.
Remove Drive Belt (Refer to Figures 201, 202 and 204). (1) Loosen bolt at bottom of compressor. (2) Remove and discard turnbuckle clip from turnbuckle. (3) Loosen turnbuckle enough to pass belt over compressor pulley. (4) Remove bolts securing spacer between end cap and support assembly. (5) Remove belt through opening where spacer was removed.
B.
Install Drive Belt (Refer to Figures 201, 202 and 204 ). (1) Insert belt through opening between end cap and support housing. (2) Reinstall spacer. Secure spacer between end cap and support assembly using bolts and washers. (3) Lift upward on compressor far enough to allow belt to slip over compressor pulley. (4) Connect clevis end of turnbuckle to compressor. (5) Adjust compressor drive belt. Refer to Adjust Drive Belt.
C.
Adjust Drive Belt (Refer to Figure 204). (1) Tension can be checked by using either of the two following methods: (a) A spring scale hooked under the belt at a point midway between compressor drive unit pulley and compressor clutch pulley, pulling perpendicular to the belt. (b) Using a Gates 150 tensiometer. (2) Correct belt tension is a 0.12-inch deflection when a load force of 3.6 to 4.4 pounds is applied to the belt. (3) If belt tension is not correct, adjust as follows: (a) Loosen bolt at bottom of compressor to allow compressor to pivot. (b) Remove and discard clips on turnbuckle. (c) Adjust turnbuckle in or out to obtain correct belt tension. NOTE:
A maximum of three threads must be exposed on adjustment arm clevis. Replace MS21252-5RS clevis with MS21252-5RL. Refer to the following table for turnbuckle adjustment ranges.
ADJUSTMENT RANGE
NORMAL
MAX (REF)
MS21252-5LL & MS21252-5RS
4.55 to 5.55 inch
5.70 inch
MS21252-5LL & MS21252-5RL
5.40 to 6.40 inch
6.60 inch
(d) (e) 5.
Install new clip on turnbuckle. Tighten bolt at bottom of compressor.
Condenser Removal/Installation A.
Remove Condenser (Refer to Figure 205). (1) Remove lower left engine cowl. Refer to Chapter 71, Engine Cowling and Nose-cap Maintenance Practices. (2) Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (3) Loosen clamps and remove hoses leading into condenser. Cap all hoses and fittings.
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Compressor Drive Belt Adjustment Figure 204 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (4) (5) (6) (7) (8) B.
6.
Remove bolt securing aft end of condenser to condenser support bracket. Remove bolts, clamps and spacers securing compressor to engine mount. Remove inlet duct and condenser from airplane. Remove bolts and washers securing inlet duct to condenser. Separate inlet duct from condenser. If required, remove seal assemblies from condenser.
Install Condenser (Refer to Figure 205). (1) If required, install seal assemblies to condenser. (2) Attach condenser to inlet duct using bolts and washers. (3) Attach condenser to engine mount using clamps, spacers and hardware as required. Do not tighten at this time. (4) Align holes in right aft corner of condenser with holes in condenser support bracket. Attach using washers and bolts. (5) Tighten clamps, spacers and hardware on engine mount. (6) Reinstall hoses to condenser. Tighten with clamps. (7) Charge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (8) Install lower left engine cowl. Refer to Chapter 71, Engine Cowling and Nose-cap - Maintenance Practices.
Receiver/Dryer Removal/Installation A.
Remove Receiver/Dryer (Refer to Figure 206). NOTE:
(1) (2) (3) (4) (5) (6) (7) B.
7.
Anytime the air conditioning system has been exposed to atmosphere for any length of time, or when any major components of the system have been replaced, the receiver/dryer should also be replaced.
Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. Disconnect fitting at manifold pressure switch housing. Remove sta-straps and disconnect electrical connector. Remove pressure switch from manifold pressure switch housing. Discard packing and cap open lines. Disconnect fitting from OUT end of receiver/dryer. Loosen clamps and remove receiver/dryer from engine mount. Remove unions from both ends of receiver/dryer. Discard packing and receiver/dryer.
Install Receiver/Dryer (Refer to Figure 206). (1) Install union fittings (with new packing) to both ends of new receiver/dryer. (2) Attach receiver/dryer to engine mount and secure clamps. (3) Attach fittings to both ends of receiver/dryer unions. (4) Attach pressure switch with new packing to manifold pressure switch housing. (5) Connect housing cap to housing plug and secure wire using sta-straps. (6) Charge system. Refer to Chapter 12, Freon Air Conditioning - Servicing.
Pressure Switch Removal/Installation A.
Remove Pressure Switch (Refer to Figure 206). (1) Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (2) Remove sta-straps from electrical wiring. (3) Disconnect housing plug from housing cap. (4) Remove pressure switch and packing from manifold pressure switch housing. Discard packing. (5) Cap manifold pressure switch housing to preclude entry of moisture and/or contaminants into system. (6) Check pressure switch for proper operation. Refer to Pressure Switch Functional Test.
B.
Install Pressure Switch (Refer to Figure 206). (1) Install pressure switch to manifold pressure switch housing using new packing. (2) Connect housing plug to housing cap. (3) Secure electrical wiring using sta-straps. (4) Charge system. Refer to Chapter 12, Freon Air Conditioning - Servicing.
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Condenser Installation Figure 205 (Sheet 1)
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Receiver/Dryer Installation Figure 206 (Sheet 1)
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8.
Air Conditioning Plumbing Removal/Installation A.
Remove Air Conditioning Plumbing (Refer to Figure 207). (1) Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. NOTE:
(2) (3) (4) B.
Refrigerant lines in the engine compartment, under the floorboards and in the fuselage sidewalls interconnect the compressor, condenser, receiver/dryer and evaporators.
Remove interior equipment and access panels as required to gain access to refrigerant lines. Disconnect plumbing and remove as necessary. Cap all lines and fittings to preclude entry of moisture and/or foreign particles into system.
Install Air Conditioning Plumbing (Refer to Figure 207).
CAUTION: The use of other thread lubricants is positively prohibited, including “Lock-Tite” or other commercial refrigerant lubricants such as “LeakLock.” (1)
Remove previously installed caps from lines and install plumbing. NOTE:
(2)
It is recommended that all straight thread fittings and O-rings be lubricated with clean refrigerant oil and all taper (pipe) threads be serviced with Teflon tape. Use care to ensure Teflon tape does not get closer than one to one-half threads from end of fitting. Should a piece of tape get into system, it can cause blockage of small orifices.
Torque lines to valves listed in table below. NOTE:
All plumbing fittings must be torqued to prevent Freon leakage and shall be rechecked after performing an air conditioning leak test.
TUBE DIAMETER
TORQUE VALUE
0.250 inch
55 to 65 inch-pounds
0.375 inch
100 to 125 inch-pounds
0.500 inch
200 to 250 inch-pounds
0.750 inch
400 to 500 inch-pounds
(3) (4) (5) (6) 9.
Perform leak test of system. Refer to Chapter 12, Freon Air Conditioning - Servicing. Charge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. Perform an operational test of the system. Refer to System Operational Test. Reinstall removed floor boards, panels and interior equipment.
Wing Mounted Evaporator Removal/Installation NOTE: A.
Evaporator removal and installation are typical for both left and right wing evaporator.
Remove Wing-Mounted Evaporators (Refer to Figure 208). (1) Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (2) Remove wing root access panel 511AB/611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Disconnect electrical housing cap from housing plug. (4) Disconnect evaporator drain hose from drain tube. (5) Disconnect elbow fitting from bottom of evaporator and cap line (6) Disconnect expansion valve from evaporator and cap line. (7) Disconnect duct at blower assembly. (8) Remove four bolts securing evaporator assembly to duct.
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Air Conditioning Plumbing Installation Figure 207 (Sheet 1)
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Air Conditioning Plumbing Installation Figure 207 (Sheet 2)
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Wing Mounted Return Air Check Valve Assembly Figure 208 (Sheet 1)
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Wing Mounted Return Air Check Valve Assembly Figure 208 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (9) B.
10.
11.
12.
Pull evaporator assembly far enough aft to allow studs to clear duct. Remove evaporator assembly from airplane.
Install Wing-Mounted Evaporators (Refer to Figure 208). (1) Position evaporator assembly in wing root area with forward studs through holes in duct. Secure evaporator assembly to duct using nuts and bolts. (2) Reconnect and tighten duct at blower assembly. (3) Connect expansion valve to evaporator. (4) Connect elbow fitting to bottom of evaporator. (5) Connect evaporator drain hose to drain tube. (6) Connect electrical housing cap to housing plug. (7) Charge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (8) Install wing root access panel 511AB/611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Wing Mounted Return Air Check Valve Removal/Installation A.
Remove and Disassemble Wing-Mounted Return Air Check Valve (Refer to Figure 208). (1) Remove wing root access panel 511AB/611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Remove clamp and flexible duct from outboard duct assembly. (3) Remove screws securing outboard duct assembly (with check valve and seal) to inboard duct assembly. (4) Remove outboard duct assembly from airplane. (5) Disassemble check valve in the following steps: (a) Remove nut at bottom of hinge pin and withdraw hinge pin from outboard duct assembly. This will allow check valve halves and spring to be removed from outboard duct assembly. (b) Remove nut at bottom of pin and withdraw pin from outboard duct assembly.
B.
Assemble and Install Wing-Mounted Return Air Check Valve (Refer to Figure 208). (1) Reassemble check valve in the following steps: (a) Assemble check valve halves and spring in outboard duct assembly. Insert hinge pin through duct, valve halves and spring. Secure hinge pin using nut. (b) Insert pin through outboard duct assembly and secure using nut. (c) Ensure check valve operates smoothly and seats fully. (2) Install outboard duct assembly (with check valve and seal) to inboard duct assembly using screws. (3) Attach flexible duct to outboard duct assembly using clamp. (4) Reinstall wing root access panel 511AB/611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Forward Evaporator Return Air Grill A.
Remove Return Air Grill (Refer to Figure 208). (1) From cabin area, remove screws securing grill to inboard duct assembly.
B.
Install Return Air Grill (Refer to Figure 208). (1) Align holes in grill with holes in headliner and inboard duct assembly. (2) Install screws to secure grill to inboard duct assembly.
Tailcone Mounted Evaporator Removal/Installation A.
Remove Aft Evaporator (Refer to Figure 209). (1) Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (2) Remove aft cabin partition to gain access to evaporator unit. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices. (3) Disconnect electrical housing plug from housing cap. (4) Disconnect evaporator drain hose from bottom of evaporator. (5) Remove recirculated air ducts from duct assembly. (6) Remove fitting from expansion valve. Cap open line.
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Tailcone Mounted Evaporator Installation Figure 209 (Sheet 1)
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13.
14.
Remove fitting from union on suction line leading into evaporator. Cap open line. Remove screws securing evaporator to brackets. Remove flexible distribution duct from blower motor and remove evaporator assembly from airplane.
Install Aft Evaporator (Refer to Figure 209 ). (1) Install evaporator to aft cabin area using screws and washers as required. (2) Attach flexible distribution duct to blower motor. (3) Install Freon lines to evaporator. (4) Connect drain line to evaporator. (5) Attach recirculated air ducts to duct assembly. (6) Connect electrical connector housing plug to housing cap. (7) Recharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (8) Install aft cabin partition. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices.
Aft Evaporator Distribution and Return Air System Removal/Installation A.
Remove Aft Evaporator Distribution and Return Air Ducts (Refer to Figure 209). (1) Remove aft cabin partition to gain access to evaporator unit. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices. (2) Loosen clamps securing recirculated air ducts to elbow assemblies. (3) Remove recirculated air ducts from airplane. (4) Loosen clamp securing flexible distribution duct to wye duct. (5) Remove flexible distribution duct from wye duct. (6) Remove screws securing wye duct to distribution duct and remove wye duct from airplane. (7) Remove screws securing distribution duct to airplane and remove duct from airplane.
B.
Install Aft Evaporator Distribution and Return Air System (Refer to Figure 209). (1) Install distribution duct to airplane using screws. (2) Attach wye duct to distribution duct. (3) Attach flexible distribution duct to wye duct using clamp. (4) Attach recirculating air ducts to elbow assemblies using clamps. (5) Install aft cabin partition. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices.
System Operational Test A.
Air Conditioning System Operational Test. NOTE: (1) (2)
(3) (4) (5)
15.
Perform system check at ambient temperatures of 55°F or higher.
Start airplane engine and run at a minimum 54% Ng. Under extremely hot outside air temperature it may be necessary to run engine at 60 to 65% Ng. Engage the following circuit breakers: (a) LEFT VENT BLWR (b) RIGHT BENT BLWR (c) AFT VENT BLWR (d) AIR COND CONT Move fan switches from HIGH to LOW and note a change in evaporator fan speed. Place the air conditioner switch to COOL and activate compressor. Temperature differential across evaporators should be at least 20°F. Measure temperatures at evaporators with dial-type thermometers. If evaporators do not cool, refer to Freon Air Conditioning - Troubleshooting.
Pressure Switch Functional Test A.
Testing Pressure Switch (Refer to Figure 206 ). (1) Discharge system. Refer to Chapter 12, Freon Air Conditioning - Servicing. (2) Remove pressure switch and packing from manifold pressure switch housing.
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6) (7) (8)
Check for electrical continuity through the switch. Switch (continuity) should be closed. Apply 355.0 PSIG dry nitrogen pressure to pressure switch. At 355.0 PSIG, +5 or -5 PSIG, switch should open (no continuity). Decrease pressure on switch. At approximately 330.0 PSIG switch should close (continuity). Replace the switch if it is not within these parameters. Install the pressure switch, with new packing, to the manifold pressure switch housing Charge system. Refer to Chapter 12, Freon Air Conditioning - Servicing.
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MODEL 208 MAINTENANCE MANUAL R134A AIR CONDITIONING SYSTEM - DESCRIPTION AND OPERATION (Airplanes 20800274 And On, and 208B0655 And On) 1.
2.
General A.
The air conditioning system provides cool air for cockpit and cabin area. This system uses a compressor in conjunction with three evaporators to distribute R134a cooled air through overhead ducts. (1) The compressor is mounted on the engine accessory section and driven by a V-belt from a drive unit assembly. (2) Two evaporator units with integral blowers are located in the wing root areas (left and right). The third evaporator unit is located in the tailcone behind the aft cabin bulkhead.
B.
An air conditioning schematic is provided to aid maintenance technicians in system understanding. Refer to Figure 1.
Description A.
Component Descriptions are as follows: (1) Compressor - The air conditioning compressor is mounted on the aft left side of engine and driven by a V-belt from a drive unit mounted on engine accessory section. Service valves are located on the suction and discharge ports labeled SUC and DIS respectively on the compressor. The compressor has a dual function. It is a means of moving refrigerant through the system and compresses gaseous refrigerant, raising pressure and temperature simultaneously. A temperature rise is desired outcome and is accomplished with a corresponding rise in pressure. (2) Condenser - The condenser is a flat tube fin coil located in the lower left section of the engine compartment. The condenser receives hot, high pressure gaseous refrigerant and converts it to a cooler, high pressure liquid. Ambient air is blown across the condenser coil. Heat from hot gas passes into the cooler air stream, and changes, back to a liquid. The liquid refrigerant is routed to the receiver/dryer for recycling. (3) Receiver/Dryer - The receiver/dryer is installed in the lower right side of the engine compartment.The receiver/dryer is a canister type using a desiccant to remove moisture and a filter to remove larger particles of impurities and hold desiccant in place. The unit also stores liquid refrigerant during the operation cycle. (4) Pressure Switch - A binary High/Low pressure safety switch is threaded into the top of the receiver/dryer. This switch disengages the compressor clutch and stops system operation should the system become overloaded. Compressor damage could occur if the system pressure becomes either too high or too low. The system will cycle on again when the pressure returns to a safe operating condition. (5) Evaporators - There are three evaporators in the system, one in each wing root and a third in the tailcone. Each evaporator consists of the evaporator coil with an expansion valve and an electrically powered squirrel cage blower. The two wing-mounted evaporators are connected into the ventilation system duct in the wing root. Air for these evaporators may be fresh air from the outside when the ventilation duct valves are closed. The rear evaporator operates on recirculated air only. Refrigerant to each of the evaporators is metered through the expansion valves. (6) Service Valves - Quick disconnect service valves in low the pressure (vapor) and high pressure (liquid) lines are located beneath the floorboard between pilot and copilot seats. (7) Controls - Controls for the air conditioning system consist of a air conditioning switch and three fan switches. located at the lower edge of the instrument panel directly above control pedestal, and two ventilation system controls in the overhead console. Figure 1 shows the cockpit control panel. (a) The air conditioner control switch has three positions, OFF, VENTILATE and COOL. All electrical operations of the air conditioning system are controlled by this switch.The OFF position prevents power from going to any component in the system. The COOL position starts the evaporator fans and makes power available to the remaining components in
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R134a Air Conditioning System Schematic Figure 1 (Sheet 1)
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R134a Air Conditioning System Schematic Figure 1 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL the system. Placing the switch in the VENTILATE position provides power only to the evaporator blowers. The AC FAN switches are two position toggle switches providing only a high or low speed operation of the fan motors. (b) The ventilation system controls in the overhead console operate control valves in the ventilation ducts located in the wing to cause either fresh air to be drawn across the wing mounted evaporators when the valve is open, or cabin air to be recirculated across the evaporators when the valves are closed. (8) Circuit Breakers - Four pull-off circuit breakers are provided and located on the left sidewall circuit breaker panel. They are labeled LEFT VENT BLWR, RIGHT VENT BLWR, AFT VENT BLWR and AIR COND CONT. (9) Refrigerant Lines - Refrigerant lines in the engine compartment, under floorboards and fuselage side walls, interconnect the compressor, condenser, receiver/dryer and evaporators. The fittings shall follow the SAE standard guidelines for special fitting to avoid cross contamination with refrigerant R12. (10) Air Distribution System - The air distribution system is the same as the standard 10 outlet, 13 outlet for 208B, ventilation air distribution system except for an added outlet outboard of the two forward seats. The wing mounted evaporators are connected to this system. (a) A separate distribution system is provided for the tailcone mounted evaporator.When the air conditioning is operating, cooled air is supplied to the cabin through 16 overhead adjustable outlets (two each above the pilot and front passenger, one above each rear passenger seat and two directing air forward and two directing air downward from the aft cabin bulkhead area), or 19 outlets for 208B. (b) The pilot’s, front passenger’s and the 8 rear passenger seat outlets (11 for 208B) are the swivel type for optimum positioning. Airflow volume is controlled by rotating the outlet nozzle which controls an internal valve. Air flow volume and direction may be controlled through the 4 aft cabin outlets via the air outlet grills (11) Return Air Check Valve - A check valve is installed in the fuselage root rib and ties into the ducting feeding into the wing mounted evaporators. The check valve allows air to exit cabin for recirculation over the evaporator but prevents air from entering the cabin through the return air duct and forces all air into the plenum distribution system. 3.
Operation A.
R134a refrigerant is pumped through a system that alternately evaporates and condenses the refrigerant to move heat from one location to another. In this case, heat is removed from the cabin through evaporators and is expelled to the outside air through the condenser. (1) On both the 208 and 208B system, the compressor is driven by a V-belt from a drive unit mounted on the engine accessory section. The compressor compresses the low pressure gas to a hot high pressure gas. Hot high pressure gas is then passed through the condenser where it rejects heat picked up from the cabin along with the heat of compression and then condenses the gas into a warm high pressure liquid. This liquid is then passed through a receiver/dryer where the remaining gas is separated from the liquid and any moisture is removed by a desiccant. (2) The receiver/dryer also acts as a reservoir for the liquid refrigerant. The warm high pressure liquid then travels to the constant pressure expansion valves where warm high pressure liquid is expanded to a low pressure, low temperature liquid/vapor mixture. This mixture then travels through the evaportaors and absorbs heat from the cabin air which evaporates the remaining liquid refrigerant. The low pressure gas then returns to the compressor to repeat the process. Refer to Figure I.
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MODEL 208 MAINTENANCE MANUAL R134A AIR CONDITIONING SYSTEM - TROUBLESHOOTING (Airplanes 20800274 and On, and 208B0655 and On) 1.
2.
General A.
An understanding of the system is required in order to troubleshoot a vapor cycle cooling system.
B.
Troubleshooting should be performed at an ambient temperature above 50° F. At temperatures below 50° F, system will not function due to low refrigerant presssure.
Tools and Equipment A.
3.
Refer to Air Conditioning - General, for a list of tools and equipment.
Troubleshooting A.
Troubleshooting charts are provided to aid maintenance technicians in system diagnosis of the R134a air conditioning system. Refer to Figure 101.
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 1)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 2)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 3)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 4)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 5)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 6)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 7)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 8)
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R134a Cooling System Troubleshooting Chart Figure 101 (Sheet 9)
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MODEL 208 MAINTENANCE MANUAL R134A AIR CONDITIONING SYSTEM - MAINTENANCE PRACTICES (Airplanes 20800274 And On, and 208B0655 And On) 1.
General Safety Precautions A.
Handling R134a Refrigerant.
WARNING: Care must be taken to minimize the release of refrigerant into the atmosphere. The Environmental Protection Agency (EPA) requires recycling/recovery of R134a as of 11/15/95. All reclamation and recovery equipment must be EPA- and UL-listed. Use the R134a reclamation system per manufacturer's instruction. WARNING: Observe safety precautions when handling refrigerant or servicing and performing maintenance on the air conditioning system. WARNING: Liquid refrigerants at normal atmospheric pressure and temperature will expand and absorb heat. As a result, the refrigerant will freeze anything it contacts. Use of protective clothing, gloves, and goggles will protect the skin and eyes. The eyes are especially susceptible to damage, so safety glasses are an absolute minimum protection. Goggles are the preferred method of protection and must be worn at all times when servicing the system. WARNING: If any liquid gets into the eyes, follow these instructions. Do not rub eye. Splash large quantities of cool water into the eye to raise the temperature. Apply a few drops of mineral oil to eye to wash it out, followed by a weak solution of boric acid to flush out all of the oil. Seek the aid of a doctor immediately. Do not attempt to treat it yourself. B.
General System Notes. NOTE:
C.
D.
Cleanliness is of the utmost importance to avoid system contamination and useless wear to the compressor and other equipment items. All plumbing and hoses shall be cleaned and capped after fabrication and shall remain capped during storage and installation until connected to their mating components. All ports shall also be capped with clean caps or plugs. When components are open, extreme care shall be exercised to assure that no contaminating matter enters the parts or system. The receiver-dryer is easily contaminated with moisture from the atmosphere. All care shall be exercised to prevent moisture from entering the receiver-dryer.
Removing Hoses Under Pressure. NOTE:
Discharge system and recover any refrigerant prior to removing hoses. Removing hoses under pressure is not recommended. Hoses removed with the system charged will spew vigorously and will whip end of hose if not restrained.
NOTE:
The compressor assembly is shipped with a slight amount of internal pressure. Remove caps and vent slowly.
Use of Intense Heat.
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WARNING: To avoid an explosion, never weld, use a blow torch, or use excessive amounts of heat on or in the immediate area of any part of the air conditioning system or a refrigerant supply tank, while they are closed to atmosphere, charged or not. E.
Proper Equipment Connection.
WARNING: Connection of low pressure equipment gages and refrigerant bottles to the high side of the compressor can result in personal injury or equipment damage. Always ensure valves on gages are closed when connecting gages and that hoses are properly connected. F.
Equipment and Materials.
WARNING: A mercury thermometer cannot be used in airplanes due to hazard of possible mercury reaction with aluminum. G.
Use of Nitrogen. NOTE:
All nitrogen pressure checks are to be made only with regulated nitrogen.
CAUTION: Do not connect nitrogen cart while service unit is installed. Damage to service unit could occur. 2.
Compressor Removal/Installation A.
Remove Compressor (Refer to Figure 201). (1) Discharge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. (2) Disconnect electrical connector from compressor. (3) Unscrew discharge service valve from compressor. (4) Cap discharge service valve and compressor fitting. (5) Unscrew suction service valve from compressor. (6) Cap suction service valve and compressor fitting. (7) Release tension on compressor by loosening nut and bolt at bottom of support plate. (8) Remove clips from turnbuckle and loosen turnbuckle. (9) Remove turnbuckle from compressor. (10) Remove nut, bolt and washer from bottom of support plate. (11) Remove belt from compressor. (12) Remove compressor from airplane. (13) If compressor is being replaced, perform following steps: (a) Drain oil from old compressor and fill new compressor with an amount of oil equal to that drained from the old compressor plus one ounce.
B.
Install Compressor (Refer to Figure 201). (1) Attach compressor to support assembly using nut, bolt and washer. Do not tighten. (2) Connect turnbuckle to adjuster plate using nut, washer and bolt. (3) Adjust compressor belt tension. Refer to Compressor Drive Belt Removal/Adjustment. (4) Remove protective caps from discharge and suction service valves and reconnect lines to compressor. (5) Connect electrical connector to compressor. (6) Charge system. Refer to Chapter 12, R134a Air Conditioning - Servicing.
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R134a Compressor Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL C.
If new compressor is being installed, perform following steps: (1) Drain oil from new compressor. NOTE:
Compressors are shipped from the factory with approximately 6.0 ounces of fluid. This fluid must be drained, discarded and replaced before compressor is attached to airplane.
CAUTION: Do not leave compressor oil containers uncapped. Refrigerant oil in open containers absorb moisture rapidly. CAUTION: Do not operate system without refrigerant oil in compressor. (2)
Determine amount of oil removed from old compressor and add 1.0 ounce to this measurement. Add this amount of new, uncontaminated compressor oil to new compressor. For a list of approved compressor oils, refer to Air Conditioning - General. (3) Reinstall drain plug. (4) Attach compressor to support assembly using nut, bolt and washer. Do not tighten. (5) Position compressor and install belt around compressor pulley. (6) Connect turnbuckle to adjuster plate using nut, washer and bolt. (7) Adjust compressor belt tension. Refer to Compressor Drive Belt Removal/Adjustment. (8) Remove protective caps from discharge and suction service valves and reconnect lines to compressor. (9) Connect electrical connector to compressor. (10) Charge system. Refer to Chapter 12, R134a Air Conditioning - Servicing.
3.
Compressor Drive Unit Removal/Installation A.
Remove the Compressor Drive Unit (Refer to Figure 202 and Figure 203). (1) Remove the drive belt. Refer to Compressor Drive Belt Removal/Installation. (2) Remove the compressor. Refer to Compressor Removal/Installation. (3) Cut the safety wire. (4) Disconnect the drain hose from the elbow at the bottom of the support assembly. (5) Remove the bolts, washers and shim(s) that attach the support assembly to the engine. (6) Carefully pull the support assembly aft to disengage the drive shaft from the engine. (7) Discard the gasket.
B.
Install the Compressor Drive Unit (Refer to Figure 202 and Figure 203). NOTE: (1) (2)
Install the compressor drive unit before you install the compressor.
Put a new flange gasket in place. Apply Plastilube (MIL W-G-632) lubricant to the forward splines of the Compressor Unit compressor drive shaft.
CAUTION: Apply Plastilube only to metal-to-metal surfaces. (3) (4)
Align the drive shaft on the support assembly with the engine drive pad. Attach the support assembly with the bolts, washers and shim(s) as follows: (a) Tighten the four bolts that attach the support housing flange and gasket to the engine with your hand. NOTE: (b) (c) (d)
Do not tighten the bolts fully.
Insert the two bolts that secure the support assembly to the compressor drive support, and engage two or three threads. Fully torque the four bolts that attach the support housing and gasket to the engine, in a diagonal step pattern. Remove the two bolts that attach the support housing to the compressor drive support.
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R134a Compressor Drive Unit Installation Figure 202 (Sheet 1)
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R134a Compressor Drive Unit Cutaway View Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (e)
Install the shim(s) with the procedures that follow: 1 Measure the space between the support assembly and the compressor drive support. 2 If the space is less than 0.062 inches (1.57 mm), remove the necessary laminate layers from the shim to make the shim thinner. NOTE:
(5) (6) (7) (8) 4.
(f) When you determine the correct shim thickness, loosen all bolts enough to insert the shim. (g) Install the shim and install bolts and washers. (h) Fully step tighten the bolts that attach the support housing to the engine. Use wire to safety the bolts on the support assembly. Refer to Chapter 20, Safetying Maintenance Practices. Connect the drain hose to the support assembly elbow with the shims in position. Install the compressor. Refer to Compressor Removal/Installation. Install the drive belt. Refer to Compressor Drive Belt Removal/Adjustment.
Compressor Drive Unit Disassembly/Assembly A.
Disassemble Compressor Drive Unit (Refer to Figure 202 and Figure 203). (1) Remove the end cap. (2) Remove the drive shaft, bearings and pulley from the support assembly. (3) Remove the outer bearing from the pulley. (4) Remove the retaining rings that hold the drive shaft and splined coupling outboard of the pulley. (5) Remove the pulley from the splined coupling. (6) Remove the retaining rings that hold the splined coupling to the drive shaft and retain the pulley. (7) Remove the oil seal from the support housing.
B.
Assemble Compressor Drive Unit (Refer to Figure 202 and Figure 203). (1) Apply turbine engine oil to the new oil seal and to the drive shaft surface that the seal is against. (2) Install the new oil seal into the support housing. (3) Install one bearing into the support housing and another in the end cap. (4) Install a retaining ring on the inner groove of the drive shaft. (5) Install a retaining ring on one end of the splined coupling and put that end on the drive shaft. (6) Install the pulley on the splined coupling, then secure the splined coupling and pulley with the retaining rings. (7) Install the end cap with the bolt that does not pass through the spacer. (8) Remove the lamination from the shim(s) as necessary to install the closure disc. NOTE:
5.
Each ply of the shim is 0.002 inch thick. It is important that the shim be no more than 0.002 inch less than the measured gap.
It is necessary to have the shim(s) to keep the closure disc in the correct position.
Compressor Drive Belt Removal/Installation. A.
Remove the Drive Belt (Refer to Figure 201, Figure 202 and Figure 204). (1) Loosen bolt at bottom of compressor. (2) Remove and discard turnbuckle clip from turnbuckle. (3) Loosen turnbuckle enough to pass belt over compressor pulley. (4) Remove bolts securing spacer between end cap and support assembly. (5) Remove belt through opening where spacer was removed.
B.
Install the Drive Belt (Refer to Figures 201, Figure 202 and Figure 204). (1) Insert belt through opening between end cap and support housing. (2) Reinstall spacer. Secure spacer between end cap and support assembly using bolts and washers. (3) Position compressor to allow belt to slip over compressor pulley. (4) Connect clevis end of turnbuckle to compressor. (5) Adjust compressor drive belt. Refer to Adjust Drive Belt.
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R134a Compressor Drive Belt Adjustment Figure 204 (Sheet 1)
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6.
Drive Belt Tension Adjustment A.
Adjust the Drive Belt Tension (Refer to Figure 204). (1) Tension can be checked by using either of the two following methods: (a) A spring scale hooked under the belt at a point midway between compressor drive unit pulley and compressor clutch pulley, pulling perpendicular to the belt. (b) Using a Gates 150 tensiometer. (2) Correct belt tension is a 0.12-inch deflection when a load force of 3.6 to 4.4 pounds is applied to the belt. (3) If belt tension is not correct, adjust as follows: (a) Loosen bolt at bottom of compressor to allow compressor to pivot. (b) Remove and discard clips on turnbuckle.
Table 201. Turnbuckle Adjustment Length Range ADJUSTMENT RANGE
NORMAL
MAX (REF)
MS21252-5LL and MS21252-5RS
4.55 To 5.55 Inch
5.70 Inch
MS21252-5LL and MS21252-5RL
5.40 To 6.40 Inch
6.60 Inch
(c)
Adjust turnbuckle in or out to obtain correct belt tension. NOTE:
(d) (e) 7.
A maximum of three threads must be exposed on adjustment arm clevis. Replace MS21252-5RS clevis with MS21252-5RL. Refer to Table 201 for turnbuckle adjustment ranges.
Install new clip on turnbuckle. Tighten bolt at bottom of compressor.
Condenser Removal/Installation A.
Remove Condenser (Refer to Figure 205). (1) Remove lower left engine cowl. Refer to Chapter 71, Engine Cowling and Nosecap Maintenance Practices. (2) Discharge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. (3) Loosen clamps and remove hoses leading into condenser. Cap all hoses and fittings. (4) Remove bolt securing aft end of condenser to condenser support bracket. (5) Remove bolts, clamps and spacers securing compressor to engine mount. (6) Remove inlet duct and condenser from airplane. (7) Remove bolts and washers securing inlet duct to condenser. Separate inlet duct from condenser. (8) If required, remove seal assemblies from condenser.
B.
Install Condenser (Refer to Figure 205). (1) If required, install seal assemblies to condenser. (2) Attach condenser to inlet duct using bolts and washers. (3) Attach condenser to engine mount using clamps, spacers and hardware as required. Do not tighten at this time. (4) Align holes in right aft corner of condenser with holes in condenser support bracket. Attach using washers and bolts. (5) Tighten clamps, spacers and hardware on engine mount. (6) Reinstall hoses to condenser. Tighten with clamps. (7) Charge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. (8) Install lower left engine cowl. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices.
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R134a Condenser Installation Figure 205 (Sheet 1)
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8.
Receiver-Dryer Removal/Installation A.
Remove Receiver-Dryer (Refer to Figure 206). NOTE:
(1) (2) (3) (4) (5) (6) (7) B.
9.
10.
Anytime air conditioning system has been exposed to atmosphere for any length of time, or when any major components of the system have been replaced, the receiver-dryer must also be replaced.
Discharge system. Refer to Chapter 12, 134A Air Conditioning - Servicing. Disconnect fitting at manifold pressure switch housing. Remove tie-straps and disconnect electrical connector. Remove pressure switch from receiver-dryer. Discard packing and cap open lines. Disconnect fitting from OUT end of receiver-dryer. Loosen clamps and remove receiver-dryer from engine mount. Remove unions from both ends of receiver-dryer. Discard packing and receiver-dryer.
Install receiver-dryer (Refer to Figure 206). (1) Install union fittings and new packing to both ends of new receiver-dryer. (2) Attach receiver-dryer to engine mount and secure clamps. (3) Attach fittings to both ends of receiver-dryer unions. (4) Attach pressure switch with new packing to receiver-dryer. (5) Connect housing cap to housing plug and secure wire using tie-straps. (6) Charge system. Refer to Chapter 12, 134A Air Conditioning - Servicing.
Pressure Switch Removal/Installation A.
Remove Pressure Switch (Refer to Figure 206). (1) Discharge system. Refer to Chapter 12, 134A Air Conditioning - Servicing. (2) Remove tie-straps from electrical wiring. (3) Disconnect electrical connector. (4) Remove pressure switch and packing from receiver-dryer. Discard packing. (5) Cap manifold pressure switch housing to preclude entry of moisture and/or contaminants into system. (6) Check pressure switch for proper operation. Refer to Pressure Switch Functional Test.
B.
Install Pressure Switch (Refer to Figure 206). (1) Install pressure switch in receiver-dryer, using new packing. (2) Connect electrical connector. (3) Secure electrical wiring using tie-straps. (4) Charge system. Refer to Chapter 12, R134a Air Conditioning - Servicing.
Air Conditioning Plumbing Removal/Installation A.
Remove Air Conditioning Plumbing (Refer to Figure 207). (1) Discharge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. NOTE:
(2) (3) (4)
Refrigerant lines in the engine compartment, under the floorboards and in the fuselage sidewalls interconnect the compressor, condenser, receiver-dryer and evaporators.
Gain access to high and low pressure service valves by removing access panel 232AC located between the pilots seats. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. Disconnect plumbing and remove as necessary. Cap all lines and fittings to preclude entry of moisture and/or foreign particles into system.
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R134a Receiver-Dryer Installation Figure 206 (Sheet 1)
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Air Conditioning Plumbing Installation Figure 207 (Sheet 1)
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Air Conditioning Plumbing Installation Figure 207 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL B.
Install Air Conditioning Plumbing (Refer to Figure 207). (1) Remove previously installed caps from lines and install plumbing. NOTE:
It is recommended that all straight thread fittings and O-rings be lubricated with clean refrigerant oil and all taper (pipe) threads be serviced with Teflon tape. Make sure that Teflon tape does not get closer than one to one-half threads from the end of the fitting. If a piece of Teflon tape gets into the system, it can block small orifices.
CAUTION: The use of other thread lubricants, including “Lock-Tite” or other commercial refrigerant lubricants such as “Leak-Lock, is positively prohibited.” C.
Torque lines to valves listed in Table 202 below. NOTE:
All plumbing fittings must be torqued to prevent R134a leakage and must be rechecked after performing an air conditioning leak test.
Table 202. Valve Plumbing Torque Specifications TUBE DIAMETER
TORQUE VALUE
0.250 inch
55 to 65 inch-pounds
0.375 inch
100 to 125 inch-pounds
0.500 inch
200 to 250 inch-pounds
0.750 inch
400 to 500 inch-pounds
11.
D.
Perform leak test of system. Refer to Chapter 12, R134a Air Conditioning - Servicing.
E.
Charge system. Refer to Chapter 12, R134a Air Conditioning - Servicing.
F.
Perform an operational test of the system. Refer to System Operational Test.
G.
Reinstall removed floorboards and panels. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
H.
Reinstall interior equipment. Maintenance Practices.
Refer to Chapter 25, Floor Covering/Control Column Cover -
Wing-Mounted Evaporator Removal/Installation NOTE:
Evaporator removal and installation for both the left and right wing evaporators are typical.
A.
Remove Wing-Mounted Evaporators (Refer to Figure 208). (1) Discharge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. (2) Remove wing root access panel 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Disconnect electrical connector. (4) Disconnect evaporator drain hose from drain tube. (5) Disconnect elbow fitting from bottom of evaporator and cap line. (6) Disconnect expansion valve from evaporator and cap line. (7) Disconnect duct at blower assembly. (8) Remove four bolts and nuts securing evaporator assembly to duct. (9) Pull evaporator assembly far enough aft to allow studs to clear duct. Remove evaporator assembly from airplane.
B.
Install Wing-Mounted Evaporators (Refer to Figure 208). (1) Position evaporator assembly in wing root area with forward studs through holes in duct. Secure evaporator assembly to duct using nuts and bolts. (2) Reconnect and tighten duct at blower assembly.
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R134a Wing Mounted Return Air Check Valve Assembly Figure 208 (Sheet 1)
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R134a Wing Mounted Return Air Check Valve Assembly Figure 208 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6) (7) (8) 12.
13.
14.
Connect expansion valve to evaporator. Connect elbow fitting to bottom of evaporator. Connect evaporator drain hose to drain tube. Connect electrical housing cap to housing plug. Charge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. Install wing root access panel 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Wing Mounted Return Air Check Valve Removal/Installation A.
Remove and Disassemble Wing-Mounted Return Air Check Valve (Refer to Figure 208). (1) Remove wing root access panel 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Remove clamp and flexible duct from outboard duct assembly. (3) Remove screws securing outboard duct assembly to inboard duct assembly. (4) Remove outboard duct assembly from airplane. (5) Disassemble check valve in the following steps: (a) Remove nut at bottom of hinge pin and withdraw hinge pin from outboard duct assembly. This will allow check valve halves and spring to be removed from outboard duct assembly. (b) Remove nut at bottom of pin and withdraw pin from outboard duct assembly.
B.
Assemble and Install Wing Mounted Return Air Check Valve (Refer to Figure 208). (1) Reassemble check valve in the following steps: (a) Assemble check valve halves and spring in outboard duct assembly. Insert hinge pin through duct, valve halves and spring. Secure hinge pin using nut. (b) Insert pin through outboard duct assembly and secure using nut. (c) Make sure check valve operates smoothly and seats fully. (2) Install outboard duct assembly to inboard duct assembly using screws. (3) Attach flexible duct to outboard duct assembly using clamp. (4) Reinstall wing root access panel 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Forward Evaporator Return Air Grill A.
Remove Grill Assembly (Refer to Figure 208). (1) From cabin area, remove screws securing grill to inboard duct assembly.
B.
Install Grill Assembly (Refer to Figure 208). (1) Align holes in grill with holes in headliner and inboard duct assembly. (2) Install screws to secure grill to inboard duct assembly.
Tailcone Mounted Evaporator Removal/Installation A.
Remove Aft Evaporator (Refer to Figure 209). (1) Discharge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. (2) Remove aft cabin partition to gain access to evaporator unit. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices. (3) Disconnect electrical connector. (4) Disconnect evaporator drain hose from bottom of evaporator. (5) Remove recirculated air ducts from duct assembly. (6) Remove fitting from expansion valve. Cap open line. (7) Remove fitting from union on suction hose leading into evaporator. Cap open line. (8) Remove screws securing evaporator to brackets. (9) Remove flexible distribution duct from blower motor and remove evaporator assembly from airplane.
B.
Install Aft Evaporator (Refer to Figure 209). (1) Install evaporator to aft cabin area using screws and washers as required. (2) Attach flexible distribution duct to blower motor. (3) Install suction hose to evaporator.
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R134a Tailcone Mounted Evaporator Installation Figure 209 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (4) (5) (6) (7) (8) 15.
16.
Connect drain line to evaporator. Attach recirculated air ducts to duct assembly. Connect electrical connector. Recharge system. Refer to Chapter 12, R134a Air Conditioning - Servicing. Install aft cabin partition. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices.
Aft Evaporator Distribution and Return Air System Removal/Installation A.
Remove Aft Evaporator Distribution and Return Air Ducts (Refer to Figure 209). (1) Remove the aft cabin partition to get to the evaporator unit. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices. (2) Loosen the clamps that attach the recirculated air ducts to the elbow assemblies. (3) Remove recirculated air ducts from airplane. (4) Loosen the clamp that attaches the flexible distribution duct to the wye duct. (5) Remove the flexible distribution duct from wye duct. (6) Remove the screws that attach the wye duct to distribution duct and remove wye duct from airplane. (7) Remove the screws that attach the distribution duct to the airplane and remove the duct from the airplane.
B.
Install Aft Evaporator Distribution and Return Air System (Refer to Figure 209). (1) Install distribution duct to airplane using screws. (2) Attach wye duct to distribution duct. (3) Attach flexible distribution duct to wye duct using clamp. (4) Attach recirculating air ducts to elbow assemblies using clamps. (5) Install aft cabin partition. Refer to Chapter 25, Rear Cargo Compartment Wall - Maintenance Practices.
System Operational Test A.
Air Conditioning System Operational Test. NOTE: (1)
(2) (3) (4)
17.
Perform system check at ambient temperatures of 55°F or higher.
Engage the following circuit breakers: (a) LEFT VENT BLWR. (b) RIGHT VENT BLWR. (c) AFT VENT BLWR. (d) AIR COND CONT. Move fan switches from HIGH to LOW and note a change in evaporator fan speed. Place air conditioner switch to COOL and activate compressor. Temperature differential across evaporators must be at least 20°F. Measure temperatures at evaporators with dial type thermometers. If evaporators do not cool, refer to 134a Air Conditioning - Troubleshooting.
Pressure Switch Functional Test A.
Test the Pressure Switch (Refer to Figure 206). (1) Discharge the system. Refer to Chapter 12, R134a Air Conditioning - Servicing. (2) Remove the pressure switch and packing from the top of the receiver-dryer. Refer to Pressure Switch Removal/Installation. (3) Do a check for the electrical continuity through the switch. (a) The switch must open between a rising pressure of 350 PSIG to 412 PSIG. (b) When the pressure is lowered 265 PSIG +30 or -30 PSIG below the opening pressure, the switch must close and engage the compressor clutch. (4) Remove the switch and packing from the pressure switch manifold.
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(6)
Apply 355 PSIG to 412 PSIG dry nitrogen pressure to the pressure switch. The switch must open fully. (a) If the switch is closed or does not operate correctly, replace the switch. (b) If the switch opens with the correct pressure, decrease the pressure and make sure the switch closes at 265 PSIG +30 or -30 PSIG. Do a check of the low pressure setting. NOTE:
This is the low pressure setting that protects the compressor.
(a) (b)
(7) (8)
Decrease the pressure and make sure the switch opens with a pressure of 25 to 35 PSIG. If the switch opens, then slowly increase the pressure. Make sure the switch is closed by 35 PSIG. (c) If the switch does not operate correctly at the low pressure setting, replace the switch. Install the pressure switch with a new packing in the top of the receiver-dryer. Refer to Pressure Switch Removal/Installation. Charge the system. Refer to Chapter 12, R134a Air Conditioning - Servicing.
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MODEL 208 MAINTENANCE MANUAL TEMPERATURE CONTROL - DESCRIPTION AND OPERATION 1.
General A.
A cabin heat control panel is located below the instrument. The following description is applicable to all model 208 airplanes: (1) TEMP HOT - The temperature control valve knob is used to control the temperature of air entering the cabin. Turning the knob clockwise raises temperature of the air. (2) BLEED AIR HEAT - This two-position toggle switch is connected electrically to the flow control valve. When placed in the ON position, the flow control valve is opened, diverting a portion of compressor bleed air into the heating system. (3) MIXING AIR - This push-pull knob is used to divert heated air to the cabin. Pull the knob to GRD-PULL position for ground operations. Push the knob in to FLT-PUSH position for flight operations. (4) AFT CABIN/FWD CABIN - This push-pull knob is utilized to divert heated air to outlets on the aft firewall or plenums on the left and right sidewalls. The knob may be moved to any intermediate position to blend air between forward and aft cabin. (5) DEFROST/FWD CABIN - This push-pull knob is utilized to divert heated air to the windshield or to the forward cabin.
B.
Refer to Figure 1 for an illustration of the heating and defrosting system controls.
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Heating and Defrosting System Controls Figure 1 (Sheet 1)
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CHAPTER
AUTO FLIGHT
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
22-00-00
Pages 1-2
Nov 3/2003
22-10-00
Pages 201-211
Apr 1/2010
22-11-00
Pages 401-402
Mar 1/2008
22-11-05
Pages 401-404
Apr 1/2010
22-12-00
Pages 201-209
Jun 1/2011
22-12-00
Pages 601-603
Jun 1/2011
22-12-01
Pages 201-202
Apr 1/2010
22-Title 22-List of Effective Pages 22-Record of Temporary Revisions 22-Table of Contents 22-List of Tasks
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS AUTO FLIGHT - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22-00-00 22-00-00 22-00-00 22-00-00 22-00-00
Page 1 Page 1 Page 1 Page 2 Page 2
KFC-225 AUTOPILOT - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Roll Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Yaw Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Trim Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Roll Servo Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Servo Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . YAW Servo Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Trim Servo Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Trim Rigging Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servo Capstan Clutch Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Set the Autopilot Roll Null . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22-10-00 Page 201 22-10-00 Page 201 22-10-00 Page 201 22-10-00 Page 201 22-10-00 Page 204 22-10-00 Page 204 22-10-00 Page 207 22-10-00 Page 207 22-10-00 Page 208 22-10-00 Page 208 22-10-00 Page 209 22-10-00 Page 210 22-10-00 Page 211 22-10-00 Page 211
KING KFC-225 AUTOPILOT COMPUTER - REMOVAL/INSTALLATION . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Autopilot Computer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22-11-00 Page 401 22-11-00 Page 401 22-11-00 Page 401
KING KAP/KFC-150 AUTOPILOT COMPUTER - REMOVAL/INSTALLATION. . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Autopilot Computer (Installed in the Center Console) Removal/Installation. . . . . . . . Autopilot Computer (Installed in the Radio Panel) Removal/Installation . . . . . . . . . . .
22-11-05 Page 401 22-11-05 Page 401 22-11-05 Page 401 22-11-05 Page 401
GFC 700 AUTOPILOT - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Roll Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Yaw Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Trim Servo Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitch Trim Rigging Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servo Capstan Clutch Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22-12-00 Page 201 22-12-00 Page 201 22-12-00 Page 201 22-12-00 Page 204 22-12-00 Page 204 22-12-00 Page 207 22-12-00 Page 209 22-12-00 Page 209
GFC 700 AUTOPILOT - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Autopilot Servos Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin Autopilot (GFC 700) Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22-12-00 Page 601 22-12-00 Page 601 22-12-00 Page 601 22-12-00 Page 602
GARMIN GMC-710 AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GMC-710 AFCS Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22-12-01 Page 201 22-12-01 Page 201 22-12-01 Page 201
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LIST OF TASKS 22-12-00-640
Autopilot Servos Lubrication
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22-12-00-720
Garmin Autopilot (GFC 700) Functional Check
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MODEL 208 MAINTENANCE MANUAL AUTO FLIGHT - GENERAL 1.
2.
General A.
The model 208 and 208 Cargomaster can have the Sperry 400B Navomatic Autopilot or the Sperry 400B Integrated Flight Control System. The Model 208 and 208B can also have the King (KAP/KFC150 or KFC-225) Autopilot/Flight Control System or the King (KFC-250) Flight Control System.
B.
These systems provide a means of automatically or manually controlling the flight of the airplane. Included are the following components which provide for tracking of any magnetic heading, automatic intercept and tracking of VOR radials or ILS localizer and glide slope beams, and includes automatic pitch synchronization and trim, manual turn (400B only) and pitch command, altitude hold, back course switching, NAV 1 or NAV 2 receiver selection, pitch attitude disengagement with an associated warning tone, autopilot annunciator lights, and A/P ROLL TRIM indicator (400B only) to indicate any adjustments necessary to neutralize autopilot roll effort and prior-to-flight test function.
Description A.
Sperry (Type AF-550A) 400B Autopilot (Model 208 only). (1) This autopilot system consists of the autopilot controller, accessory unit, flux detector, directional and horizontal gyros, roll, pitch and pitch trim actuators, slaving accessory, computer amplifier, altitude sensor, warning horn, airspeed switch, roll trim indicator, and autopilot annunciator lights.
B.
Sperry (Type IF-550A) 400B Integrated Flight Control System (Model 208 only). (1) This optional system incorporates go-around and pitch synchronization functions and a mode selector in addition to the components of the 400B Autopilot (Type AF-550A) System. This system utilizes a flight director indicator instead of an attitude gyro and a slaved HSI installed to replace the standard directional gyro.
C.
King (KFC-250) Flight Control System (Airplanes 20800007 Thru 20800083, and 208B0044 Thru 208B0147). (1) This autopilot/flight director system consists of a mode controller, a mode annunciator panel, an attitude flight command indicator, a slaved pictorial navigation indicator, a slaving accessory and compensator unit, control wheel switches for autopilot disconnect/trim interrupt, control wheel steering and manual electric trim control switches, a go-around button is mounted on the power lever, remote mounted roll trim and pitch actuators, A/P computer, air data unit and inverter, and a panel mounted flight control system switch panel incorporating attitude gyro fast erect switch, inverter selector switch, trim test switch, autopilot roll rate monitor test switch and a flight director/autopilot NAV 1/NAV 2 selector switch.
D.
King (KAP/KFC-150) Autopilot/Flight Control System. (1) This system has a mode annunciator, attitude/flight command indicator, slaved pictorial navigation indicator, slaving accessory and compensator unit, and a combined computer/ controller unit. The combined computer/controller unit contains computer functions, vertical modes, mode control buttons, and an altitude sensor. The control wheel switches supply the autopilot disconnect/trim interrupt, control wheel steering, and manual electric trim control. The roll, pitch, and trim actuators are installed in different locations in the airplane.
E.
King (KFC-225) Autopilot/Flight Control System. (1) This system has a flight command indicator, slaved pictorial navigation indicator, go-around mode, slaving accessory and compensator unit, external configuration module, data plug and a combined computer/controller/annunciator unit. The system can also have a mode annunciator. The combined computer/controller/annunciator unit contains computer functions, vertical modes, yaw damper, mode control buttons, annunciator lights and an altitude sensor. The control wheel switches supply autopilot disconnect/trim interrupt, control wheel steering, and manual electric trim control. The roll, pitch, and trim actuators are installed in different locations in the airplane.
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3.
Operation A.
4.
For operation of the various systems, refer to the Approved Airplane Flight Manual and Pilot’s Operating Handbook.
Maintenance Practices A.
For (1) (2) (3)
maintenance practices related to the above systems, refer to the following: Chapter 34, Navigation - General. Chapter 76, Quadrant Assembly and Controls - Maintenance Practices. Model 208 Avionic Installations Service/Parts Manual (Airplanes 20800001 thru 20800347 and Airplanes 208B0001 thru 208B0899). (4) Model 208 customized avionic wiring diagram that was supplied with the airplane (Airplanes 20800348 and On and Airplanes 208B0900 and On). (5) Specific vendor publications listed in Introduction - List of Publications.
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MODEL 208 MAINTENANCE MANUAL KFC-225 AUTOPILOT - MAINTENANCE PRACTICES 1.
General A.
2.
3.
A three-axis autopilot with heading hold is installed as standard equipment on the airplane. The dual-axis system gives both vertical speed and altitude hold selection. Altitude alerting and altitude preselection are optional features with the two-axis autopilot system.
Roll Servo Removal/Installation A.
Remove the Roll Servo (Refer to Figure 201). (1) Make sure the MASTER and AVIONICS switches are in the off position. (2) Remove the access panel (232DR). Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (3) Disconnect the electrical connector from the roll servo. (4) Release the control cable tension and loosen the roll servo control cable at the turnbuckle. (5) Remove the bolts and washers that attach the roll servo to the bracket assembly. (6) Remove the roll servo from the airplane. (7) Do an inspection of the roll servo. Refer to Roll Servo Inspection.
B.
Install the Roll Servo (Refer to Figure 201). (1) Put the roll servo in position on the bracket assembly and attach with the bolts and washers. (2) Connect the electrical connector to the roll servo. (3) Install the roll servo control cable on the roll servo. (4) Make sure the aileron and bell crank are in the neutral position. (5) Wind the control cable around the servo drum approximately 1.25 turns in each direction from the swaged ball (drum ball detent inboard). (6) Make sure the flanges of the control cable guard do not touch the control cable. (7) Make sure the flanges of the control cable guard are on each side of the notches around the outer edge of the mount. (8) Adjust the roll servo control cable tension to 12 pounds, +2 or -2 pounds. (9) Tighten bolts on cable clamps to 50 inch pounds, + 5 or - 5 inch pounds (5.64 N-m, +0.564 or 0.564 N-m). (10) Install the access panel. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (11) Put the MASTER and AVIONICS switches in the ON position. (12) Do a operational test of the autopilot. Refer to Introduction in the List of Manufacturers Technical Publications for the manufacturer's installation manual. Refer to Chapter 6, Access/Inspection Plates - Description and Operation.
Pitch Servo Removal/Installation A.
Remove Pitch Servo (Refer to Figure 202). (1) Make sure the MASTER and AVIONICS switches are in the off position. (2) Remove the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices. (3) Disconnect the electrical connector from the pitch servo. (4) Release the control cable tension and loosen the pitch servo control cable at the turnbuckle. (5) Remove the bolts and washers that attach the pitch servo to the bracket assembly. (6) Remove the pitch servo from the airplane. (7) Do an inspection of the pitch servo. Refer to Pitch Servo Inspection.
B.
Install the Pitch Servo (Refer to Figure 202). (1) Put the pitch servo in position on the bracket assembly and attach with the bolts and washers. (2) Connect the electrical connector to the pitch servo. (3) Install the pitch servo control cable on the pitch servo actuator. (4) Make sure the elevator and bell crank are in the neutral position. (5) Wind the control cable around the servo drum approximately 1.25 turns in each direction from the swaged ball (drum ball detent inboard). (6) Make sure the flanges of the control cable guard do not touch the control cable.
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MODEL 208 MAINTENANCE MANUAL
Autopilot Roll Servo Installation Figure 201 (Sheet 1)
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Autopilot Pitch Servo Installation Figure 202 (Sheet 1)
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Make sure the flanges of the control cable guard are on each side of the notches around the outer edge of the mount. (8) Make sure that the main cables tension is set correctly. (9) Use the turnbuckle to adjust the pitch servo cable tension to 20 pounds, +5 or -5 pounds. (10) Install the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. (11) Put the MASTER and AVIONICS switches in the ON position. (12) Do a operational test of the autopilot. Refer to Introduction in the List of Manufacturers Technical Publications for the manufacturer's installation manual. 4.
5.
Yaw Servo Removal/Installation A.
Remove Yaw Servo (Refer to Figure 203). (1) Make sure the MASTER and AVIONICS switches are in the off position. (2) Remove the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices. (3) Disconnect the electrical connector from the yaw servo. (4) Release the control cable tension and loosen the yaw servo control cable at the turnbuckle. (5) Remove the bolts and washers that attach the yaw servo to the bracket assembly. (6) Remove the yaw servo from the airplane. (7) Do an inspection of the yaw servo. Refer to Yaw Servo Inspection.
B.
Install the Yaw Servo (Refer to Figure 203). (1) Put the yaw servo in position on the bracket assembly and attach with the bolts and washers. (2) Connect the electrical connector to the yaw servo. (3) Install the yaw servo control cable on the yaw servo actuator. (4) Make sure the rudder and bell crank are in the neutral position. (5) Wind the control cable around the servo drum approximately 1.25 turns in each direction from the swaged ball (drum ball detent inboard). (6) Make sure the flanges of the control cable guard do not touch the control cable. (7) Make sure the flanges of the control cable guard are on each side of the notches around the outer edge of the mount. (8) Make sure that the main cables tension is set correctly. (9) Use the turnbuckle to adjust the yaw servo cable tension to 20 pounds, +5 or -5 pounds. (10) Install the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. (11) Put the MASTER and AVIONICS switches in the ON position. (12) Do a operational test of the autopilot. Refer to Introduction in the List of Manufacturers Technical Publications for the manufacturer's installation manual.
Pitch Trim Servo Removal/Installation A.
Remove the Pitch Trim Servo (Refer to Figure 204). (1) Make sure the MASTER and AVIONICS switches are in the off position. (2) Get access to the pitch trim servo. Refer to Chapter 27, Electric Elevator Trim - Removal/ Installation. NOTE:
(3) (4) (5)
The Electric Elevator Trim - Removal/Installation section gives the method necessary to remove and install the electric elevator trim motor that is installed on some models. This same method is valid to remove and install the pitch trim servo.
Disconnect the electrical connector from the pitch trim servo. Remove the pitch trim servo from the airplane. Do an inspection of the pitch trim servo. Refer to Pitch Trim Servo Inspection.
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MODEL 208 MAINTENANCE MANUAL
Autopilot Yaw Servo Installation Figure 203 (Sheet 1)
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Autopilot Pitch Trim Servo Installation Figure 204 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
Install the Pitch Trim Servo (Refer to Figure 204). (1) Install the pitch trim servo in the airplane. Refer to Chapter 27, Electric Elevator Trim - Removal/ Installation. NOTE:
(2) (3) (4) (5) 6.
The Electric Elevator Trim - Removal/Installation section gives the method necessary to remove and install the electric elevator trim motor that is installed on some models. This same method is valid to remove and install the pitch trim servo.
Connect the electrical connector to the pitch trim servo. Close access to the pitch trim servo. Refer to Chapter 27, Electric Elevator Trim - Removal/ Installation. Put the MASTER and AVIONICS switches in the ON position. Do an operational test of the autopilot. Refer to Introduction of the List of Manufacturers Technical Publications, for the manufacturer's installation manual.
Roll Servo Inspection A.
Do an Inspection of the Roll Servo (Refer to Figure 201). (1) Remove the servo cover.
CAUTION: Make sure the maintenance personnel and the table are electrically grounded. Do disassembly or assembly of the servo at an electrostatic-safe area.
(2)
(3) (4) (5) 7.
(a) Put an electrical ground on the maintenance personnel and table. (b) Remove the two screws that attach the cover to the unit. (c) Carefully remove the cover over the wiring harness. (d) Put the servo on the table so the inner parts of the unit will not be damaged. Do inspection of the solenoid and clutch. (a) Make sure the solenoid shaft moves freely in and out of the solenoid body. (b) Make sure there is no dirt, contamination or corrosion around the solenoid shaft. (c) Make sure the release spring freely pulls the shaft out of the solenoid and against the stop fitting. (d) Make sure the pinion gear turns and does not touch the clutch gears. Do a general inspection of the roll servo. (a) Examine the electrical wiring for indication of wear or damage of the insulation. (b) Examine the servo for any loose hardware or other defects. Install the cover. (a) Carefully put the cover in position. (b) Install the screws with Loctite 222 or Loctite 242. Remove the servo capstan assembly and do a check of the slip-clutch torque setting (Refer to Servo Capstan Clutch Adjustment).
Pitch Servo Inspection A.
Do an Inspection of the Pitch Servo (Refer to Figure 202). (1) Remove the servo cover.
CAUTION: Make sure the maintenance personnel and the table are electrically grounded. Do disassembly or assembly of the servo at an electrostatic-safe area. (a) (b) (c)
Put an electrical ground on the maintenance personnel and table. Remove the two screws that attach the cover to the unit. Carefully remove the cover from the wiring harness.
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CAUTION: Do not move any wires, tie wraps or the spring clamp. The position of each is set by the manufacturer and is necessary for correct operation. (2)
(3) (4)
(5) (6) 8.
(d) Put the servo on the table so the inner parts of the unit will not be damaged. Do inspection of the solenoid and clutch. (a) Make sure the solenoid shaft moves freely in and out of the solenoid body. (b) Make sure there is no dirt, contamination or corrosion around the solenoid shaft. (c) Make sure the release spring freely pulls the shaft out of the solenoid and against the stop fitting. (d) Make sure the pinion gear turns and does not touch the clutch gears. Do a general inspection. (a) Examine the electrical wiring for indication of wear or damage of the insulation. (b) Examine the servo for any loose hardware or other defects. Do an inspection of the pitch servo motor. (a) Put the servo in position so the baseplate is on the bottom side of the unit. (b) Hold the top section of the motor and carefully turn the motor shaft. (c) The motor shaft must turn freely from side to side a small quantity. Install the cover. (a) Carefully put the cover in position. (b) Install the screws with Loctite 222 or Loctite 242. Remove the servo capstan assembly and do a check of the slip-clutch torque setting (Refer to Servo Capstan Clutch Adjustment).
YAW Servo Inspection A.
Do an Inspection of the Pitch Trim Servo (Refer to Figure 203). (1) Remove the servo cover.
CAUTION: Make sure the maintenance personnel and the table are electrically grounded. Do disassembly or assembly of the servo at an electrostatic-safe area.
(2)
(3) (4) (5) 9.
(a) Put an electrical ground on the maintenance personnel and table. (b) Remove the two screws that attach the cover to the unit. (c) Carefully remove the cover over the wiring harness. (d) Put the servo on the table so the inner parts of the unit will not be damaged. Do inspection of the solenoid and clutch. (a) Make sure the solenoid shaft moves freely in and out of the solenoid body. (b) Make sure there is no dirt, contamination or corrosion around the solenoid shaft. (c) Make sure the release spring freely pulls the shaft out of the solenoid and against the stop fitting. (d) Make sure the pinion gear turns and does not touch the clutch gears. Do a general inspection. (a) Examine the electrical wiring for indication of wear or damage of the insulation. (b) Examine the servo for any loose hardware or other defects. Install the cover. (a) Carefully put the cover in position. (b) Install the screws with Loctite 222 or Loctite 242. Remove the servo capstan assembly and check the slip-clutch torque setting (Refer to Servo Capstan Clutch Adjustment).
Pitch Trim Servo Inspection A.
Do an Inspection of the Pitch Trim Servo (Refer to Figure 204). (1) Remove the servo cover.
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CAUTION: Make sure the maintenance personnel and the table are electrically grounded. Do disassembly or assembly of the servo at an electrostatic-safe area.
(2)
(3) (4) (5) 10.
(a) Put an electrical ground on the maintenance personnel and table. (b) Remove the two screws that attach the cover to the unit. (c) Carefully remove the cover over the wiring harness. (d) Put the servo on the table so the inner parts of the unit will not be damaged. Do inspection of the solenoid and clutch. (a) Make sure the solenoid shaft moves freely in and out of the solenoid body. (b) Make sure there is no dirt, contamination or corrosion around the solenoid shaft. (c) Make sure the release spring freely pulls the shaft out of the solenoid and against the stop fitting. (d) Make sure the pinion gear turns and does not touch the clutch gears. Do a general inspection. (a) Examine the electrical wiring for indication of wear or damage of the insulation. (b) Examine the servo for any loose hardware or other defects. Install the cover. (a) Carefully put the cover in position. (b) Install the screws with Loctite 222 or Loctite 242. Remove the servo capstan assembly and check the slip-clutch torque setting (Refer to Servo Capstan Clutch Adjustment).
Pitch Trim Rigging Inspection A.
Do a check of the pitch trim rigging. (1) Attach an inclinometer to the trim tab. (2) Put the trim tab in the 0 degree position. (3) Manually operate the trim tab to the up and down limits. (4) Record the limits of travel. (5) Put an observer at the right-hand access opening of the tailcone. (6) Put the electrical trim to the full nose-up position until the observer sees the clutch slip. (7) Turn the manual trim wheel nose-up (test load condition) 1/4 turn more while the clutch slips. (8) Make sure the swaged ball on the control cable assembly does not turn aft of the tangent point. (9) Release the trim wheel and disengage the autopilot. (10) Manually operate the trim to the full nose-up position. (11) Do a check of the trim tab position with an inclinometer. (12) Trim tab position that is greater than the limits of travel values recorded is an indication that the stop blocks slipped. (a) Do the trim system rigging again. (b) Make sure the stop block bolts torque is correct. (c) Repeat the check of the pitch trim rigging. (13) If necessary, make adjustments to the swaged ball position. (a) Put the control cable assembly chain in the applicable position on the gear teeth of the actuator sprocket. NOTE:
One chain link adjustment is related to approximately 17 degrees of travel on the capstan.
Apply the applicable tension to the control cable and repeat the check of the pitch trim rigging. (14) Do the procedure again for the full nose-down trim condition. (b)
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11.
Servo Capstan Clutch Adjustment A.
Do a check of the clutch torque setting. Refer to Chapter 27, Aileron and Spoiler System - Adjustment/ Test, Servo Slip Clutch Torque Setting to determine the servo mount part number and setting required for each axis of the airplane. (1) Remove the servo capstan. (2) Remove the control cable guard from the servo capstan. (3) Attach the servo capstan on the capstan test stand. Refer to Aileron and Spoiler System Adjustment/Test, Servo Slip Clutch Torque Setting. (4) Place the adapter tool over the servo capstan. (5) Insert the adapter pin from the straight up position to attach the adapter tool. (6) Insert the torque wrench. (7) Apply 28 VDC (1 amp maximum) electrical power to the test stand. (8) Do a check of the torque reading with the test stand motor in the clockwise operation. NOTE:
(9)
The check of the torque reading will be done three times.
(a) Put the capstan switch in the clockwise position. (b) Record the torque reading of the torque wrench. (c) Put the switch in the off position. Do a check of the torque reading with the test stand motor in the counterclockwise operation. NOTE:
The check of the torque reading will be done three times.
(a) Put the capstan switch in the counterclockwise position. (b) Record the torque reading of the torque wrench. (c) Put the switch in the off position. (10) Average the six torque readings. NOTE:
The torque reading to be used is the average of the six torque readings.
(11) Refer to Table 201 for the correct torque reading of the servo capstan. Table 201. KAP-150 Autopilot Servo Clutch Torque Setting 208
208B
Roll Servo Capstan
33, +3 or -3 inch-pounds (3.7, +0.33 or -0.33 N-m)
38, +4 or -4 inch-pounds (4.3, +0.45 or -0.45 N-m)
Pitch Servo Capstan
43, +4 or -4 inch-pounds (4.9, +0.45 or -0.45 N-m)
43, +4 or -4 inch-pounds (4.9, +0.45 or -0.45 N-m)
Pitch Trim Servo Capstan
45, +5 or -5 inch-pounds (5.1, +0.56 or -0.56 N-m)
45, +5 or -5 inch-pounds (5.1, +0.56 or -0.56 N-m)
Yaw Servo Capstan
50, +5or -5 onch-pounds (5.6, +0.56 or -0.56 N-m)
50, +5or -5 onch-pounds (5.6, +0.56 or -0.56 N-m)
(a)
If the torque indication is below the value given in Table 201, rotate the clutch adjust nut clockwise and do the check of the torque readings again. (b) If the torque indication is above the value given in Table 201, rotate the clutch adjust nut counterclockwise and do the check of the torque readings again. (12) Record the slip clutch torque indication, airplane type, axis, and date on the decal attached to the servo mount body. (13) Install the control cable guard on the servo capstan. (14) Install the servo capstan.
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12.
Set the Autopilot Roll Null A.
Set (1) (2) (3) (4)
the Autopilot Roll Null (If the Autopilot is Installed). Make sure the autopilot flight computer completes the preflight test. Disconnect the roll servo connector from the airplane harness. Apply a ground to pin K of the harness connector. Connect a digital multimeter across the harness connector at pins D and L to monitor the servo drive voltage. (5) Push the autopilot AP button on the autopilot flight computer to engage it. (a) Make sure the default ROL mode is set. NOTE: (b) (c) (d)
For example, the HDG, NAV or APR modes are not engaged.
Use a DMM to measure the DC voltage across pins D and L of the roll servo harness connector. Adjust the pot until a value of 0 volts, +0.020 or -0.020 volts are measured. 1 If the end of the pot movement is reached before the servo drive is nulled, disengage the autopilot, turn the pot fully to the opposite stop and then engage the autopilot. The roll null adjustment range emulates a four turn pot that lets the method of the pot adjustment range to be set. NOTE:
This adjustment lets offsets be in the roll axes. coordinator.
This includes the turn
(e) Continue to turn the pot to null the voltage. (6) Connect the airplane roll servo harness connector to the servo connector. Tools, Equipment and Materials NOTE:
Equivalent alternatives can be used for the items that follow.
NAME
NUMBER
MANUFACTURER
USE
Test Stand
071-06028-0000
Honeywell International, Inc. 1 Technology Center Olathe, KS 66061
To hold the servo mount in position while the servo clutch torque setting is adjusted.
Adapter Tool
071-06021-0003
Honeywell International, Inc.
To adjust the servo clutch torque setting.
Adapter Pin
071-06021-0002
Honeywell International, Inc.
To adjust the servo clutch torque setting.
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MODEL 208 MAINTENANCE MANUAL KING KFC-225 AUTOPILOT COMPUTER - REMOVAL/INSTALLATION 1.
General A.
2.
This section gives procedures for the removal and installation of the King KFC-225 Autopilot Computer.
Autopilot Computer Removal/Installation A.
Remove the Autopilot Computer (Refer to Figure 401). (1) Disengage the A/P CONT circuit breaker on the left circuit breaker panel. (2) Put a 3/32 hex wrench through the front panel of the autopilot computer and engage the lock mechanism. (3) Turn the lock mechanism counterclockwise to loosen and unlock the autopilot computer from the mounting rack. (4) Remove the hex wrench from the autopilot computer. (5) Pull the autopilot computer from the mounting rack. (6) Remove the autopilot computer from the airplane.
B.
Install the Autopilot Computer (Refer to Figure 401). (1) Put the autopilot computer in position in the mounting rack. (2) Put a 3/32 hex wrench through the front panel of the autopilot computer and engage the lock mechanism. (3) Turn the lock mechanism counterclockwise and move the autopilot computer forward into the mounting rack until it stops. (4) Turn the lock mechanism clockwise until the autopilot computer bezel is ßush with the radio panel. NOTE: (5) (6) (7)
This will lock the autopilot computer in the mounting rack and engage the connectors.
Remove the hex wrench from the autopilot computer. Engage the A/P CONT circuit breaker on the left breaker panel. If a new unit is installed or the unit is calibrated, do a system alignment. Refer to Introduction, the List of Manufacturers Technical Publications for the manufacturer's installation manual.
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Autopilot Computer Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL KING KAP/KFC-150 AUTOPILOT COMPUTER - REMOVAL/INSTALLATION 1.
2.
General A.
This section gives procedures for the removal and installation of the King KAP/KFC-150 Autopilot Computer and the yaw damper switch.
B.
The autopilot computer and optional yaw damper switch is found in the center console or in the radio panel.
Autopilot Computer (Installed in the Center Console) Removal/Installation A.
Remove the Autopilot Computer (Refer to Figure 401). (1) Disengage the A/P CONT circuit breaker on the left circuit breaker panel. (2) Loosen the hex nut that holds the autopilot computer in the center console. (a) Use a long hex wrench through the autopilot front panel to get access to the autopilot hex nut. (3) Lift the autopilot computer from the center console. (4) Disconnect the static hose. (5) Disengage the yaw damper switch from the top panel of the center console. (6) Remove the yaw damper switch from the center console. (7) Remove the autopilot computer and yaw damper switch from the airplane.
B.
Install the Autopilot Computer (Refer to Figure 401). (1) Install the yaw damper switch in the top panel of the center console. (a) Connect the electrical connector from the autopilot computer to the yaw damper switch. (2) Connect the static hose to the autopilot computer. (3) Put the autopilot computer in place in the mounting rack. (4) Use a long hex wrench and tighten the autopilot computer hex screw.
CAUTION: Make sure that the hex screw is not torqued too much. This will help prevent damage to the hex screw. (5) 3.
Engage the A/P CONT circuit breaker on the left breaker panel.
Autopilot Computer (Installed in the Radio Panel) Removal/Installation A.
Remove the Autopilot Computer (Refer to Figure 402). (1) Disengage the A/P CONT circuit breaker on the left circuit breaker panel. (2) Put a 3/32 hex wrench through the front panel of the autopilot computer and engage the lock mechanism. (3) Turn the lock mechanism counterclockwise to loosen and unlock the autopilot computer from the mounting rack. (4) Remove the hex wrench from the autopilot computer. (5) Pull the autopilot computer from the mounting rack. (6) Loosen the clamp that holds the static hose to the autopilot computer. (7) Disconnect the static hose from the autopilot computer. (8) Remove the autopilot computer from the airplane. (9) If necessary, remove the yaw damper switch from the radio panel. (a) Disconnect the electrical connector from the yaw damper switch. (b) Remove the yaw damper switch from the radio panel. (c) Remove the yaw damper switch from the airplane.
B.
Install the Autopilot Computer (Refer to Figure 402). (1) If necessary, install the yaw damper switch in the radio panel. (a) Connect the electrical connector from the autopilot computer to the yaw damper switch. (2) Connect the static hose to the autopilot computer. (3) Tighten the clamp that holds the static hose to the autopilot computer. (4) Put the autopilot computer in position in the mounting rack.
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Autopilot Computer with Yaw Damper Switch Figure 401 (Sheet 1)
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Autopilot Computer Figure 402 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7)
Put a 3/32 hex wrench through the front panel of the autopilot computer and engage the lock mechanism. Turn the lock mechanism counterclockwise and move the autopilot computer forward into the mounting rack until it stops. Turn the lock mechanism clockwise until the autopilot computer bezel is flush with the radio panel. NOTE:
(8) (9)
This will lock the autopilot computer in the mounting rack and engage the connectors.
Remove the hex wrench from the autopilot computer. Engage the A/P CONT circuit breaker on the left breaker panel.
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MODEL 208 MAINTENANCE MANUAL GFC 700 AUTOPILOT - MAINTENANCE PRACTICES 1.
General A.
2.
A three-axis autopilot with heading hold is installed as standard equipment on the airplane.
Roll Servo Removal/Installation A.
Remove the Roll Servo (Refer to Figure 201). (1) Disconnect electrical power from the aircraft. (2) Remove the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (3) Remove the access panel 232DR. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (4) Disconnect the electrical connector from the roll servo. (5) Release the control cable tension and loosen the roll servo bridle cable at the turnbuckle. (6) For airplanes 20800416, 20800514, 20800518 thru 20800523 and airplanes 208B2090 and 208B2168 thru 208B2230, remove the screws and the nuts that attach the roll servo fairlead cover to the fairlead guard assembly. (a) Remove the roll servo fairlead cover from the fairlead guard assembly and the airplane. (7) For airplanes 20800524 and On, 208B2231 and On, and airplanes that incorporate CAB10-9, remove the screws that attach the roll servo fairlead cover to the fairlead guard assembly. (a) Remove the roll servo fairlead cover from the fairlead guard assembly and the airplane. (8) Remove the bolts and washers that attach the roll servo to the bracket assembly. (9) Remove the roll servo from the airplane. (10) Do an inspection of the roll servo. Refer to Roll Servo Inspection.
B.
Install the Roll Servo (Refer to Figure 201). (1) Put the roll servo actuator in position on the torque mount and attach with bolts and washers. (a) Torque the bolts to 45 inch-pounds, +5 or -5 inch-pounds. (2) Connect the electrical connector to the roll servo. (3) Install the roll servo bridle cable on the roll servo. (4) Make sure the aileron and bell crank are in the neutral position. (5) Wind the bridle cable around the servo drum approximately 1.25 turns in each direction from the swaged ball (drum ball detent inboard). (6) Make sure the flanges of the bridle cable guard do not touch the bridle cable. (7) Make sure the flanges of the bridle cable guard are on each side of the notches around the outer edge of the mount. (8) Make sure that the primary control cables tension is correct before checking or adjusting bridle cable tension (Refer to Chapter 27, Aileron and Control Column - Maintenance Practices, Rigging Aileron System). (9) Make sure that the roll servo bridle cable tension is 12 pounds, +2 or -2 pounds. (10) If the bridle cable tension is not correct do the steps that follow: (a) Remove panel 232AC (Refer to Chapter 6, Access/Inspection Plates - Description and Operation.) (b) Set the control wheels with the ailerons in neutral position. (c) Put a bar across the control wheels and tape the bar to the control wheels. NOTE: (d) (e) (f) (g) (h) (i)
The bar connects the wheels and locks them in the neutral position.
Make sure that the roll servo drum is oriented with the swagged ball on the roll bridle cable at the 12 o'clock position. If either end of the bridle cable is slack, loosen the clamp at that end and move it away from the servo until the cable is no longer slack. Torque the three screws on the right bridle clamp to 25 to 30 inch pounds. Loosen the left bridle clamp just enough to allow it to move. Hold and pull the left bridle cable clamp until the bridle cable tension is 12 pounds, +2 or -2 pounds and torque the three screws on the left bridle cable clamp to 25 to 30 inch pounds. Remove the bars from the control wheels.
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Autopilot Roll Servo Installation Figure 201 (Sheet 1)
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Autopilot Roll Servo Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (11) For airplanes 20800416, 20800514, 20800518 thru 20800523, and airplanes 208B2090 and 208B2168 thru 208B2230, put the roll servo fairlead cover in its position on the fairlead guard assembly. (a) Install the screws and the nuts. (12) For airplanes 20800524 and On, 208B2231 and On, and airplanes that incorporate CAB10-9, put the roll servo fairlead cover in its position on the fairlead guard assembly. (a) Install the screws. (13) Install access panels 232AC and 232DR. Refer to Chapter 6, Access/Inspection Plates Description and Operation. (14) Connect electrical power from the aircraft. (15) Do a operational test of the autopilot. Refer to Introduction in the List of Manufacturers Technical Publications for the manufacturer's installation manual. 3.
4.
Pitch Servo Removal/Installation A.
Remove Pitch Servo (Refer to Figure 202). (1) Disconnect electrical power from the aircraft. (2) Remove the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices. (3) Disconnect the electrical connector from the pitch servo. (4) Release the bridle cable tension and loosen the pitch servo bridle cable at the turnbuckle. (5) Remove the bolts and washers that attach the pitch servo to the bracket assembly. (6) Remove the pitch servo from the airplane. (7) Do an inspection of the pitch servo. Refer to Pitch Servo Inspection.
B.
Install the Pitch Servo (Refer to Figure 202). (1) Put the pitch servo in position on the bracket assembly and attach with the bolts and washers. (a) Torque the bolts to 45 inch-pounds, +5 or -5 inch-pounds. (2) Connect the electrical connector to the pitch servo. (3) Install the pitch servo bridle cable on the pitch servo actuator. (4) Make sure the elevator and bell crank are in the neutral position. (5) Wind the bridle cable around the servo drum approximately 1.25 turns in each direction from the swaged ball (drum ball detent inboard). (6) Make sure the flanges of the bridle cable guard do not touch the bridle cable. (7) Make sure the flanges of the bridle cable guard are on each side of the notches around the outer edge of the mount. (8) Use the turnbuckle to adjust the pitch servo cable tension to 20 pounds, +5 or -5 pounds. (9) Install the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. (10) Connect electrical power to the aircraft. (11) Do an operational test of the autopilot. Refer to Introduction in the List of Manufacturers Technical Publications for the manufacturer's installation manual.
Yaw Servo Removal/Installation A.
Remove Yaw Servo (Refer to Figure 203). (1) Disconnect electrical power from the aircraft. (2) Remove the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices. (3) Disconnect the electrical connector from the yaw servo. (4) Release the bridle cable tension and loosen the yaw servo bridle cable at the turnbuckle. (5) Remove the bolts and washers that attach the yaw servo to the bracket. (6) Remove the yaw servo from the airplane. (7) Do an inspection of the yaw servo. Refer to Yaw Servo Inspection.
B.
Install the Yaw Servo (Refer to Figure 203). (1) Put the yaw servo in position on the bracket and attach with the bolts and washers. (a) Torque the bolts to 45 inch-pounds, +5 or -5 inch-pounds. (2) Connect the electrical connector to the yaw servo.
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Autopilot Pitch Servo Installation Figure 202 (Sheet 1)
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Autopilot Yaw Servo Installation Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6) (7) (8) (9) (10)
(11) (12) (13) 5.
Install the yaw servo bridle cable on the yaw servo actuator. Make sure the rudder and bell crank are in the neutral position. Wind the bridle cable around the servo drum approximately 1.25 turns in each direction from the swaged ball (drum ball detent inboard). Make sure the flanges of the bridle cable guard do not touch the bridle cable. Make sure the flanges of the bridle cable guard are on each side of the notches around the outer edge of the mount. Make sure that the primary control cables tension is correct before checking or adjusting bridle cable tension (Refer to Chapter 27, Rudder - Maintenance Practices, Rudder System Rigging). Make sure that the yaw servo bridle cable tension is 20 pounds, +2 or -2 pounds. If the bridle cable tension is not correct do the steps that follow: (a) Make sure that the rudder is in the neutral position. (b) Make sure that the swagged ball on the yaw bridle cable is positioned on the forward side of the yaw servo drum, centered between the two forward yaw servo attachment bolts. (c) If either end of the bridle cable is slack, loosen the clamp at that end and move it away from the servo until the cable is no longer slack. (d) Torque the three screws on each bridle cable clamp to 25 to 30 inch pounds. (e) Remove the turnbuckle clips and use the turnbuckle to set the yaw bridle cable tension to 20 pounds +5 or -5 pounds. (f) Install the turnbuckle clips on the turnbuckle. Install the Rear Compartment Wall. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. Connect electrical power to the aircraft. Do an operational test of the autopilot. Refer to Introduction in the List of Manufacturers Technical Publications for the manufacturer's installation manual.
Pitch Trim Servo Removal/Installation A.
Remove the Pitch Trim Servo (Refer to Figure 204). (1) Disconnect electrical power from the aircraft. (2) Get access to the pitch trim servo. Refer to Chapter 27, Electric Elevator Trim - Removal/ Installation. NOTE:
(3) (4) (5) B.
The Electric Elevator Trim - Removal/Installation section gives the method necessary to remove and install the electric elevator trim motor that is installed on some models. This same method is valid to remove and install the pitch trim servo.
Disconnect the electrical connector from the pitch trim servo. Remove the pitch trim servo from the airplane. Do an inspection of the pitch trim servo. Refer to Pitch Trim Servo Inspection.
Install the Pitch Trim Servo (Refer to Figure 204). (1) Install the pitch trim servo in the airplane. Refer to Chapter 27, Electric Elevator Trim - Removal/ Installation. (a) Torque the bolts to 45 inch-pounds, +5 or -5 inch-pounds. NOTE:
(2) (3) (4) (5)
The Electric Elevator Trim - Removal/Installation section gives the method necessary to remove and install the electric elevator trim motor that is installed on some models. This same method is valid to remove and install the pitch trim servo.
Connect the electrical connector to the pitch trim servo. Close access to the pitch trim servo. Refer to Chapter 27, Electric Elevator Trim - Removal/ Installation. Connect electrical power to the aircraft. Do an operational test of the autopilot. Refer to Introduction of the List of Manufacturers Technical Publications, for the manufacturer's installation manual.
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Autopilot Pitch Trim Servo Installation Figure 204 (Sheet 1)
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6.
Pitch Trim Rigging Inspection A.
Do a check of the pitch trim rigging. (1) Attach an inclinometer to the trim tab. (2) Put the trim tab in the 0 degree position. (3) Manually operate the trim tab to the up and down limits. (a) Record the limits of travel. (4) Have an observer at the right-hand access opening of the tailcone. (5) Put the electrical trim to the full nose-up position until the observer sees the clutch slip. (6) Turn the manual trim wheel nose-up (test load condition) 1/4 turn more while the clutch slips. (7) Make sure the swaged ball on the bridle cable assembly does not turn aft of the tangent point. (8) Release the trim wheel and disengage the autopilot. (9) Manually operate the trim to the full nose-up position. (10) Do a check of the trim tab position with an inclinometer. NOTE:
Trim tab position that is greater than the limits of travel values recorded is an indication that the stop blocks slipped.
(11) If the stop blocks slip, do the steps that follow. (a) Do the trim system rigging again. (b) Make sure the stop block bolts torque is correct. (c) Do a check of the pitch trim rigging again. (12) If necessary, make adjustments to the swaged ball position. (a) Put the bridle cable assembly chain in the applicable position on the gear teeth of the actuator sprocket. NOTE:
One chain link adjustment is related to approximately 17 degrees of travel on the capstan.
(b)
Apply the applicable tension to the bridle cable and do a check of the pitch trim rigging again. (13) Do the procedure again for the full nose-down trim condition. 7.
Servo Capstan Clutch Adjustment A.
Adjust the servo capstan clutch in accordance with the G1000 Caravan Line Maintenance Manual and servo mount installation manual. Refer to Introduction, the List of Manufacturers Technical Publications for the manufacturer's installation manual.
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MODEL 208 MAINTENANCE MANUAL GFC 700 AUTOPILOT - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the GFC 700 autopilot in a serviceable condition.
Task 22-12-00-640 2.
Autopilot Servos Lubrication A.
General (1) This task gives the procedures to do a lubrication of the Garmin Roll, Pitch, Yaw, and Pitch Trim Servo output gears. NOTE:
This task is only applicable for the Model 208 airplanes with the Garmin 1000 and the GFC 700 autopilot system installed.
B.
Special Tools (1) Aeroshell 33MS Grease (preferred), or Aeroshell 17 Grease.
C.
Access (1) Remove the copilot's seat to get access to the roll servo. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (2) Remove access panel 232DR to get access to the roll servo. Refer to Chapter 6, Access/ Inspection Plates - Description and Operation. (3) Remove the rear compartment partition or unzip the canvas wall to get access to the yaw and pitch servo. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. (4) Remove access panels 226A and 226D from the pedestal to get access to the pitch trim servo. Refer to Chapter 6, Access/Inspection Plates - Description and Operation.
D.
Do a lubrication of the Autopilot Servos. (1) Remove the roll, pitch, yaw, and pitch trim servos (Refer to GFC 700 Autopilot - Maintenance Practices).
CAUTION: Do not use solvents to clean the output gears. (2) (3) (4) (5) (6)
Use a lint-free cloth to remove the excess grease build-up from the output gears for the different servos. Apply Aeroshell 33MS (preferred), or Aeroshell 17 grease to the output gear. Install the roll, pitch, yaw, and pitch trim servos (Refer to GFC 700 Autopilot - Maintenance Practices). Operate all of the control surfaces through their full range of travel. Make a maintenance log entry that shows that the GSA 8X servo output gears have been lubricated with Aeroshell 33MS or Aeroshell 17 grease.
E.
Restore Access (1) Install the rear compartment partition or zip the canvas wall. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. (2) Install access panels 226A and 226D to the pedestal. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (3) Install access panel 232DR. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (4) Install the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL Task 22-12-00-720 3.
Garmin Autopilot (GFC 700) Functional Check A.
General (1) This task gives the procedures to do a Functional Check of the Garmin Autopilot (GFC 700).
B.
Special Tools (1) External Electrical Power Unit (2) Cable Tensionmeter
C.
Access (1) Remove the copilot's seat to get access to the roll servo. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (2) Remove access panel 232DR to get access to the roll servo. Refer to Chapter 6, Access/ Inspection Plates - Description and Operation. (3) Remove the rear compartment partition or unzip the canvas wall to get access to the yaw and pitch servo. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. (4) Remove access panels 226A and 226D from the pedestal to get access to the pitch trim servo. Refer to Chapter 6, Access/Inspection Plates - Description and Operation.
D.
Do a Functional Check of the Garmin Autopilot (GFC 700). (1) Examine the servos, connectors, support structures, and control cables for corrosion, chaffing, cracks, or other damage. (2) Have a helper manually move the ailerons, (for roll servo), elevators (for pitch servo), elevator trim wheel (for pitch trim servo), and rudder pedals (for yaw servo) from stop to stop and examine the servo, capstan, and control surface operation. (a) Make sure there are no binds in the control cables and that the capstan pulleys turn freely. (3) Examine the servo control cables to make sure there is no fraying, corrosion, or other damage. (a) If the condition of the cable is unknown, replace it with a new one. (4) Examine the tension of each servo control cable. Refer to GFC 700 Autopilot - Maintenance Practices. (5) Examine the GFC 700 autopilot system wiring and make sure there is no chaffing, wear, or other damage.
E.
Do a GSM Servo Slip Clutch Check. (1) Apply external electrical power to the airplane. (2) Set the external power switch to BUS. (3) Set the battery switch to ON. (4) Set the avionics switches 1 and 2 to ON. (5) Make sure that the A/P SERVOS circuit breaker and A/P CONT circuit breaker on the lower left circuit breaker panel are engaged. (6) Push the AP key on the GMC 710 AFCS controller to engage the autopilot. NOTE:
(7)
Apply force to the control yoke to find if the autopilot clutches can be overpowered in pitch and roll. (a) If the autopilot clutches cannot be overpowered, examine the servo clutch torque settings. Refer to G1000 Caravan Line Maintenance Manual. NOTE:
(8)
The GFC 700 uses electronic torque limiting as well as mechanical slip clutches to limit the maximum servo effort. When the system is on the ground, the electronic torque limiting is removed, to allow manual checks of the slip-clutch settings.
There is an overpowered condition if the control surfaces can be moved by applying force to the control wheel or the rudder pedals against the resistance of the engaged autopilot.
Apply force to the rudder pedals to find if the autopilot clutches can be overpowered in yaw. (a) If the autopilot clutches cannot be overpowered, examine the servo clutch torque settings. Refer to G1000 Caravan Line Maintenance Manual.
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MODEL 208 MAINTENANCE MANUAL (9)
(10) (11) (12) (13)
Use the Pitch MET (Manual Electric Trim) switch to initiate an autopilot disconnect. (a) While the trim is running, grasp the aircraft pitch trim wheel and make sure that the trim clutch can be overpowered by preventing the trim wheel from moving. 1 If it cannot be overpowered, examine the servo clutch torque setting. Refer to G1000 Caravan Line Maintenance Manual. (b) Make sure that the trim wheel moves smoothly in both directions through the full trim range during the Pitch MET (Manual Electric Trim) operation. 1 If the trim wheel does not move in 2 seconds, this can show that the pitch trim clutch is slipping. (c) Make sure that the clutch setting and cable tensions are correct. 1 If the clutch setting and cable tensions are in tolerance, examine the aircraft pitch trim system for too much friction. Set the avionics switches 1 and 2 to OFF. Set the battery switch to OFF. Set the external power switch to OFF. Remove external electrical power to the airplane.
F.
Restore Access (1) Install access panels 226A and 226D to the pedestal. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (2) Install the rear compartment partition or zip the canvas wall. Refer to Chapter 25, Rear Compartment Wall - Maintenance Practices. (3) Install access panel 232DR. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (4) Install the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL GARMIN GMC-710 AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) - MAINTENANCE PRACTICES 1.
General A.
2.
The Garmin GMC-710 AFCS is the control interface for the Garmin GFC-700 autopilot system.
GMC-710 AFCS Removal/Installation A.
Remove the AFCS Controller (Refer to Figure 201). (1) Remove the electrical power from the airplane. (2) Disengage the A/P CONT and the A/P SERVOS & AP DISC circuit breakers on the avionics circuit breaker panel. (3) Turn the captive screws on the AFCS controller counterclockwise 1/4 turn. (4) Move the controller aft to get access to the rear of the controller. (5) Disconnect the connector from the controller. (6) Remove the controller from the airplane.
B.
Install the AFCS Controller (Refer to Figure 201). (1) Put the AFCS controller in position in the instrument panel and connect the connector. (2) Install the controller fully in the instrument panel. (3) Turn the captive screws on the controller clockwise 1/4 turn to attach it to the autopilot tray assembly. (4) Engage the A/P CONT and the A/P SERVOS & AP DISC circuit breakers. (5) Make sure the AFCS operates correctly. (a) If a new unit is installed, load the software and configuration. Refer to the Garmin G1000 Line Maintenance Manual. (b) Do a check to make sure the AFCS operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
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MODEL 208 MAINTENANCE MANUAL
GMC-710 AFCS Installation Figure 201 (Sheet 1)
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23 CHAPTER
COMMUNICATIONS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
23-00-00
Pages 1-2
Mar 1/2008
23-00-10
Pages 201-202
Nov 3/2003
23-10-00
Page 201
Mar 1/2008
23-10-01
Pages 401-402
Mar 1/2001
23-50-01
Pages 401-402
Apr 1/2010
23-51-00
Pages 201-203
Apr 1/2010
23-60-00
Pages 201-203
Jun 1/2011
23-60-00
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Jun 1/2011
23-70-00
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Jun 3/2002
23-70-00
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Apr 1/2010
23-70-00
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Jun 3/2002
23-Title 23-List of Effective Pages 23-Record of Temporary Revisions 23-Table of Contents 23-List of Tasks
23 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS COMMUNICATIONS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-00-00 23-00-00 23-00-00 23-00-00 23-00-00
Page 1 Page 1 Page 1 Page 2 Page 2
COMMUNICATION - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Wheel Pushbutton Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . Control Wheel Pushbutton Switch Button Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-00-10 Page 201 23-00-10 Page 201 23-00-10 Page 201 23-00-10 Page 201
GIA 63 INSTALLATION - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-10-00 Page 201 23-10-00 Page 201
FLEXCOMM 1 ANTENNA - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLEXCOMM 1 Antenna Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-10-01 Page 401 23-10-01 Page 401 23-10-01 Page 401
PASSENGER ADDRESS AND COCKPIT SPEAKER AMPLIFIER - REMOVAL/ INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Address Headphone And Cockpit Speaker Amplifier Removal/ Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-50-01 Page 401 23-50-01 Page 401 23-50-01 Page 401
GARMIN GMA 1347 AUDIO PANEL - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin GMA 1347 Audio Panel Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . Intercom Jacks Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-51-00 Page 201 23-51-00 Page 201 23-51-00 Page 201 23-51-00 Page 201
STATIC DISCHARGING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Static Discharger Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Discharger Base Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Bonding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-60-00 Page 201 23-60-00 Page 201 23-60-00 Page 201 23-60-00 Page 201 23-60-00 Page 201 23-60-00 Page 203
STATIC DISCHARGING - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Static Discharge System Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-60-00 Page 601 23-60-00 Page 601 23-60-00 Page 601
L3 COMMUNICATIONS FA2100 COCKPIT VOICE RECORDER - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FA 2100 Cockpit Voice Recorder Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FA2100 Cockpit Voice Recorder Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . L3 COMMUNICATIONS FA2100 COCKPIT VOICE RECORDER - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FA2100 Cockpit Voice Recorder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . Impact Switch (5G) Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cockpit Voice Recorder Control Panel Removal/Installation . . . . . . . . . . . . . . . . . . . . . CVR Adapter Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23-70-00 23-70-00 23-70-00 23-70-00
23-70-00 Page 201 23-70-00 Page 201 23-70-00 Page 201 23-70-00 Page 201 23-70-00 Page 201 23-70-00 Page 203 23-70-00 Page 203
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MODEL 208 MAINTENANCE MANUAL L3 COMMUNICATIONS FA2100 COCKPIT VOICE RECORDER - ADJUSTMENT/ TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FA2100 Cockpit Voice Recorder (CVR) Operational Test. . . . . . . . . . . . . . . . . . . . . . . . Impact Switch Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Underwater Locator Device (ULD) Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FA2100 CVR Bulk Erase. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 23-60-00-720
Static Discharge System Functional Check
23-60-00 Page 601
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MODEL 208 MAINTENANCE MANUAL COMMUNICATIONS - GENERAL 1.
2.
General A.
This chapter provides a brief description of various communication systems, units and components which provide a means of communicating from one part of the airplane to another, from airplane-toairplane or airplane-to-ground locations. This includes voice and continuous wave communicating components.
B.
This chapter also provides maintenance practices for static wicks. Refer to Static Discharging Maintenance Practices .
Description A.
Sperry (Type RT-385A) NAV/COMM. (1) This standard system consists of a panel-mounted receiver-transmitter, a single or dual-pointer remote deviation indicator, a VHF COMM antenna, a balanced loop omni/glide slope antenna and interconnecting cables. A DME receiver or a glideslope receiver, or both, may be interconnected with the NAV/COMM set for automatic selection of the associated DME or glideslope frequency.
B.
Sperry (Type RT-385B) NAV/COMM. (1) This optional system consists of a panel-mounted receiver/transmitter, a single or dual-pointer remote 300 or 400 series course deviation indicator, a VHF COMM antenna, a balanced loop omni/glideslope antenna and interconnecting cables. A DME receiver or a glide slope receiver, or both, may be interconnected with the NAV/COMM set for automatic selection of the associated DME or glideslope frequency.
C.
King (Type KX-165) NAV/COMM with Integral Glide Slope. (1) This optional system consists of a panel-mounted receiver/transmitter, a slaved IG-832A Horizontal Situation Indicator (HSI), a VHF COMM antenna, a balanced loop omni/glideslope antenna and interconnecting cables. A DME receiver may be interconnected with the NAV/COMM set for automatic selection of the associated DME frequency.
D.
King (Type KY-196) Digital COMM. (1) This optional system consists of a panel-mounted receiver-transmitter, a VHF COMM antenna and interconnecting cables.
E.
King (Type KHF-950) HF SSB Transceiver. (1) This optional panel-mounted, solid-state HF single sideband transceiver system is controlled by a KCU-951 Dzus rail-mounted control display unit. The system also incorporates a KAC952 power ampliÞer/antenna coupler, a KTR-953 receiver/exciter, an MF and HF antenna and interconnecting cables.
F.
Sperry (Type F-490A) Audio Control Panel. (1) This standard system provides for ampliÞcation of audio signals for speaker system and allows audio switching for cabin speaker, headset(s), intercom and microphone(s). The audio control panel will accommodate two transceivers, an ADF, DME and marker beacon. (2) The audio control panel incorporates a pilot and copilot intercom phone system. The system incorporates its own audio ampliÞer with a volume control (labeled INT) and a hot mike feature. The intercom is used with headphones only.
G.
King (Type KMA-24) Audio Console. (1) Two King Audio Control Systems are available. The only difference between the two systems is the choice of the third MIC function that can be either HF functions (to accommodate a HF radio installation) or TEL functions (to accommodate the airborne radio telephone installation). (2) Both systems have a combination audio ampliÞer, an audio distribution panel and a marker beacon receiver. The audio ampliÞer is for ampliÞcation of the audio signals for the speaker system. All receiver audio distribution functions are controlled by two rows of alternate-action push buttons. Both rows are completely independent of each other, allowing simultaneous use of speaker and/or headphones. A rotary selector switch on the right side of the console connects the microphone to either telephone, HF radio, COMM 1 or COMM 2.
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MODEL 208 MAINTENANCE MANUAL H.
3.
Operation A.
4.
King (Type KHF-1050) Communication System (1) This optional system is a solid-state high frequency single sideband transceiver system for voice communication in the 2- to 29.9999-MHz band with 100 Hz resolutions. (2) The KHF-1050 as installed in the 208/208B Caravan has a PS440 HF Control Head, KAC-1052 Antenna Coupler, KPA-1052 Power AmpliÞer, KRX-1053 Receiver/Exciter, and an HF Antenna.
For operation of the various systems, refer to the FAA Approved Airplane Flight Manual and Pilot’s Operating Handbook.
Maintenance Practices A.
For maintenance practices related to the above systems, refer to the following manuals: (1) Model 208 Avionic Installations Service/Parts Manual. (2) SpeciÞc vendor publications listed in Introduction - List of Publications.
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MODEL 208 MAINTENANCE MANUAL COMMUNICATION - MAINTENANCE PRACTICES 1.
General A.
2.
3.
This section gives procedures for the removal and installation of the control wheel pushbutton switches. This section also gives procedures to clean the control wheel pushbutton switches.
Control Wheel Pushbutton Switch Removal/Installation A.
Remove the Control Wheel Pushbutton Switch (Refer to Figure 201). (1) Remove the nut that attaches the control wheel pushbutton switch (SA03, SA04, SU03) to the escutcheon. (2) Remove the screw that attaches the escutcheon to the control wheel. (3) Lift up the escutcheon to get access to the control wheel pushbutton switch and disconnect the switch from the control wheel connection.
B.
Install the Control Wheel Pushbutton Switch (Refer to Figure 201). (1) Connect the control wheel pushbutton switch (SA03, SA04, SU03) to the connection in the control wheel. (2) Attach the control wheel pushbutton switch to the escutcheon with the nut. (3) Set the escutcheon in position and install the screw in the escutcheon.
Control Wheel Pushbutton Switch Button Cleaning A.
Clean the Control Wheel Pushbutton Switch (Refer to Figure 201). NOTE:
(1)
Continued use of the button can cause oil and dirt to collect on the internal electrical contacts of the switch. This can cause a malfunction with the control wheel pushbutton switch.
Apply electrical contact cleaning spray around the full edge of the button so that it will soak down into the control wheel pushbutton switch. NOTE:
(2) (3) (4) (5)
The electrical cleaner will help to remove oil and dirt from the internal electrical contacts of the control wheel pushbutton switch. The recommended contact cleaner is Electro Contact Cleaner 03116 or equivalent, which can be supplied by LPS Laboratories, Inc. The LPS Laboratories, Inc. telephone number is 1-800-241-8334.
Push the button many times to make sure that the cleaner gets into the internal electrical contacts of the control wheel pushbutton switch. Complete an operational check of the control wheel pushbutton switch. If the button does not operate after the first application of the electric cleaner, apply more cleaner. If the button continues to have a malfunction, replace the control wheel pushbutton switch. Refer to Control Wheel Pushbutton Switch - Removal/Installation.
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Control Wheel Pushbutton Switch Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GIA 63 INSTALLATION - MAINTENANCE PRACTICES 1.
General A.
For information on the GIA 63 installation, go to Chapter 34, Garmin Integrated Avionics Unit (GIA 63).
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MODEL 208 MAINTENANCE MANUAL FLEXCOMM 1 ANTENNA - REMOVAL/INSTALLATION 1.
2.
General A.
This section covers removal and installation procedures for the FLEXCOMM 1 Antenna.
B.
The FLEXCOMM 1 Antenna is located on the cargo pod.
FLEXCOMM 1 Antenna Removal/Installation A.
Remove the FLEXCOMM 1 antenna (Refer to Figure 401). NOTE: (1) (2) (3) (4) (5) (6)
B.
Access to FLEXCOMM 1 antenna is gained through the third cargo pod door compartment and on left side of the forward compartment in that compartment.
Disengage COM 1 circuit breaker located on left circuit breaker panel. Remove protective cover assembly mounted on third cargo pod compartment partition by removing and retaining screws holding it in place. Disconnect antenna connector. Remove screws holding antenna in place. Carefully remove antenna from cargo pod and clean sealant from antenna and cargo pod. Remove antenna from airplane.
Install FLEXCOMM 1 Antenna (Refer to Figure 401). (1) Clean antenna and cargo pod mounting surface using an approved solvent. Refer to Chapter 20, General Solvents/Cleaners - Maintenance Practices. (2) Place antenna into position and secure with screws. (3) Connect coax cable assembly to antenna. (4) Fillet seal around antenna base with Type I, Class B sealant. Refer to Chapter 20, Fuel, Weather, Pressure and High-Temperature Sealing - Maintenance Practices. (5) Install protective cover assembly on cargo pod compartment partition using screws previously removed. (6) Engage COM 1 circuit breaker.
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FLEXCOMM 1 Antenna Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PASSENGER ADDRESS AND COCKPIT SPEAKER AMPLIFIER - REMOVAL/INSTALLATION 1.
2.
General A.
This section covers removal and installation procedures for the passenger address headphone amplifier and cockpit speaker amplifier.
B.
The passenger address headphone amplifier and cockpit speaker amplifier is located aft of the copilot position on lower fuselage at FS 175, underneath the right side cabin floor board.
Passenger Address Headphone And Cockpit Speaker Amplifier Removal/Installation A.
Remove the passenger address headphone and cockpit speaker amplifiers (Refer to Figure 401). NOTE:
(1) (2) (3) (4) (5) (6) (7) (8) B.
Access to the passenger address headphone amplifier and cockpit speaker amplifier is gained through the cabin floorboard of the airplane at access panel 252AR . Refer to Figure 401.
Disengage AUDIO 1, AUDIO 2, PA AMP, PKR AMP and ICS circuit breakers. Remove power from airplane. Gain access to cabin floorboard panel cover. Reference Chapter 6, Access Plates And Panels Identification - Description And Operation. Remove center cabin floorboard to gain access to the passenger address headphone amplifier and cockpit speaker amplifier. Disconnect cable from desired amplifiers. Remove 4 each installation screws from each unit. Carefully remove desired unit. Engage AUDIO 1, AUDIO 2, PA AMP, SPKR AMP and ICS circuit breakers.
Install the Passenger Address System and Cockpit Speaker Amplifier (Refer to Figure 401). (1) Gain access to cabin floorboard panel cover. Reference Chapter 6, Access Plates And Panels Identification - Description And Operation. (2) Disengage AUDIO 1, AUDIO 2, PA AMP, PKR AMP and ICS circuit breakers. (3) Remove power from airplane. (4) Gain access to cabin floorboard panel cover. Reference Chapter 6, Access Plates And Panels Identification - Description And Operation. (5) Remove center cabin floorboard to gain access to passenger address headphone and speaker amplifiers. (6) Install passenger address headphone and cockpit speaker amplifiers with four each screws. Refer to Figure 401. (7) Connect coax cables. (8) Engage AUDIO 1, AUDIO 2, PA AMP, SPKR AMP and ICS circuit breakers.
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Passenger Address and Cockpit Speaker Amplifier Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GARMIN GMA 1347 AUDIO PANEL - MAINTENANCE PRACTICES 1.
2.
3.
General A.
The audio panel is installed in the middle of the instrument panel. It has an audio function, an intercom function, and a marker beacon function in one unit.
B.
Maintenance practices for the audio panel have procedures for the removal and installation of the audio panel and intercom jacks. Some airplanes can have more intercom jacks in the passenger compartment.
Garmin GMA 1347 Audio Panel Removal/Installation A.
Remove the Audio Panel (Refer to Figure 201). (1) Make sure the AVIONICS and MASTER switches are in the off position. (2) Turn the captive screw on the face of the audio panel counterclockwise until it releases from the mounting tray. (3) Carefully pull the audio panel out of the mounting tray.
B.
Install the Audio Panel (Refer to Figure 201). (1) Put the audio panel in position and move it forward into the mounting tray. (2) Turn the captive screw on the face of the audio panel clockwise until the audio panel is attached to the mounting tray. (3) Do a check to make sure the transponder operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
Intercom Jacks Removal/Installation NOTE:
A.
Installations in some airplanes are installed in an escutcheon assembly. Installations in some airplanes are installed flush with the interior panels with a plate installation. Refer to the type installation on the airplane you do maintenance on.
Remove the Pilot/Copilot and Passenger Intercom Jacks (Escutcheon Installation) ( Refer to Figure 202). NOTE: (1) (2) (3) (4) (5) (6) (7) (8) (9)
B.
The removal and installation of the microphone jack (small plug) and the headphone jack (large plug) are identical.
Remove external electrical power from the airplane. Make sure that the BATTERY switch is in the off position. Remove the escutcheon to get access to the passenger jacks. Remove the LED panel to get access to the pilot jacks. Note the arrangement of the insulating washers under the jack and nut. Remove the jamnut and washer that attaches the jack to the escutcheon. Put a label on the applicable wires of the microphone jack (small plug) or the headphone jack (large plug). Cut the wires near the soldered joint of the applicable jack. Remove the jack , bottom washer and insulating washer if applicable, from the escutcheon or instrument panel.
Install the Pilot/Copilot and Passenger Intercom Jacks (Refer to Figure 202). (1) Solder the applicable wires to the jack. Refer to Model 208 Wiring Diagram Manual, Chapter 20, Soldering - Maintenance Practices. (2) Put the jack, insulating washer, and bottom washer in their position in the escutcheon or in instrument panel. (3) Install the jamnut, washer, and insulating washer that attaches the jack to the escutcheon. (4) If removed, install the LED panel.
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GMA 1347 Audio Panel Installation Figure 201 (Sheet 1)
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Intercom Jack Installation (Escutcheon) Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL STATIC DISCHARGING - MAINTENANCE PRACTICES 1.
General A.
2.
Trailing edge static dischargers are installed at airplane control surface extremities. These dischargers are used to reduce stored potential voltage that is the result of electrostatic charging from flying through haze, dust, rain, snow or ice crystals. Reduction of stored voltage is necessary to prevent undesirable electrostatic currents that could cause unacceptable radio noise or electrical insulation failures.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Megohmeter
Model 2850
Inotek 9902 E. 42nd Street Tulsa, OK 47146
To measure high resistance.
Bonding Meter
Shallcross Type 670
Eaton Corp. Operation and Technical Center 4201 N. 27th Street Milwakee, WI 53200-0000
To check electrical bonding connections.
3.
Static Discharger Removal/Installation A.
Remove Static Discharger (Refer to Figure 201). (1) Remove static discharger and lock washer from base. NOTE:
B. 4.
Static dischargers showing deterioration should be replaced.
Install Static Discharger (Refer to Figure 201). (1) Position lock washer on static discharger and screw into base.
Discharger Base Removal/Installation A.
Remove Discharger Base (Refer to Figure 201). (1) Drill out blind rivets securing discharger base to airplane. (2) Remove discharger base from airplane.
B.
Install Discharger Base (Refer to Figure 201). (1) Use fine grit sandpaper and remove any paint around attaching holes or under discharger base footprint. (2) Using a 500 or 600 grit emery cloth, break aluminum oxide in the footprint area. NOTE: (3) (4) (5)
Do not delay performing steps 4.B.(3) and 4.B.(4), as new oxide will form within minutes.
Clean mating surface of airplane skin with solvent. Brush cleaned skin with Aluma Prep 1201 alodine and wait until it is dry before proceeding. Install new base using an appropriate sized blind rivet.
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Static Discharging Figure 201 (Sheet 1)
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Primer and paint may be used as desired on airplane skin. NOTE:
5.
Cover any attached static discharger and base with paper or rag. Do not use tape. Screw threads in base should be protected with a lightly inserted wooden plug.
Electrical Bonding A.
Individual electrically conductive components and structures of the airplane must be electrically bonded together. This bonding is necessary to ensure that all conductive materials on the airplane are at the same electric potential. If electrical bonding is not maintained, crew members or passengers may encounter electrical shocks, radio and other avionic system interference or even damage will result and corrosion between dissimilar materials may occur.
B.
Bond resistance between structures should not exceed 0.003 ohms unless otherwise specified in specific installations. After major repair and/or replacement of components or control surfaces, an electrical bonding check is required.
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MODEL 208 MAINTENANCE MANUAL STATIC DISCHARGING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the static discharging system in a serviceable condition.
Task 23-60-00-720 2.
Static Discharge System Functional Check A.
General (1) This task gives the information needed to complete the inspection procedures for the static discharge system.
B.
Special Tools (1) Digital Ohmmeter (2) Megohmmeter
C.
Access (1) None
D.
Do the Static Discharge System Functional Check. (1) Visually examine the static dischargers for lightning damage and erosion of the airplane skin at the attach points. (a) If the static discharger shows signs of a lightning strike, replace the static discharger and examine the entire aircraft for lightning strike damage. Refer to Chapter 5, Unscheduled Maintenance Checks. (2) Visually examine between the tips of the static dischargers and the base assemblies for erosion. (3) Visually examine the static dischargers for condition and security. (4) Replace the damaged or the missing static dischargers. (5) Make sure that all static dischargers are tight.
E.
Do a Functional Check of the Static Discharge System. (1) Use an ohmmeter (bonding meter) to do a check of the resistance between the base assemblies and a good airplane ground. (a) Make sure that the resistance between the base assembly and the metal surface is 0.5 ohms or less. (b) Make sure there is a good ground before you do the next step.
WARNING: Use precaution when you use a high voltage megohmmeter to prevent an electrical shock. (2)
Use a megohmmeter set to 500 volts to do a check of the resistance between the base assemblies and the static dischargers. (a) Make sure that the resistance between the base assembly and the static discharger is 1 to 100 megohms. (b) If the resistance between the base assembly and the static discharger is not in tolerance, replace the static discharger.
F.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL L3 COMMUNICATIONS FA2100 COCKPIT VOICE RECORDER - DESCRIPTION AND OPERATION 1.
General A.
2.
3.
The L3 Communications FA2100 Cockpit Voice Recorder (CVR) System is a 120 minute recorder that accepts the four cockpit audio input streams and processes these four audio input streams into six internal audio streams. Four of these audio streams, Channel 1 through Channel 4, are high quality audio streams with a minimum of 30 minutes recording time. (1) The fifth and sixth audio streams are internal to the recorder and provide a minimum of 120 minutes of standard quality recording time. The fifth audio stream is a summation of the inputs on Channels 1, 2 and 3. The sixth audio stream is a standard quality recording of the cockpit area microphone (CAM) input on Channel 4. The six audio streams are then changed into a digital format for storage in the crash protected solid-state memory.
FA 2100 Cockpit Voice Recorder Description A.
The FA2100 Cockpit Voice Recorder System consists of a tailcone mounted recorder unit, underwater locator beacon (ULD) and a control panel mounted with the radio stack in the instrument panel.
B.
The FA2100 Cockpit Voice Recorder System is designed to operate with an area microphone. The area microphone is permanently mounted as part of the S162 Control Panel. The area microphone senses conversation from the airplane cockpit.
C.
An Underwater Locator Device (ULD) is mounted on the front panel of the FA2100 CVR. It is designed to transmit a locator signal, should the aircraft ever become submerged in water. The battery replacement decal is also located on the front panel. Battery must be replaced every six years.
D.
The impact switch is located under the aft avionics shelf near the cockpit voice recorder. It is designed to remove electrical power with an impact of 5 G's or more. This action is designed to open the electrical circuit to the voice recorder to preserve the recorded data. The data could be erased automatically if recorder operation were to continue.
FA2100 Cockpit Voice Recorder Operation A.
The FA2100 Cockpit Voice Recorder System automatically comes on when electrical power is applied to the airplane. All data in and out of the solid-state memory is controlled by a microprocessor.
B.
A self-test function is started by pushing the TEST button located on the CVR control panel. The CVR control panel is normally located at the bottom of the radio stack but location may vary, depending upon specific customer orders.
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MODEL 208 MAINTENANCE MANUAL L3 COMMUNICATIONS FA2100 COCKPIT VOICE RECORDER - MAINTENANCE PRACTICES 1.
General A.
2.
Tools and Equipment A.
3.
Maintenance of the optional L3 Communications FA2100 Cockpit Voice Recorder System (SSCVR) consists of removal/installation of the aft avionics shelf mounted components and removal/installation of the instrument panel mounted SSCVR control panel. The system consists of a mounted recorder unit, CVR Adapter, panel mounted S162 Control Panel and an impact switch (5G) located under the cockpit voice recorder on the aft avionics shelf.
For a list of tools and equipment, refer to Communications - General.
FA2100 Cockpit Voice Recorder Removal/Installation A.
Remove FA2100 Cockpit Voice Recorder (Refer to Figure 201). NOTE:
(1) (2) (3) (4) (5) (6) B.
4.
The voice recorder unit is tray-mounted on the avionics shelf in the tailcone at RBL 8.80, WL 136.92 between FS 370.80 and FS 385.60. Access to the CVR is from the aft side of the avionics shelf.
Disengage the cockpit circuit breaker located on the left circuit breaker panel. Get access to cockpit voice recorder through the aft tailcone entrance. Cut and remove the safety wire between the knurl nuts located on the forward side of the mounting tray. Loosen the knurl nuts and the unlatching locking mechanism located on the front of the mounting tray. Carefully move the FA2100 cockpit voice recorder straight forward from the mounting tray. Remove the FA2100 cockpit voice recorder from airplane.
Install FA2100 Cockpit Voice Recorder (Refer to Figure 201). (1) From the aft side of the avionics shelf, put the FA2100 cockpit voice recorder on the mounting tray and slide in until the electrical connectors are firmly engaged. (2) Secure the FA2100 cockpit voice recorder to tray by engaging the locking mechanism. (3) Engage the tightening mechanism and tighten and safety wire the knurl nuts together on forward side of the mounting tray. Refer to Chapter 20, Safetying - Maintenance Practices.
Impact Switch (5G) Removal/Installation A.
Remove Impact Switch (5G) (Refer to Figure 201). (1) Remove electrical power from the airplane.
CAUTION: The CVR and FDR are linked to the same impact switch circuit. If one circuit breaker is disengaged, the other circuit breaker must be disengaged. (2) (3) (4) (5) (6) B.
Disengage cockpit voice recorder circuit breaker and if installed, the FDR circuit breaker located on the left circuit breaker panel. Get access to the impact switch (5G), located under the aft avionics shelf, near the cockpit voice recorder. Disconnect electrical connector from impact switch (5G). Remove the screws that attach the impact switch (5G) to the shelf. Remove impact switch (5G) from airplane.
Install Impact Switch (5G) (Refer to Figure 201). (1) Install impact switch (5G) to the shelf with screws. (2) Connect electrical connector to impact switch (5G).
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Cockpit Voice Recorder and Impact Switch Installation Figure 201 (Sheet 1)
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5.
Cockpit Voice Recorder Control Panel Removal/Installation A.
Remove Cockpit Voice Recorder Control Panel (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5)
B.
6.
The cockpit voice recorder S162 Control Panel is normally located at the bottom of the right radio panel on the instrument panel.
Disengage the cockpit voice recorder circuit breaker, located on the left circuit breaker panel. Release the Dzus fasteners that attach the cockpit voice recorder control panel assembly to the instrument panel. Remove the cockpit voice recorder control panel from the instrument panel. Disconnect electrical connector from the cockpit voice recorder control panel. Remove the cockpit voice recorder control panel from airplane.
Install Cockpit Voice Recorder Control Panel (Refer to Figure 201). (1) Position the cockpit voice recorder control panel in the mounting position in the instrument panel. (2) Connect electrical connector to the cockpit voice recorder control panel. (3) Install the cockpit voice recorder control panel in the instrument panel, attaching with Dzus fasteners. (4) Engage the cockpit voice recorder circuit breaker, located on the left circuit breaker panel.
CVR Adapter Removal/Installation NOTE:
Removal of both pilot's and copilot's CVR adapter is typical. The CVR adapter is located on the forward side of the pedestal.
A.
Remove CVR Adapter (Refer to Figure 201). (1) Disconnect the CVR circuit breaker. (2) Disconnect electrical connector from the CVR adapter. (3) Remove screws that attach the CVR adapter to the bracket. (4) Remove the CVR adapter from the airplane.
B.
Install CVR Adapter (Refer to Figure 201). (1) Put the CVR adapter in place and secure with screws. (2) Connect electrical connector to the CVR adapter. (3) Connect the CVR circuit breaker.
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MODEL 208 MAINTENANCE MANUAL L3 COMMUNICATIONS FA2100 COCKPIT VOICE RECORDER - ADJUSTMENT/TEST 1.
General A.
2.
This section provides information necessary to do an operational test for the L3 Communications FA2100 Cockpit Voice Recorder System.
Tools and Equipment
Table 501. Tools and Equipment NAME
NUMBER
MANUFACTURER
USE
Portable Interface Unit
17TES0043
L3 Communications Box 3041 Sarasota, FL 34230
To test FA2100 CVR.
Available Locally
To test FA2100 audio.
Dukane Corp. Ultrasonics Division 2900 Dukane St. Charles, IL 60174
To test ULD.
Headset Pinglite Tester
3.
PL-3
FA2100 Cockpit Voice Recorder (CVR) Operational Test A.
Do an FA2100 CVR Operational Test.
CAUTION: The actual CVR test for Channel 1 through Channel 4 must be completed within 30 minutes of the start recorder time, in order to receive the crew audio on the specified channels. If at any time the recorder time is longer than one half hour, a restart of this functional test procedure is required. The time that is recorded will be monitored from the portable interface unit (piu) tester which indicates the actual elapsed time of the recorded audio. NOTE:
Channel 1 through 4 audio sources are provided for a high quality (HQ) encoder for half hour continuous recording which records over the previous recording every 30 minutes. Channel 5 and 6 audio sources are provided for a standard quality (SQ) encoder for 120 minutes continuous recording.
(1) (2) (3)
Make sure the POWER switch, located on the left switch panel, is OFF. Make sure the AVIONICS POWER switch, located on the left switch panel, is OFF. Connect the FA2100 Portable Interface Unit (PIU) tester to the FA2100 CVR with the test cable provided. (4) Connect the external power cart to the external power receptacle and adjust to 29.0 VDC, +0.25 or -0.25 VDC. (5) Engage avionics and electrical circuit breakers as required. (6) Make sure the POWER switch, located on the left circuit breaker panel, is in the ON position. (7) Make sure the AVIONICS switch is in the ON position. (8) Make sure all necessary avionics power switches are ON. (9) Make sure the CENTER PANEL LIGHTS circuit breaker, located on the left circuit breaker panel, is engaged. (10) Rotate the appropriate Dimming Knob to full clockwise position. (11) Make sure the cockpit voice recorder control panel legend and nomenclature are at full bright and constant in intensity. (12) Adjust for the desired level of brightness.
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MODEL 208 MAINTENANCE MANUAL B.
Do a Power Compliance Verification Test. (1) Make sure the POWER switch, located on the left circuit breaker panel, is in ON position. (2) Make sure the AVIONICS switch, located on the left circuit breaker panel, is the OFF position. (3) Engage the cockpit voice recorder circuit breaker, located on the left circuit breaker panel. (4) Disengage the AUDIO 1 and AUDIO 2 circuit breakers, located on the right circuit breaker panel. (5) Make sure the FA2100 CVR is operational when the DC POWER switch is in ON position, only when the cockpit voice recorder circuit breaker is engaged.
C.
Do an FA2100 CVR Ground/Self-Test. NOTE: (1) (2) (3) (4)
This test checks operation of the FA2100 CVR self-test and observation of the COCKPIT VOICE RECORDER annunciator response.
Bulk erase the FA2100 CVR. Refer to FA2100 Bulk Erase. Attach a headset to the HEADSET jack on the CVR control panel. Push and hold the CVR TEST button located on the CVR control panel for a minimum of five seconds. Listen for a 640 Hz tone in the headset. NOTE:
(5)
Make sure the green cockpit voice recorder TEST annunciator, located on the cockpit voice recorder control panel, comes on, which indicates a successful test. NOTE:
4.
An internal 650 Hz test tone generator is keyed and sequentially switches through the six channels. Each channel receives the test tone for approximately 0.5 seconds.
A tone is generated every four seconds on Channel 1 only and is heard throughout the entire recorded tape.
D.
Do a Pilot's Received Audio Test. (1) Set the Portable Interface Unit (PIU) to channel 3. (2) Monitor the audio through the PIU speaker. (3) Make sure audio is heard seperately from the pilot's boom microphone, pilot headset, cockpit speaker and hand mic audio.
E.
Do a Copilot's Received Audio Test. (1) Set the Portable Interface Unit (PIU) to channel 2. (2) Monitor the audio through the PIU speaker. (3) Make sure audio is heard seperately from the pilot's boom microphone, pilot headset, cockpit speaker and hand mic audio.
F.
Do a Cockpit Area Microphone Audio Test. (1) Set the Portable Interface Unit (PIU) to channel 4. (2) Monitor the audio through the PIU speaker. (3) Make sure audio is heard from the cockpit area microphone, located on the CVR control panel.
G.
Do a Cabin PA Audio Test (If Installed). (1) Set the Portable Interface Unit (PIU) to channel 1. (2) Monitor the audio through the PIU speaker. (3) Make sure both pilot and copilot cabin PA audio can be heard.
Impact Switch Test A.
Do an Impact Switch Test. (1) Make sure the impact switch, located under the tailcone aft avionics shelf is tripped OPEN and the impact switch lamp comes on. (2) Make sure the CVR is not receiving power.
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If not tripped OPEN, manually trip it to the OPEN position. NOTE:
5.
Underwater Locator Device (ULD) Test A.
Do an Underwater Locator Device (ULD) Test. (1) Remove the Velcro collar from the Pinglite Model PL-3 Tester. (2) Apply the spring end of the Pinglite Model PL-3 Tester firmly to the ULD. NOTE: (3)
6.
A manual reset operation of the impact switch causes the switched output to be turned ON and the impact switch lamp to extinguish. Upon impact, the impact causes the output to be turned OFF and the impact switch lamp to come on.
The ULD is physically attached to the FA2100 CVR Locator Beacon.
Make sure the center spring is making contact with the center pin of the water switch. (a) With the center pin shorted to the center pin of the water switch, depress and hold the remaining three springs of the Pinglite PL-3 Tester against the ULD body. (b) Listen for audible pinging and make sure the LED flashes with each output pulse indicating its operation. (c) If audible pinging is not heard, or the LED does not flash, remove the FA2100 CVR and return to the Radio Lab for retest and/or rejection of the Underwater Acoustic Beacon.
FA2100 CVR Bulk Erase A.
Bulk (1) (2) (3) (4) (5) (6) (7)
Erase the FA2100 CVR. Open one or both aft doors. Make sure the impact switch (5G) is set to CLOSED. Make sure the impact switch annunciator goes off. Push and hold the ERASE button on the FA2100 CVR control panel for a minimum of 2 seconds and then release. Make sure the erase function worked by listening for a loud 400 Hz tone lasting for 5 to 10 seconds in the PIU headset. Remove all test equipment. Return aircraft to original configuration.
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24 CHAPTER
ELECTRICAL POWER
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
24-00-00
Pages 1-3
Apr 1/2010
24-00-01
Pages 501-513
Apr 1/2010
24-30-00
Page 1
Aug 1/1995
24-31-00
Pages 201-205
Aug 1/1995
24-32-00
Pages 101-107
Aug 1/1995
24-32-00
Pages 201-206
Apr 1/2010
24-32-00
Page 601
Jun 1/2011
24-33-00
Pages 201-203
Mar 1/1999
24-33-00
Page 601
Jun 1/2011
24-34-00
Page 1
Aug 1/1995
24-34-00
Pages 101-103
Aug 1/1995
24-34-00
Pages 401-407
Apr 1/1996
24-34-00
Pages 501-509
Mar 1/1999
24-34-00
Page 601
Jun 1/2011
24-34-00
Pages 701-702
Aug 1/1995
24-35-00
Pages 1-3
Aug 1/1995
24-35-00
Pages 101-103
Aug 1/1995
24-35-00
Pages 201-202
Aug 1/1995
24-36-00
Page 1
Aug 1/1995
24-36-00
Pages 101-103
Aug 1/1995
24-36-00
Pages 201-210
Jun 1/2011
24-36-00
Pages 601-603
Jun 1/2011
24-37-00
Pages 101-103
Aug 1/1995
24-37-00
Pages 201-202
Aug 1/1995
24-38-00
Pages 201-202
Aug 1/1995
24-39-00
Pages 1-2
Jan 2/2006
24-39-00
Pages 201-203
Jan 2/2006
24-40-00
Pages 101-102
Aug 1/1995
24-40-00
Pages 201-208
Mar 3/1997
24-50-00
Pages 201-205
Mar 1/2001
24-50-00
Pages 601-604
Jun 1/2011
24-52-01
Pages 1-15
Apr 1/2010
24-Title 24-List of Effective Pages 24-Record of Temporary Revisions 24-Table of Contents 24-List of Tasks
24 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS ELECTRICAL POWER - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-00-00 24-00-00 24-00-00 24-00-00
Page 1 Page 1 Page 1 Page 3
ELECTRICAL POWER - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fabrication of GCU Test Box . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical System Adjustment/Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-00-01 Page 501 24-00-01 Page 501 24-00-01 Page 501 24-00-01 Page 502
DIRECT CURRENT GENERATION - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-30-00 Page 1 24-30-00 Page 1
GENERATOR CONTROL UNIT - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GCU Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GCU Adjustment/Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-31-00 Page 201 24-31-00 Page 201 24-31-00 Page 201 24-31-00 Page 201
FLOODED LEAD-ACID BATTERY - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-32-00 Page 101 24-32-00 Page 101
FLOODED LEAD-ACID BATTERY - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Precautions and Notes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Initial Charge for New Batteries. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Quick-Disconnect Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Receptacle Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing Battery. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Adjustment/Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-32-00 Page 201 24-32-00 Page 201 24-32-00 Page 201 24-32-00 Page 202 24-32-00 Page 202 24-32-00 Page 202 24-32-00 Page 203 24-32-00 Page 203 24-32-00 Page 203 24-32-00 Page 203 24-32-00 Page 205
FLOODED LEAD-ACID BATTERY - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Gill Flooded Lead-Acid Battery Functional Check (Capacity Check) . . . . . . . . . . . . . .
24-32-00 Page 601 24-32-00 Page 601 24-32-00 Page 601
SEALED LEAD ACID BATTERY - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Receptacle Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Quick-Disconnect Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-33-00 Page 201 24-33-00 Page 201 24-33-00 Page 201 24-33-00 Page 201 24-33-00 Page 201
SEALED LEAD ACID BATTERY - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Concord Sealed Lead Acid Battery Functional Check (Capacity Check) . . . . . . . . . .
24-33-00 Page 601 24-33-00 Page 601 24-33-00 Page 601
NI-CAD BATTERY - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-34-00 Page 1 24-34-00 Page 1 24-34-00 Page 1
NI-CAD BATTERY - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-34-00 Page 101 24-34-00 Page 101
NI-CAD BATTERY - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Notes and Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Quick-Disconnect Receptacle Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Intercell Connector Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Cell Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-34-00 Page 401 24-34-00 Page 401 24-34-00 Page 401 24-34-00 Page 401 24-34-00 Page 403 24-34-00 Page 403 24-34-00 Page 403
24 - CONTENTS © Cessna Aircraft Company
Page 1 of 3 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Notes and Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Reconditioning Using the Marathon PCA-131-50/60 Charger/Analyzer . . . . Battery Reconditioning Using Alternate Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Leak Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Receptacle Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Quick-Disconnect Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-34-00 Page 501 24-34-00 Page 501 24-34-00 Page 501 24-34-00 Page 502 24-34-00 Page 502 24-34-00 Page 504 24-34-00 Page 508 24-34-00 Page 508 24-34-00 Page 509
NI-CAD BATTERY - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Marathon Ni-Cad Battery Functional Check (Capacity Check) . . . . . . . . . . . . . . . . . . .
24-34-00 Page 601 24-34-00 Page 601 24-34-00 Page 601
NI-CAD BATTERY - CLEANING/PAINTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning/Painting Battery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-34-00 Page 701 24-34-00 Page 701 24-34-00 Page 701 24-34-00 Page 701
BATTERY OVERHEAT WARNING - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-35-00 Page 24-35-00 Page 24-35-00 Page 24-35-00 Page
BATTERY OVERHEAT WARNING - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-35-00 Page 101 24-35-00 Page 101
BATTERY OVERHEAT WARNING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Temperature Sensor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment/Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-35-00 Page 201 24-35-00 Page 201 24-35-00 Page 201 24-35-00 Page 201
STANDBY ELECTRICAL SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-36-00 Page 1 24-36-00 Page 1 24-36-00 Page 1
STANDBY ELECTRICAL SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-36-00 Page 101 24-36-00 Page 101
STANDBY ELECTRICAL SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Alternator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Alternator Control Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Relay Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Output Voltage Adjustment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-36-00 Page 201 24-36-00 Page 201 24-36-00 Page 201 24-36-00 Page 201 24-36-00 Page 207 24-36-00 Page 207 24-36-00 Page 210
STANDBY ELECTRICAL SYSTEM - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drive Pulley Assembly/Bearing Housing Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Standby Alternator Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-36-00 Page 601 24-36-00 Page 601 24-36-00 Page 601 24-36-00 Page 601
ELECTRICAL MONITORING - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-37-00 Page 101 24-37-00 Page 101
ELECTRICAL MONITORING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Volt-Ammeter Gage Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Selector Switch Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-37-00 Page 201 24-37-00 Page 201 24-37-00 Page 201 24-37-00 Page 201
BATTERY CONTACTOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Contactor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-38-00 Page 201 24-38-00 Page 201 24-38-00 Page 201
24 - CONTENTS © Cessna Aircraft Company
1 1 1 1
Page 2 of 3 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL 12-VOLT DIRECT CURRENT POWER OUTLET SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-39-00 24-39-00 24-39-00 24-39-00
Page 1 Page 1 Page 1 Page 1
12-VOLT DIRECT CURRENT POWER OUTLET SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Outlet Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Converter Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12VDC Power Outlet System Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-39-00 Page 201 24-39-00 Page 201 24-39-00 Page 201 24-39-00 Page 201 24-39-00 Page 203
EXTERNAL POWER - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-40-00 Page 101 24-40-00 Page 101
EXTERNAL POWER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Components Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Power Box Fireproof Coating Inspection and Repair. . . . . . . . . . . . . . . . . . .
24-40-00 Page 201 24-40-00 Page 201 24-40-00 Page 201 24-40-00 Page 208
ELECTRICAL DISTRIBUTION SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Circuit Breaker Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bus Bar Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-50-00 Page 201 24-50-00 Page 201 24-50-00 Page 201 24-50-00 Page 201 24-50-00 Page 205
ELECTRICAL DISTRIBUTION SYSTEM - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Distribution Boxes Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-50-00 Page 601 24-50-00 Page 601 24-50-00 Page 601
ELECTRICAL LOAD ANALYSIS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
24-52-01 Page 1 24-52-01 Page 1 24-52-01 Page 1
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 24-32-00-720
Gill Flooded Lead-Acid Battery Functional Check (Capacity Check)
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24-33-00-720
Concord Sealed Lead Acid Battery Functional Check (Capacity Check)
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24-34-00-720
Marathon Ni-Cad Battery Functional Check (Capacity Check)
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24-36-00-220
Standby Alternator Detailed Inspection
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Power Distribution Boxes Detailed Inspection
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL POWER - GENERAL 1.
2.
Scope A.
This chapter gives information about the electrical units and components which generate, control, and supply electrical power for the airplane systems. This includes such items as generators, relays, batteries, and power outlets.
B.
The system is equipped with an external power receptacle for ground operation of the electrical equipment and battery conservation when starting the engine.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Digital Voltmeter
Model 87
John Fluke Mfg. Co. 6920 Seaway Blvd. Everett, WA 98206
General electrical use.
Available Commercially
Supply external power.
Variable Power Supply GCU Test Box
9870008-1
Fabricate Locally
Used during electrical power system functional check.
Alignment Screwdriver
5000 GC
GC Electronics 1801 Morgan Street Rockford, IL 61102
Used to adjust generator control unit.
Battery Charger
RF80H
Christie Electric Corp. 20665 Manhattan Place Torrance, CA 90501
Charge battery.
Hydrometer (1.100 to 1.310 specific gravity range)
Available Commercially
Test specific gravity of electrolyte.
Small syringe
Available Commercially
Service battery.
Nonmetallic Brush (Acid Resistant)
Available Commercially
Cleaning battery cells.
GO/NO-GO Gage
TBP 2132
Teledyne Battery Products 840 West Brockton Redlands, CA 92373
Used to test connector.
Adhesive
41-30
Mid-West Industrial Chemical Co. 1509 Sublette St. Louis, MO 63110
Used to secure battery vent drain tubes to battery case elbows.
Rubber Gloves, Rubber Apron, and Protective Goggles
Available Commercially
Clean battery.
Cleaning Cloth
Available Commercially
Clean battery.
Penn Central Corp. Marathon Battery Co. 8301 Imperial Drive Waco, TX 76714
Charge ni-cad battery.
Battery Charger/ Analyzer
PCA-131-50/60
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
Torque Wrench 0 to 150 inch-pounds
MANUFACTURER
USE
Available Commercially
Torque components.
Fill Cap Vent Plug Wrench (Nylon)
16515
Marathon Battery Co.
Servicing battery.
S1694-9-54.00 Inch Tubing
FIT-105-3/4
Alpha Wire Corp. 711 Lidgerwood Ave. Elizabeth, NJ 07207
Protective wire sheath used on GCU Test Box.
Feed-Thru
Lapp-SL13
Lapp Inc. 30 Plymouth Road Fairfield, NJ 07006
Wire feed-thru used on GCU test box.
Pin Jack
105-0803-001
E.F. Johnson Company 299 10th Ave. Southwest Waseca, MN 56093
Used in GCU test box.
Switch
30-1
Grayhill Inc. P.O. Box 373 La Grange, IL 60525
Used in GCU test box.
S392-3 Switch
8824K8
Cutler-Hammer Inc. Aerospace Switch Plant Manatee, FL
Used in GCU test box.
Placard
9870008-4
Cessna Aircraft Company
Placard used for GCU test box.
Box
9870008-2
Newark Electronics 500 N. Pulaski Road Chicago, IL 60624
GCU test box shell.
Lamp Holder
920-401X330RN
Newark Electronics
Lamp holder used in GCU test box.
Lamp
327
Newark Electronics
Lamp used in GCU test box.
Sealer
Ablative RTV
Dow Corning 3901 S Saginaw Rd. P. O. Box 997 Midland, MI 48640
To seal standby electrical system relay cover.
Anti-Static Adjustment Tool
758ie8608
Techni-Tool 1547 N. Trooper Rd. P. O. Box 1117 Worchester, PA 19490-1117
To adjust the Alternator Control Unit output voltage of the standby electrical system.
Power Outlet Tester
VEC008 LCD Voltage Meter (Power Outlet Tester) or a Digital Multi-Meter
Direct Depot 109 Wyndham Way Wilmington, NC 28411
To test for the correct voltage of the 12VDC power outlet system.
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3.
Definition A.
The chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. Brief definitions of the sections incorporated in this chapter are as follows: (1) The section on DC generation covers that portion of the system used to generate, regulate, control and indicate DC electrical power. Included are such items as the DC generator system, battery system and DC indicating systems. (2) The section on external power covers that portion of the system which connects external electrical power to the airplane electrical system. (3) The section on load distribution covers that portion of the system which provides connection of DC power to the using systems.
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL POWER - ADJUSTMENT/TEST 1.
General A.
2.
This adjustment/test procedure is designed to furnish maintenance technicians with tools and information needed to test and troubleshoot various functions of the electrical system (voltage, over voltage protection, generator control unit operation, etc.). Included in this section are the following topics: (1) Fabrication of Generator Control Unit (GCU) Test Box. (a) This topic covers fabrication of the GCU test box and provides a detailed list of parts needed to build the box. (2) Electrical System Adjustment/Test. (a) This topic provides information needed to test various electrical systems using the GCU test box. Included within this topic are matrix tables, a legend used in conjunction with the matrix tables and a starter/generator wiring diagram.
Fabrication of GCU Test Box A.
Fabrication Procedures (Refer to Figure 501). (1) Refer to Figure 501 for fabrication and wiring of the text box. (2) For fabrication of GCU test box, refer to Table 501 for a listing of wire gage and terminals.
Table 501. GCU Test Box Material Requirements WIRE CODE
GA
MATERIAL
-50 thru -54
20
-20-9
MS25036-101
SOLDER
-55
16
-16-9
MS25036- 106
SOLDER
-56 thru -73
20
-20-9
MS25036-101
SOLDER
-74
20
-20-9
MS25036-101
M39029/30-217 (NOTE 1)
-75
20
-20-9
MS25036-101
M39029/30-217 (NOTE 1)
-76 thru -78
20
-20-9
M39029/30-217 (NOTE 1)
MS25036-101
-79
16
-16-9
M39029/30-218 (NOTE 1)
MS25036-106
-80 thru -97
20
-20-9
M39029/30-217 (NOTE 1)
MS25036-101
-98 thru -115
20
-20-9
MS25036-101
SOLDER
-116
16
-16-9
MS25036-106
SOLDER
-117 thru -121
20
-20-9
MS25036- 101
SOLDER
-122
20
-20-9
SOLDER
SOLDER
-123
20
-20-9
SOLDER
SOLDER
-124 thru -140
16
-16-9
SOLDER
SOLDER
-141
20
-20-9
SOLDER
MS25036-101
-142 thru -152
20
-20-9
SOLDER
SOLDER
-153
20
-20-9
SOLDER
MS25036-101
-154
20
-20-9
MS25036-101
MS25036-101
TERMINALS
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MODEL 208 MAINTENANCE MANUAL Table 501. GCU Test Box Material Requirements (continued) WIRE CODE
GA
MATERIAL
-155
16
-16-9
MS25036-106
MS25036-101
-187
20
-20-9
SOLDER
SOLDER
-196
20
-20-9
SOLDER
SOLDER
-198
16
-16-9
MS25036-106
MS25036-110
TERMINALS
NOTE 1: Typical. Use this part number to order if necessary. 3.
Electrical System Adjustment/Test A.
Prepare Airplane For Test. (1) Electrical system components shall be installed per factory installation. (2) Fuel tanks shall be closed. (3) Airplane shall be parked and grounded. (4) All switches shall be in the off or normal position. (5) Ensure that battery has a good state of charge. NOTE: (6)
(7) (8) B.
The following procedure is performed with the GCU test box, an analog display multimeter and applicable airplane wiring diagrams.
Disengage the following circuit breakers on the forward firewall power junction box: (a) KEEP ALIVE 1. (b) KEEP ALIVE 2. (c) CLOCK. Disengage all circuit breakers on the circuit breaker panel with exception of START CONT, GEN CONT, GEN FIELD, and all circuit breakers labeled BUS 1 PWR and BUS 2 PWR . Disconnect battery connector from battery.
Generator Control Unit Check. (Refer to Figure 502). (1) Disconnect electrical connector from GCU and connect to GCU test box receptacle. (2) Place OHMS/LIGHTS switch on the GCU test box to the OHMS position.
CAUTION: Insulate or secure power wires PB1 and PB40 to avoid inadvertent contact with airframe. (3)
Disconnect power wires PB1 and PB40 from large terminal of the starter/generator. Refer to Chapter 80, Starter/Generator - Maintenance Practices. (4) Disconnect speed sensor connector from starter/generator. Verify high resistance of over 1000 ohms between pins X and Y on GCU test box, and between pins X and Y and airframe ground. (5) Connect speed sensor connector to starter/generator. Verify low resistance of 200 ohms or less between pins X and Y on GCU test box. (6) Measure resistance between pin G and airplane ground. Resistance shall be less than 1.0 ohm. (7) Measure resistance between pin B and airplane ground. Resistance shall be approximately 2.0 ohms, but not greater than 4.0 ohms. (8) Visually verify the following wire connections on the starter/generator: (a) Wire PB6 to terminal A. (b) Wire PB5 to terminal D. (c) Wires PB2, PB3, PB4 and PB25 to terminal E. (9) Connect a well-regulated external power supply to airplane and set at 28.0 VDC, +0.5 or -0.5 VDC. (10) Place airplane’s external power switch in the BUS position. (11) Place VOLT/AMMETER selector switch to the VOLT position. Voltmeter shall read 28.0 VDC. (12) Move selector switch on GCU Test Box from OHMS to LIGHTS.
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GCU Test Box Fabrication Figure 501 (Sheet 1)
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GCU Test Box Fabrication Figure 501 (Sheet 2)
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GCU Test Box Fabrication Figure 501 (Sheet 3)
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GCU Test Box Fabrication Figure 501 (Sheet 4)
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Engine Not Running Matrix Table Figure 502 (Sheet 1)
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Engine Not Running Matrix Table Figure 502 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (13) Using the Engine Not Running Matrix Table, verify proper GCU operation and light indications. (a) If GCU test box lights indicate a fault, refer to Table 502, Matrix Table Fault Analysis to determine cause. (14) Verify battery switch is ON and start switch is in the START position. Verify voltage at starter/ generator wire PB1 and PB40 to ground is 28.0 VDC. (15) Turn start switch to the OFF position. (16) Place battery switch to OFF position. Turn external power unit off. Do not reconnect starter/ generator. C.
External Power Check. (1) Place airplane external power switch to the BUS position and airplane battery switch to OFF. (2) Ensure all circuit breakers are disengaged on the circuit breaker panel with exception of START CONT, GEN CONT, GEN FIELD, and all circuit breakers labeled BUS 1 PWR and BUS 2 PWR. (3) Apply external power to the airplane and increase voltage until airplane power is cut off. This cutoff shall occur at 31.5 VDC, +0.5 or -0.5 VDC. (4) Decrease external power voltage to 0.0 VDC. Ensure all airplane power remains off. (5) Slowly increase external power voltage until airplane system is reactivated. This shall occur at 22.0 VDC, +1.0 or -1.0 VDC. (6) Set external power voltage to 28.5 VDC or 10 amp maximum. (7) Place airplane battery switch to ON position. With ammeter selector switch placed to BATT position, check to verify battery is being charged. (8) Turn BATT switch to OFF position. Turn external power switch to OFF position. (9) Disconnect external power unit. (10) Reconnect wiring to starter/generator.
D.
GCU Check with Engine Running. (Refer to Figure 503 and Figure 504). (1) Start engine. Refer to Pilot’s Operating Handbook and Approved Airplane Flight Manual. (2) Using the Matrix Table With Engine Running, verify proper GCU operation and light indications. (a) If GCU test box lights indicate a fault, refer to Table 502 and Figure 504 to determine cause and remedy of fault indication. (3) Shut down airplane. (4) Turn battery and external power switches OFF. (5) Remove GCU Test Box. (6) Reconnect airplane wiring to GCU.
Table 502. Matrix Table Fault Analysis TROUBLE
PROBABLE CAUSE
REMEDY
ANY LAMP ON GCU TEST BOX EXCEPT "S" AND "E" ILLUMINATES BEFORE GCU IS CONNECTED.
Faulty wiring or switch.
Check wiring to associated switch. Clear all faults before connecting GCU.
LAMP "A" INOPERATIVE.
Faulty wiring between GEN FLD breaker and current limiter in Power Box.
Check wires. Repair or replace.
LAMP "S" INOPERATIVE.
Faulty wiring between START CONT breaker and current limiter on Bus in Power Box.
Check wires. Repair or replace.
LAMP "E" INOPERATIVE.
GEN CONT SENSE breaker open or faulty.
Reset breaker or replace.
LAMP "Z" INOPERATIVE. (WITH LAMP "W" ILLUMINATED, EXCEPT ON O/V TEST).
Faulty GCU.
Check lamp "Z". Replace lamp "Z". If ok, replace GCU.
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MODEL 208 MAINTENANCE MANUAL Table 502. Matrix Table Fault Analysis (continued) TROUBLE
PROBABLE CAUSE
REMEDY
LAMP "W" INOPERATIVE.
Faulty wiring through GEN FLD breaker GEN switch, BAT switch, STR switch, IGN switch, and START CONT breaker.
Check wiring. Repair or replace wiring.
LAMP "H" INOPERATIVE. AFTER START.
Start switch ON or faulty GCU.
Verify start switch OFF, Check wiring. Repair or replace GCU.
LAMP "T" INOPERATIVE.
Faulty GEN switch wiring or faulty GCU.
Check GEN switch, generator wiring or anti-cycle breaker.
LAMP "V" INOPERATIVE.
Faulty GEN switch wiring or faulty GCU.
Check GEN switch wiring. Check power from pin "V" when lamp "A" is illuminated. If not, replace GCU.
LAMP "J" INOPERATIVE.
Faulty GEN switch wiring or faulty GCU.
Check GEN switch wiring. Check power from pin "V" when lamp "A" is illuminated. If not, replace GCU.
LAMP "D" INOPERATIVE.
Faulty wiring or GEN CONT breaker open.
Check GEN CONT breaker connections through limiter.
LAMP "B" ILLUMINATES BEFORE ENGINE START.
Faulty wiring.
Check wire from generator to GCU for proper connections and continuity.
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Matrix Table with Engine Running Figure 503 (Sheet 1)
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Starter/Generator Wiring Diagram Figure 504 (Sheet 1)
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Starter/Generator Wiring Diagram Figure 504 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL DIRECT CURRENT GENERATION - DESCRIPTION AND OPERATION 1.
General A.
Electrical power for the airplane consists of a 24 VDC, 45 ampere-hour lead acid battery or optional ni-cad battery. A 200-ampere, engine-driven starter/generator is installed in the airplane. The starter/ generator functions as a starter motor for engine starting and as a generator after start, when the starter switch is placed in the OFF position above 41 percent Ng (gas generator) speed.
B.
This section provides maintenance information for the generator control unit (GCU). For maintenance of the starter/generator (including generator brushes), refer to Chapter 80 Starter/Generator - Maintenance Practices. For removal/installation of the blast tube, refer to Chapter 80, Starter/Generator Cooling Air Blast Tube - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL GENERATOR CONTROL UNIT - MAINTENANCE PRACTICES 1.
2.
Description A.
The generator control unit (GCU) is a solid-state unit installed on the left cabin sidewall just forward of FS 118.00. The unit monitors and controls the electrical power system, sequencing system operation from starting the engine through generator operation. The unit is equipped with overvoltage sensing. This maintenance practice provides removal/installation and voltage adjustment procedures for the GCU.
B.
For GCU (and electrical power) troubleshooting, refer to Electrical Power - Adjustment/Test.
GCU Removal/Installation A.
Remove GCU (Refer to Figure 201). (1) Disconnect airplane battery. Place maintenance tag on instrument panel with following statement: NOTE: (2) (3) (4) (5)
B.
3.
Do not turn BATT switch on. Battery disconnected from airplane.
Remove substrate panel forward of left circuit breaker panel to gain access to GCU. Disconnect electrical connector from GCU. Remove screws securing GCU to airplane. Remove GCU from airplane.
Install GCU (Refer to Figure 201). (1) Position GCU on bracket and secure using screws. (2) Connect electrical connector to GCU. (3) Install substrate panel. (4) Reconnect airplane battery and remove maintenance tag from instrument panel.
GCU Adjustment/Test
CAUTION: Do not disconnect electrical connector from GCU while power is applied to the airplane. A.
Adjustment Procedures. (1) Remove substrate panel forward of circuit breaker panel on lower left cabin sidewall to access GCU. (2) Connect voltmeter to voltage test jacks on GCU. (3) Start engine. Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. (4) Ensure battery switch is in the ON position. (5) With start switch in the start position and engine operating at 80 percent Ng, check voltage at test jacks. Voltage should indicate 28.5 VDC, +0.1 or -0.1 VDC. If not, actuate the generator switch to TRIP and then to RESET. If generator still does not build up voltage, check GCU circuits.
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Generator Control Unit Installation Figure 201 (Sheet 1)
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Generator Control Unit Installation Figure 201 (Sheet 2)
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Generator Control Unit Installation Figure 201 (Sheet 3)
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CAUTION: If a metallic screwdriver is used, a short circuit may occur, damaging the GCU. CAUTION: If the GCU regulated voltage is changed, then the alternator control unit for the standby alternator (if installed) must also be changed. This voltage should be set 1.0 VDC below the GCU setting. (6)
Loosen dust cover over the GCU voltage adjustment screw. Slowly rotate the voltage adjustment screw using a nonmetallic screwdriver until voltage at test jack reads 28.5 VDC, +0.1 or -0.1 VDC. For recommended voltage settings due to temperature variations, refer to Table 201 or Table 202. NOTE:
The recommended voltage settings are an approximation only. A general rule to follow in setting voltage is to lower the adjustment when the battery is using too much water (indicating too high voltage) or to increase the adjustment when the battery does not remain charged.
Table 201. Voltage Settings With Lead Acid Battery OUTSIDE AIR TEMPERATURE
GCU VOLTAGE
120°F and Above
27.5 VDC, +0.1 or -0.1 VDC
60°F
28.5 VDC, +0.1 or -0.1 VDC
0°F and Below
29.5 VDC, +0.1 or -0.1 VDC
Table 202. Voltage Settings With Ni-Cad Battery OUTSIDE AIR TEMPERATURE
GCU VOLTAGE
Operation in regions where outside air temperature does not get above 60°F.
Increase GCU setting by 0.5 VDC
Operation in regions where outside air temperature does not get below 60°F.
Decrease GCU setting by 0.5 VDC
NOTE 1: Bus voltage may vary up to 0.5 volts (plus or minus) over the range from zero to full load and from 65 percent to 100 percent Ng engine speed. For engine speeds below 65 percent, bus voltage may drop off considerably with higher loads. (7) (8)
Stop engine and remove voltmeter from GCU test jack. adjustment screw. Install substrate panel.
Secure dust cover overvoltage
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MODEL 208 MAINTENANCE MANUAL FLOODED LEAD-ACID BATTERY - TROUBLESHOOTING 1.
General A.
A troubleshooting chart is included to aid the maintenance technician in system understanding. Refer to Figure 101.
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Flooded Lead-Acid Battery Troubleshooting Chart Figure 101 (Sheet 1)
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Flooded Lead-Acid Battery Troubleshooting Chart Figure 101 (Sheet 2)
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Flooded Lead-Acid Battery Troubleshooting Chart Figure 101 (Sheet 3)
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Flooded Lead-Acid Battery Troubleshooting Chart Figure 101 (Sheet 4)
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Flooded Lead-Acid Battery Troubleshooting Chart Figure 101 (Sheet 5)
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Flooded Lead-Acid Battery Troubleshooting Chart Figure 101 (Sheet 6)
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MODEL 208 MAINTENANCE MANUAL FLOODED LEAD-ACID BATTERY - MAINTENANCE PRACTICES 1.
Description A.
2.
Standard airplane battery is a 24 VDC, 45 ampere-hour (12 cell) lead acid battery, located on the right forward side of the fire wall and is comprised of lead compound plates immersed in a diluted solution of sulfuric acid and water (electrolyte). Each cell, connected in series, has a nominal voltage of approximately 2.0 volts when fully charged. State-of-battery charge can be determined by checking specific gravity with a hydrometer and voltage test.
General Precautions and Notes A.
Proper maintenance is essential if the battery is to achieve maximum life and performance. To assure these goals, periodic inspection in the airplane and periodic maintenance is a must.
WARNING: National Electric Code forbids charging batteries installed in airplanes or within 10 feet of fuel tank areas. CAUTION: Separate lead acid and nickel-cadmium battery facilities, including separate shops and service tools, must be used to prevent electrolyte contamination of acid and alkaline batteries. CAUTION: To minimize battery discharge during airplane storage or periods of low airplane utilization (more than 5 days), battery should be disconnected and/or circuit breakers disengaged on all items on hot battery bus bar. B.
A battery should never be allowed to remain in a discharged condition for any appreciable time. If allowed to remain in a discharged condition, the lead sulfate will grow into a hard, white crystalline formation known as sulfation. This condition closes the pores in the active material and destroys the plates.
C.
Electrolyte level must be maintained above the plates. Failure to do so leaves the plates exposed to air and causes rapid sulfation. Regular checking of the electrolyte level is a necessity and, if low, should be filled with distilled water.
D.
When placed on a charge, some lead sulfate, instead of reverting to spongy lead or lead oxide, dislodges from plates in small particles and drops to the bottom as sediment, resulting in irreparable damage. This material is lost for active use. In normal operation, all cells shed a small amount of active material; however, this process is quickened in the case of a sulfated battery, and its life is greatly reduced.
E.
A greenish deposit of copper salts may form on terminals and connectors. This corrosion is caused by normal venting or spilled electrolyte and should be removed using a stiff brush, followed by thoroughly washing the area with ammonia and water or a five percent solution of baking soda and water to neutralize any remaining electrolyte. Areas shall then be coated with a thin film of grease or preventive compound to prevent corrosion. Using a voltmeter from the negative terminal to the outside surface, check exterior of battery for acid bridges. A voltage reading indicates an acid bridge which must be removed by a thorough washing and drying.
F.
Electrolyte shall be added to an older battery only if electrolyte is lost as a result of spillage. A fully charged battery may have a specific gravity of 1.285 to 1.295 when new, whereas, it may have a fully charged specific gravity of 1.260 to 1.275 when near the end of its life. In this case, electrolyte shall not be added. The plates may be slightly sulfated and the addition of a higher specific gravity electrolyte will only aggravate this condition.
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3.
Initial Charge for New Batteries
CAUTION: Keep sparks, flames, burning cigarettes or other ignition sources away from battery at all times, and always shield eyes when working near batteries. A.
Preparation for Initial Charge. NOTE:
(1)
Although new batteries are received dry-charged and will deliver 75 percent of rated capacity after initial filling of electrolyte, it is essential that battery be given an initial charge to full capacity to ensure its airworthiness before installing in airplane.
Remove seals from battery cells.
CAUTION: Do not use automotive electrolyte to service battery. (2) (3) (4) (5) (6) B.
Fill each cell with 1.285 specific gravity electrolyte to bottom of split ring. Using care not to spill electrolyte, gently rock battery from side to side to release any trapped air. Readjust electrolyte as necessary. Allow battery stand for one hour. Readjust by adding electrolyte to proper level. Install vent plugs tightly into each cell.
Initial Battery Charge.
CAUTION: Do not allow battery to stand longer than 10 hours before beginning charge. (1) (2) (3) (4) (5) 4.
Battery Inspection A.
5.
Charge battery at 3.5 amps until initial gassing begins. Continue charge at 2.5 amps until all cells are gassing freely and charge voltage and specific gravity of electrolyte are constant over three successive readings taken at one hour intervals. During period of charging, electrolyte temperatures shall be maintained between 60°F and 110°F. When battery is completely charged, verify specific gravity is between 1.285 and 1.295. Adjust electrolyte level by removing or adding electrolyte to bottom of split ring, as required. After charge is complete, neutralize and remove any electrolyte spilled on battery.
Visual inspection of battery in airplane should be done in accordance with time limits set forth in Chapter 5, Inspection Time Limits.
Battery Quick-Disconnect Inspection A.
Inspect Battery and Components. (1) Check for excessively loose handle and locking assembly. (2) Check for pitted or corroded mating surfaces. (3) Check for burn marks caused when battery is disconnected under load. (4) Test for resiliency of mating surfaces of Elcon connector oversized pin. (a) Insert larger 0.385 inch diameter probe of GO/NO-GO gage into each helix or sleeve to maximum depth. Ensure a snug fit with a removal force greater than one pound. (b) If connector fails to pass resiliency test, replace connector. (5) Insert smaller 0.370 inch diameter probe of GO/NO-GO gage into connector to ensure adequate contact is present. (a) Elcon connector, ensure a snug fit with a nominal removal force of one pound. (b) Rebling connector, ensure each socket exerts sufficient pressure on pin to hold 0.370 inch diameter GO/NO-GO gage securely when quick-disconnect is inverted with gage pointed downward. (c) If connector exhibits excessive wear or damage, replace connector.
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6.
Battery Receptacle Inspection A.
7.
Servicing Battery A.
8.
9.
Inspect Connector Pins. (1) Connector pins shall be inspected for corrosion, pitting or burn marks. If any conditions prevent total electrical contact, surface shall be cleaned. (2) If cleaning process reduces pin diameter less than 0.370 inch, battery shall be replaced.
For servicing of battery, refer to Chapter 12, Flooded Lead Acid Battery - Servicing.
Battery Removal/Installation A.
Remove Battery (Refer to Figure 201). (1) Ensure battery switch is positioned to OFF. (2) Open right cowl door. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices. (3) Disconnect battery connector from battery. (4) Pull lever to release battery tray from latch on fire wall. (5) Swing battery tray away from fire wall. (6) Cut and remove safety wire from wing nuts. (7) Remove wing nuts and washers from battery cover; remove battery cover. (8) Remove vent lines from elbows. (9) Clean adhesive from elbows and vent lines using isopropyl alcohol. (10) Remove battery from airplane.
B.
Install Battery (Refer to Figure 201). (1) Clean battery support and battery tray as necessary to ensure proper installation. (2) Position battery on battery tray, but do not secure. (3) Connect battery connector to battery and hand tighten. (4) Install battery cover using washers and wing nuts. (5) Safety wire wing nuts. Refer to Chapter 20, Safetying - Maintenance Practices. (6) Apply 14-30 adhesive to vent lines and elbows. (7) Install vent lines to elbows. (8) Swing battery aft until lever engages latch on fire wall. (9) Close right cowl door. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices.
Battery Adjustment/Test A.
General.
CAUTION: Tools or equipment used for servicing nickel-cadmium batteries shall not be used for servicing lead acid batteries. Lead acid batteries shall be completely removed from nickel-cadmium battery service areas. The slightest acid contamination will deteriorate nickel-cadmium batteries. CAUTION: Do not charge batteries installed in airplane or within 10 feet of fuel tank areas. (1)
All procedures are to be accomplished in a designated service area away from airplane.
B.
Tools, Equipment and Materials. (1) For a list of tools, equipment and materials, refer to Electrical Power - General.
C.
Battery Test.
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Flooded Lead Acid Battery Installation Figure 201 (Sheet 1)
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CAUTION: If electrolyte is removed from a discharged cell and replaced with electrolyte of high specific gravity, cell will be in a discharged condition even though hydrometer test indicates a full charge. (1)
Specific gravity of battery may be measured with a hydrometer to determine the state of battery charge. If hydrometer reading is low, charge battery and retest. Hydrometer readings of electrolyte must be compensated for temperature of electrolyte. Refer to Table 201 for various hydrometer readings with an electrolyte temperature of 80°F.
Table 201. Battery Hydrometer Reading at 80°F
D.
READINGS
BATTERY CONDITION
1.280 Specific Gravity
100 Percent Charged
1.250 Specific Gravity
75 Percent Charged
1.220 Specific Gravity
50 Percent Charged
1.190 Specific Gravity
25 Percent Charged
1.160 Specific Gravity
Discharged
NOTE:
All readings shown are for an electrolyte temperature of 80°F. For higher temperatures, readings will be slightly lower. For cooler temperatures, readings will be slightly higher. Some hydrometers have a built-in temperature conversion chart and a thermometer. Corrected readings shall agree with Table 201.
NOTE:
If a specific gravity indicates battery is not fully charged, charge battery atapproximately 29.0 VDC for 30 minutes or until battery voltage rises to 28.0 VDC. After charging, a load tester will provide more accurateresults. A specific gravity check can be used for charging, but cannot identify cells which short under load or have broken connectors between plates or cells.
Battery Charging.
WARNING: When a battery is being charged, hydrogen and oxygen gases are generated. Accumulation of these gases can create a hazardous explosive condition. Always keep sparks and open flame away from battery. Allow unrestricted ventilation of battery area during charging. (1)
(2)
10.
Remove battery from airplane and place in a well ventilated area for charging. When battery is fully charged, electrolyte level must be checked and adjusted by adding distilled water at a level even with horizontal baffle plate or split ring at bottom of filler holes. If battery is extremely cold, allow battery to warm before adding water as level will rise with warming. Main points of consideration during a battery charge are excessive battery temperature and violent gassing. Test battery with a hydrometer to determine amount of charge. Decrease charging rate or stop charging temporarily if electrolyte temperature exceeds 115°F.
Battery Cleaning A.
Tools, Equipment and Materials. (1) For a list for required tools, equipment and materials, refer to Electrical Power - General.
B.
Cleaning Procedures. (1) Remove battery. Refer to Battery Removal/Installation. (2) Tighten battery cell filler caps to prevent cleaning solution from entering cells.
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Wipe battery cable ends, battery terminals and entire surface of battery with a clean cloth moistened with a solution of bicarbonate of soda (baking soda) and water. Rinse with clear water, wipe away excess water and allow battery to dry. Examine vent plugs to ensure gas escape holes are clear of obstruction. Brighten cable ends and battery terminals using emery cloth or a wire brush. Coat battery terminals with petroleum jelly or an ignition spray product to reduce corrosion. Install battery. Refer to Battery Removal/Installation.
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MODEL 208 MAINTENANCE MANUAL FLOODED LEAD-ACID BATTERY - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the battery in a serviceable condition.
Task 24-32-00-720 2.
Gill Flooded Lead-Acid Battery Functional Check (Capacity Check) A.
General (1) This section gives the information needed to do the functional check of the flooded lead-acid battery.
B.
Special Tools (1) Elcon Inspection Gauge #029 or Equivalent.
C.
Access (1) Open the right cowling door to get access to the battery. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do the Gill Flooded Lead-Acid Battery Functional Check (Capacity Check). (1) Remove the battery from the airplane. Refer to Flooded Lead-Acid Battery - Maintenance Practices. (2) Visually examine the lead acid battery for general condition. (3) Examine the terminals for an overheat indication, burns, or signs of arcing. (4) Examine the vent tubes for deterioration or wear. (5) Make sure that the battery tray is clean. (6) Do a capacity check of the flooded lead-acid battery. Refer to the Gill Flooded Lead-Acid Main Battery Maintenance Manual Supplement. (7) Install the battery in the airplane. Refer to Flooded Lead-Acid Battery - Maintenance Practices.
E.
Restore Access (1) Close the right cowling door. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL SEALED LEAD ACID BATTERY - MAINTENANCE PRACTICES 1.
General A.
2.
3.
NOTE:
Battery is serviced and charged at the factory. Although the battery is a maintenance free battery, to ensure airworthiness, battery capacity must be checked periodically. Refer to Chapter 5, Inspection Time Limits.
NOTE:
When replacing Ni- Cad battery with lead acid battery, refer to Ni-Cad Battery - Removal/ Installation for modification of Battery Overheat Warning System.
Battery Removal/Installation A.
Remove Battery (Refer to Figure 201). (1) Ensure battery switch is in OFF position. (2) Open right side cowl door. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (3) Disconnect battery connector from battery. (4) Pull lever to release battery tray from latch on fire wall. (5) Swing battery tray away from fire wall. (6) Cut and remove safety wire from wing nuts. (7) Remove wing nuts and washers from strap securing battery cover to battery and tray. (8) Remove battery cover. (9) Remove vent lines from elbows. (10) Clean adhesive from vent lines and elbows using isopropyl alcohol. (11) Remove battery from airplane.
B.
Install Battery (Refer to Figure 201). (1) Clean battery support and battery tray as necessary. (2) Position battery on battery tray, but do not secure . (3) Connect battery connector to battery and hand tighten. (4) Position battery cover on battery and secure with strap, washers and wing nuts. (5) Safety wire wing nuts. Refer to Chapter 20, Safetying - Maintenance Practices. (6) Apply 14-30 adhesive to vent lines and elbows. (7) Install vent lines to elbows. (8) Swing battery aft until lever engages latch on fire wall. (9) Close right cowl door. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices.
Battery Receptacle Inspection A.
4.
The battery is a maintenance free, 24 VDC, 40 ampere-hour sealed lead acid battery. The battery is a recombinant gas (RG) absorbed electrolyte battery. Because the electrolyte is absorbed in glass mat (AGM) separators, no leakage will occur, even if the case is cracked or damaged through mishandling. The battery is equipped with overboard vent lines, which connect to the vent fittings on the battery case. The battery is located on the right side of the forward fire wall.
Items To Inspect. (1) Connect pins should be inspected for corrosion, pitting or burn marks. If any of these defects are evident to the extent that total electrical contact could be prevented, the surface shall be cleaned. (2) If cleaning process reduces pin diameter below 0.370 inch, the battery shall be replaced.
Battery Quick-Disconnect Inspection A.
Items To Inspect (1) Check for excessively loose handle and locking assembly. (2) Check for pitted or corroded mating surfaces. (3) Check for burn marks caused when battery is disconnected under load.
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Sealed Lead Acid Battery Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (4) (5)
(6)
Test for resiliency of mating surfaces to an oversized pin (Elcon connector only). Insert larger diameter probe (0.385 inch diameter) of a GO/NO-GO gage into helix or sleeve to maximum depth. Ensure a snug fit with a removal force greater than one pound. To assure contact is adequate for a worn battery pin, insert small diameter end (0.370 inch diameter) of GO/NO-GO gage. (a) Elcon connector, ensure a snug fit with a nominal removal force of one pound. (b) Rebling connector, ensure each socket exerts sufficient pressure on the pin to hold the 0.370 inch diameter GO/NO-GO gage when the quick-disconnect is inverted to a position where the gage is pointed downward. If the connector fails to pass resiliency test or shows excessive wear or damage, replace connector.
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MODEL 208 MAINTENANCE MANUAL SEALED LEAD ACID BATTERY - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the battery in a serviceable condition.
Task 24-33-00-720 2.
Concord Sealed Lead Acid Battery Functional Check (Capacity Check) A.
General (1) This section gives the information needed to do the functional check of the sealed lead acid battery.
B.
Special Tools (1) None
C.
Access (1) Open the right cowling door to get access to the battery. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do the Concord Sealed Lead Acid Battery Functional Check (Capacity Check) (1) For airplanes that operate less than 1000 hours per year: (a) Do the battery capacity check at 12 calendar months after the initial battery installation (plus or minus one month). (b) As long as the battery capacity check is above 90 percent, then you should do subsequent battery capacity checks every six months (plus or minus one month). If the battery capacity check is between 85 and 90 percent, then you should do 1 subsequent battery capacity checks every 3 months (plus or minus one month). If the battery capacity check is less than 85 percent, install a new battery. 2 (2) For airplanes that operate 1000 or more hours per year: (a) Do the battery capacity check at 1000 hours after the initial battery installation (plus or minus 100 hours). (b) As long as the battery capacity check is above 90 percent, then you should do subsequent battery capacity checks every 500 hours (plus or minus 100 hours). If the battery capacity check is between 85 and 90 percent, then you should do 1 subsequent battery capacity checks every 250 hours (plus or minus 100 hours). If the battery capacity check is less than 85 percent, install a new battery. 2 (3) Remove the battery from the airplane. Refer to Sealed Lead Acid Battery - Maintenance Practices. (4) Visually examine the lead acid battery for general condition. (5) Examine the terminals for an overheat indication, burns, or signs of arcing. (6) Examine the vent tubes for deterioration or wear. (7) Make sure that the battery tray is clean. (8) Do a capacity check of the lead acid battery. Refer to the Concorde Valve Regulated Lead Acid Main Battery Maintenance Manual Supplement. (9) Install the battery in the airplane. Refer to Sealed Lead Acid Battery - Maintenance Practices.
E.
Restore Access (1) Close the right cowling door. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - DESCRIPTION AND OPERATION 1.
General A.
2.
The Model 208 may incorporate a 25.2 VDC, 40 ampere hour (20 cell) nickel-cadmium battery, located on the right forward side of the firewall.
Description A.
Battery. (1) The electrolyte in a nickel-cadmium battery is a solution of distilled water and potassium hydroxide. The electrolyte is used only as a conductor and does not react with the plates as does the electrolyte in a lead-acid battery. The state of battery charge cannot readily be determined by a specific gravity reading, since the electrolyte does not change appreciably. For this reason, it is not possible to determine the charge state of a nickel-cadmium battery by checking the electrolyte with a hydrometer. Neither can the charge be determined by a voltage test, because of the inherent characteristic that the voltage remains constant during 90 percent of the discharge cycle. However, a visual indication is beneficial because the plates are porous and absorb the electrolyte while discharging and expel the electrolyte while charging. (2) The negative plates in the battery are cadmium hydroxide, the positive plates are nickel hydroxide. During charging, all oxygen is driven out of the negative plates and only metallic cadmium remains. The oxygen dispelled from negative plates is picked up by the positive plates to form nickel dioxide. Toward the end of the charging process, the electrolyte will gas due to electrolysis taking place in the electrolyte. A slight amount of gassing is necessary to completely charge the battery. (3) During discharge, the reverse chemical action takes place. The negative plates gradually gain back the oxygen, as the positive plates lose oxygen. Due to this interchange of oxygen, the chemical energy of the plates is converted into electrical energy and the electrolyte is absorbed by the plates. For this reason, the level of the electrolyte should be checked only when the battery is fully charged.
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MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been prepared to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Ni-Cad Battery Troubleshooting Chart Figure 101 (Sheet 1)
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Ni-Cad Battery Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - REMOVAL/INSTALLATION 1.
General A.
2.
The following removal and installation procedures are for the battery, quick-disconnect receptacle, intercell connector and individual battery cell.
Notes and Precautions A.
Proper maintenance is essential if the battery is to achieve maximum life and performance. To ensure these goals, periodic inspection in the airplane and periodic maintenance is a must.
WARNING: The electrolyte used in nickel-cadmium batteries is a caustic solution of potassium hydroxide. Serious burns will result if it comes in contact with any part of the body. Use rubber gloves, rubber apron and protective goggles when handling this solution. If electrolyte gets on skin, wash affected areas thoroughly with water, and neutralize with three-percent acetic acid, vinegar or lemon juice. If electrolyte gets into eyes, flush with water and get immediate medical attention. WARNING: Rings, metal watchbands and other metallic jewelry should be removed before working around the battery. Should such metallic objects contact intercell connectors of opposing polarity, they may fuse themselves to the connectors and cause severe skin burns. CAUTION: Tools or equipment used for servicing lead-acid batteries shall not be used for servicing ni-cad batteries. Ni-Cad batteries should be completely removed from lead-acid battery service area. The slightest acid contamination will deteriorate Ni-Cad batteries. 3.
Battery Removal/Installation A.
Remove Battery (Refer to Figure 401). (1) Ensure battery switch is positioned to OFF. (2) Open right cowl door. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices. (3) Disconnect battery connector and temperature connector from battery. (4) Pull lever to release battery tray from latch on firewall. (5) Swing battery tray away from firewall. (6) Remove vent lines from elbows. (7) Cut safety wire from wing nuts and remove cover from battery tray. (8) Remove battery from airplane. (9) If battery is to be replaced with a lead acid battery, perform the following steps. (Refer to Battery Overheat Warning - Description and Operation). (a) Jumper pin A to pin B and pin C to pin D. (b) Stow the heat sensor electrical connector. (c) Replace the BATTERY OVERHEAT and BATTERY HOT annunciator lenses with blank ones (P/N 25-0890-89 or equivalent). (d) Install lead acid battery in accordance with Sealed Lead-Acid Battery - Maintenance Practices.
B.
Install Battery (Refer to Figure 401). (1) Clean battery support and battery tray as necessary to ensure proper installation.
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Ni-Cad Battery Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) (6) 4.
Install battery to battery tray. Secure with hold-down rods, washers and wing nuts as required. Safety wire wing nuts. Refer to Chapter 20, Safetying - Maintenance Practices. Connect vent lines to elbows. Connect battery connector and temperature connector to battery. Swing battery aft until handle engages latch on firewall. Close right cowl door. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices.
Quick-Disconnect Receptacle Removal/Installation A.
Remove Quick-Disconnect Receptacle (Refer to Figure 402).
WARNING: Do not drop tools or other metallic objects onto the intercell connectors; severe arcing will occur, resulting in possible injury to personnel and damage to the battery. Only insulated tools should be used for servicing ni-cad batteries. (1)
Remove positive and negative intercell connectors attached to quick-disconnect receptacle. Note position and placement of all hardware for later reinstallation. NOTE:
(2) (3) B.
Care should be taken in removal of quick-disconnect receptacle to preserve all hardware and gaskets, if possible, so that new part may be installed properly.
Remove screws securing quick-disconnect receptacle to battery case. Remove quick-disconnect receptacle and gasket from battery case.
Install Quick-Disconnect Receptacle (Refer to Figure 402). (1) Install quick- disconnect receptacle and gasket to battery case. Secure using screws.
CAUTION: Do not fabricate intercell connectors. Connectors are designed to carry particular electrical loads. If replacement parts are needed, contact marathon battery company for replacement parts. (2) (3) 5.
Install positive and negative connectors to quick-disconnect receptacle. Torque connectors. Refer to Ni-Cad Battery - Adjustment/Test, Table 501 for torque values.
Intercell Connector Replacement
CAUTION: Do not fabricate intercell connectors. Intercell connectors are designed to carry specific electrical loads and should be replaced using only marathon battery parts. A.
6.
Battery cells are connected to each other using (intercell) connectors. Refer to Figure 402 for an illustration of typical connectors and their hardware. Torque intercell connectors to one another using torque values found in Ni-Cad Battery - Adjustment/Test, Table 501.
Battery Cell Removal/Installation A.
Remove Cells. (1) Remove battery from airplane. Refer to Battery Removal/Installation. (2) Clean battery. Refer to Ni-Cad Battery - Cleaning/Painting. (3) Remove (intercell) connectors. Save all hardware for reinstallation. (4) Remove all vent plugs using vent wrench. Refer to Electrical Power - General for vent wrench part number. (5) Remove enough intercell connectors to permit individual cells to be withdrawn from battery case.
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Ni-Cad Connector Installation Figure 402 (Sheet 1)
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Ni-Cad Connector Installation Figure 402 (Sheet 2)
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CAUTION: Do not withdraw cell from battery unless replacement cell is available immediately. (6) B.
Withdraw cell(s) using cell puller tool. Refer to Figure 403 for fabrication details.
Install Cells (Refer to Figure 402). (1) Replace cell, ensuring that cell polarity symbols are oriented correctly. NOTE:
All cells are connected plus to minus.
NOTE:
If cell is difficult to insert, apply a light coat of petroleum jelly or silicone grease to sides of cell case before assembly.
CAUTION: Do not fabricate intercell connectors. Connectors are designed to carry particular electrical loads. If replacement parts are needed, contact marathon battery company for replacement parts. (2) (3) (4) (5)
Install intercell connectors between cells. Tighten finger tight. Torque intercell connectors. Refer to Ni-Cad Battery - Adjustment/Test, Table 501 for torque values. Recharge and test the battery. Refer to Ni-Cad Battery - Adjustment/Test, Battery Reconditioning. The battery is ready to be returned to service.
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Ni-Cad Battery Strap Fabrication Figure 403 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - ADJUSTMENT/TEST 1.
General A.
2.
This adjustment/test procedure is designed to provide maintenance technicians with information needed to recondition Marathon brand Ni-Cad batteries and to provide a method to check for electrical leakage. (1) Reconditioning of batteries consists of charging, discharging, deep cycling and recharging procedures. (a) The preferred method of reconditioning a Ni-Cad battery is to utilize a Marathon PCA-13150/60 battery charger/analyzer. (b) Alternate methods of reconditioning a Ni-Cad battery are also provided in this section. These alternate methods include procedures for constant current, stepped constant current, float charging and constant voltage.
Notes and Precautions A.
A new battery is shipped discharged and contains the proper amount of electrolyte. It does not require leveling even though the battery may appear to have insufficient electrolyte.
B.
The electrolyte, which is a 30 percent solution (by weight) of potassium hydroxide in distilled water, does not take an active part in the chemical reaction. It is used only to provide a path for the current flow. At 70°F, the specific gravity of the solution should remain within the range of 1.24 to 1.30.
C.
An unusual characteristic of theNi-Cad battery is that, when the battery is completely discharged, some cells will reach zero potential and charge in the reverse polarity. This action will adversely affect the battery, such that it will not retain a full capacity charge. As a result, it becomes the equivalent of a much smaller rated battery. The cure for this problem is to deep cycle the battery and short- circuit each cell to obtain a cell balance at zero potential. This process is known as equalization. Battery must be sent to an authorized battery service shop for cell equalization.
WARNING: The electrolyte used in Ni-Cad batteries is a caustic solution of potassium hydroxide. Serious burns will result if it comes in contact with any part of the body. Use rubber gloves, rubber apron and protective goggles when handling this solution. If solution gets on the skin, flush the affected area thoroughly with water, neutralize with three percent acetic acid, vinegar or lemon juice. If electrolyte gets into the eyes, flush with water. Get immediate medical attention. D.
The battery electrolyte is corrosive and should not be serviced in the airplane. Any amount of electrolyte that is expelled will react with carbon dioxide and form white crystals of potassium carbonate. The white crystals are a noncorrosive, nontoxic substance which are easily wiped away with a clean, damp cloth.
E.
The battery may be charged at a temperature of less than -20°F. However, charging is more efficient at battery temperatures between +40°F and +80°F.
F.
The length of time required to fully charge a battery is dependent on the condition of the battery and the method of charging.
G.
The battery requires reconditioning service when the BATTERY OVERHEAT warning light illuminates. Perform a minimum reconditioning program.
H.
Battery reconditioning under a maintenance inspection program or preventive maintenance requires shop inspection and testing.
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MODEL 208 MAINTENANCE MANUAL I.
3.
Tools, Equipment and Materials A.
4.
The following characteristics apply to Ni-Cad batteries: (1) Apparent Loss or Temporary Loss of Capacity - When this temporary loss occurs, the battery capacity will be lower than the rated capacity. This effect is more common when recharging across a constant potential bus. The loss of capacity is normally an indication of imbalance between cells because of differences in temperature, charge efficiency and self-discharge rate in the cells. (2) Minimal Capacity Loss with Age - A loss of capacity is a warning and should not be treated lightly. The only way to accurately check the capacity (state-of- charge) is by measured discharge.
For a list of required tools, equipment and materials, refer to Electrical Power - General.
Battery Reconditioning Using the Marathon PCA-131-50/60 Charger/Analyzer A.
Battery Reconditioning Procedures. NOTE:
(1) (2) (3) (4) (5) (6)
(7) (8)
The Marathon Model PCA-131-50/60 series charger/analyzer is designed to provide maximum service from ni-cad batteries. It features a GO/NO-GO indication of battery conditions. The correct charge and discharge current is preselected with setting of switch position The battery can be left unattended during charge and automatically adjusts for changes in line voltages. It will automatically terminate discharge if the average battery voltage falls below a preselected end voltage. The actual discharge time can be determined from the running timer.
Clean battery. Refer to Ni-Cad Battery - Cleaning/Painting. Perform electrical leak check. Refer to Electrical Leak Check in this section. Connect battery to PCA-131-50/60 charger/analyzer. Rotate SELECTOR knob to MA-5 position. Move AUTOMATIC CYCLE/MANUAL DISCHARGE switch to AUTOMATIC CYCLE position and AC POWER switch to the ON position. Rotate timer control to TOPPING position. Top charge battery until voltage reaches 30.0 VDC minimum. If battery voltage does not rise to 30.0 VDC within 30 minutes, check voltage of each cell and ensure no cells are shorted. If shorted cell is found, replace cell. Refer to Ni-Cad Battery - Removal/Installation. Check electrolyte level in cells and adjust as required. Refer to Chapter 12, Ni-Cad Battery Servicing . Rotate timer control to DISCHARGE position. NOTE:
(9) (10) (11) (12) (13) (14)
Battery will discharge for approximately two hours in the DISCHARGE cycle. If battery conditions are normal, the automatic cycle will continue. If discharge time is less than the 120 minutes, BATTERY LOW light will illuminate on charger/analyzer. If this occurs, set AUTOMATIC CYCLE/MANUAL DISCHARGE switch to MANUAL DISCHARGE position and continue battery discharge.
Fabricate 14 shorting straps and six 1.0 ohm, two-watt resistors. Refer to Figure 501 for fabrication details. As each individual cell reaches 0.5 VDC or less, place shorting strap across both terminals while the load is applied. Ensure switch remains in MANUAL DISCHARGE position. Continue the discharge until 14 of the cells are shorted out with shorting straps. Place the 1.0 ohm, two-watt resistor across each of the remaining cells. Disconnect battery from charger/analyzer and allow shorting straps and resistors to remain attached to battery cells for a minimum of three hours. If the discharge time of step 4.A.(8) was more than 100 minutes, proceed to final charge as detailed in step 4.A.(15). If the discharge time of step 4.A.(8) was less than 100 minutes, the battery must be charged per step 4.A.(15), discharged per step 4.A.(8) through step 4.A.(12), and then charged per step 4.A.(15).
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Shorting Strap and Resistor Fabrication Figure 501 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (15) Set the automatic/manual switch to AUTOMATIC CYCLE position, turn the PCA- 131-50/60 power on, and rotate the clock to the MAIN CHARGE mode. The charge will continue until both the main and topping charges are completed. (16) During the final five minutes of the TOPPING charge (in the event the charger has turned off, reset the charger to TOPPING charge for 10 minutes), check the voltage of each cell (with charger current flowing to the battery). With the battery at room temperature (60° to 90°F), the minimum voltage should be 1.55 VDC per cell, and the maximum voltage should be 1.75 VDC per cell. NOTE:
If any cell fails to rise to a minimum of 1.55 VDC, reset the timer to TOPPING charge for one hour and recheck the voltage per step 4.A.(16).
NOTE:
Any cell that fails to rise to 1.55 VDC or peaks above 1.55 VDC and then decreases below 1.50 VDC must be replaced. Any cell that has a voltage rising above 1.75 VDC should also be replaced. For battery cell replacement, refer to Ni-Cad Battery - Removal/Installation . If five or more cells are found to be defective, it is recommended that the entire battery be replaced. If the battery discharge time is less than 100 minutes after three charge/discharge cycles, it should be removed from service.
(17) If the battery has passed all the requirements of steps 4.A.(1) through (16), proceed to step (18). (18) Recheck electrolyte level and adjust as required. Refer to Chapter 12, Ni-Cad Battery Servicing. (19) Torque all loose nuts and screws that attach the intercell connectors to the cell terminals. Refer to Table 501. (20) Perform a second electrical leak check. Refer to Electrical Leak Check in this section. (21) If the battery has passed all the preceding requirements, it is ready for installation or storage. Table 501. Battery Cell Torque Values
5.
Fasteners Thread Size
Socket Head Cap Screw
Torque Values
1/4”-28
3/16”
100 to 125 inch-pounds.
Battery Reconditioning Using Alternate Methods A.
Alternate Reconditioning Methods. (1) Clean battery. Refer to Ni-Cad Battery - Cleaning/Painting. (2) Preform electrical leak check. Refer to Electrical Leak Check in this section. (3) Place battery on a constant current charger at five-hour charge rate (8.0 amps) until the battery voltage reaches a minimum of 31.0 VDC (an average of 1.55 VDC per cell). Refer to Table 502 for other charge rates. (4) If the battery voltage does not rise to 31.0 VDC within 30 minutes, check voltage of each cell for short. Replace shorted cell(s). Refer to Ni-Cad Battery - Removal/Installation for cell replacement procedures. (5) Check the electrolyte level. Refer to Chapter 12, Ni-Cad Battery - Servicing.
CAUTION: Do not attempt to discharge a battery at an excessively high rate and then attempt to short-circuit each cell at the end of this discharge. Batteries which have been discharged at high rates are not fully discharged. Application of shorting devices to individual cells at the end of such a high discharge will produce severe arcing and intense heat. (6)
Record the time it takes to accomplish the following: (a) Discharge battery at a rate of one hour with a 40 amp load or two hours with a 20 amp load.
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(7)
Continue to discharge the battery. As each individual cell reaches 0.5 VDC or less, place a shorting strap across its terminals while still discharging battery. (Refer to Figure 501.) (c) When 14 of the cells are shorted out with shorting straps, place a 1.0 ohm two-watt resistor across each of the remaining cells. Allow the battery to remain shorted for at least three hours. After three hours, remove the resistance load bank, shorting straps and 1.0 ohm resistors. If the discharge time recorded in step 5.A.(6) was less than 100 minutes, charge the battery per step 5.A.(3), discharge the battery per step 5.A.(6) and charge battery again per step 5.A.(3).
Table 502. Constant Current Charge Rates Battery Type
1-Hour Charge Rate
Fast 4-Hour Charge Rate
14-Hour Charge Rate
Marathon TSP-410
40.0 Amperes for 1 hour
8.4 Amperes first 2 hours 8.4 Amperes second 2 hours
4.3 Ampers for 14 hours
(8)
Recharge the battery using one of the following methods:
CAUTION: When operating at higher battery temperatures, battery is subjected to danger of thermal runaway due to overcharge. This condition is characterized by continuously increasing current and rising battery temperature during constant current charging. CAUTION: The continual charging above gassing potential results in excess water loss due to decomposition and heat generation. At high ambient temperatures, heat loss of battery through radiation and conduction is lower than the heat-generating rate. This results in a net increase in battery temperature. This increase causes a higher charge current during constant current charge. This higher charge current further increases temperature which continues to drive off water through decomposition. (a)
Constant Current Charging - The length of time required to charge a battery by constant current charging depends on the capacity of battery and the state of charge current. In order to maintain a constant current, it should be noted that voltage of system will vary from 1.4 VDC per cell at beginning of charge to a maximum of 1.75 VDC per cell at end of charge. The constant current method employs a five hour rate. The five hour rate is arrived at by dividing five hours into 40 ampere-hours for an 8.0 ampere rate. This 8.0 ampere rate should be sustained for seven hours to provide a 140 percent charge to the battery. Constant current charging can be accomplished manually by monitoring charging the apparatus and adjusting voltage periodically to maintain a constant charged rate. NOTE:
It may not be necessary to charge battery for entire time. To determine whether battery has reached full charge state as indicated in Figure 502 after a sharp voltage rise, it is then necessary to permit battery to charge for an additional two hours. During this two hour period, it is possible to check cell voltages and determine whether all cells are rising evenly. Should some cells indicate a voltage lower than 1.55 VDC, it is advisable to leave battery on charge until these cells come up to a minimum of 1.55 VDC.
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Ni-Cad Battery Charging Charts Figure 502 (Sheet 1)
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CAUTION: Stepped constant current charging may cause a loss of water, resulting in higher concentration of electrolyte and possible battery cell damage due to overheating. (b)
(c)
Stepped Constant Current Charging - If it is necessary to obtain a rapid charge, battery may be permitted to charge at a current equal to its capacity (a 40 ampere-hour battery, for example, may be charged at a rate of 40 amperes). At this high rate, battery should be monitored and when total battery voltage averages 1.50 VDC per cell, charging rate shall be reduced to a five-hour rate (8.0 ampere) for remainder of charging cycle. At time of reduction, a conditioned battery will have completed approximately 70 percent of charge cycle if it was completely discharged to begin with. This method of charging must be monitored because the high rate will eventually use all water in battery if permitted to continue beyond gassing point. Float Charging - Fully-charged batteries may be floated across a line voltage of approximately 1.42 VDC per cell. The battery may be trickle-charged at 28.4 VDC. A battery floated in this manner will draw about 0.003 amps per ampere-hour of capacity. NOTE:
Voltage setting of 1.42 VDC will vary slightly with ambient temperature of operation.
CAUTION: When using constant voltage charging at high ambient temperatures, adjustments must be made in charging voltage to ensure that a thermal runaway condition does not exist. If battery temperature exceeds 120°F, no charging should be attempted as a thermal runaway will occur, destroying battery. Constant Voltage Charging - A voltage produced by generator permits current flow to battery. In a discharged battery, maximum surge of current will be approximately ten times rated capacity of battery. High-surge currents are due to low internal resistance of battery. Ensure charging source is protected against overload in event of a marginal power supply. Recommended voltage setting for battery charger voltage regulator is 30.0 VDC at 70°F to 80°F. Multiply recommended voltage per cell by number of cells to obtain correct voltage setting for charging from a constant voltage source. At lower temperatures, battery will accept charge if proper adjustments are made to regulating source. If voltage is not corrected, battery will not deliver its rated capacity. During final five minutes of charge (with current still flowing), measure voltage of each individual cell. Maximum voltage for each cell should be 1.55 VDC, and maximum voltage for each cell should be 1.75 VDC at room temperature (60°F to 90°F). If any cell fails to rise to at least 1.55 VDC continue current charge for an additional hour. During the final five minutes of charge, measure voltage of each cell again. If any cell fails to rise to 1.55 VDC or peaks above 1.55 VDC and then decreases below 1.50 VDC or exceeds 1.75 VDC, it must be removed from battery and replaced with another cell. Refer to Ni-Cad Battery - Removal/Installation. This will also require another battery discharge per step 5.A.(6) and charge per step 5.A.(7). After completion of the charge, check the electrolyte level again. Refer to Chapter 12, Ni-Cad Battery - Servicing. If battery discharge time recorded in step 5.A.(6) is less than 100 minutes, battery must be recharged per step 5.A.(7) and then discharged per step 5.A.(6). Record discharge time. Recharged per step 5.A.(7). If battery fails three charge/discharge cycles, it should be removed from service. Tighten loose nuts or screws that attach intercell connectors to cell terminals; refer to Table 501 for torque values. Perform electrical leak check. Refer to Electrical Leak Check. (d)
(9) (10) (11)
(12) (13) (14) (15) (16) B.
If battery has passed preceding requirements, it is ready for installation or for storage.
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6.
Electrical Leak Check A.
Perform an electrical leak check of battery. NOTE:
(1)
Before performing electrical leak check, ensure all intercell connectors are installed and that nuts on screws attaching intercell connector to cell terminals are torqued per Table 501.
Using a multimeter, set the range selector to 500 milliampere range or higher and measure the battery external leakage as follows:
CAUTION: Do not touch the multimeter negative lead to the negative terminal of the battery receptacle. (a)
(b)
(c) (d)
7.
Place the multimeter positive lead on the positive terminal of the battery receptacle and momentarily touch the multimeter negative lead to one of the screws that are used to mount the battery receptacle to the battery case. Record the current (ampere) reading shown on the multimeter scale. 1 Remove the multimeter positive lead from the positive terminal of the battery 2 receptacle. Place the multimeter negative lead on the negative terminal of the battery receptacle and momentarily touch the multimeter positive lead to one of the screws that is used to mount the battery receptacle to the battery case. Record the current (ampere) reading shown on the multimeter scale. If the current reading in either steps 6.A.(1)(a) or (b) exceeds 50 milliamperes, the tops of the cells shall be flushed with water and dried with clean absorbent toweling or with dry compressed air. Repeat steps 6.A.(1)(a) and (b). If the current is still greater than 50 milliamperes, one or more of the cells may be leaking. To determine which cell is leaking, measure each cell voltage as follows: 1 Using a voltmeter of 1000 ohms per volt or greater, place one of the meter leads on either the negative or positive terminal of the battery and the other meter lead on one of the screws used to mount the battery receptacle to the battery case. If the meter reads anything other than zero, there is a leak in one or more cells. If no leaky cells are found, the electrical leakage path may be due to electrolyte a along the outside of the cells and at the bottom of the battery case. Clean the battery. Refer to Ni-Cad Battery - Cleaning/Painting. If a leaky cell is found, replace and perform step 6.A.(1)(d)1 and 2 again. b With one meter lead on one of the screws used to mount the battery receptacle to the 2 battery case, move the other lead from one cell terminal to another in a sequential order, noting the voltage readings. Voltage readings will decrease and eventually indicate negative, disclosing location of the path and possibly a leaky cell. If a leaky cell is found, replace and perform step 6.A.(1)(d)1 and 2 again. a If no leaky cells are found, the electrical leakage path may be due to electrolyte b along the outside of the cells and at the bottom of the battery case. Clean the battery. Refer to Ni-Cad Battery - Cleaning/Painting.
Battery Receptacle Inspection A.
Items To Inspect. (1) Connect pins should inspected for corrosion, pitting or burn marks. If any of these defects are evident to the extent that total electrical contact could be prevented, the surface should be cleaned. (2) If cleaning process reduces pin diameter below 0.370 inch, the battery shall be replaced.
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8.
Battery Quick-Disconnect Inspection A.
Items To Inspect (1) Check for excessively loose handle and locking assembly. (2) Check for pitted or corroded mating surfaces. (3) Check for burn marks caused when battery is disconnected under load. (4) Test for resiliency of mating surfaces to an oversized pin (Elcon connector only). Insert larger diameter probe (0.385 inch diameter) of a GO/NO-GO gage into helix or sleeve to maximum depth. Ensure a snug fit with a removal force greater that one pound. (5) To assure contact is adequate for a worn battery pin, insert small diameter end (0.370 inch diameter) of GO/NO-GO gage. (a) Rebling connector, ensure each socket exerts sufficient pressure on the pin to hold the 0.370 inch diameter GO/NO-GO gage when the quick-disconnect is inverted to a position where the gage is pointed downward. (b) Elcon connector, ensure a snug fit with a nominal removal force of one pound. (6) If the connector fails to pass resiliency test or shows excessive wear of damage, replace connector.
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MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the battery in a serviceable condition.
Task 24-34-00-720 2.
Marathon Ni-Cad Battery Functional Check (Capacity Check) A.
General (1) This section gives the information needed to complete the inspection procedures for the ni-cad battery.
B.
Special Tools (1) None
C.
Access (1) Open the right cowling door to get access to the battery. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do the Marathon Ni-Cad Battery Functional Check (Capacity Check). (1) Remove the battery from the airplane. Refer to Ni-Cad Battery - Removal/Installation. (2) Visually examine the nickel-cadmium battery for its general condition. (3) Examine the connectors for an overheat indication, burns, or signs of arcing. (4) Examine the vent tubes for deterioration, rubs, and wear. (5) Examine the battery support structure and the adjacent areas for corrosion, cracks, and rubs. (6) Do a capacity check and an electrolyte level adjustment of the Marathon Ni-Cad battery. Refer to the MarathonNorco Aerospace Operating and Maintenance Manual. (7) Install the battery in the airplane. Refer to Ni-Cad Battery - Removal/Installation.
E.
Restore Access (1) Close the right cowling door. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL NI-CAD BATTERY - CLEANING/PAINTING 1.
General A.
2.
Tools, Equipment and Materials A.
3.
The battery should be removed from the airplane and cleaned in an authorized service area. If a white deposit has formed on top of the cells, it is potassium carbonate which is harmless and can be removed by brushing with a nonmetallic, acid-resistant brush, or with a clean cloth.
Refer to Electrical Power - General for a list of required tools, equipment and materials.
Cleaning/Painting Battery
WARNING: The electrolyte used in nickel-cadmium batteries is a caustic solution of potassium hydroxide. Serious burns will result if it comes in contact with any part of the body. Use rubber gloves, rubber apron and protective goggles when handling this solution. If electrolyte gets on the skin, wash the affected areas with large quantities of water, neutralize with three percent acetic acid, vinegar or lemon juice. If electrolyte gets into the eyes, flush with water and get immediate medical attention. CAUTION: Do not use solvents to clean battery. Damage to battery case liner and cover gasket may result. A.
Cleaning Battery. (1) Wipe off battery case and cover with cleaning cloth. (2) Remove cover.
CAUTION: Do not use wire brush as shorting will occur. Damage to cell cases, filler cap vent plugs and battery terminal links may also result. (3) (4) (5) (6)
Using filtered compressed air with a nonmetallic nozzle, blow around the tops of the cell to remove any dust and/or salt crystals that may have been deposited on the battery case. If intercell connectors and cell tops are excessively corroded, brush top of cell cases, filler cap vent plugs and battery terminal links with a nonmetallic brush to loosen all white deposits. Wipe top of cell cases, filler cap vent plugs and battery terminal links with a cleaning cloth to remove all foreign material. When the top of the battery case and cell tops are wet from minor spewage, proceed as follows: (a) Check tightness of vent plugs. Tip battery at about a 45 degree angle with its receptacle facing upward in a downward direction to prevent any water from entering the battery case. (b) Disperse excess water with shop air. (c) Using a multimeter, measure the current between each battery terminal and the battery case. If current flow is measurable, the battery should be cleaned. NOTE:
(7)
Take reading from battery terminals of the receptacle to the snaps that restrain the cover.
When excessive electrolyte spewage has occurred to the extent that it has run down between the cells, the following procedure shall be performed. (a) Perform the same cleaning procedure described in step (6) minor spewage. (b) The battery shall be completely discharged and disassembled.
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CAUTION: Do not remove cells from the case unless reinstallation can be accomplished immediately. (c) (d) (e) (f)
(g) (h)
(i) (j) (k)
Inspect each cell for cracks, holes, or other defective conditions. Defective cells shall be replaced with new or rebuilt cells. The cells shall be washed under running water. Do not allow the water to enter the cells interior. The cells shall be dried with clean, absorbent toweling or with compressed dry air. Remove accumulated dirt, carbonate deposits, and corrosion prevention silicone from connectors, screws, nuts and washers (after they are removed from the cells) by wiping with a dry cloth. Heavy deposits shall be removed by scrubbing with a stiff bristle, nonmetallic, acid-resistant brush. Ensure all parts are thoroughly dry before reassembling. Inspect all parts. Damaged or heavily corroded parts shall be replaced. Connecting straps that are burned, bent, or have defective nickel plating shall be repaired or replaced. Tarnished connecting straps shall be polished with a fine emery cloth, being careful not to remove plating. Inspect the battery power receptacle for burns, cracks, and bent or pitted terminals. Defective receptacles shall be replaced. (Defective receptacles can overheat, arc, depress battery voltage, and cause premature battery failure.) Bent battery cases and covers and loose or damaged battery cover gaskets shall be repaired or replaced. Broken or cracked intercell connectors shall be replaced. Reassemble the battery components in the battery case. Refer to Ni-Cad Battery Removal/Installation. NOTE:
(l) B.
If a cell is difficult to insert, apply a light coat of petroleum jelly or silicone grease to the sides of the cell case before reassembly.
Capacity check the battery.
Cleaning Battery Filler Cap Vent Plug.
WARNING: Electrolyte will cause serious burns if allowed to contact the skin. (1) (2) (3) (4) (5) (6) C.
Remove filler cap vent plug using nylon wrench. Inspect cell vent caps, O-rings and vent sleeve for obstructions, cracks, or damaged seals. Damaged parts shall be replaced. Wash cap under running water. Remove white deposits from filler cap vent plug using nonmetallic, acid-resistant brush. Dry filler cap vent plug with clean cloth or dry compressed air. Install filler cap vent plug.
Painting. (1) The battery case and cells are covered with an epoxy coating. This epoxy coating is designed to improve the insulation between the cells and the battery container. (2) If the epoxy coating is damaged, contact Marathon Battery Company, Customer Service Department.
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MODEL 208 MAINTENANCE MANUAL BATTERY OVERHEAT WARNING - DESCRIPTION AND OPERATION 1.
General A.
2.
Description A.
3.
The battery overheat warning system consists of two thermostatic switches. The system warns pilot when battery temperature reaches predetermined temperatures. When BATTERY HOT warning light illuminates, the battery switch, located on left sidewall switch panel, should be switched to OFF position to prevent battery from a thermal runaway and from being destroyed. When BATTERY OVERHEAT warning light illuminates, battery switch should be switched to "OFF" and airplane landed as soon as possible.
The battery temperature sensor is installed between cells of battery to measure temperature of cells from top to bottom.
Operation A.
The battery overheat warning system illuminates a warning light on annunciator panel when battery is overheated. Battery annunciation will change from an amber annunciator to a red annunciation dependent on battery temperature. Refer to Figure 1 for an illustration of annunciator and switch panel locations. Refer to Figure 2 for a wiring diagram of the battery overheat warning system.
WARNING: If either amber caution light or red warning light illuminates, the battery switch should be moved to the off position to prevent battery from a thermal runaway and from being destroyed. If red light illuminates, the flight should be terminated as soon as possible. (1) (2) B.
When battery temperature is 140°F to 160°F, the amber BATTERY HOT light will illuminate on the annunciator panel. When battery temperature is over 160°F, the red BATTERY OVERHEAT light will illuminate on the annunciator panel.
The annunciator test switch on the instrument panel may be used to check system for proper operation.
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Battery Overheat Switch Panel Figure 1 (Sheet 1)
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Battery Overheat Wiring Diagram Figure 2 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL BATTERY OVERHEAT WARNING - TROUBLESHOOTING 1.
General A.
A troubleshooting chart is included to aid the maintenance technician in system troubleshooting. Refer to Figure 101. This chart can be used in conjunction with Battery Overheat Warning - Description and Operation (Figure 2) to troubleshoot the system.
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Battery Overheat Warning System Troubleshooting Chart Figure 101 (Sheet 1)
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Battery Overheat Warning System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL BATTERY OVERHEAT WARNING - MAINTENANCE PRACTICES 1.
General A.
2.
3.
When removing and installing temperature sensor in battery, battery should be reconditioned. Temperature sensor may be removed after battery has completed deep cycle.
Temperature Sensor Removal/Installation A.
Remove Temperature Sensor (Refer to Figure 201). (1) Remove battery cover and battery from airplane. Refer to Ni-Cad Battery - Removal/Installation. (2) Recondition battery. Refer to Ni-Cad Battery - Adjustment/Test for reconditioning procedures. (3) Remove center cells as required. Refer to Ni-Cad Battery - Removal/Installation. (4) Remove battery temperature sensor electrical connector. (5) Remove battery temperature sensor.
B.
Install Temperature Sensor (Refer to Figure 201). (1) Install battery temperature sensor between battery cells. (2) Install sensor and electrical connector. (3) Install removed cells. Refer to Ni-Cad Battery - Removal/Installation. (4) Install battery in airplane. Refer to Ni-Cad Battery - Removal/Installation.
Adjustment/Test A.
Test (1) (2) (3) (4) (5) (6) (7)
Procedures. Turn battery switch to ON position. Ensure ANNUNCIATOR PANEL circuit breaker is engaged. Push annunciator light test switch. Amber BATTERY HOT and red BATTERY OVERHEAT lights should illuminate on annunciator panel. Release test switch. Lights shall extinguish. Disconnect battery temperature sensor electrical connector from battery. Amber BATTERY HOT and red BATTERY OVERHEAT lights should illuminate on annunciator panel. Reconnect battery temperature sensor electrical connector to battery. Turn battery switch to OFF position.
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Temperature Sensor Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL STANDBY ELECTRICAL SYSTEM - DESCRIPTION AND OPERATION 1.
2.
Description A.
An optional standby electrical system may be installed. The standby electrical system is designed to automatically supply power to the main buses if the system voltage drops below a preset level.
B.
The standby electrical system consists of the following components: (1) Alternator - A belt-driven, 95-amp alternator, operated at 75-amp capacity, is mounted at the rear of the engine and utilizes a rear engine accessory pad to drive the alternator. (2) Alternator Control Unit (ACU) - An alternator control unit is mounted forward of the left circuit breaker panel to control the system. Field excitation to the alternator control unit is supplied through diode logic from a circuit breaker in the standby alternator relay assembly or the KEEP ALIVE NO. 2 circuit breaker on the electrical power relay box. (3) Relay Assembly - A standby alternator relay assembly is mounted on the upper left forward side of the firewall (4) Switches - Two switches are installed on the left sidewall switch panel. The switches are two position toggle switches, labeled ON/OFF/STBY PWR, and a guarded two-position toggle type breaker/switch, labeled AVIONICS STBY PWR. The guard covering the avionics standby power switch must be lifted to select the ON position. (5) Circuit Protection - Circuit protection and isolation are provided by two circuit breakers labeled STBY PWR. These circuit breakers are located on the left circuit breaker panel. (6) Monitoring Lights - System operation is monitored by two amber lights labeled STBY ELECT PWR ON and STBY ELECT PWR INOP, located in the annunciator panel. Total amperage supplied from the standby electrical system can be monitored on the airplane volt-ammeter with the selector switch in the ALT position.
Operation A.
To operate the standby electrical system, follow the starting procedures in the Pilot’s Operating Handbook and FAA approved Airplane Flight Manual. When engine is started, place the ON/OFF/STBY PWR switch to the ON position. The system is now engaged to automatically supply the electrical load if the bus voltage drops below a preset level. Anytime the STBY ELECT PWR ON light in the annunciator panel illuminates, the standby electrical system is supplying power to the main buses. If the drop in voltage is temporary, such as just after engine start, the STBY ELECT PWR ON light will go out, indicating that system voltage is normal and the main generating system is carrying all the load. If the STBY ELECT PWR ON light illuminates continuously, it would indicate a malfunction in the main generating system has occurred, and steps should be initiated to isolate the problem. If the STBY ELECT PWR INOP light is illuminated, the standby alternator is inoperative.
B.
The ON/OFF/STBY PWR switch should be placed in the OFF position when the airplane is not in use, to remove the ACU drain from the battery. The annunciator panel will remain ON with the master switch OFF until the STBY PWR switch is turned OFF.
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MODEL 208 MAINTENANCE MANUAL STANDBY ELECTRICAL SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been included to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Standby Electrical System Troubleshooting Chart Figure 101 (Sheet 1)
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Standby Electrical System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL STANDBY ELECTRICAL SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Tools and Equipment A.
3.
Maintenance of the standby electrical system includes the removal and installation of system components, alternator belt tension adjustment, and the adjustment of the output voltage.
For a list of tools and equipment, refer to Electrical Power - General.
Alternator Removal/Installation A.
Remove Alternator (Refer to Figure 201). (1) Make sure the airplane power is OFF. (2) Disconnect the battery connector. (3) Remove the electrical connector from the alternator. (4) Record the position of the resistor, washers, and nuts for reinstallation and disconnect the electrical wires that remain from the alternator. (5) Carefully cut the tie strap from around protective sheathing. (6) Remove the nut from the pivot bolt. (7) Cut the safety wire from the tensioning bolt. (8) Remove the bolt and washer from the alternator. (9) Remove the drive belt. (10) Hold the alternator and remove the pivot bolt from the pivot bracket assembly. Record the position of the bushing for reinstallation. (11) Remove the alternator from airplane. NOTE:
B.
For removal of the standby alternator tension bracket assembly, refer to Chapter 71, Engine Equipment Attach Brackets - Maintenance Practices.
Install the Alternator (Refer to Figure 201). (1) Make sure the bushing location is correct. (2) Install the alternator to the pivot bracket with the bolt, washer, and new locking nut. Do not apply final torque to the nut. (3) Align the top of alternator to the tension bracket. NOTE:
It can be necessary to loosen the tension bracket and the pivot bracket to get the correct alignment with the top of the alternator. If necessary the tension bracket is loosened first. If the top of alternator still cannot be aligned to the tension bracket, the pivot bracket is loosened. Refer to Chapter 71, Engine Equipment Attach Brackets Maintenance Practices.
(a)
(4) (5) (6)
If necessary to get the correct alignment of the tension bracket to the alternator, remove the safety wire and loosen the nuts that attach the tension bracket. 1 If you still cannot align the top of alternator to the tension bracket, remove the safety wire and loosen the nut and bolts that attach the pivot bracket. (b) Attach the alternator to the tension bracket with the belt tensioning bolt and washer, but do not tighten. (c) If necessary, step torque the pivot bracket fasteners. Use 65 - 85 inch-pounds. 1 Safety the pivot bracket bolts with wire. (d) If necessary, step torque the tension bracket fasteners. Use 65 - 85 inch-pounds. 1 Safety the tension bracket nuts with wire. Install the drive belt. Adjust the belt tension by applying 170 inch-pounds, +10 inch-pounds or -10 inch-pounds to the belt tensioning boss on the alternator. While the belt is correctly tensioned, tighten the upper bolt in position. (a) Safety the bolt with wire.
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Alternator and Drive Pulley Installation Figure 201 (Sheet 1)
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Alternator and Drive Pulley Installation Figure 201 (Sheet 2)
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Alternator and Drive Pulley Installation Figure 201 (Sheet 3)
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Alternator and Drive Pulley Installation Figure 201 (Sheet 4)
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Alternator and Drive Pulley Installation Figure 201 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL (7)
Make sure the pivot bolt nut is correctly tightened. (a) Torque the nut on the pivot bolt between 450 inch-pounds to 500 inch-pounds. (8) Install all electrical wires and resistors to the alternator. (a) Make sure all the washers are installed correctly and tighten the hardware. (9) Install rubber boots over all the exposed electric wires to protect the connections from arcing. (10) Connect the electrical connector to the alternator. (11) Attach the protective sheathing to the alternator using a tie strap. (12) Apply electrical power. C.
Remove the Drive Pulley Assembly (Refer to Figure 201).
CAUTION: Use care to make sure the splined coupling and gaskets are not damaged during removal. (1) D.
Remove the nuts and washers that attach the mounting flange of the drive pulley assembly to the studs on the accessory pad at the engine oil scavenge pump.
Install the Drive Pulley Assembly (Refer to Figure 201). (1) Put the new gasket between the alternator tension bracket and the engine oil scavenge pump.
CAUTION: Use care to make sure the splined coupling and gaskets are not damaged during the installation. (2) (3) (4) (5) (6)
4.
5.
Put the new gasket between the tension bracket and the alternator drive pulley assembly. Put the alternator drive pulley assembly mounting flange on the accessory pad studs. Carefully put the splined coupling in the scavenge pump shaft. It can be necessary to turn the engine slightly to engage the splined coupling to the scavenge pump shaft. Make sure the gasket between the tension bracket and the scavenge pump has not been moved. Attach the pulley assembly to the accessory pad with the washers and the new locking nuts, but do not apply final torque to the nuts. Final torque will be applied after all the mounting holes are aligned to the alternator. Refer to Alternator - Removal/Installation.
Alternator Control Unit Removal/Installation A.
Remove Alternator Control Unit (ACU) (Refer to Figure 202). (1) Remove substrate panel, located forward of the left circuit breaker panel, to gain access to the Alternator Control Unit. (2) Disconnect electrical connector from ACU. (3) Remove screws securing ACU to nutplates in bracket. (4) Remove ACU from airplane.
B.
Install Alternator Control Unit (ACU) (Refer to Figure 202). (1) Align ACU with nutplates in bracket. (2) Install screws. (3) Reconnect electrical connector to ACU. (4) Adjust the output voltage. Refer to Output Voltage Adjustment, in this section. (5) Install substrate panel.
Relay Assembly Removal/Installation A.
Remove Relay Assembly (Refer to Figure 202). (1) Open left side of engine cowling. Refer to Chapter 71, Engine Cowling and Nosecap Maintenance Practices. (2) Disconnect electrical connector from relay assembly. (3) Remove screws attaching relay assembly to firewall. (4) Remove relay assembly from airplane.
B.
Install Relay Assembly (Refer to Figure 202). (1) Fillet seal opening between cover and base with Ablative RTV If relay cover is removed. (2) Align relay assembly to nutplates in firewall and secure using screws.
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Standby Electrical System Installation Figure 202 (Sheet 1)
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Standby Electrical System Installation Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (3) (4) 6.
Connect electrical connector to relay assembly. Close and secure left engine cowling. Refer to Chapter 71, Engine Cowling and Nosecap Maintenance Practices.
Output Voltage Adjustment A.
Do a check of the Standby Alternator Voltage for the ACU. (1) Open the left side of the engine cowling. Refer to Chapter 71, Engine Cowling and Nosecap Maintenance Practices. (2) Remove the substrate panel, found forward of the left circuit breaker panel, to get access to the ACU. (3) Connect the voltmeter to the alternator (A+) terminal and to the airplane ground. (4) Start the airplane. Refer to the Pilot’s Operating Handbook and Approved Airplane Flight Manual. (5) First, make sure the starter/generator is off-line, and the engine is at 52 percent Ng. Then make sure the output voltage is 27.5 VDC, +0.1 or -0.1 VDC, with an airplane load of approximately 14 amps.
(6) (7)
NOTE:
Voltages are calculated at a typical 70°F ambient temperature. Alternator voltage will decrease by approximately 10 percent at the alternator's maximum operating temperature. The maximum operating temperature is influenced by demand and altitude, and may be achieved regardless of the ambient temperature and altitude.
NOTE:
The 27.5 VDC value for the ACU is based on a generator control unit voltage setting of 28.5 VDC. If the generator control unit regulated voltage is changed, you must change the ACU setting accordingly to keep a value of 1.0 VDC less than the generator control unit setting.
Use the procedures in the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual in order to increase the engine speed to takeoff power, and then load to the maximum rated amps (approximately 75 amps). Gradually decrease the load to 20.0 amps and make sure the output voltage is 27.5 VDC, +0.25 or -0.4 VDC, throughout the load range.
CAUTION: Use of a ferrous metal tool will cause damage to the ACU. If adjustment is necessary, use the anti-static adjustment tool to adjust the ACU as follows: (a) If the airplane load is approximately 14 amps adjust the ACU to 27.5 VDC, +0.1 or -0.1 VDC. (b) If the airplane load is 20 to 75 amps, adjust the ACU to 27.5 VDC, +0.25 or -0.4 VDC. (9) Stop the engine. (10) Remove the voltmeter from the alternator. (11) Close the cowling. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices. (12) Install the substrate panel.
(8)
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MODEL 208 MAINTENANCE MANUAL STANDBY ELECTRICAL SYSTEM - INSPECTION/CHECK 1.
General A.
2.
This section has the inspections and checks necessary to keep the standby electrical system in a serviceable condition.
Drive Pulley Assembly/Bearing Housing Inspection NOTE:
Dripping grease at the pulley drive shaft may show that the oil scavenge pump seal leaks into the standby alternator drive assembly. If this occurs, the standby alternator drive assembly must be replaced with a serviceable unit.
NOTE:
The maximum allowable leakage of engine oil from the engine scavenge pump/standby alternator drive access pad is 2cc per hour. The engine scavenge pump must be inspected for oil leakage that drips past the drive shaft seal. Repairs must be done in accordance with the P&WC Engine Maintenance Manual.
A.
Do an inspection of the Drive Pulley Assembly/Bearing Housing. (Refer to Figure 601). (1) Remove the drive pulley assembly. Refer to Alternator and Drive Pulley Removal/Installation. (2) Fully clean the housing with Methyl n-Propyl Ketone or equivalent. (3) Visually do an inspection of the mounting holes for cracks. (4) Do an inspection of the mounting face for flatness with a straight edge or equivalent. (a) Do a check that any gaps between straight edge and surface of mounting face do not exceed 0.010 inch. (5) Visually do an inspection of the seal land for damage and/or cracks. (6) Visually do an inspection of the bearing lands for evidence of galling and/or cracks. NOTE:
Scratches related to the installation and removal of bearing is acceptable.
(7)
Measure the bearing lands for out-of-round condition. (a) Check to make sure that the inner diameter from out-of-round measurement is 1.5746 inch, +0.0002 or -0.0002 inch. (b) Check to make sure that the outer diameter form out-of-round measurement is 1.497 inch, +0.010 or -0.010 inch. (8) If the housing has drain lines, visually do an inspection for cracked, plugged, collapsed or outof-round lines. (a) If damage exists and is repairable, repair as necessary. (9) On part number 2601270-4, visually do an inspection for plugged vent holes. (a) If the vent holes are clogged, clean as necessary. Use a piece of safety wire or equivalent. (10) Install the drive pulley assembly. Refer to Alternator and Drive Pulley Removal/Installation. Task 24-36-00-220 3.
Standby Alternator Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the standby alternator.
B.
Special Tools (1) None
C.
Access (1) Open the left side of the engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices.
D.
Do the Standby Alternator Detailed Inspection. (1) Examine the alternator for condition, security, correct safety of the mount bolts, and correct installation.
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Drive Pulley Assembly/Bearing Housing Inspection Figure 601 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2)
(3) (4) (5)
(6) (7)
Examine the alternator drive for condition, leaks, and security. (a) Move the top and the bottom of the drive pulley in and out by hand to examine for free play. 1 If there are signs of free play, remove the drive and do a detailed inspection of the drive splines and the coupling. Examine the mount for condition, cracks, corrosion, and security. Examine the ground strap for condition and security. Examine the alternator drive belt for condition and correct tension. Refer to the Standby Electrical System - Maintenance Practices, Alternator Removal/Installation. (a) If it is necessary to adjust the tension, refer to Alternator Removal/Installation, Standby Electrical Systems - Maintenance Practices. Examine the alternator electrical boots, and components for condition and security. Examine the wiring for condition and security of the connectors at the alternator terminals. NOTE:
(8)
(9)
Make sure that the in-line resistor is attached with a plastic screw, washers, and a nut.
Examine the drive drain (if installed) for condition and security. (a) Examine the drain hose for correct routing, kinks, and security. (b) Examine the drain can for condition and security. (c) Drain the contents of the can. Examine the alternator filter capacitor, mounting bracket, and wire lead for security and condition.
E.
Restore Access (1) Close the left side of the engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL MONITORING - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101. This chart covers the volt- ammeter and selector switch.
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Volt-Ammeter and Selector Switch Troubleshooting Chart Figure 101 (Sheet 1)
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Volt-Ammeter and Selector Switch Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL MONITORING - MAINTENANCE PRACTICES 1.
2.
General A.
A volt-ammeter gage and a selector switch, mounted on the left side of the instrument panel, are provided to monitor the electrical system operation. The ammeter selector switch is labeled GEN, ALT, BATT and VOLT. By positioning the selector switch to GEN, ALT or BATT position, the respective generator, alternator or battery amperage can be monitored. Positioning the switch to VOLT allows electrical system bus voltage to be monitored.
B.
Maintenance practices consist of removal/installation of the volt-ammeter gage and selector switch.
Volt-Ammeter Gage Removal/Installation A.
Remove Volt-Ammeter Gage (Refer to Figure 201). (1) Ensure airplane electrical power is off. (2) Remove nuts, lockwashers and wires from volt-ammeter gage. Tag wires to ensure correct polarity when reinstalling. (3) On front side of instrument panel, remove screws securing volt-ammeter gage to instrument panel.
B.
Install Volt- Ammeter Gage (Refer to Figure 201). (1) Position volt-ammeter gage to instrument panel from back side. (2) Secure volt-ammeter gage to instrument panel using screws. (3) Attach wires to volt-ammeter gage using nuts and lockwashers. NOTE: (4)
3.
Ensure polarity is correct before attaching wires to volt-ammeter gage.
Restore airplane electrical power.
Selector Switch Removal/Installation A.
Remove Selector Switch (Refer to Figure 201). (1) Ensure airplane electrical power is off. (2) Loosen setscrew on knob and remove knob from instrument panel. (3) Remove nut on front side of instrument panel, disconnect electrical connector and remove selector switch from airplane.
B.
Install Selector Switch (Refer to Figure 201). (1) Install selector switch through hole in instrument panel from back side. (2) Install nut to secure selector switch to instrument panel. (3) Connect electrical connector. (4) Install knob using setscrew. Ensure selector switch rotates to all four positions. (5) Restore airplane electrical power.
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Electrical Monitoring Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL BATTERY CONTACTOR - MAINTENANCE PRACTICES 1.
2.
Description A.
The battery connector is mounted in the electrical power box. The contactor’s function is to connect battery power to the power panel bus bar. The contactor is controlled by the battery switch.
B.
Maintenance to the battery contactor is limited to removal/installation of the contactor.
Contactor Removal/Installation A.
Remove Contactor (Refer to Figure 201). (1) Open left and right cowling. Refer to Chapter 71, Engine Cowling and Nosecap - Maintenance Practices. (2) Disconnect battery. (3) Remove screws securing cover on electrical power box and remove cover. (4) Remove nuts, lockwashers, and washers securing wires and bus bar to contactor. (5) Remove bolts securing contactor to electrical power box and remove contactor.
B.
Install Contactor (Refer to Figure 201). (1) Position contactor with terminal post through bus bar and install contactor mounting bolts. (2) Install all wires, washers, Iockwashers, and nuts removed. (3) Install cover on electrical power box. (4) Connect battery and close cowling. Refer to Chapter 71, Engine Cowling and Nosecap Maintenance Practices.
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Battery Contactor Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL 12-VOLT DIRECT CURRENT POWER OUTLET SYSTEM - DESCRIPTION AND OPERATION 1.
2.
General A.
This section gives information about the components of the 12-Volt Direct Current (VDC) power outlet system. These components include a 10-amp circuit breaker, 3-position switch, an Astron N2412-12 DC-DC converter, and optional second converter and a 12VDC power outlet with an optional two more power outlets in the cabin.
B.
The actual voltage output of the system is 13.8VDC. The DC-DC converter changes the airplane 28VDC to 13.8VDC. The system voltage will be referred to as 12VDC in this manual.
Description
CAUTION: Not all devices with a cigarette lighter style plug are approved for use aboard an airplane. The airplane operator is responsible to know which Personal Electronic Devices (PED) will not interfere with the safe operation of the airplane in which the PED is operated. Refer to AC 91.21-A, Use of Portable Electronic Devices Aboard Aircraft for guidance material about FAA approved devices.
3.
A.
The function of the 12VDC Power Outlet System is to provide power for Portable Electronic Devices (PED) that are compatible with the automotive-style 12VDC power outlets at an equivalent voltage.
B.
There are two subsystems that make up the 12VDC system. These systems are divided into the power outlet for the crew, and the optional second converter and power outlets for the passengers. (1) The crew power outlet is found in the pedestal between the pilot and copilot. (2) The two passenger-use power outlets are found on the two sidewalls, approximately halfway to the rear of the airplane. The power outlet on the 208 will be found at FS 206.15, and the power outlet on the 208B will be found at FS 226.12.
C.
The DC-DC converters are found in the floor compartment under the access panel (232AC). Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. A single DC-DC converter can supply 138 Watts continuously, or 165 Watts at maximum output. There is one converter that supplies power to the crew power outlet in the pedestal and an optional second converter that supplies power to the passenger-use power outlets found approximately halfway to the rear of the airplane. The two passenger-use outlets are found along the sidewalls of the cabin, and are connected to the DC-DC converter in parallel. The passenger-use power outlets must share the 138 Watts continuous output.
Operation A.
The 12VDC Power Outlet System components include a 10-amp circuit breaker, 3-position switch, the Astron N2412-12 DC-DC converter(s) and the 12VDC power outlet(s). (1) The 10-amp circuit breaker is found in the pilot's circuit breaker panel and is placarded as AUX 12VDC PWR. The 10-amp circuit breaker directly supplies power to the 3-position switch. The 10-amp circuit breaker lets the total output power of the 12VDC power outlet system get to only a maximum of 250 Watts, which will create a limit on the DC-DC converters of approximately 88% efficiency. NOTE:
(2) (3)
The circuit breaker is sized to trip just before the maximum continuous load of 276 Watts. The maximum wattage is calculated as 138 Watts x 2, for both converters operating at the same time.
The DC-DC converters supply power to the power outlets that are installed in the airplane. The 28VDC airplane voltage is converted by the DC-DC converters to 13.8VDC at the power outlets. The 3-position switch (SZ10) controls power output to the individual power outlets installed in the airplane. Put the 3-position switch in the OFF position in order to disengage power to the 12VDC system. Put the 3-position switch in the CREW position to supply 28-Volt power to the
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MODEL 208 MAINTENANCE MANUAL DC-DC converter which supplies 13.8VDC to the crew power outlet only. The crew power outlet is found in the pedestal which is accessible to the pilot and copilot. Put the 3-position switch in the ALL position to supply power to both the standard DC-DC converter and the optional DC-DC converter, which supplies power to the power outlets in the cabin for passenger-use. NOTE:
If the optional passenger-use power outlets are not installed, the crew power outlet in the pedestal will still supply power for the pilot and copilot when the 3-position switch is in the ALL position.
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MODEL 208 MAINTENANCE MANUAL 12-VOLT DIRECT CURRENT POWER OUTLET SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives information about the removal and installation of the power outlets and power converter assembly for the 12-Volt Direct Current (VDC) power outlet system. This installation is standard for Model 208 Airplanes 20800396 and On, and 208B1171 and On.
Power Outlet Removal/Installation NOTE:
3.
The removal and installation procedures for the power outlet in the pedestal, are different than the power outlet(s) in the cabin. Procedures are typical for the pilot and copilot side power outlets.
A.
Remove the Power Outlet (Refer to Figure 201). (1) Disengage the AUX 12VDC PWR circuit breaker on the pilot's circuit breaker panel. (2) For the power outlets found in the cabin, do the steps that follow: (a) Remove the screws that attach the power outlet mounting plate to the airplane. (b) Carefully pull the power outlet away from the airplane to get access to the electrical connector. (3) For the power outlet found in the pedestal, remove the access panel (226D) to get access to the power outlet wires. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (4) Disconnect the electrical connector. (5) Turn the backshell counterclockwise to remove it from the power outlet. (6) Remove the power outlet from the mounting plate.
B.
Install the Power Outlet (Refer to Figure 201). (1) For the power outlets found in the cabin, do the steps that follow: (a) Put the power outlet in place, then turn the backshell clockwise to attach the power outlet to the mounting plate. (b) Connect the electrical connector. (c) Put the power outlet and mounting plate in place, then attach the mounting plate with the screws. (d) Make sure that the retaining strap on the cap is in position between the outlet and the mounting plate. (2) For the power outlet found in the pedestal, do the steps that follow: (a) Put the power outlet in place, then turn the backshell clockwise to attach the power outlet to the pedestal. (b) Connect the electrical connector. (c) Install the access panel (226D). Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Engage the AUX 12VDC PWR circuit breaker on the pilot's circuit breaker panel. (4) Do a check of the 12VDC power outlet system. Refer to 12VDC Power Outlet System Check.
Power Converter Assembly Removal/Installation NOTE:
The removal and installation procedures are typical for both Power Converter Assemblies.
A.
Remove the Power Converter Assembly (Refer to Figure 201). (1) Disengage the AUX 12VDC PWR circuit breaker on the pilot's circuit breaker panel. (2) Remove access panel (232AC) to get access to the DC-DC converter. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Disconnect the electrical connector (J1) from the wire bundle. (4) Remove the screw(s) that attach(es) the ground wire(s) to the airplane. (5) Remove the screws that attach the DC-DC converter to the airplane. (6) Remove the DC-DC converter from the airplane.
B.
Install the Power Converter Assembly (Refer to Figure 201). (1) Set the DC-DC converter in position in the floor compartment aft of the pedestal.
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12VDC Power Outlet Components Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) (6) (7) 4.
Attach the DC-DC converter in the airplane with the screws. Install the screw(s) that attach(es) the ground wire(s) to the airplane. Connect the electrical connector to the wire bundle. Install the access panel (232AC) in the airplane. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Engage the AUX 12VDC PWR circuit breaker on the pilot's circuit breaker panel. Do a check of the 12VDC power outlet system. Refer to 12VDC Power Outlet System Check.
12VDC Power Outlet System Check A.
Do a check of the 12VDC power outlet system. (1) Do a preliminary visual check of the connectors and components to make sure they are attached correctly. (2) Connect the Ground Power Unit (GPU) to the airplane. NOTE: (3) (4) (5)
Put the EXTERNAL POWER switch in the BUS position. Engage the AUX 12VDC PWR circuit breaker on the pilot's circuit breaker panel. Put the VEC008 LCD Voltage Meter (power outlet tester) into each of the installed power outlets. NOTE:
(6) (7) (8)
Do the power outlet system check with the GPU set to the airplane voltage of +28.5VDC +0.5 or -0.5VDC, with a current capacity of at least 400 amps.
If a power outlet tester is not available, use a Digital Multi-Meter (DMM). The red (+) lead must be put in the center terminal of the power outlet.
Operate the POWER OUTLET switch to each of its three positions. Make sure the voltage polarity is positive for all the test steps followed in the power outlet system check. Make sure the voltage output is correct. Refer to Table 201.
Table 201. Switch Positions Switch Position
Pedestal Outlet (standard)
Passenger Outlet - Left Side (optional)
Passenger Outlet - Right Side (optional)
OFF
0VDC
0VDC
0VDC
CREW
12.5 - 14.5VDC
0VDC
0VDC
ALL
12.5 - 14.5VDC
12.5 - 14.5VDC
12.5 - 14.5VDC
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MODEL 208 MAINTENANCE MANUAL EXTERNAL POWER - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system understanding. Refer to Figure 101.
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External Power Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL EXTERNAL POWER - MAINTENANCE PRACTICES 1.
Description and Operation A.
The external power system supplies electrical power for starting the engine and for ground operation of the airplane’s electrical system. The external power system consists of the following components: (1) Receptacle and Shield Assembly - This receptacle is mounted on an engine mount bracket and provides the means to connect external electrical power to the airplane. (2) Ground Power Monitor - This component is mounted in the upper right portion of the electrical power box and has over-voltage and polarity sensing. The external power monitor will open the circuit if any of the following conditions occur: (a) The external power source connected to the receptacle is above 31.5 VDC, +0.5 or -0.5 VDC. (b) The external power source connected to the receptacle is below 24.5 VDC, +0.5 or -0.5 VDC. (c) External power is connected with the external power switch actuated. (d) Polarity is reversed. (3) External Power Contactor - This relay is mounted in the electrical power box and opens or closes to supply external power to the airplane upon command from the control switch. (4) Switch - The external power switch is a three-position switch (OFF/STARTER/BUS) located above the cockpit circuit breaker panel. In the STARTER position, electrical power is supplied to external start contactor, but external start contactor will remain open until starter switch is placed in MOTOR or START position. With external start contactor closed, electrical power is then supplied to starter/generator for starting engine. The external power contactor will not close with external power switch in STARTER position. When the switch is placed in BUS position, the external power contactor will close, supplying electrical power for operation of the airplane’s electrical systems. NOTE:
B. 2.
External power must be plugged in and voltage set before external power switch is actuated.
Maintenance practices for the external power system consists of removal/installation of components, and ensuring all connections are clean and tight.
Components Removal/Installation A.
Remove External Power Contactor (Refer to Figure 201). (1) Open left and right engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. (2) Ensure that battery switch is in the OFF position. (3) Disconnect battery. (4) Remove screws securing cover and remove cover from electrical power box. (5) Remove nuts, lock washers, washers, wires, and diode assembly from contactor. Tag wires for identification. (6) Remove bolts securing contactor to electrical power box. (7) Lower contactor down and out of bus bars and remove.
B.
Install External Power Contactor (Refer to Figure 201). (1) Position contactor with terminal posts through bus bars and install mounting bolts. (2) Install wires and diode assembly on contactor terminals using, washers, lock washers, and nuts. (3) Remove tags from wires installed for identification. (4) Install cover on electrical power box. (5) Connect battery. (6) Close left and right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
C.
Remove Ground Power Monitor (Refer to Figure 201). (1) Open left and right engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. (2) Ensure battery switch is in the OFF position.
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Electrical Components Installation Figure 201 (Sheet 1)
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Electrical Components Installation Figure 201 (Sheet 2)
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Electrical Components Installation Figure 201 (Sheet 3)
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Electrical Components Installation Figure 201 (Sheet 4)
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Electrical Components Installation Figure 201 (Sheet 5)
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Electrical Components Installation Figure 201 (Sheet 6)
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3.
Disconnect battery. Disconnect ground power monitor electrical leads where possible. Remaining leads may be cut and spliced on installation. Remove screws and washers securingground power monitor to electrical power box and remove ground power monitor.
D.
Install Ground Power Monitor (Refer to Figure 201). (1) Position ground power monitor in electrical power box and secure with washers and screws. (2) Connect electrical leads or splice as necessary. (3) Install electrical power box cover. (4) Connect battery. (5) Close left and right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
E.
Remove External Power Receptacle (Refer to Figure 201). (1) Open left and right cowling doors. Refer to Chapter 71, Engine Cowling And Nose Cap Maintenance Practices. (2) Slide protective rubber boot from small terminal and remove hardware securing single wire to terminal. Remove wire from terminal. (3) Remove cover, nut and star washer securing large wires to terminal. Remove large wires from terminal. (4) Remove hardware securing ground strap to terminal. (5) Remove nuts, screws and washers securing shield assembly and external power receptacle to engine mount bracket. Remove from airplane.
F.
Install External Power Receptacle (Refer to Figure 201). (1) Install shield assembly and external power receptacle to engine mount bracket using nuts, washers and screws. (2) Reattach ground strap to terminal. (3) Install large wires to terminal and secure using nut and star washer. Install cover over terminal end to prevent possible arcing. (4) Reattach single wire to terminal terminal using lock washer and nut. Slide protective rubber boot over terminal end to prevent possible arcing. (5) Close left and right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
Electrical Power Box Fireproof Coating Inspection and Repair A.
Fireproof Coating Inspection and Repair. (1) Inspect exterior of electrical power box for paint which has peeled, blistered, or seperated. (2) If loose areas are found, repair as follows: (a) Sand out damaged area of coating. (b) Clean sanded area with methyl propyl ketone. (c) Use Scotchbrite pad and roughen surface to improve coating adhesion. (d) Apply No.173 Intumescent paint; 0.030 inch minimum thickness and cover with Gray paint. (e) Finish repair with two coats of Seal Gray modified urethane paint. Refer to Chapter 20, Interior and Exterior Finish - Cleaning and Painting
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL DISTRIBUTION SYSTEM - MAINTENANCE PRACTICES 1.
2.
Description and Operation A.
The electrical distribution system consists of a main power panel bus bar, two systems bus bars (No. 1 and No. 2) and two avionics bus bars (No. 1 and No. 2). The system bus bars are each divided into three segments with a tie bus connecting the segments.
B.
The system bus bars are connected to the power panel bus by three wires in parallel. Each wire has a current limiter at the power panel bus, and a pull-type circuit breaker at the system bus. The current limiters protect wires between the power panel bus and the system bus. Should one voltage limiter open, the system bus can still function on the remaining two wires. Should two limiters open, the airplane systems on that bus bar must be operated at a limited capacity.
C.
The avionics bus bars are each connected to a power panel bus by one wire with a current limiter at power panel bus, and a switch/circuit breaker at the avionics bus bar. A switch/circuit breaker between the two avionics bus bars enables the bus bars to operate on one limiter should the other limiter open.
D.
The airplane systems are controlled by individual switches for each system. Individual circuit breakers on the bus bars protect the individual systems.
E.
Maintenance of the distribution system consists of removal/installation of the circuit breakers, bus bars and switches.
Circuit Breaker Removal/Installation A.
Remove Circuit Breaker (Refer to Figure 201). (1) Disconnect battery terminals. Display a maintenance warning tag stating:
WARNING: Do not connect battery terminals; maintenance in progress. (2) (3) (4) (5) (6) (7) B.
3.
Remove screws securing circuit breaker panel to left side of airplane. Identify, tag and disconnect electrical lead from circuit breaker. Remove screw and washer securing bus bar to circuit breaker. Remove hex nut securing circuit breaker to panel and retain for reinstallation. Remove and retain lock washer and knurl nut from circuit breaker. Remove circuit breaker from panel.
Install Circuit Breaker (Refer to Figure 201). (1) Assemble knurl nut and lock washer on circuit breaker. (2) Push circuit breaker thru panel and secure using hex nut. (3) Secure bus bar to circuit breaker using screw and washer. (4) Identify and connect electrical lead to circuit breaker using screw and washer. (5) Position circuit breaker panel to side of airplane and secure using screws. (6) Reconnect battery and remove maintenance warning tag from battery connector.
Bus Bar Removal/Installation A.
Remove Bus Bar (Refer to Figure 201). (1) Disconnect battery terminals. Display a maintenance warning tag stating:
WARNING: Do not connect battery terminals; maintenance in progress. (2) (3) B.
Remove screws securing circuit breaker panel to left side of airplane. Remove screws securing bus bar to circuit breakers and remove bus bar.
Install Bus Bar (Refer to Figure 201). (1) Position bus bar on circuit breakers. (2) Install screws securing bus bar to circuit breakers. (3) Position circuit breaker panel to side of airplane and secure using screws. (4) Reconnect battery and remove maintenance warning tag from battery connector.
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Electrical System Distribution Installation Figure 201 (Sheet 1)
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4.
Switch Removal/Installation A.
Remove Switch (Refer to Figure 201). (1) Disconnect battery terminals. Display a maintenance warning tag stating:
WARNING: Do not connect battery terminals; maintenance in progress. (2) (3) (4) (5) B.
Remove screws securing circuit breaker panel to left side of airplane. Remove screws securing switch panel to top of circuit breaker panel. Identify, tag and disconnect electrical lead from switch. Remove hex nut and guard, if installed, from switch. Remove switch from airplane.
Install Switch (Refer to Figure 201). (1) Position switch and guard, if required, through switch panel and secure using hex nut. (2) Connect electrical leads to switch. Ensure wires are connected properly. (3) Position switch panel on circuit breaker panel an install screws. (4) Position circuit breaker panel to side of airplane and secure using screws. (5) Reconnect battery and remove maintenance warning tag from battery connector.
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL DISTRIBUTION SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the electrical distribution system components in a serviceable condition.
Task 24-50-00-220 2.
Power Distribution Boxes Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the power distribution boxes.
B.
Special Tools (1) None
C.
Access (1) Remove the upper left and right cowling doors to get access to the battery and the power distribution boxes. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a Detailed Inspection of the Power Distribution Box (Electrical Power). Refer to Figure 601. (1) Set the BATTERY switch to the OFF position. (2) Remove the external electrical power from the airplane. (3) Disconnect the battery terminals. (4) Attach a warning tag to the battery and the external power receptacle that have the statement that follows:
WARNING: Do Not Connect or Apply Electrical Power to the Airplane Maintenance in Progress. (5)
Remove the screws that attach the cover to the electrical power distribution box. (a) Remove the cover from the box. (6) Examine all electrical components for condition and security. (7) Examine all electrical wires and cables for correct routing, support, chafing, and security of the connectors. (8) Examine the box and the cover for condition and security. (9) Examine the sealant between the box and firewall for condition. (a) If the seal is broken, loose, or deteriorated, replace it with a new fillet seal using Type II, Class B-4 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices. (10) Examine all current limiters for signs of an open link. (a) If the condition is unknown, remove the current limiter(s) and do a resistance test with an ohmmeter. 1 The resistance must be less than 1 ohm. (11) Examine the sealant on the firewall electrical connectors for condition. (a) If the seal is broken, loose, or deteriorated, replace it with new silicone sealant (part number Q3-6077). Refer to Chapter 20, Electrical Bonding - Maintenance Practices. (12) Put the cover in its position on the electrical power distribution box. (a) Install the screws. E.
Do a Detailed Inspection of the Power Distribution Box (Standby Electrical Power). Refer to Figure 602. (1) Remove the screws that attach the cover to the standby electrical power distribution box. (a) Remove the cover from the box. (2) Examine the box and the cover for condition and security.
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Electrical Power Distribution Box Figure 601 (Sheet 1)
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Standby Electrical Power Distribution Box Figure 602 (Sheet 1)
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Examine all electrical components for condition, contamination, and security. (a) If there are signs of contamination, remove the contamination and apply Type II, Class B-4 sealant across the top of the cover and the relay base assembly. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (4) Examine all electrical wires and cables for correct routing, support, chafing, and security of the connectors. (5) Examine the current limiters for signs of an open link. (a) If the condition is unknown, remove the current limiter(s) and do a resistance test with an ohmmeter. 1 The resistance must be less than 1 ohm. (6) Examine the sealant between the base of the box and the firewall for condition. (a) If the seal is broken, loose, or deteriorated, replace it with a new fillet seal using Type II, Class B-4 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices. (7) Put the cover in its position on the standby electrical power distribution box. (a) Install the screws. (8) Apply a new fillet seal between the cover and the base using Type II, Class B-4 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (9) Remove the warning tag from the battery and the external power receptacle. (10) Connect the battery. F.
Restore Access (1) Install the upper left and right cowling doors. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL LOAD ANALYSIS - DESCRIPTION AND OPERATION 1.
General A.
2.
This section describes the electrical load in chart form for the Models 208/208B.
Description A.
For airplanes without G1000, the electrical load analysis depicts a standard airplane plus available optional equipment. Only the maximum loads of the individual systems are recorded in the load charts. Refer to (Figure 1). (1) Two flight regimes were analyzed; a typical night flight under instrument conditions using deicing equipment, and a night flight under instrument conditions at 90°F using the air conditioner. A 100-minute duration was selected for both flights. The flight profile phases were landing, start, taxi, take off/climb and cruise/descent. (2) The charging system was assumed to be regulated at 28.5 volts DC. It is further assumed that the full 200 ampere capacity of the generator is available (Ng RPM above 66%) immediately after starting. Under this condition, the aircraft electrical loads, plus battery charging current, never exceeds generator capacity and the battery is always in a constant voltage charging mode.
B.
For airplanes with G1000, the electrical load analysis depicts the highest load configuration for a standard airplane plus available optional equipment. Nominal load values are shown for individual systems. Refer to (Figure 2). (1) Maximum possible and average total loads are shown for each bus, both for normal generator operation, and for emergency operation, using the standby alternator only. (2) The charging system was assumed to be regulated at 28.5 volts DC. It is further assumed that the full 200 ampere capacity of the generator is available (Ng RPM above 64.2%) immediately after starting. Under this condition, the aircraft electrical loads, plus battery charging current, never exceeds generator capacity and the battery is always in a constant voltage charging mode.
Table 1. Airplanes without G1000 Avionics Note
Definition
A
275 amps average for 24.68 seconds during 0° start, 292 amps for 13.79 seconds during 75° start.
B
External continuous operation is mutually exclusive with internal continuous operation.
C
Load during start only.
D
Used for periodic maintenance test only.
E
Require no bus power.
F
For the purpose of calculation, all dimmable lights are de-rated to 50% during night flight.
G
Off during icing conditions.
H
On during icing conditions.
I
De-rated for 10% duty cycle.
J
Secondary actuator operation is mutually exclusive with primary actuator operation.
K
327 amps average for 14.78 seconds during 0° start.
L
De-rated to 50% duty cycle on timer (not manual).
M
Powered by other units.
N
In a dual installation, only 1 unit can transmit at a time, assumed to be the #1 unit.
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Table 2. Airplanes with G1000 Avionics Note
Definition
A
In auto mode duty cycle is 90 sec on, 90 sec off.
B
On during icing conditions.
C
In normal mode, only draws current in fuel low pressure condition.
D
Used for periodic maintenance testing only.
E
The primary and standby flap motors can not be operated at the same time. The larger load is considered.
F
For the purpose of calculations, half of the interior cabin lights are assumed on.
G
Average battery charge.
H
In normal anti-ice mode, 2 pumps run 20 sec on, 100 sec off (duty cycle is 17%). When the max anti-ice mode is selected, BOTH pumps run continuously for 2 minutes each time max anti-ice is selected.
I
In high anti-ice mode, one pump runs continuously. When the max anti-ice mode is selected, the second pump also runs continuously for 2 minutes each time max anti-ice is selected.
J
Backup anti-ice will be on only if primary anti-ice is not available.
K
For the purpose of calculations, 1 minute of transmission for every 10 minutes during taxi and flight, and 3 minutes of transmission for every 10 minutes during takeoff and landing.
L
Between COM 1, COM 2, only one can transmit at same time. The largest load is included.
M
Horn draws load only during impending stall condition.
N
Only one transponder selected at a time, the second is assumed to be in standby mode.
O
For purpose of calculation assume one extention or one retraction per flight phase.
P
On only during engine start.
Q
Warning on only in abnormal switch position.
R
For the purpose of calculations, auxiliary 12 Vdc outputs are assumed operating at 50% capability.
S
Dimmers on MAX.
T
Standby mode. 0.20 amps with alarm on.
U
Intermittent use. For purpose of calculation use 75% duty cycle per flight phase.
V
Average demand on the listed bus in italics.
W
The Model 208 has 11 reading lights. The model 208B has 14, and the Super Cargomaster has no reading lights.
X
Electric prop anti-ice is mutually exclusive with the TKS option. The larger load is considered.
Y
Left and right vent blower load not considered, as that total load is lower than air condtioning total load.
Z
ELT switch light not used during normal operation.
AA
For ground operations, heaters are limited to 2 minutes on, 2 minutes off.
AB
Air conditioning is assumed not to be on in icing conditions.
NOTE LS stands for loads that were shed in standby alternator only configuration or battery only operation after the loadshed.
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Model 208 and Model 208B Electrical Load Analysis without G1000 Figure 1 (Sheet 2)
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Model 208 and Model 208B Electrical Load Analysis without G1000 Figure 1 (Sheet 3)
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Model 208 and Model 208B Electrical Load Analysis without G1000 Figure 1 (Sheet 4)
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Model 208 and Model 208B Electrical Load Analysis with G1000 Figure 2 (Sheet 1)
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Model 208 and Model 208B Electrical Load Analysis with G1000 Figure 2 (Sheet 2)
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Model 208 and Model 208B Electrical Load Analysis with G1000 Figure 2 (Sheet 3)
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Model 208 and Model 208B Electrical Load Analysis with G1000 Figure 2 (Sheet 4)
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Model 208 and Model 208B Electrical Load Analysis with G1000 Figure 2 (Sheet 5)
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Model 208 and Model 208B Electrical Load Analysis with G1000 Figure 2 (Sheet 6)
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25 CHAPTER
EQUIPMENT/ FURNISHINGS
CESSNA AIRCRAFT COMPANY
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LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
25-00-00
Pages 1-2
Jun 1/2011
25-10-00
Pages 201-208
Mar 1/1999
25-10-00
Pages 601-602
Jun 1/2011
25-10-01
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Jan 2/2006
25-10-05
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Jan 2/2006
25-21-00
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Dec 1/2006
25-21-00
Page 601
Jun 1/2011
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Aug 1/1995
25-25-00
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Aug 1/1995
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Pages 201-203
Sep 2/1997
25-51-00
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Mar 1/2000
25-51-00
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Jun 1/2011
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Apr 1/2010
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Jun 1/2011
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Mar 1/2000
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Jun 1/2011
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Aug 1/1995
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Apr 1/2010
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Mar 1/2008
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Jun 1/2011
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Aug 1/1995
25-Title 25-List of Effective Pages 25-Record of Temporary Revisions 25-Table of Contents 25-List of Tasks
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS EQUIPMENT/FURNISHINGS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-00-00 25-00-00 25-00-00 25-00-00
Page 1 Page 1 Page 1 Page 2
FLIGHT COMPARTMENT - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pilot and Optional Copilot (Right) Crew Seat Removal/Installation . . . . . . . . . . . . . . . Seat Belt and Shoulder Harness Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . Armrest Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-10-00 Page 201 25-10-00 Page 201 25-10-00 Page 201 25-10-00 Page 208 25-10-00 Page 208
FLIGHT COMPARTMENT - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Smoke Goggle General Visual Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Seats Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inertia Reel Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-10-00 Page 601 25-10-00 Page 601 25-10-00 Page 601 25-10-00 Page 601 25-10-00 Page 602
SUNVISORS - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sunvisor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-10-01 Page 201 25-10-01 Page 201 25-10-01 Page 201
BEVERAGE CUP HOLDER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ring Style Cup Holder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Molded Cup Holder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-10-05 Page 201 25-10-05 Page 201 25-10-05 Page 201 25-10-05 Page 201
PASSENGER SEATS - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Utility Passenger Seats and Restraint System Removal/Installation . . . . . . . . . . . . . . Commuter Passenger Seats and Restraint System Removal//Installation. . . . . . . . . Restraint System Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Three-Place Bench Seat and Restraint System Removal/Installation . . . . . . . . . . . . .
25-21-00 Page 201 25-21-00 Page 201 25-21-00 Page 201 25-21-00 Page 201 25-21-00 Page 216 25-21-00 Page 217 25-21-00 Page 217
PASSENGER SEATS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Seats Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-21-00 Page 601 25-21-00 Page 601 25-21-00 Page 601
CABIN UPHOLSTERY - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upholstery Removal/Installation (Model 208 and 208B Passenger) . . . . . . . . . . . . . . Upholstery Removal/Installation (Cargomaster and 208B) . . . . . . . . . . . . . . . . . . . . . .
25-23-00 Page 201 25-23-00 Page 201 25-23-00 Page 201 25-23-00 Page 210
REAR COMPARTMENT WALL - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Canvas Wall Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Padded Rear Wall Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-25-00 Page 201 25-25-00 Page 201 25-25-00 Page 201 25-25-00 Page 201
FLOOR COVERING/CONTROL COLUMN COVER - MAINTENANCE PRACTICES . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Column Cover Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Floor Covering Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-28-00 Page 201 25-28-00 Page 201 25-28-00 Page 201 25-28-00 Page 201
CARGO COMPARTMENT COMPONENTS - MAINTENANCE PRACTICES . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Tie-Down Straps Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Barrier and Access Nets Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Partition Nets Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Door Restraint Net Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-51-00 Page 201 25-51-00 Page 201 25-51-00 Page 201 25-51-00 Page 201 25-51-00 Page 208 25-51-00 Page 212
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MODEL 208 MAINTENANCE MANUAL CARGO COMPARTMENT COMPONENTS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Nets Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-51-00 Page 601 25-51-00 Page 601 25-51-00 Page 601
CARGO POD - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Pod Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Pod Components Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Modification for Flight Without the Cargo Pod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reinstalling the Cargo Pod After Flight Without the Cargo Pod . . . . . . . . . . . . . . . . . .
25-52-00 Page 201 25-52-00 Page 201 25-52-00 Page 201 25-52-00 Page 208 25-52-00 Page 210 25-52-00 Page 213
CARGO POD - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Pod Zonal Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Pod Drains Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-52-00 Page 601 25-52-00 Page 601 25-52-00 Page 601 25-52-00 Page 602
CARGO POD - APPROVED REPAIRS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Repair Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-52-00 Page 801 25-52-00 Page 801 25-52-00 Page 801 25-52-00 Page 801
CARGO POD HEAT SHIELDS - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shield Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-52-01 Page 201 25-52-01 Page 201 25-52-01 Page 201
LOADING ZONE WEIGHT LIMITS - CLEANING/PAINTING. . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-53-00 Page 701 25-53-00 Page 701
EMERGENCY EQUIPMENT - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Artex C406-2 Emergency Locator Transmitter (ELT) Functional Check . . . . . . . . . . . Artex ME406 Emergency Locator Transmitter (ELT) Functional Check . . . . . . . . . . . ARTEX C406-N Emergency Locator Transmitter (ELT) Functional Check. . . . . . . . . Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 Emergency Locator Transmitter (ELT) Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Locator Transmitter Battery Discard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-60-00 Page 601 25-60-00 Page 601 25-60-00 Page 601 25-60-00 Page 604 25-60-00 Page 610
EMERGENCY LOCATOR TRANSMITTER - DESCRIPTION AND OPERATION . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-61-00 Page 25-61-00 Page 25-61-00 Page 25-61-00 Page
EMERGENCY LOCATOR TRANSMITTER - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-61-00 Page 101 25-61-00 Page 101
EMERGENCY LOCATOR TRANSMITTER - MAINTENANCE PRACTICES . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transmitter Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Panel Mounted Rocker Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Antenna Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Pack Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Locator Transmitter Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-61-00 Page 201 25-61-00 Page 201 25-61-00 Page 201 25-61-00 Page 201 25-61-00 Page 207 25-61-00 Page 207 25-61-00 Page 209
ARTEX C406-2 EMERGENCY LOCATOR TRANSMITTER SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-62-00 Page 25-62-00 Page 25-62-00 Page 25-62-00 Page
25-60-00 Page 615 25-60-00 Page 617
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ARTEX C406-2 EMERGENCY LOCATOR TRANSMITTER SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Locator Transmitter (ELT) Removal/Installation . . . . . . . . . . . . . . . . . . . . . ELT Antenna Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELT/NAV Interface Unit Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mode S Program Module Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ARTEX ME406 EMERGENCY LOCATOR TRANSMITTER SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-62-00 Page 201 25-62-00 Page 201 25-62-00 Page 201 25-62-00 Page 201 25-62-00 Page 205 25-62-00 Page 205 25-63-00 25-63-00 25-63-00 25-63-00
Page 1 Page 1 Page 1 Page 1
ARTEX ME406 EMERGENCY LOCATOR TRANSMITTER SYSTEM TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ME406 Emergency Locator Transmitter (ELT) Self Test Preparation . . . . . . . . . . . . .
25-63-00 Page 101 25-63-00 Page 101 25-63-00 Page 101 25-63-00 Page 101
ARTEX ME406 EMERGENCY LOCATOR TRANSMITTER SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Locator Transmitter (ELT) Removal/Installation . . . . . . . . . . . . . . . . . . . . . ELT Buzzer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remote Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELT Antenna Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-63-00 Page 201 25-63-00 Page 201 25-63-00 Page 201 25-63-00 Page 201 25-63-00 Page 204 25-63-00 Page 204
ARTEX C406-N EMERGENCY LOCATOR TRANSMITTER SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-64-00 25-64-00 25-64-00 25-64-00
Page 1 Page 1 Page 1 Page 1
ARTEX C406-N EMERGENCY LOCATOR TRANSMITTER SYSTEM TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C406-N Emergency Locator Transmitter (ELT) Self Test Preparation . . . . . . . . . . . . .
25-64-00 Page 101 25-64-00 Page 101 25-64-00 Page 101 25-64-00 Page 101
ARTEX C406-N EMERGENCY LOCATOR TRANSMITTER SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Locator Transmitter (ELT) Removal and/or Installation . . . . . . . . . . . . . . . ELT Buzzer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remote Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELT Antenna Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-64-00 Page 201 25-64-00 Page 201 25-64-00 Page 201 25-64-00 Page 204 25-64-00 Page 204 25-64-00 Page 204
SOUNDPROOFING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Soundproofing Damping Foam Panel Removal/Installation. . . . . . . . . . . . . . . . . . . . . . Fiberglass Batting Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Therma-Sil Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
25-80-00 Page 201 25-80-00 Page 201 25-80-00 Page 201 25-80-00 Page 201 25-80-00 Page 209
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 25-10-00-210
Smoke Goggle General Visual Inspection
25-10-00 Page 601
25-10-00-220
Crew Seats Detailed Inspection
25-10-00 Page 601
25-10-00-710
Inertia Reel Operational Check
25-10-00 Page 602
25-21-00-220
Passenger Seats Detailed Inspection
25-21-00 Page 601
25-51-00-220
Cargo Nets Detailed Inspection
25-51-00 Page 601
25-52-00-210
Cargo Pod Zonal Inspection
25-52-00 Page 601
25-52-00-710
Cargo Pod Drains Operational Check
25-52-00 Page 602
25-60-00-720
Artex C406-2 Emergency Locator Transmitter (ELT) Functional Check
25-60-00 Page 601
25-60-00-721
Artex ME406 Emergency Locator Transmitter (ELT) Functional Check
25-60-00 Page 604
25-60-00-722
ARTEX C406-N Emergency Locator Transmitter (ELT) Functional Check
25-60-00 Page 610
25-60-00-723
Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 Emergency Locator Transmitter (ELT) Functional Check
25-60-00 Page 615
25-60-00-960
Emergency Locator Transmitter Battery Discard
25-60-00 Page 617
25 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL EQUIPMENT/FURNISHINGS - GENERAL 1.
Scope A.
2.
This chapter provides information on the equipment and furnishings of the flight compartment, passenger compartment, cargo and accessory compartment, and emergency locator transmitters (ELTs). This chapter does not include structures, or equipment and systems specifically assigned to other chapters.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Aeroflex Communications Test Set
IFR 4000
Aeroflex, Wichita Division 10200 West York Street Wichita, KS 67215-8935
To complete the functional test of the Artex ELT 110-406 Emergency Locator Transmitter.
Attenuator (30 dB)
500-3000
Artex Aircraft Supplies, Incorporated 14405 Keil Road NE Aurora, OR 97002
To do the emergency locator transmitter functional check.
Pro-Seal
890/890A
PRC-DeSoto International 5454 SanFernando Rd. Glendale, CA 91203
To seal cargo pod at various points.
M. C. Gill Corporation 4056 Easy Street El Monte, CA 91771
To tape between seams on hard shelled interior.
3M Masking and Packaging Div. 3M Center Building 220-8W-01 St. Paul, MN 55144-1000
To tape between seams on hard shelled interior.
M.C. Gill Corporation
To repair minor damage to hard shell interior panels.
Gilltape
3M Tape
367FR
Gillpatch Industrial Cement
EC1300L
3M
To secure fiberglass batting to fuselage.
Fiberglass Cloth
Hexcell 181, 1581 or 7781 Style
Commercially available
To make patch for repairing cargo pod.
Filler
EC3524 Syntatic Foam
Commercially available
To replace core of cargo pod.
Plastic Sheet
Wrightlon 7400 IPP Nylon
Commercially available
To make vacuum bag.
Seal Tape
GS213 Peal Ply Release Film Tacky Tape
Commercially available
To seal vacuum bag.
Commercially available
To remove air from vacuum bag and patch.
Vacuum Pump (Capable of pulling 20.0 inches of vacuum)
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NAME
NUMBER
Plastic Squeegee
MANUFACTURER
USE
Commercially available
To work syntatic foam and adhesive.
Adhesive
EA956 Hysol
Commercially available
To bond patch.
Sealant Tape
SF-4291
H. B. Fuller Company Assembly Products Division 5220 Main St., N.E. Minneapolis, MN 55421
To seal area between cargo pod and fuselage.
SARSAT Beacon Test Set
453-0131
Artex P.O. Box 1270 Canby, OR 97013
To complete the functional test of the Artex ELT C406-2 Emergency Locator Transmitter.
50 Ohm Dummy Load
5 Watt or Greater
3.
To test ELT.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief definition of the sections incorporated in this chapter is as follows: (1) The section on flight compartment includes the description, operation and maintenance practices for equipment and furnishings used by the crew. (2) The section on passenger compartment includes the description, operation and maintenance practices for the equipment and furnishings used in this compartment. (3) The section on cargo compartment includes the description, operation and maintenance practices for equipment and furnishings used in the cargo compartment and cargo versions of the airplane. (4) The section on emergency locator transmitters provides information and maintenance practices related to the various systems used on the airplane.
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MODEL 208 MAINTENANCE MANUAL FLIGHT COMPARTMENT - MAINTENANCE PRACTICES 1.
General A.
2.
Flight compartment equipment and furnishings consist of the following items: (1) Pilot’s Seat - Pilot’s seat is standard equipment. It may be adjusted forward or aft, up or down and the seat angle changed. (2) Optional Copilot’s (right) Seat - Optional copilot’s (right) seat contains the same operating features as the pilot’s seat. (3) Seat Belts and Shoulder Harness - On the 208 and 208B Passenger, the pilot and optional copilot (right) crew seats are equipped with a dual strap shoulder harness, lap belt and inertia reel which is fabricated as one assembly. Shoulder harness and lap belts are bolted to the lower seat frame, and the inertia reel is mounted at the bottom of the seat back frame. On the 208B, the pilot and optional copilot (right) crew seats are equipped with a five-point restraint system. This system consists of seat belts, a crotch strap and an inertia reel with double strap shoulder harness. (4) Removable Right Side Controls - Optional copilot (right) controls consist of the right side control wheel and right side rudder pedals. These controls are easily removed to allow for greater storage capacity. Refer to Chapter 27, Flight Controls - General , for topics related to control removal and installation. (5) Sunvisor - Sunvisors are provided as standard equipment for the pilot and optional copilot (right) crew seat positions. The sunvisors are fully adjustable and can be positioned as desired for maximum utilization. Removal and installation of the sunvisors is covered in Sunvisors Maintenance Practices. (6) Beverage Cup Holder - Two beverage cup holders are mounted underneath the forward edge of the instrument panel. Removal and installation of the cup holders is covered in Beverage Cup Holder - Maintenance Practices.
Pilot and Optional Copilot (Right) Crew Seat Removal/Installation A.
Remove Seats (Refer to Figure 201). (1) Remove forward and aft seat rail stop bolts by removing nuts and two spacers from each position. (2) Pull seat release assembly handle forward to disengage seat locking pin. (3) Exert upward pressure on the seat and slide seat fore and aft until rollers and roller housings are clear of seat rail. (4) Lift seat up and away from seat rail. (5) Remove seat from airplane.
B.
Install Seats (Refer to Figure 201). (1) Position seat on seat rail and slide fore and aft until rollers and roller housings engage seat rails.
WARNING: It is extremely important that seat stop bolts are securely installed and seat locking pins are securely engaged. Acceleration and deceleration could permit seat to become disengaged from rails, creating a hazardous situation, especially during takeoff and landing. (2) (3)
Install seat stop bolt, two spacers and nut in forward and aft holes of each seat rail. Ensure locking pin mechanism functions correctly and that seat locking pin engages in seat rail. NOTE:
Airplanes 20800021 and On and 20800001 thru 20800021 incorporating SK208-2 have an adjustable link installed on one side to enable full engagement of both locking pins.
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Crew Seat Installation Figure 201 (Sheet 1)
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Crew Seat Installation Figure 201 (Sheet 2)
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Crew Seat Installation Figure 201 (Sheet 3)
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Crew Seat Installation Figure 201 (Sheet 4)
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Crew Seat Installation Figure 201 (Sheet 5)
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Crew Seat Installation Figure 201 (Sheet 6)
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3.
Seat Belt and Shoulder Harness Removal/Installation A.
Remove Lap-Belt Style Restraint System (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7)
4.
Shoulder harness, lap belt and inertia reel are one assembly and must be removed as a unit.
Separate the upper section of seat back assembly from Velcro fasteners sufficiently to remove two screws attaching headrest to seat frame. Remove headrest from seat frame. Bend bottom flanges of seat back panel down and aft sufficiently to clear seat frame. Pull seat back panel up sufficiently to gain access to inertia reel. Remove bolts and nuts attaching inertia reel to seat back frame. Remove nuts, washers, spacers and bolts attaching stirrups to seat frame. Remove belts, stirrups and inertia reel from seat.
B.
Install Lap-Belt Style Restraint System (Refer to Figure 201). (1) Install belts, stirrups and inertia reel on seats. (2) Attach stirrups to seat frame using spacers, bolts, washers and nuts. (3) Attach inertia reel to seat frame using bolts and nuts. (4) Push seat back panel down until bottom flanges can be located below seat frame. (5) Attach headrest to seat frame and install screws. (6) Locate and attach upper section of seat back to Velcro fasteners on seat frame.
C.
Remove Five-Point Restraint System (Refer to Figure 201). (1) Separate seat back cushion from Velcro fasteners and remove. (2) Carefully bend inboard flanges of seat back panel out and around seat back frame so panel may be pulled back to gain access to inertia reel. (3) Remove nuts, washers and bolts securing inertia reel to seat frame. (4) Pull inertia reel out while feeding shoulder harness through slot beneath headrest. (5) Remove nuts, washers, spacers and bolts securing stirrups to seat frame. (6) Remove, nut, washer and bolt securing crotch strap to seat frame, and remove crotch strap.
D.
Install Five-Point Restraint System (Refer to Figure 201). (1) With seat back panel pulled out away from seat, feed shoulder harness up and through slot beneath headrest. (2) Position inertia reel to seat and secure using bolts, washers and nuts. (3) Slide seat belt stirrups between seat back and seat cushion. Attach to seat frame using bolts, spacers, washers and nuts. Ensure buckle side of seat belt is installed on right side of seat. (4) Slide stirrup end of crotch strap down through slot at forward end of seat cushion and attach to seat frame using bolt, washer and nut. (5) Carefully bend inboard flanges of seat back panel out and around seat back frame to install panel. (6) Install seat back cushion on Velcro fasteners.
Armrest Removal/Installation NOTE:
Left and right armrest removal/installation procedures are typical.
A.
Remove Armrest (Refer to Figure 201). (1) Remove bolt, washers and armrest from crew seat.
B.
Install Armrest (Refer to Figure 201). (1) Position washers and armrest to crew seat and secure using bolt.
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MODEL 208 MAINTENANCE MANUAL FLIGHT COMPARTMENT - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the flight compartment equipment and furnishings in a serviceable condition.
Task 25-10-00-210 2.
Smoke Goggle General Visual Inspection A.
General (1) This task gives the information needed to complete the general visual inspection of the smoke goggles (if installed).
B.
Tools and Equipment (1) None
C.
Access (1) None
D.
Complete a General Visual Inspection of the Smoke Goggle. (1) Get access to the smoke goggle . (2) Examine the smoke goggle for scratches or cracks. (3) Make sure the smoke goggle is clean. (4) Install the smoke goggle in its correct location.
E.
Restore Access (1) None End of task Task 25-10-00-220 3.
Crew Seats Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the crew seats.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do a Detailed Inspection of the Crew Seats. (1) Remove the crew seats from airplane. Refer to Flight Compartment - Maintenance Practices. (2) Examine the crew seat assembly for rips, tears, cleanliness, security of attached components, and other signs or damage. (3) Examine the crew seat headrest and armrest assemblies (if equipped) for security and correct movement. (4) Examine the seat belts, shoulder straps, restraint straps, and retainers for security of installation, rips, cuts, tears, frayed edges, cleanliness, and general condition. (a) Replace frayed and/or cut belts. (5) Examine all belts for legibility of the certification label. (6) Examine the belt length adjustments for correct operation. (7) Examine all belt attach points for security and damage. (8) Examine the buckle assemblies for signs of wear, cleanliness, security, and general condition. (9) Examine the belt latching mechanism for correct operation. (10) Examine seat belt inertia reels for security and general condition.
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MODEL 208 MAINTENANCE MANUAL (11) Examine the inertia reel for freedom and smoothness of operation by slowly pulling out on the shoulder harness belt. (a) Make sure that the shoulder harness belt retracts onto the inertia reel when it is released. (12) Pull the shoulder harness belt out quickly from the inertia reel and make sure that the inertia reel locks. (a) Make sure that the locking mechanism releases when the shoulder harness is relaxed. (13) Turn the seat over and examine the rollers, pins, adjustment levers, guides, and other components for security of installation, cleanliness, corrosion, wear, interference, or other damage. (14) Examine the seat adjustment levers and jackscrews for freedom and smoothness of operation. (15) Examine the seat release mechanism for freedom and smoothness of operation. (16) Examine the seat rollers for flat spots. (17) Examine the springs that keep the lock pins in position in the seat track holes for security and general condition. (18) Examine seat frame and structure for evidence of damage or corrosion. (19) Examine seat back stop bolts for loose or missing nuts. (20) Examine the crew seat rails and seat rail holes for cleanliness, cracks, corrosion, loose or failed fasteners, or other damage. (21) Install crew seats. Refer to Flight Compartment - Maintenance Practices. E.
Restore Access (1) None End of task Task 25-10-00-710 4.
Inertia Reel Operational Check A.
General (1) This task gives the information needed to complete the operational check of the inertia reel.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Inertia Reel Operational Check. (1) Examine the inertia reel locking mechanism. (a) Pull the shoulder harness belt slowly to make sure that the belt extends smoothly. (b) Make sure that the shoulder harness belt correctly retracts on the inertia reel when the belt is released. (c) Pull the shoulder harness belt out rapidly from inertia reel and make sure that the inertia reel locks. (d) Release tension on the belt and make sure that the locking mechanism releases the shoulder harness belt.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL SUNVISORS - MAINTENANCE PRACTICES 1.
2.
General A.
Maintenance includes the removal and installation of the sunvisors.
B.
There are two types of sun visors for the 208 Caravan. One type of sun visor has components that include an adjustable friction nut, and the other has a mounting plate that is adjustable.
Sunvisor Removal/Installation A.
Remove the Sunvisor (Refer to Figure 201). (1) Remove the screws that attach the sunvisor to the airplane.
B.
Install the Sunvisor (Refer to Figure 201). (1) Install the sunvisor with the screws. NOTE:
You can loosen or tighten the screws to adjust the sunvisors. Also, if installed on the sunvisor, the friction nut can be used to adjust the sunvisor.
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Sunvisor Installation Figure 201 (Sheet 1)
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Sunvisor Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL BEVERAGE CUP HOLDER - MAINTENANCE PRACTICES 1.
2.
3.
General A.
This section includes the removal and installation procedures of the cup holders.
B.
There are two styles of cup holders that can be installed in the airplane. The original cup holder is a ring-style cup holder that can be found attached to the instrument panel. The new style of molded cup holder can hold a wider range of cups, and can be found attached to the control column cover.
Ring Style Cup Holder Removal/Installation A.
Remove the Cup Holder (Refer to Figure 201). (1) Remove the screw, washers, spacer and nut that attach the cup holder to the instrument panel. (2) Remove the cup holder from the airplane.
B.
Install the Cup Holder (Refer to Figure 201). (1) Put the cup holder in position to the instrument panel, then attach with the screw, washers, spacer and nut.
Molded Cup Holder Removal/Installation A.
Remove the Molded Cup Holder (Refer to Figure 201). (1) Remove the screws and washers that attach the molded cup holder to the control column cover. (2) Remove the molded cup holder from the airplane.
B.
Install the Molded Cup Holder (Refer to Figure 201). (1) Put the cup holder in position to the control column cover, then attach with the screws and washers.
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Beverage Cup Holder Figure 201 (Sheet 1)
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Molded Cup Holder Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PASSENGER SEATS - MAINTENANCE PRACTICES 1.
2.
General A.
Maintenance practice describes various passenger seat conÞgurations and provides removal and installation procedures for various seating types.
B.
For various passenger seating conÞgurations, refer to Figure 201.
Description A.
Utility Passenger Seats (Model 208 and 208B Passenger) (Refer to Figure 202). (1) Optional utility seats are constructed of canvas on tubular frames, and are positioned behind each other (four seats on each side of the airplane). Seats may be folded into a compact space for stowage in the aft baggage compartment when not in use. When desired, the seats can be unfolded and installed in the passenger area. The seats are Þxed position and nonadjustable. Seat belts and shoulder harness are attached to each seat.
B.
Commuter Passenger Seats (Model 208 and 208B Passenger) (Refer to Figure 203). (1) Optional commuter seats are constructed of tubular frames with foam padding and fabric covering. Seats are Þxed position with nonadjustable backs and may be either single passenger or two-place passenger types. The seats are easily removed to facilitate cargo hauling. All seats are equipped with seat belts and shoulder harnesses attached to seat frames. (a) Seats with manual adjust shoulder harness restraints are installed in Airplanes 20800001 thru 20800403, and in Airplanes 208B0001 thru 208B1231. (b) Seats with inertial reel automatic adjust shoulder harness 3-point restraints are installed in Airplanes 20800404 and On, and in Airplanes 208B1232 and On. (2) Model 208B airplanes with 10-Place and 13-Place Commuter Passenger Seating may install left seats on right side of airplane and right seats on left side of airplane to provide additional space between seat and sidewall.
C.
Passenger Seats (14-Place Model 208 and 208B Passenger) (Refer to Figure 204).
WARNING: The 14-place seating provides increased passenger capacity to support certain international commuter operations. Utilization is limited to those nations where approval is received. The United States does not authorize more than nine passenger seats (excluding crew) under Provisions of 14 CFR 23.1. (1)
3.
Optional 14-place seating arrangement utilizes commuter seating, as well as a three-place, Þxedposition bench seat, located on the raised ßoorboard area immediately aft of passenger and cargo doors. The seat is easily removable to facilitate cargo hauling. Seat belts and shoulder harnesses, furnished with seats, attach to the airplane structure.
Utility Passenger Seats and Restraint System Removal/Installation A.
Remove Seats (Refer to Figure 202). (1) Raise lock on each seat leg attach Þtting (four total). (2) Move seat forward or aft slightly until attach Þttings are aligned with openings in seat track. (3) Lift seat out.
B.
Install Seats (Refer to Figure 202). (1) Position seat on seat track and insert attach Þttings in seat track openings. (2) Move seat forward or aft until locks are aligned with opening in seat track.
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Passenger Seat ConÞgurations Figure 201 (Sheet 1)
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Passenger Seat ConÞgurations Figure 201 (Sheet 2)
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Passenger Seat ConÞgurations Figure 201 (Sheet 3)
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Passenger Seat ConÞgurations Figure 201 (Sheet 4)
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Passenger Seat ConÞgurations Figure 201 (Sheet 5)
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Optional Utility Seat Installation Figure 202 (Sheet 1)
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Passenger Commuter Seat Installation Figure 203 (Sheet 1)
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Passenger Commuter Seat Installation Figure 203 (Sheet 2)
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Passenger Commuter Seat Installation Figure 203 (Sheet 3)
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Passenger Commuter Seat Installation Figure 203 (Sheet 4)
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Passenger Commuter Seat Installation Figure 203 (Sheet 5)
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Passenger Commuter Seat Installation Figure 203 (Sheet 6)
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Optional Bench Seat Installation Figure 204 (Sheet 1)
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Optional Bench Seat Installation Figure 204 (Sheet 2)
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WARNING: It is extremely important that passenger seat locks are securely engaged. Acceleration and deceleration could permit seat to become disengaged from rails, creating a hazardous situation, especially during takeoff and landing. (3)
4.
Push locks down into seat track openings. Ensure locks are fully engaged in seat track.
C.
Remove Seat Belts and Shoulder Harness (Refer to Figure 202). (1) Remove bolts, washers and nuts attaching ends of seat belt to seat frame. (2) Remove bolt, washer and nut attaching shoulder harness to top of seat back frame. (3) Remove seat belt and shoulder harness as a unit.
D.
Install Seat Belts and Shoulder Harness (Refer to Figure 202). (1) Position seat belt and shoulder harness on seat. Ensure correct side is up and no twists are present in belt. (2) Attach ends of seat belt and shoulder harness to seat frame and to top of seat back frame using bolts, washers and nuts. (3) Ensure all hardware is tight.
E.
Utility Seat Folding/Unfolding and Removal/Installation (Refer to Figure 202 ). (1) Fold and remove utility seat. (a) Pull pins out of each seat back brace and fold seat forward. (b) Pull Þve pins out of leg braces and fold all legs toward center of seat. (c) Remove and stow seat in aft cargo compartment. Refer to Remove Seats, located in this section. (2) Install and unfold utility seat. (a) Position utility seat in airplane and unfold legs. (b) Insert pins in each seat back bracket. (c) Insert Þve pins in leg braces. (d) Ensure all pins are inserted correctly. (e) Install seats in airplane. Refer to Install Seats, located in this section.
Commuter Passenger Seats and Restraint System Removal//Installation NOTE:
Removal/Installation procedures are typical for both the single passenger and two-place passenger seats.
NOTE:
Model 208B airplanes with 10-Place and 13-Place Commuter Passenger Seating may install left seats on right side of airplane and right seats on left side of airplane to provide additional space between seat and sidewall.
A.
Remove Seats (Refer to Figure 203). (1) Raise seat track lock on each attach Þtting. (2) Move seat forward or aft slightly until track attach Þttings are lined up with openings in seat. (3) Lift seat out.
B.
Install Seats (Refer to Figure 203). (1) Position seat on seat track and insert attach Þttings in seat track openings. (2) Move seat forward or aft until locks are lined up with opening in seat track.
WARNING: It is extremely important that passenger seat locks are securely engaged. Acceleration and deceleration could permit seat to become disengaged from rails and create a hazardous situation, especially during takeoff and landing. (3)
Push locks down into seat track openings. Ensure locks are fully engaged in seat track.
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5.
Restraint System Removal and Installation A.
Remove and Install the Manual Adjust Seat Belt and Shoulder Harness (Refer to Figure 203). (1) Remove the manual adjust seat belt/shoulder harness. (a) Remove the bolts, washers, spacers and nuts that connect the attach stirrups to the frame. (b) Remove the screws that attach the seat panel to the frame and remove the seat panel. (c) Remove the headrest and the shoulder harness stirrup. (2) Install the manual adjust seat belt and shoulder harness. NOTE: (a) (b) (c)
B.
Install the shoulder harness stirrup and the headrest to the seat frame. Attach the seat panel to the frame with screws. Install the attach stirrups to the seat frame with bolts, washers, spacers and nuts.
Remove and Install the Inertial Reel Automatic Adjust Seat Belt and Shoulder Harness (Refer to Figure 203). (1) Remove the Inertial Reel Automatic Adjust Seat Belt/Shoulder Harness. (a) Remove and discard the bolts, washers, spacers, and nuts that attach the lap belt stirrups to both sides of the frame. (b) Remove the lap belt from the seat frame. (c) Remove and keep the screws that attach the seat back panel to the seat back frame and remove seat back panel. (d) Remove the webbing from the web guide at the top of the seat back. (e) Remove the inertial real hardware and the inertial reel from the seat back. Discard the hardware. (2) Install the Inertial Reel Automatic Adjust Seat Belt/Shoulder Harness. NOTE: (a) (b) (c) (d)
6.
The seat belt and shoulder harness must not be twisted when you install them.
The seat belt and the shoulder harness must not be twisted when you install them.
Install the inertial reel to the seat back with new hardware. Install the webbing through the web guide at the top of the seat back. Install the seat back panel to the seat back frame with the kept screws. Install the lap belt connectors to both sides of the seat frame with new hardware.
Three-Place Bench Seat and Restraint System Removal/Installation NOTE:
A 3-place bench seat is installed in the 14-Place passenger airplane.
A.
Remove Bench Seats (Refer to Figure 204). (1) Grasp front edge of lower cushion and lift to disengage fabric fastener. Remove cushion. (2) Lift and remove seat back.
B.
Install Bench Seats (Refer to Figure 204). (1) Locate lower seat cushion on raised ßoorboard area and press Þrmly to ßoor to fully engage fabric fastener. (2) Place seat belts across lower cushion. (3) Wedge seat back between lower cushion and aft cabin partition.
C.
Remove Seat Belts and Shoulder Harness (Refer to Figure 204). (1) Disconnect seat belt assembly quick-release stirrups from inboard and outboard attach brackets located on ßoor. (2) Remove bolt and spacer attaching shoulder harness assembly to aft partition.
D.
Install Seat Belts and Shoulder Harness (Refer to Figure 204). (1) Attach seat belt assembly quick-release stirrups to inboard and outboard attach brackets located on ßoor. (2) Position shoulder harness stirrups to aft partition and secure using bolt and spacer.
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MODEL 208 MAINTENANCE MANUAL PASSENGER SEATS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the passenger seats in a serviceable condition.
Task 25-21-00-220 2.
Passenger Seats Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the passenger seats.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do a Detailed Inspection of the Passenger Seats. (1) Examine all passenger seat assemblies for rips, tears, cleanliness, security of attached components, and other signs or damage. (2) Examine the seat belts, shoulder straps, restraint straps, and retainers for security of installation, rips, cuts, tears, frayed edges, cleanliness, and general condition. (a) Replace frayed and/or cut belts. (3) Examine all belts for legibility of the certification label. (4) Examine the belt length adjustments for correct operation. (5) Examine all belt attach points for security and damage. (6) Examine the buckle assemblies for signs of wear, cleanliness, security, and general condition. (7) Examine the belt latching mechanism for correct operation. (8) Examine the seat belt inertia reels for security and general condition. (9) Check the inertia reel for freedom and smoothness of operation by slowly pulling out on the shoulder harness belt. (a) Make sure that the shoulder harness belt retracts onto the inertia reel when it is released. (10) Pull the shoulder harness belt out quickly from the inertia reel and make sure that the inertia reel locks. (a) Make sure the locking mechanism releases when the shoulder harness is relaxed. (11) Examine the seat frame and the attach brackets for cracks, corrosion, and signs of damage. (12) Examine the seat rails for cleanliness, cracks, corrosion, loose or failed fasteners, or other damage.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL CABIN UPHOLSTERY - MAINTENANCE PRACTICES 1.
General A.
2.
Cabin upholstery consists of a headliner, side panels, window trim, floor covering and a control column cover. Maintenance practices consist of removal and replacement of the various panels.
Upholstery Removal/Installation (Model 208 and 208B Passenger) A.
Remove Fabric Headliner (Refer to Figure 201). (1) Remove window moldings, light consoles, fresh air inlet covers and any other visible retainers securing headliner. (2) Carefully separate edges from channels at top of windows.
(3) (4) B.
NOTE:
Always work front to rear when removing headliner.
NOTE:
Due to difference in length and contour of wire bows, each bow should be tagged to ensure proper location in headliner.
Starting at front of headliner, work headliner down, removing bows from metal tabs which hold bows to cabin top. Detach each bow in succession. Remove headliner assembly and bows from airplane.
Install Fabric Headliner (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5)
Before installation, check all items concealed by headliner for security. Use a wide cloth tape to secure loose wires to fuselage. Straighten tabs bent during removal of headliner.
Insert wire bows into headliner seams and secure rear most edges of headliner after positioning two bows at rear of headliner. Stretch material along edges to ensure it is properly centered, but do not stretch enough to destroy ceiling contours or distort wire bows. Glue edges of headliner under channel along cabin at top of windows and above doors. Work headliner forward, installing each wire bow in place with metal tabs. Stretch headliner just taut enough to avoid wrinkles and maintain a smooth contour. When all bows are in place and fabric edges are secured, trim off excess fabric and reinstall all items removed.
C.
Remove Upholstery Side Panels And Trim (Refer to Figure 202). (1) Remove seats to gain access to side panels Refer to Passenger Seats - Maintenance Practices. (2) Remove retaining strips and window trim as required. (3) Remove machine screws to free panels as required.
D.
Install Upholstery Side Panels And Trim (Refer to Figure 202). (1) Position panels on areas to be installed. (2) Locate nutplates and install screws. (3) Install retaining strips and window trim. (4) Install seats. Refer to Passenger Seats - Maintenance Practices.
E.
Remove Crew Door Panel Upholstery (Refer to Figure 203). (1) Open door. (2) Remove all loose items in door pouch. (3) Remove foul weather window from pilot crew door. Refer to Chapter 56, Flight Compartment Windows - Removal/Installation. (4) Remove door assist strap from door. (5) Remove window molding. (6) Remove inner door latch handle, escutcheon, and lock knob. Refer to Chapter 52, Passenger and Crew Doors - Maintenance Practices. (7) Remove crew door trim panel. (8) If applicable, remove fire extinguisher bracket from crew door trim panel.
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Fabric Headliner Installation Figure 201 (Sheet 1)
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Fabric Headliner Installation Figure 201 (Sheet 2)
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Upholstery Side Panels and Trim Installation Figure 202 (Sheet 1)
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Upholstery Side Panels and Trim Installation Figure 202 (Sheet 2)
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Upholstery Side Panels and Trim Installation Figure 202 (Sheet 3)
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Crew Door Panel Installation Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL F.
Install Crew Door Panel Upholstery (Refer to Figure 203). (1) If applicable, attach fire extinguisher bracket to door using screws. (2) Position crew door trim panel to door and secure with screws. (3) Install inner door latch escutcheon, handle and lock knob. Refer to Chapter 52, Passenger and Crew Doors - Maintenance Practices. (4) Position window trim around window and secure with screws. (5) Attach door assist strap to door using screws. (6) Install foul weather window on pilot crew door. Refer to Chapter 56, Flight Compartment Windows - Removal/Installation.
G.
Remove Cargo Door Panels And Trim (Refer to Figure 204). (1) Open door. (2) Remove inner door latch handle, escutcheon and lockplate. Refer to Chapter 52, Cargo Door Maintenance Practices. (3) Disconnect door pull cable from both the lower cargo door and the upper cargo door. (4) Remove door pull cable hardware from both the lower cargo door and the upper cargo door. (5) Remove screws securing both ends of door restraint latch assembly to lower door. (6) Remove lower cargo door trim panel from lower door.
CAUTION: It is necessary to provide support for the upper door when gas springs are removed. (7) (8) H.
Detach gas springs from upper door. Refer to Chapter 52, Cargo Door - Maintenance Practices. Remove screws securing upper cargo door window molding to door and remove molding.
Install Cargo Door Panels And Trim (Refer to Figure 204). (1) Position upper cargo door window molding to door and secure using screws.
CAUTION: It is necessary to provide support for the upper door during the following steps. (2) (3) (4) (5) (6) (7)
Attach gas springs to upper door. Refer to Chapter 52, Cargo Door - Maintenance Practices. Position lower cargo door trim panel to lower door and secure using screws. Install both ends of door restraint latch assembly to lower door. Attach door pull cable hardware to both the lower cargo door and the upper cargo door. Connect door pull cable between the doors. Install inner door latch escutcheon, handle and lockplate. Refer to Chapter 52, Cargo Door Maintenance Practices. Close doors.
I.
Remove Passenger Door Panels And Trim (Refer to Figure 204). (1) Open door. (2) Remove inner door latch handle, escutcheon and lockplate. Refer to Chapter 52, Passenger And Crew Doors - Maintenance Practices. (3) On upper door, remove pins attaching door pull cable to eyebolt and remove eyebolt from door. (4) On lower door, remove step assembly. Refer to Chapter 52, Passenger And Crew Doors Maintenance Practices. (5) Detach gas springs from upper door. Refer to Chapter 52, Passenger And Crew Doors Maintenance Practices. (6) Remove screws securing upper cargo door window molding to door and remove molding.
J.
Install Passenger Door Panels And Trim (Refer to Figure 204). (1) Position window molding to upper door and secure using screws.
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Passenger and Cargo Door Upholstery Installation Figure 204 (Sheet 1)
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CAUTION: It is necessary to provide support for the upper door during the following steps. (2) (3) (4) (5) (6) (7) 3.
Attach gas springs to upper door. Refer to Chapter 52, Passenger And Crew Doors Maintenance Practices. Install eyebolt to upper door and attach door pull cable to eyebolt. Install step assembly to lower door. Refer to Chapter 52, Passenger And Crew Doors Maintenance Practices. Install eyebolt to lower door and attach door pull cable to eyebolt. Install inner door latch handle, escutcheon and lockplate. Refer to Chapter 52, Passenger And Crew Doors - Maintenance Practices. Close doors.
Upholstery Removal/Installation (Cargomaster and 208B) A.
Remove Hard Shelled Model 208 Headliner (Refer to Figure 205). NOTE: (1)
(2) (3) (4)
The hard shell consists of a fabric-type headliner in section 1 and rigid shells in the remaining five sections.
Section 1. (a) Remove windshield retainer molding, speaker covers, warning horn cover, light consoles, fresh air outlet covers and any other visible retainers securing headliner. (b) Carefully separate edges from channels at top of windows. (c) Remove headliner assembly and bow from airplane. Section 2. (a) Remove screws securing headliner and pull headliner down away from Velcro fastener strips on cabin top. Section 4 and 5. (a) Remove screws securing interior light escutcheons and remove escutcheons. Section 3, 4, 5 and 6. (a) Remove screws and pop rivets securing headliner, and pull headliner down away from cabin top.
B.
Install Hard Shelled Model 208 Headliner (Refer to Figure 205 ). (1) Section 1. (a) Check all items concealed by headliner for security. Use a wide cloth tape to secure loose wires to fuselage and straighten tabs bent during removal of headliner. (b) Insert wire bow into headliner seam and secure bow to fuselage attach points. (c) Stretch material edges to ensure it is properly centered. No not stretch enough to destroy ceiling contours or distort wire bow. (d) Glue edges of headliner under channel along cabin at top of windows and above doors. (e) Install windshield retainer molding, speaker covers, warning horn cover, light consoles, fresh air outlet covers and any other visible retainers securing headliner. (2) Section 2. (a) Flex headliner and position up against Velcro strips on top of fuselage. Secure section to airplane using screws. (3) Section 3, 4, 5 and 6. (a) Flex headliner and position against cabin top. Secure section to airplane using screws and pop rivets. (4) Position interior light escutcheons to cabin top and secure using screws.
C.
Remove Hard Shelled Model 208B Headliner (Refer to Figure 205). (1) Section 1. (a) Remove windshield retainer molding, speaker covers, warning horn cover, light consoles, fresh air outlet covers and any other visible retainers securing headliner. (b) Carefully separate edges from channels at top of windows. (c) Remove headliner assembly and bow from airplane.
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Hard Shell Headliner Installation Figure 205 (Sheet 1)
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Hard Shell Headliner Installation Figure 205 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5)
Section 2. (a) Remove screws securing headliner and pull headliner down away from Velcro fastener strips on cabin top. Section 3 and 7. (a) Remove screws securing interior light escutcheons and remove escutcheons. Section 5, 6, 7 and 8. (a) Remove screws securing anchor plates and remove anchor plates. Section 3, 4, 5, 6, 7 and 8. (a) Remove screws and pop rivets securing headliner and pull headliner down away from cabin top.
D.
Install Hard Shelled Model 208B Headliner (Refer to Figure 205). (1) Section 1. (a) Check all items concealed by headliner for security. Use a wide cloth tape to secure loose wires to fuselage and straighten tabs bent during removal of headliner. (b) Insert wire bow into headliner seam and secure bow to fuselage attach points. (c) Stretch material edges to ensure it is properly centered. Do not stretch enough to destroy ceiling contours or distort wire bow. (d) Glue edges of headliner under channel along cabin at top of windows and above doors. (e) Install windshield retainer molding, speaker covers, warning horn cover, light consoles, fresh air outlet covers and any other visible retainers securing headliner. (2) Section 2. (a) Flex headliner and position up against Velcro strips on top of fuselage. Secure section to airplane using screws. (3) Section 3, 4, 5, 6, 7 and 8. (a) Flex headliner and position against cabin top. Secure section to airplane using screws and pop rivets. (4) Position interior light escutcheons to cabin top and secure using screws. (5) Position anchor plates and install screws.
E.
Remove Upholstery Hard Shell Side Panels and Trim (Refer to Figure 206). NOTE: (1)
Superficial damage to side panels may be repaired using a Gillpatch. Major damage requires replacement of panels.
Remove tape from seams. NOTE:
(2)
Airplanes 20800001 Thru 20800116 and 208B0001 Thru 208B0027 use Gilltape on panel seams. Airplanes 20800117 and On and 208B0028 and On use 3M tape on panel seams.
Remove machine screws attaching panels to airplane.
F.
Install Upholstery Hard Shell Side Panels and Trim (Refer to Figure 206). (1) Position panels on areas to be covered. (2) Locate nutplates behind panel and secure panels using machine screws. (3) Apply tape to all panel seams. Center tape over joint and use three inch lap at intersections.
G.
Remove Crew Door Panel Upholstery (Refer to Figure 203). (1) Open door. (2) Remove all loose items in door pouch. (3) Remove foul weather window from pilot crew door. Refer to Chapter 56, Flight Compartment Windows - Removal/Installation. (4) Remove door assist strap from door. (5) Remove window molding. (6) Remove inner door latch handle, escutcheon, and lock knob. Refer to Chapter 52, Passenger and Crew Doors - Maintenance Practices. (7) Remove crew door trim panel. (8) If applicable, remove fire extinguisher bracket from crew door trim panel.
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Hard Shell Side Panels and Trim Installation Figure 206 (Sheet 1)
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Hard Shell Side Panels and Trim Installation Figure 206 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL H.
Install Crew Door Panel Upholstery (Refer to Figure 203). (1) If applicable, attach fire extinguisher bracket to door using screws. (2) Position crew door trim panel to door and secure with screws. (3) Install inner door latch escutcheon, handle and lock knob. Refer to Chapter 52, Passenger and Crew Doors - Maintenance Practices. (4) Position window trim around window and secure with screws. (5) Attach door assist strap to door using screws. (6) Install foul weather window on pilot crew door. Refer to Chapter 56, Flight Compartment Windows - Removal/Installation.
I.
Remove Cargo Door Panels And Trim (Refer to Figure 204). (1) Open door. (2) Remove inner door latch handle, escutcheon and lockplate. Refer to Chapter 52, Cargo Door Maintenance Practices. (3) Disconnect door pull cable from both the lower cargo door and the upper cargo door. (4) Remove door pull cable hardware from both the lower cargo door and the upper cargo door. (5) Remove screws securing both ends of door restraint latch assembly to lower door. (6) Remove lower cargo door trim panel from lower door.
CAUTION: It is necessary to provide support for the upper door when gas springs are removed. (7) (8) J.
Detach gas springs from upper door. Refer to Chapter 52, Cargo Door - Maintenance Practices. Remove screws securing upper cargo door window molding to door and remove molding.
Install Cargo Door Panels And Trim (Refer to Figure 204). (1) Position upper cargo door window molding to door and secure using screws.
CAUTION: It is necessary to provide support for the upper door during the following steps. (2) (3) (4) (5) (6) (7)
Attach gas springs to upper door. Refer to Chapter 52, Cargo Door - Maintenance Practices. Position lower cargo door trim panel to lower door and secure using screws. Install both ends of door restraint latch assembly to lower door. Attach door pull cable hardware to both the lower cargo door and the upper cargo door. Connect door pull cable between the doors. Install inner door latch escutcheon, handle and lockplate. Refer to Chapter 52, Cargo Door Maintenance Practices. Close doors.
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MODEL 208 MAINTENANCE MANUAL REAR COMPARTMENT WALL - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Airplanes may be equipped with a canvas rear wall or a one- piece, padded rear wall. (1) On airplanes with canvas rear walls, zippers are installed on the left and right side which allow the canvas rear wall to be rolled up and secured near the top of the airplane structure. Strap assemblies are located behind the canvas rear wall and add support to the wall. (2) On airplanes with padded rear walls, the panel is secured to the structure using screws.
Canvas Wall Removal/Installation A.
Remove Canvas Wall (Refer to Figure 201). (1) Release Camloc fasteners at bottom of canvas rear wall. (2) Open zippers and roll canvas up toward top of structure. (3) Remove nuts, bolts washers and spacers securing strap assemblies to structure. (4) Remove aft substrate trim panels. Refer to Cabin Upholstery - Maintenance Practices. (5) Carefully separate glued canvas from airplane structure at left, right and top attach points.
B.
Install Canvas Wall (Refer to Figure 201). (1) Remove all traces of glue from canvas and from airplane structure. (2) Attach canvas to structure at left, right and top positions. Allow to dry. (3) Install aft substrate trim panels. Refer to Cabin Upholstery - Maintenance Practices. (4) Install strap assemblies to structure using nuts, bolts, washers and spacers. (5) Close zippers. (6) Fasten Camlocs.
Padded Rear Wall Removal/Installation A.
Remove Rear Wall. (1) Remove screws and wide-area flat washers securing padded rear wall to structure. (2) Grasp straps at bottom of wall and pull forward. Remove padded wall from airplane.
B.
Install Rear Wall. (1) Position padded rear wall to structure and secure using screws and washers.
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Rear Wall Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FLOOR COVERING/CONTROL COLUMN COVER - MAINTENANCE PRACTICES 1.
General A.
2.
3.
This maintenance practice covers control column cover removal/installation and vinyl floor covering removal/installation. For maintenance practices related to airplanes equipped with plywood floorboards, refer to Chapter 53, Floorboards - Maintenance Practices.
Control Column Cover Removal/Installation A.
Remove Control Column Cover (Refer to Figure 201). (1) Remove screws securing top of control column cover to structure. (2) Remove screws securing bottom of control column cover to floor. (3) Remove bolts and washers on aft side of control column cover and remove control column cover from airplane.
B.
Install Control Column Cover (Refer to Figure 201). (1) Position control column cover around control column. (2) Secure aft end of control column cover together using bolts and washers. (3) Attach control column cover to structure using screws.
Floor Covering Removal/Installation A.
Remove Floor Covering on Airplanes 20800001 Thru 20800242 and 208B0001 Thru 208B0458 (Refer to Figure 201). (1) Using a dull putty knife between floor covering and floorboard, carefully separate floor covering from floorboard. (2) Remove floor covering from floorboard.
B.
Install Floor Covering on Airplanes 20800001 Thru 20800242 and 208B0001 Thru 208B0458 (Refer to Figure 201). (1) Position floor covering on floorboard. (2) Apply industrial grade rubber cement to outer edges of floor covering to hold in place.
C.
Floor Covering on Airplanes 20800243 and On and 208B0459 and On (Refer to Figure 201). NOTE:
The floor covering is held in place by a hook type fastener stitched to the floor covering and a hook attached to the floor panels using pressure sensitive adhesive backed fasteners. After removal of the outer edge strips, secured with bolts, removal and/or installation of the floor covering is simply a matter of pulling off the covering or pressing back in place.
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Cabin Floor Covering and Control Column Cover Figure 201 (Sheet 1)
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Cabin Floor Covering and Control Column Cover Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL CARGO COMPARTMENT COMPONENTS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Cargo airplanes may be equipped with a variety of nets, barriers and hold-down devices designed to secure and restrain cargo at various points in the airplane. The following devices may be installed on the airplanes: (1) Cargo Tie-Down Straps - Cargo tie-down straps are used to secure cargo against movement. Tie-down straps are typically secured to the airplane fuselage by either seat rail tracks or anchor plates. (a) Beginning at Airplane 20800093 and On, a system of tie-down straps and anchors are offered that can be utilized to both tie down cargo within the airplane and to tie down the airplane itself. The standard tie-down configuration, with a 3,000 pound rating, may be utilized to 1 restrain cargo at any attach point within the airplane. Additionally, this strap may be used to tie down the airplane at approved mooring points. The heavy duty tie-down configuration, with a 5,000 pound rating, may be utilized 2 only at the aft passenger seat tracks to restrain cargo. Additionally, this strap may be used to tie down the airplane at approved mooring points. (2) Cargo Barrier - The cargo barrier is located aft of the flight compartment at station 166.45. Brackets on the lower edge of the barrier are bolted to pilot and copilot seat rails. Brackets on the top edge of the barrier are bolted to the bulkhead. Pilot and copilot can gain access to cargo compartment through access nets located outboard, and on top of, the cargo barrier. Access nets are secured with four to six quick-disconnect fasteners, anchored to the cargo barrier and the bulkhead. (3) Cargo Partition Nets - The airplane may be equipped with two canvas partition nets. When in use, partition nets are secured with quick-disconnect fasteners engaged with anchor plates. The anchor plates are secured to seat rails and bulkheads located at FS168.70, 181.50, 208.00, 234.00, 259.00 and 284.00 (Model 208); or FS188.70, 246.80, 282.00, 307.00, 332.00 and 356.00 (Model 208B). When not in use, partitions are stored with the airplane loose equipment. (4) Cargo Door Restraint Nets - The airplane may be equipped with cargo restraint nets installed at the cargo door entrance to protect personnel from shifted cargo when cargo door is opened. Rings, attached to the top edges of restraint nets, encompass a rod which is secured to fuselage structure. Nets are joined together from top to bottom, at center of cargo door entrance, with snap fasteners. Edges of nets, opposite to snaps, are secured to fuselage with screws.
Cargo Tie-Down Straps Removal/Installation A.
Remove Cargo Tie-Down Straps on Airplanes 20800001 Thru 20800092 (Refer to Figure 201). (1) Loosen tie-down straps. (2) Press down on plunger and slide fittings forward to remove from seat rail or anchor plate.
B.
Install Cargo Tie-Down Strap on Airplanes 20800001 Thru 20800092 (Refer to Figure 201). (1) Position fitting over top of seat rail (or anchor plate). Push down on plunger and slide fitting rearward until fitting seats fully in place. (2) Grasp tie-down strap and pull firmly to ensure a positive lock in the seat rail or anchor.
C.
Remove Cargo Tie-Down Straps on Airplanes 20800093 and On (Refer to Figure 201 ). (1) Loosen tie-down straps. (2) Remove fittings from seat rails or anchor plates.
D.
Install Cargo Tie-Down Strap on Airplanes 20800001 Thru 20800092 (Refer to Figure 201). (1) Position fittings on top of seat rail (or anchor plate). Push down (or pull up) on release and slide fitting rearward until fitting seats fully in place. (2) Grasp tie-down strap and pull firmly to ensure a positive lock in the seat rail or anchor.
Cargo Barrier and Access Nets Removal/Installation A.
Remove Access Nets (Refer to Figure 202). (1) Position thumb on plunger.
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Cargo Tie-Downs Figure 201 (Sheet 1)
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Cargo Tie-Downs Figure 201 (Sheet 2)
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Cargo Tie-Downs Figure 201 (Sheet 3)
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Cargo Barrier and Access Nets Installation Figure 202 (Sheet 1)
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Cargo Barrier and Access Nets Installation Figure 202 (Sheet 2)
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Cargo Barrier and Access Nets Installation Figure 202 (Sheet 3)
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4.
Position fitting between forefinger and middle finger and pull on fitting. Slide plunger out of anchor plate. Repeat steps (1) thru (3) until access net is free of all anchor plates.
B.
Install Access Nets (Refer to Figure 202). (1) Position net in place. (2) Grasp fitting between forefinger and middle finger, and push plunger in with thumb. (3) Slide plunger into anchor plate and release fitting. Ensure all fasteners are secured.
C.
Remove Anchor Plates (Refer to Figure 202). (1) Remove cargo nets. Refer to Cargo Barrier And Access Nets Removal/Installation. (2) Remove screws securing anchor plates to barriers or bulkheads and remove anchor plates.
D.
Install Anchor Plates (Refer to Figure 202). (1) Position anchor plates to barriers or bulkheads and secure using screws. (2) Install Access Nets. Refer to Cargo Barrier And Access Nets Removal/Installation.
E.
Remove Cargo Barrier (Refer to Figure 202). (1) Remove cargo nets. Refer to Cargo Barrier And Access Nets Removal/Installation. (2) On Airplanes 20800001 Thru 20800108, remove cotter pin, washers and clevis pins at bottom of cargo barrier. (3) On Airplanes 20800109 and On, and 208B0001 and On, remove pins, nuts, washers and bolts from bottom of cargo barrier. (4) Remove bolts and washers securing top of cargo barrier to bracket. (5) Remove cargo barrier.
F.
Install Cargo Barrier (Refer to Figure 202). (1) Position cargo barrier to seat rails and bulkhead. (2) Secure top of cargo barrier to bulkhead using bolts and washers. (3) Secure bottom of cargo barrier to seat rail using cotter pins, nuts, washers and bolts or clevis pins as required. (4) Check for security of installation.
Cargo Partition Nets Removal/Installation A.
Remove Cargo Partition Nets (Refer to Figure 203 ). (1) Loosen adjustable straps. (2) Position thumb on plunger. (3) Position fitting between forefinger and middle finger and pull on fitting. (4) Slide plunger out of anchor plate. NOTE: (5)
B.
After all fasteners have been released, remove cargo partition net.
Install Cargo Partition Nets (Refer to Figure 203). (1) Position cargo partition properly so that fasteners will align with correct anchor plates. (2) Position fittings between forefinger and middle finger and push plunger with thumb. (3) Slide plunger into anchor plate and release fitting. NOTE:
(4) C.
On seat rail fasteners, slide plunger to cutout and lift plunger out of seat rail.
On seat rail fasteners, position plunger down through cutout in seat rail. Slide plunger forward or aft enough to align fitting with cutouts so that when released, fitting will drop down through cutout and secure fastener to seat rail.
When all fasteners have been installed, tighten adjustable straps as required.
Remove Anchor Plates (Refer to Figure 203). (1) Remove partition nets. Refer to Cargo Partition Nets Removal/Installation. (2) Remove screws securing anchor plate to bulkhead floorboard. (3) Remove anchor plate.
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Cargo Partition Net Installation Figure 203 (Sheet 1)
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Cargo Partition Net Installation Figure 203 (Sheet 2)
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Cargo Partition Net Installation Figure 203 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL D.
5.
Install Anchor Plates (Refer to Figure 203). (1) Position anchor plate to bulkhead floorboard and secure using screws. (2) Install partition nets. Refer to Cargo Partition Nets Removal/Installation.
Cargo Door Restraint Net Removal/Installation A.
Remove Door Restraint Net (Refer to Figure 204 ). (1) Unsnap and separate cargo net. (2) Remove screw securing nets to fuselage structure. (3) Remove screws, washers and brackets securing rod to fuselage structure. (4) Slide rod out of net rings.
B.
Install Door Restraint Net (Refer to Figure 204). (1) Slide rod through net rings and brackets. (2) Position rod to fuselage structure and secure using screws and washers. (3) Position nets to fuselage structure using screws. (4) Ensure that net rings slide easily on rod.
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Cargo Door Restraint Net Installation Figure 204 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL CARGO COMPARTMENT COMPONENTS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the cargo compartment components in a serviceable condition.
Task 25-51-00-220 2.
Cargo Nets Detailed Inspection A.
General (1) This task gives the information needed to complete the inspection procedures for the cargo net and the barrier.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Cargo Nets Detailed Inspection. (1) Cargo Nets and Vertical Partitions Nets. (a) Examine the nets for condition, security, deterioration, stitching, and correct operation of the attachment devices. (b) Examine the cargo door restraint net attachment bar for condition, security, and correct attachment of the net to the bar. (c) Examine the door restraint nets for condition and security of the stud and the socket hardware to the forward and the aft nets. (2) General Visual Inspection of the Cargo Barrier and Access Nets. (a) Examine the condition and the security of installation of the cargo barrier. (b) Use the coin tap test to examine the barrier for debonds. 1 Examine the cargo barrier and the access mounting area for signs of delamination. (3) General Visual Inspection of the Cargo Tie-Down Straps (a) Examine the straps for condition, wear, and positive locking of the attaching mechanism (b) Examine the floor anchors for condition and correct operation. (c) Examine the cargo nets for condition and security of the attachment hardware.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL CARGO POD - MAINTENANCE PRACTICES 1.
2.
General A.
A cargo pod with a 80-cubic-foot and a 700 pounds capacity can be installed on Model 208 airplanes. Screws attach the pod to the bottom of the fuselage. The pod (and doors) have a nomex inner housing, a layer of kevlar and an outer layer of fiberglass. Aluminum bulkheads divide the pod into three compartments. Each compartment has a door to load cargo installed on the left side of the pod. The doors have hinges on the bottom and latches on the top. The doors lock in the closed position when latch handles are turned to a horizontal position.
B.
A cargo pod with a 111.5-cubic-foot and a 1090 pounds capacity can be installed on Model 208B airplanes. Screws attach the pod to the bottom of the fuselage. The pod (and doors) are fabricated with a nomex inner housing, a layer of kevlar and an outer layer of fiberglass. Aluminum bulkheads divide the pod into four compartments. Each compartment has a door to load cargo installed on the left side of the pod. The doors have hinges on the bottom and latches on the top. The doors lock in the closed position when latch handles are turned to a horizontal position.
Cargo Pod Removal/Installation A.
Remove the Cargo Pod (Refer to Figure 201). (1) Make sure that the cargo pod is empty. (2) Remove the TKS anti-ice fluid tank. Refer to Chapter 30, TKS Anti-Ice Fluid Tank Removal/ Installation. (3) Remove the fuel line drain hoses. (a) Find the fuel line drain hoses on the forward side of the cargo pod compartment bulkheads. (b) Remove all brackets and hardware that attach the fuel line drain hoses to the bulkheads. (c) Loosen the clamps at the top of the fuel line drain hoses. (d) Remove the fuel line drain hoses from the pod. (4) Remove the fuselage moisture drain lines (Airplanes 20800072 and On and 208B0001 and On). (a) Find the moisture drain lines on the cargo pod bulkheads (Refer to Figure 201, Sheet 2). (b) Loosen the screws and clamps that attach the moisture drain lines to the bulkheads. (c) Move the moisture drain lines down and out of the cargo pod. (5) Find and disconnect the DME and transponder cables from the antennas. NOTE:
Cover assemblies give protection to antenna cable connectors that are on cargo pod bulkheads. Remove the cover assemblies before you disconnect the cables from the antenna.
(6)
Disconnect the antenna cables at the DME and transponder remote units: (a) Remove the floor cover and pilot’s control column cover. Refer to Floor Covering/Control Column Cover - Maintenance Practices. (b) Remove the access panels 211EL, 231BL, 231DL, 232BR, and 232DR. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation . (c) At the DME antenna and transponder remote unit cables, cut and discard tie wraps. (d) Identify and disconnect the DME antenna cable from the DME remote unit. (e) Identify and disconnect the transponder antenna cable from the transponder remote unit. (7) Find the fuel drain line sleeve on the left side of the cargo pod, a small distance above the main gear. (8) Remove the screws, washers and nuts that attach the fuel drain line sleeve to the cargo pod. (9) Carefully use a non metal scraper to disconnect the fuel drain line sleeve from the cargo pod (on the external area of the cargo pod) and from the fuel drain line cover (on the internal area of the cargo pod). (10) When the sealant bond is broken, carefully pull the fuel drain line sleeve out from the cargo pod and remove it from the airplane. (11) Remove the main gear-to-cargo pod fairings. Refer to Chapter 32, Main Landing Gear Maintenance Practices. (12) Remove the main gear inner and outer covers from the cargo pod.
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Cargo Pod Installation Figure 201 (Sheet 1)
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Cargo Pod Installation Figure 201 (Sheet 2)
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Cargo Pod Installation Figure 201 (Sheet 3)
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Cargo Pod Installation Figure 201 (Sheet 4)
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Cargo Pod Installation Figure 201 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL (13) Loosen the nose gear strut fairing and move it forward. (a) Make sure that the cargo pod does not touch the fairing when it is lowered. (14) Remove the brace cover halves that are on the front of the cargo pod. (15) Find and remove the grounding straps at FS 208.00 and 284.00. (16) Put two light capacity hydraulic jacks (with jack pads) on the lower surface of the cargo pod, forward of the aft cargo pod door. (17) Put one light capacity hydraulic jack (with jack pad) aft of the forward cargo pod door. (18) Lift the jacks until the jack pads rest tightly against the lower surface of the cargo pod. (19) Remove the screws that attach the cargo pod to the fuselage. (20) Lower all the jacks 0.50 inch at the same time. (21) Find a fuselage internal structural member at the forward edge of the cargo pod. (22) Use a thin nonmetal scrapper and perforate the seal between the fuselage and the pod. (23) Use safety wire pliers to twist the two pieces 6 to 8 foot lengths of 0.032 stainless steel lockwire together to make a seal cutting tool. (24) Put one end of the seal cutting tool through the perforation. (25) With one person in the pod and a second person external, wind the cutter wire at each of the two ends around a block of wood or equivalent tool to serve as handles. (26) Cut through the seal. (a) Put wood tongue depressors or an equivalent tool between the fuselage and the pod, at 1 to 2 foot increments. NOTE:
This helps to make sure that the seal does not attach again to the pod.
CAUTION: When you lower the cargo pod, make sure that the pod does not touch the fuel drain lines and the fuel drain valve knob on left side of the cargo pod. (27) Lower all three jacks at the same time. NOTE:
Antenna cables are to remain in the cargo pod whenever the pod is removed from the airplane.
Make sure that the DME and the transponder cables that extend up into fuselage are free to move easily out of the fuselage. (b) Make sure that the DME and the transponder cables are not damaged. (28) If the airplane is to be returned to service without the pod, refer to Modification For Flight Without Pod for further instructions. (a)
B.
Install the Cargo Pod (Refer to Figure 201). (1) Clean the fuselage skin and the cargo pod at all mating points. (2) Put the cargo pod below the fuselage and align the cargo pod with the fuselage attach points. (3) Put two light duty hydraulic jacks (with jack pads) under the lower surface of the cargo pod, forward of the aft cargo pod door. (4) Put one light duty hydraulic jack (with jack pad) under the lower surface of the cargo pod, aft of the forward cargo pod door. (5) Extend the DME and transponder antenna cables up through the holes in the fuselage. (6) Apply FS4291 sealant tape to the cargo pod mount flange surface and remove sealant tape paper liner. (7) Put personnel in the airplane to move the antenna cables up through the fuselage when you lift the cargo pod to the fuselage.
CAUTION: Make sure that the cargo pod does not touch the fuel drain lines or fuel drain valve knob on the left side of the airplane when you lift the cargo pod to the airplane fuselage. (8)
Carefully lift the cargo pod to the airplane fuselage with the jacks. (a) Make sure that the attach holes in cargo pod align with the fuselage nutplates.
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MODEL 208 MAINTENANCE MANUAL (9) Attach the cargo pod to the fuselage with the screws and washers. (10) Attach the tops of the fuel line drain hoses, found on the forward side of the compartment bulkheads, to the fuel drain lines. (11) Attach the inboard and outboard fuel drain line covers, found at the middle point of the bottom of the fuselage with screws. (12) Apply Pro-Seal 890 to the areas that follow: (a) The fuel valve sleeve where it connects with the cargo pod (the external area of the pod). (b) The fuel drain line cover (in the pod). (c) Seal where pod connects to the fuselage. (13) Put the fuel valve sleeve in the pod. (14) Use screws to attach the sleeve to the cargo pod. (15) Use clamps to attach the top of fuel drain line to the fuel drain hoses. (16) Seal the fuel drain line where it goes through the bottom of the cargo pod with Pro-Seal 890. (17) Connect the DME antenna cable to the DME remote unit in the cockpit area. (18) Connect the transponder antenna cable to the transponder remote unit in the cockpit area. (19) Install tie wraps on the DME and transponder cables where appropriate to make sure that they stay in the correct location. (20) Install access panels 211EL, 231BL, 231DL, 232BR and 232DR. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (21) Install the pilot control column cover and the floor covering. Refer to Floor Covering/Control Column Cover - Maintenance Practices. (22) Connect the DME and transponder antenna cables to the correct antennas. (23) Install the protective cover assemblies around the DME and transponder antennas. (24) Apply Pro-Seal 890 to the brace covers at front of cargo pod, and (25) Use screws to attach the brace covers to the cargo pod. (26) Move the nose gear strut fairing in its position aft against the cargo pod. (27) Tighten the fasteners to attach the nose gear strut fairing to the fuselage. (28) Install the main gear inner and outer covers to the cargo pod. (29) Install the main gear-to-cargo pod fairing. Refer to Chapter 32, Main Landing Gear Maintenance Practices. (30) Install the TKS anti-ice fluid tank assembly. Refer to Chapter 30, TKS Anti-Ice Fluid Tank. Removal/Installation (31) Spray water along the cargo pod, with a high pressure water hose, where it attaches to the fuselage and check the interior of pod for leaks. 3.
Cargo Pod Components Removal/Installation A.
Remove the Cargo Pod Door (Refer to Figure 202). (1) Open the cargo pod door. (2) Remove the screws, washers and nuts that attach the door and hinges to the cargo pod. (3) Remove the door.
B.
Install the Cargo Pod Door (Refer to Figure 202). (1) Put the door to the cargo pod and align the hinge holes to the cargo pod attach points. (2) Install the screws, washers and nuts that attach the door to the cargo pod.
C.
Remove the Door Hinges (Refer to Figure 202). (1) Remove the cargo pod door. (2) Remove the screws and washers that attach the hinge to the door, and remove the hinge.
D.
Install the Door Hinges (Refer to Figure 202). (1) Put the hinge on the door and install the washers and screws. (2) Put the hinge on the cargo pod and attach the screws, washers and nuts.
E.
Remove the Door Latch (Refer to Figure 202). (1) Open the cargo door. (2) Remove the pin from the handle, and slide the block and shaft out of the door.
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Cargo Pod Components Installation Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL F.
Install the Door Latch (Refer to Figure 202 ). (1) Put the washers on the shaft.
(2) (3) (4) 4.
NOTE:
Install the washers as necessary to bring the exterior surface of door flush with the exterior surface of the pod. Do not use more than three washers for each screw.
NOTE:
If additional adjustment is needed, loosen the screw and nut that attach the block to the shaft. Slide the block in the desired direction, and tighten the block on the shaft with a screw and a nut.
Place the spring washer on the shaft. Place the shaft through the door. Place the handle on the shaft and attach it with a pin.
Modification for Flight Without the Cargo Pod NOTE: A.
You must complete this procedure if the cargo pod has been removed and the airplane is returned to service without the pod.
Modification Procedures (Refer to Figure 203). (1) Install the screws and washers on all the fuselage nutplates where the cargo pod attaches to the fuselage. (2) Turn both fuel selectors, found in the overhead console, to the OFF position. (3) Find the fuel drain valve knob below the left main gear. (4) Pull the fuel drain valve knob out to drain the fuel reservoir. (5) Remove the screws that attach the inboard fuel drain line cover to the fuselage. (6) Remove the cover to get access to the drain valves, lines and reservoir. (7) Disconnect the fuel drain line fitting at the fuel drain valve. (8) Unscrew and remove the insert from the fuel drain valve. (9) Disconnect the cable from the fuel drain valve. (10) Loosen the clamp that attaches the cable to the fuselage. (11) Disconnect the pump seal fuel drain lines from the elbow fittings. (12) Remove the drain lines, fuel drain valve knob, cable and the outboard fuel drain line cover as one assembly. NOTE:
(13) (14) (15) (16) (17) (18) (19) (20) (21) (22) (23) (24) (25)
On Airplanes 20800001 thru 20800086, inboard and outboard fuel drain line covers were assembled as a single unit. To access the drain valve, remove all the screws and lower the covers away from the fuselage being careful not to bend or break the drain lines and cable. One-piece covers on Airplanes 20800001 thru 20800086 can be replaced with the two-piece configuration, but when you order spares, the covers must be replaced in pairs.
Remove the skin panel from the bottom of the fuselage to get access to the fuel reservoir. Disconnect the upper pump seal fuel drain line from the adapter. Remove the screws and washers that attach the access cover to the reservoir. Remove the access cover from the airplane. Cut the safety wire and remove the drain valve, O-ring, washer and nut from the access cover. (a) Discard the O-ring. Install the special drain valve (Part Number 967B-5), new O-ring, washer and the nut in the access cover. Put the access cover (with new gasket) in its position on the reservoir. Use screws to attach the access cover to the reservoir. Install the pump seal drain line to the adapter. Apply sealant tape to the skin panel, remove the paper liner, and install skin panel to fuselage. Use screws to attach the skin panel to the fuselage. Use plug buttons to close out the antenna cable holes and fuel line hole in the fuselage skin. Refer to the applicable Chapter 34 transponder section for procedures to install the transponder antenna(s) on the bottom of the fuselage
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Preparation for Flight Without Cargo Pod Figure 203 (Sheet 1)
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Preparation for Flight Without Cargo Pod Figure 203 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (26) Install cockpit access panels 211EL, 231BL, 231DL, 232BR and 232DR. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (27) Clean the sealant from the fuselage along the cargo pod attach points. (28) Make a patch plate with a 1.28 inch diameter hole to accommodate the drain valve. (29) Attach the patch plate to the airplane with screws. 5.
Reinstalling the Cargo Pod After Flight Without the Cargo Pod NOTE: A.
This procedure must be used after the airplane has been returned to service without the cargo pod, and the cargo pod is being reinstalled.
Procedures For Reinstalling the Cargo Pod After Flight Without the Pod (Refer to Figure 203). (1) Turn both fuel tank selectors, found in the overhead console, to the OFF position. (2) Push the fuel valve lever up, and rotate it 90 degrees to drain the fuel from fuel reservoir. (3) Remove the patch plate from the underside of the fuselage. (4) Remove the fuselage skin panel from the underside of the fuselage. (5) Disconnect the pump seal fuel drain line from the adapter. (6) Remove the screws and washers that attach the access cover to the fuel reservoir. (7) Remove and discard the gasket from the access cover. (8) Remove the drain valve, O-ring and nut from the access cover. (a) Discard the O-ring. (9) Install the remotely-operated fuel drain valve, new O-ring and nut on the access cover. (10) Safety wire the remotely-operated fuel drain valve to the access cover. (11) Install the access cover to the fuel reservoir with the new O-ring. (12) Turn both of the fuel tank selectors, found in the overhead console, to the ON position and check for leaks. (13) Install the upper pump seal drain valve line to the remotely-operated fuel drain valve. (14) Apply sealant tape to the skin panel, remove the paper liner, and install skin panel to fuselage. (15) Install seal plug buttons in the fuselage skin drain holes as necessary. Refer to Table 201 for exact locations.
Table 201. Fuselage Plug Locations MODEL
FS LOCATION
BL LOCATION
QUANTITY
208 Only
115.88
0.00
1
208 and 208B
116.65
9.35
2
208 and 208B
164.85
8.60 and 19.60
4
208 Only
192.80
0.25 and 15.75
3
208 Only
201.20
0.00
1
208 Only
209.60
15.75
2
208B Only
212.80
0.25 and 15.75
3
208B Only
221.20
0.00
1
208B Only
257.60
15.75
2
208 Only
283.40
15.75
2
208 Only
285.05
14.25
2
208B Only
331.40
15.75
2
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MODEL 208 MAINTENANCE MANUAL (16) Apply Pro-Seal 890 to the surfaces that connect the outboard fuel drain line cover. NOTE:
On Airplanes 20800001 thru 20800085, the inboard and outboard fuel drain line covers were assembled as a single unit. On this fuel drain line cover, it will be necessary to connect the drain line and cable before you attach the fuel drain line cover to the fuselage.
(17) (18) (19) (20) (21) (22) (23) (24) (25)
Put the outboard fuel drain line cover in its position on the fuselage. Install the screws that attach the outboard fuel drain line cover to the fuselage. Connect the cable to the drain valve. Connect the drain line to the drain valve. Attach the cable to the fuselage skin with a clamp. Connect the lower pump seal drain line to the elbow. Apply sealant to the surface that connects the inboard fuel drain line cover to the fuselage. Install the screws that attach the inboard fuel drain line cover to the fuselage. Check the fuel drain valve knob for proper operation and correct installation as follows: (a) Pull the drain valve knob out and make sure that fuel flows through the fuel drain lines. (b) Push the drain valve knob in and make sure that no fuel flows through fuel drain lines (26) Refer to the applicable Chapter 34 transponder section for procedures to remove the transponder antenna(s) from the bottom of the fuselage. (27) Refer to the applicable Chapter 34 transponder section for procedures to install the transponder antenna(s) on the bottom of the pod. (28) Install the protective cover assemblies over the antenna cables. NOTE:
Seal the antenna holes in the fuselage with fuselage plugs.
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MODEL 208 MAINTENANCE MANUAL CARGO POD - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the cargo pod components in a serviceable condition.
Task 25-52-00-210 2.
Cargo Pod Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for a zonal inspection of the cargo pod. NOTE:
Each zonal inspection includes a GV/GVI to find the general condition and security of items included in the ZIP. The zonal inspections will be completed at a distance no more than an arms length. This includes an examination for signs of degradation such as corrosion, cracks, chafing of tubing, loose duct support, wiring damage, cable and pulley wear, fluid leaks, insufficient drainage, and for other conditions which could cause corrosion or damage.
B.
Special Tools (1) None
C.
Access (1) Open all cargo pod doors. (2) Remove the cargo pod upholstery.
D.
Do a Zonal Inspection of the Cargo Pod. (1) Examine the heat shield that is installed on the right forward side of the cargo pod for security and condition. (a) If there is heat damage to the front section of the pod, Refer to CAB89-30 (SK208-69). (2) Examine the condition of sealant between the pod and the fuselage. (a) If sealing or spot sealing is needed, use the approved sealant type. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (3) If installed, examine the TKS anti-ice fluid tank and components for condition and security of installation. (4) Thoroughly clean the interior pod area. (5) Examine the interior and the exterior structure for condition and security of installation, bulges in surface skin, cuts in the exterior or the interior skins, and blistered or pealed paint. (a) If you find a cut with the fiber showing on the interior or the exterior skin, you must do an immediate repair to prevent moisture contamination to the interior structure. Refer to Cargo Pod - Approved Repairs. (6) Examine all wire bundle assemblies and the electrical components for signs of overheating, correct installation, frayed or chafed wiring insulation, electrical bonding, damage, and corrosion. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. (7) Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment, and correct installation. (8) Examine all tubing, hose, and fluid fittings for signs of leaks, damage and chafing, and correct clamp installation. (9) Examine all placards and markings for security of installation, legibility, and correct location. For the correct placards and placard locations, refer to the Model 208 Illustrated Parts Catalog or the Pilot's Operating Handbook.
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MODEL 208 MAINTENANCE MANUAL (10) Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
(11) Examine the pod drain holes for obstructions. (12) Examine the rubber door seals for condition and correct attachment. (13) Examine the cargo pod door hinges and the latches for signs of damage, wear, security, and loose or failed fasteners.. (a) Make sure that the latches operate correctly. (14) Examine the door structure for cracks, delamination, and general condition. (a) If you think there is damage or delamination, Refer to Cargo Pod - Approved Repairs. E.
Restore Access (1) Install the cargo pod upholstery. (2) Close all cargo pod doors. End of task Task 25-52-00-710 3.
Cargo Pod Drains Operational Check A.
General (1) This task gives the information needed to do an operational check of the cargo pod drains.
B.
Special Tools (1) Air Compressor (2) Approved Container
C.
Access (1) Open the cargo pod doors. Refer to Cargo Pod - Maintenance Practices.
D.
Do the Cargo Pod Drains Operational Check. (1) Examine the fuselage drain system for condition, security and evidence of water leaks. Inspect drain tube outlets for obstructions. (2) Examine the pod drain holes for obstructions. (3) Thoroughly clean the interior pod area. (4) Use shop air to blow any obstructions from the cargo pod drain holes. (5) Put an approved container under the cargo pod drain hole and pour water through drain hole to make sure there is correct drainage.
E.
Restore Access (1) Close the cargo pod doors. Refer to Cargo Pod - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL CARGO POD - APPROVED REPAIRS 1.
General A.
2.
Tools, Equipment and Materials A.
3.
The following procedure is to be used for repairing delamination of outer or inner wall skin or both. The bonding action of the replacement core (Syntatic foam) will repair the delamination of the opposite wall skin.
Refer to Equipment/Furnishings - General for a list of required tools, equipment and materials.
Repair Procedures
CAUTION: A mask, gloves and eye protection shall be worn when cutting, grinding, drilling or sanding composite materials. Repair should be made in an open area, away from other workers. A.
Refer to Figure 801 for illustrated repair steps.
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Cargo Pod Repair Figure 801 (Sheet 1)
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Cargo Pod Repair Figure 801 (Sheet 2)
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Cargo Pod Repair Figure 801 (Sheet 3)
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Cargo Pod Repair Figure 801 (Sheet 4)
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Cargo Pod Repair Figure 801 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL CARGO POD HEAT SHIELDS - MAINTENANCE PRACTICES 1.
General A.
2.
A heat shield is installed on the cargo pod on the forward right side. This shield is grounded by a grounding strap inside the cargo pod. Dependent on serialization, the shield is attached by bonding, or by bonding and screws.
Shield Removal/Installation A.
Remove Shield (Refer to Figure 201). (1) Remove screws (if installed) from heat shield. (2) Carefully break adhesive bond between heat shield and cargo pod.
B.
Install Shield (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5)
A vacuum bag must be fabricated to assist in heat shield bonding. Refer to Cargo Pod Approved Repairs, Figure 801, for vacuum bag fabrication information.
Clean mating surfaces of heat shield and cargo pod. Apply Class 1C adhesive to mating surface of heat shield and cargo pod. Refer to Chapter 20, Adhesive and Solvent Bonding - Maintenance Practices. Secure heat shield to cargo pod using screws. Apply a surface vacuum to the heat shield area for a minimum of 24 hours. Vacuum should be capable of a minimum of 20.0 inches mercury. After adhesive has cured, remove vacuum bag.
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Cargo Pod Heat Shield Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL LOADING ZONE WEIGHT LIMITS - CLEANING/PAINTING 1.
General A.
The cargo compartment of the Model 208 and 208B is divided into six loading zones. Painted on the sidewalls of each zone is the weight limit for that zone. Refer to Figure 701 for painting/repainting procedures and details.
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Loading Zone Limits - Painting Instructions Figure 701 (Sheet 1)
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Loading Zone Limits - Painting Instructions Figure 701 (Sheet 2)
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Loading Zone Limits - Painting Instructions Figure 701 (Sheet 3)
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Loading Zone Limits - Painting Instructions Figure 701 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL EMERGENCY EQUIPMENT - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the emergency equipment in a serviceable condition. The procedures for the emergency locator system are in accordance with 14 CFR 91.207.
Task 25-60-00-720 2.
Artex C406-2 Emergency Locator Transmitter (ELT) Functional Check A.
General (1) This task gives the procedures to do a functional check of the Artex C406-2 Emergency Locator Transmitter (ELT).
B.
Special Tools (1) 50 Ohm Dummy Load (2) Amplitude Modulation (AM) Receiver (3) Attenuator (30 dB) (4) SARSAT Tester
C.
Access (1) Open access panel 340A on the right side of the vertical stabilizer. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do an inspection of the ELT, mounting tray, antenna, and the ELT battery for condition and correct installation. (1) Make sure that the ELT switch, found on the forward end of the ELT, is set to the OFF position. (2) Remove the ELT from the mounting tray. Refer to ARTEX C406-2 Emergency Locator Transmitter System - Maintenance Practices.
CAUTION: Do not use solvents to clean the ELT, mounting tray, or electrical contacts. Solvents used in these areas can cause damage to the ELT housing. (3) (4) (5)
Examine the ELT and the mounting tray for correct installation, cleanliness, cracks, or other damage. Examine the ELT battery for corrosion. Look at the battery expiration date. (a) Make sure that the battery life limit is not expired. (b) Make sure that the battery expiration date is shown correctly in the maintenance records. NOTE:
The battery manufacturer puts a mark on the battery to show the battery life limit. When you install a new battery in an ELT, make sure that you make a record of the expiration date in the space given on the ELT name and data plate.
(c) (d)
(6)
If you have to replace the ELT battery, refer to ARTEX Maintenance Manual 570-5000. You must replace the ELT battery with a new battery if one or more of the conditions that follow occur: • Use of the ELT battery in an emergency • Operation for an unknown amount of time • Use for more than one hour of cumulative time • Replace the battery if the voltage under load is less than 12.0 vdc. • Replacement date shown on the battery label has expired or will expire before the next scheduled inspection. (e) Record the new battery expiration date in the maintenance log if you replaced it. Examine the ELT antenna for correct installation and cracks or other damage.
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MODEL 208 MAINTENANCE MANUAL E.
Do a G-Switch Test. NOTE:
If possible, do the test procedure for the emergency locator transmitter inside a metal hangar with the doors closed to decrease the signal transmission from the ELT unit during the test.
CAUTION: Operate the emergency locator transmitter system only during the first five minutes of each hour. If the functional test must be completed at a time other than the first five minutes of the hour, the nearest FAA tower or Flight Service Station must be told of the test in accordance with FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the emergency locator transmitter for more than 15 seconds at a time. During the first 15 seconds of transmission, the satellite system will receive the 406.025 MHz signal as a test signal. If the signal continues for more than 15 seconds, the signal will be identified as a distress signal. (1)
Install a jumper wire between pins 5 and 8 on the electrical connector of the ELT. Refer to the ARTEX web site as necessary.
CAUTION: Do this procedure with an experienced technician because of the potential physical damage that can occur if the jumper wire is not installed correctly. NOTE:
(2) (3) (4) (5) (6) (7) (8) F.
The ELT will not activate with the G-switch unless electrical pins 5 and 8 have a jumper wire installed between them (this happens automatically when the ELT is locked into the mount tray with the electrical connector in position).
Make sure the ELT switch is in the OFF position. Use a amplitude modulation (AM) and set it to 121.5 MHz to listen for the aural warning sweep tone. Hold the ELT transmitter tightly in one hand and make a throwing movement, then an opposite movement of the ELT transmitter. Make sure that the G-switch operates and that the aural warning sweep tone is heard on the AM receiver that is set to 121.5 MHz. Set the ELT switch to the ON position and then back to the OFF position to reset the G-switch. Remove the jumper wire from electrical pins 5 and 8 on the electrical connector of the ELT. Install the emergency locator transmitter in the airplane. Refer to ARTEX C406-2 Emergency Locator Transmitter System - Maintenance Practices.
Put the Airplane in the Test Configuration. NOTE: (1) (2) (3) (4) (5) (6) (7)
The ELT antenna is disconnected and the 50 ohm dummy load is installed to the airplane coax or to the ELT with coax as necessary.
Examine the ELT battery to make sure that it is not due for replacement. (a) If the battery must be replaced, follow the manufacturers instructions to replace it. Do the functional test with a 50 ohm dummy load as an alternative to the antenna. Connect the GPU to the external jack. Turn the GPU on and adjust to +28. +0.25 or -0.25 V. Make sure that the COM 1, COM 2, AUDIO AMP , and GPS circuit breakers are engaged. Put the battery switch in the BATT position. Put the avionics switch in the ON position. Make sure that the GPS position has been initialized on FMS.
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CAUTION: Operate the emergency locator transmitter system only during the first five minutes of each hour. If the functional test must be completed at a time other than the first five minutes of the hour, the nearest FAA tower or Flight Service Station must be told of the test in accordance with FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the emergency locator transmitter for more than 15 seconds at a time. During the first 15 seconds of transmission, the satellite system will receive the 406.025 MHz signal as a test signal. If the signal continues for more than 15 seconds, the signal will be identified as a distress signal. G.
Transmitter Test of the ARTEX C406-2 Emergency Locator Transmitter (ELT) System. (1) Pull the COM 1 and SPKR knobs on the audio control panel. (2) Adjust the volume to make sure that the transmissions from the radio are heard in the cockpit. (3) Adjust the COM 1 frequency to 121.50 MHz. Make sure that audio is heard through the cockpit speakers.
CAUTION: Operate the emergency locator transmitter system only during the first five minutes of each hour. If the functional test must be completed at a time other than the first five minutes of the hour, the nearest FAA tower or Flight Service Station must be told of the test in accordance with FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the emergency locator transmitter for more than 15 seconds at a time. The ELT must be OFF for at least 60 seconds between each 15-second activation cycle. During the first 15 seconds of transmission, the satellite system will receive the 406.025 MHz signal as a test signal. If the signal continues for more than 15 seconds, the signal will be identified as a distress signal. (4) (5) (6) (7) (8)
H.
Put the ELT remote switch (SZ09) in the ON position for approximately 1 second . Make sure that the ELT signal is heard through the cockpit speakers and that the LED adjacent to the switch is on. Immediately put the ELT remote switch in the ARM position. Make sure that the LED stays on for approximately 1 second before it is turned off. If the ELT system has sensed a fault in the system, the LED will flash a fault code at this time. (a) For information on the possible codes, refer to the Installation and Operation Manual for the Artex C406-2 system.
Do a NAV Interface Test. (1) If necessary put the airplane in the test configuration. Refer to Put the Airplane in the Test Configuration in this section. (2) After a minimum of 60 seconds, engage the ELT/NAV circuit breaker circuit breaker on the cockpit circuit breaker panel. (3) Use a 30-dB load to connect the SARSAT Tester to the 406.025 MHz coax cable at the base of the antenna. NOTE: (4) (5)
The SARSAT test set is held no more than six inches away from the antenna.
Turn on the SARSAT tester. Engage the receive function of the SARSAT tester. (a) Make sure that the display on the tester shows that it is searching for a signal.
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CAUTION: Operate the emergency locator transmitter system only during the first five minutes of each hour. If the functional test must be completed at a time other than the first five minutes of the hour, the nearest FAA tower or Flight Service Station must be told of the test in accordance with FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the emergency locator transmitter for more than 15 seconds at a time. Also, the ELT must be off for at least 60 seconds between each 15 second activation cycle. During the first 15 seconds of transmission, the satellite system will receive the 406.025 MHz signal as a test signal. If the signal continues for more than 15 seconds, the signal will be identified as a distress signal. (6) (7) (8) (9)
Put the ELT remote switch (SZ09), on the right switch panel, in the ON position. Within 15 seconds put the ELT remote switch in the ARM position. Monitor the SARSAT tester to see if it received a signal from the ELT system. (a) If a signal was not received, the cycle can be completed again after the 60-second off cycle. Make sure that the tail number shown on the SARSAT tester is correct. NOTE:
When ownership of an aircraft is transferred within the same country, the C406-2 ELT must be registered with the applicable authority. When an aircraft with a C406-2 ELT changes tail number or country registration, the ELT must have the new identification data entered. The ELT must be registered with the applicable authority.
(10) Make sure that the Mode S code shown on the SARSAT tester is the same as the number found on the back of the transmitter. (11) Make sure that the latitude and longitude information is the same as that shown on the FMS display. (12) Turn the SARSAT tester off. (13) Disconnect the 30 dB load and SARSAT tester from the 406.025 MHz antenna cable. (14) Connect the coaxial cable to the antenna. (15) If no other tests are necessary, do the steps that follow: (a) Put the avionics switch in the OFF position. (b) Put the battery switch in the OFF position. (c) Remove the external electrical power from the airplane. I.
Restore Access (1) Close access panel 340A on the right side of the vertical stabilizer. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task Task 25-60-00-721 3.
Artex ME406 Emergency Locator Transmitter (ELT) Functional Check A.
General (1) This task gives the procedures to do a functional check of the Artex ME406 Emergency Locator Transmitter (ELT).
B.
Special Tools (1) 50 Ohm Dummy Load (2) Amplitude Modulation (AM) Receiver (3) Attenuator (30 dB) (4) SARSAT Tester
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Access (1) Open access panel 340A on the right side of the vertical stabilizer. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do an inspection of the ELT, mounting tray, antenna, and the ELT battery for condition and correct installation. (1) Make sure that the ELT switch, found on the forward end of the ELT, is set to the ARM position. (2) Remove the ELT from the mounting tray. Refer to Artex ME406 Emergency Locator Transmitter System - Maintenance Practices.
CAUTION: Do not use solvents to clean the ELT, mounting tray, or electrical contacts. Solvents used in these areas can cause damage to the ELT housing. (3) (4) (5)
Examine the ELT and the mounting tray for correct installation, cleanliness, cracks, or other damage. Examine the ELT battery for corrosion. Look at the battery expiration date. (a) Make sure that the battery life limit is not expired. (b) Make sure that the battery expiration date is shown correctly in the Maintenance Records. NOTE:
The battery manufacturer puts a mark on the battery to show the battery life limit. When you install a new battery in an ELT, make sure a record of the expiration date is put in the space given on the ELT name and data plate.
(c) (d)
(6) E.
If you have to replace the ELT battery, refer to Artex Maintenance Manual 570-1600. You must replace the ELT battery with a new battery if one or more of the conditions that follow occur: • Use of the ELT battery in an emergency • Operation for an unknown amount of time • Use for more than one hour of cumulative time • Replacement date shown on the battery label has expired or will expire before the next scheduled inspection. (e) Record the new battery expiration date in the maintenance log if you replaced it. Examine the ELT antenna for correct installation and cracks or other damage.
Do a G-Switch Operational Test. NOTE:
If possible, do the test procedure for the emergency locator transmitter inside a metal hangar with the doors closed to decrease the signal transmission from the ELT unit during the test.
CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the first five minutes of each hour. If you must complete the functional test at a time other than the first five minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than five seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz distress signal after it is activated for approximately 50 seconds. (1)
Install a jumper wire between pins 5 and 12 on the electrical connector of the ELT.
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CAUTION: It is recommended that an experienced technician do this procedure because of the potential physical damage that can occur if the jumper wire is not installed correctly. NOTE:
(2) (3) (4) (5) (6) (7) (8) F.
The ELT will not activate with the G-switch unless electrical pins 5 and 12 have a jumper wire installed between them (this happens automatically when the ELT is locked into the mount tray with the electrical connector in position).
Make sure the ELT switch is in the ARM position. Use a amplitude modulation (AM) receiver and set it to 121.5 MHz to listen for the aural warning sweep tone. Hold the ELT transmitter tightly in one hand and make a throwing movement followed by an opposite movement of the ELT transmitter. Make sure that the G-switch operates and that the aural warning sweep tone is heard on the AM receiver set to 121.5 MHz. Set the ELT switch to the ON position and then back to the ARM position to reset the G-switch. Remove the jumper wire from electrical pins 5 and 12 on the electrical connector of the ELT. Install the emergency locator transmitter in the airplane. Refer to Artex ME406 Emergency Locator Transmitter System - Maintenance Practices.
Transmitter Test of the Artex ME406 Emergency Locator Transmitter (ELT) System.
CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the first five minutes of each hour. If you must complete the functional test at a time other than the first five minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than five seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz distress signal after it is activated for approximately 50 seconds. (1) (2) (3) (4) (5) (6) (7) (8)
Make sure the BATTERY switch and the AVIONICS switches are in the OFF position. Connect external electrical power to the airplane. Make sure that the COM/NAV 1 and AUD/MKR circuit breakers on the circuit breaker panel are engaged. Set the BATTERY switch to the ON position. Set the AVIONICS switches to the ON position. Make sure that the ELT remote switch on the right panel is in the ARM position. Set one of the communication units to receive a frequency of 121.5 MHz. Set the communication unit to the airplane speakers at an audio level loud enough to be heard. NOTE:
The SARSAT tester is used as an example to gather test information. However, other equivalent test equipment such as the Aeroflex IFR 4000 Communications Test Set is acceptable.
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MODEL 208 MAINTENANCE MANUAL (9)
Have another person use the SARSAT tester set to the RECV function. Refer to Figure 601. NOTE:
The SARSAT tester must be less than 15 feet from the ELT antenna and must have a line-of-sight between the ELT antenna and SARSAT tester.
NOTE:
The person with the SARSAT tester must make sure that the ELT buzzer is heard during the test.
NOTE:
If it is necessary to do the transmitter test after the first five minutes of the hour, the SARSAT tester is connected directly to the ELT with a coaxial cable and a 30 dB attenuator. You will not hear the sweep tone from the ELT on the airplane speakers with the attenuator installed.
(10) Install the 30 dB attenuator between the ELT and SARSAT tester if necessary. (11) Set the ELT remote switch on the right panel to the ON position. (12) Let the ELT make three sweeps on the airplane speakers. NOTE:
This will take one second. The ELT remote switch will start to flash.
(13) Set the ELT remote switch back to the ARM position and monitor the LED. NOTE:
The ELT will do a self-test. The LED will stay on for one second and the ELT sweeps are not audible on the airplane speakers if the ELT operation is normal.
NOTE:
The ELT does not transmit a 406.028 MHz test signal to the SARSAT tester until the ELT remote switch is set back to the ARM position.
(14) If the LED continues to flash, refer to Artex ME406 Emergency Locator Transmitter System Troubleshooting. (15) If the SARSAT tester did not receive a 406.028 MHz signal and the ELT remote switch LED does not show a transmitter problem, do the test again. (16) When the SARSAT tester receives a 406.028 MHz signal, scroll the pages on the tester, as necessary, and make sure of the information that follows: (a) Make sure the information shown by the SARSAT tester agrees with the placard on the ELT. NOTE:
(b)
The information that follows must match the data on the ELT placard:
• COUNTRY code • 15-digit Hex code ID • Aircraft identification number. Make sure that the SARSAT tester shows the messages that follow: • S' TEST OK • Frequency - PASS • Homing frequency • Message format (short).
NOTE:
When ownership of an aircraft is transferred within the same country, the ME406 ELT should be registered with the applicable authority. When an aircraft with a ME406 ELT changes tail number or country registration, the ELT will need to have the new identification data entered. The ELT will also need to be registered with the applicable authority.
G.
Restore Access (1) Close access panel 340A. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. End of task
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Artex ME406 Emergency Locator Transmitter (ELT) SARSAT Test Set-up Figure 601 (Sheet 1)
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Artex ME406 Emergency Locator Transmitter (ELT) SARSAT Test Set-up Figure 601 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL Task 25-60-00-722 4.
ARTEX C406-N Emergency Locator Transmitter (ELT) Functional Check A.
General (1) This task gives the procedures to do a functional check of the Artex C406-N Emergency Locator Transmitter (ELT).
B.
Special Tools (1) 50 Ohm Dummy Load (2) Amplitude Modulation (AM) Receiver (3) Attenuator (30 dB) (4) SARSAT Tester
C.
Access (1) Open access panel 340A on the right side of the vertical stabilizer. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do an inspection of the ELT, mounting tray, antenna, and the ELT battery for condition and correct installation. (1) Make sure that the ELT switch, found on the forward end of the ELT, is set to the OFF position. (2) Remove the ELT from the mounting tray. Refer to ARTEX C406-N Emergency Locator Transmitter System - Maintenance Practices.
CAUTION: Do not use solvents to clean the ELT, mounting tray, or electrical contacts. Solvents used in these areas can cause damage to the ELT housing. (3) (4) (5)
Examine the ELT and the mounting tray for correct installation, cleanliness, cracks, or other damage. Examine the ELT battery for corrosion. Look at the battery expiration date. (a) Make sure that the battery life limit is not expired. (b) Make sure that the battery expiration date is shown correctly in the maintenance records. NOTE:
The battery manufacturer puts a mark on the battery to show the battery life limit. When you install a new battery in an ELT, make sure that you make a record of the expiration date in the space given on the ELT name and data plate.
If you have to replace the ELT battery, refer to ARTEX Maintenance Manual 570-5060. You must replace the ELT battery with a new battery if one or more of the conditions that follow occur: • Use of the ELT battery in an emergency • Operation for an unknown amount of time • Use for more than one hour of cumulative time • Replace the battery if the voltage under load is less than 12.0 vdc. • Replacement date shown on the battery label has expired or will expire before the next scheduled inspection. (e) Record the new battery expiration date in the maintenance log if you replaced it. Examine the ELT antenna for correct installation and cracks or other damage. (c) (d)
(6) E.
Do a G-Switch Operational Test. NOTE:
If possible, do the test procedure for the emergency locator transmitter inside a metal hangar with the doors closed to decrease the signal transmission from the ELT unit during the test.
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CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the first five minutes of each hour. If you must complete the functional test at a time other than the first five minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular 91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than five seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz signal for 520 milliseconds approximately every 50 seconds. This transmission is an encoded digital message and is sent to a satellite as a distress signal. (1)
Install a jumper wire between pins 12 and 13 on the electrical connector of the ELT.
CAUTION: Do this procedure with an experienced technician because of the potential physical damage that can occur if the jumper wire is not installed correctly. NOTE:
(2) (3) (4) (5) (6) (7) (8) F.
The ELT will not activate with the G-switch unless electrical pins 12 and 13 have a jumper wire installed between them (this happens automatically when the ELT is locked into the mount tray with the electrical connector in position).
Make sure the ELT switch is in the OFF position. Use a amplitude modulation (AM) receiver and set it to 121.5 MHz to listen for the aural warning sweep tone. Hold the ELT transmitter tightly in one hand and make a throwing movement, then an opposite movement of the ELT transmitter. Make sure that the G-switch operates and that the aural warning sweep tone is heard on the AM receiver set to 121.5 MHz. Set the ELT switch to the ON position and then back to the OFF position to reset the G-switch. Remove the jumper wire from electrical pins 12 and 13 on the electrical connector of the ELT. Install the emergency locator transmitter in the airplane. Refer to ARTEX C406-N Emergency Locator Transmitter System - Maintenance Practices.
Transmitter Test of the ARTEX C406-N Emergency Locator Transmitter (ELT) System.
CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the first five minutes of each hour. If you must complete the functional test at a time other than the first five minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular 91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than five seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz signal for 520 milliseconds approximately every 50 seconds. This transmission is an encoded digital message and is sent to a satellite as a distress signal. (1) (2) (3) (4) (5)
Make sure that the BATTERY switch and the AVIONICS switches are in the OFF position. Connect external electrical power to the airplane. Make sure that the COM/NAV 1 and AUD/MKR circuit breakers on the circuit breaker panel are engaged. Set the BATTERY switch to the ON position. Set the AVIONICS switches to the ON position.
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MODEL 208 MAINTENANCE MANUAL (6) (7) (8)
Make sure that the ELT remote switch on the right panel is in the ARM position. Set one of the communication units to receive a frequency of 121.5 MHz. Set the communication unit to the airplane speakers at an audio level that will be heard. NOTE:
(9)
The SARSAT (Search And Rescue Satellite Aided Tracking) tester is used as an example to gather test information. Refer to Artex website for more information on testing equipment. Generally, the testing is completed with the Artex Handheld Programmer 453-1000 for all 406 Mhz ELTs. However, other equivalent test equipment such as the Aeroflex IFR 4000 Communications Test Set is acceptable.
Another person must use the SARSAT tester set to the RECV function. Refer to Figure 602. NOTE:
The SARSAT tester must be less than 15 feet from the ELT antenna and must have a line-of-sight between the ELT antenna and SARSAT tester.
NOTE:
The person with the SARSAT tester must make sure that the ELT buzzer is heard during the test.
NOTE:
If it is necessary to do the transmitter test after the first five minutes of the hour, connect the SARSAT tester directly to the ELT with a coaxial cable and a 30 dB attenuator. You will not hear the sweep tone from the ELT on the airplane speakers with the attenuator installed.
(10) Install the 30 dB attenuator between the ELT and SARSAT tester if necessary. (11) Set the ELT remote switch on the right panel to the ON position. (12) Let the ELT make three sweeps on the airplane speakers. NOTE:
This will take one second. The ELT remote switch will start to flash.
(13) Set the ELT remote switch back to the ARM position and monitor the LED. NOTE:
The ELT will do a self-test. The LED will stay on for one second and the ELT sweeps are not audible on the airplane speakers if the ELT operation is normal.
NOTE:
The ELT does not transmit a 406.028 MHz test signal to the SARSAT tester until the ELT remote switch is set back to the OFF position.
(14) If the LED continues to flash, refer to ARTEX C406-N Emergency Locator Transmitter System - Troubleshooting. (15) If the SARSAT tester did not receive a 406.028 MHz signal and the ELT remote switch LED does not show a transmitter problem, do the test again. (16) When the SARSAT tester receives a 406.028 MHz signal, scroll the pages on the tester, as necessary, and make sure of the following: (a) Make sure that the information shown by the SARSAT tester agrees with the placard on the ELT. NOTE:
(b)
The information that follows must match the data on the ELT placard:
• COUNTRY code • 15-digit Hex code ID • Aircraft identification number. Make sure that the SARSAT tester shows the messages that follow: • S' TEST OK • Frequency - PASS • Homing frequency
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ARTEX C406-N Emergency Locator Transmitter (ELT) SARSAT Test Set-up Figure 602 (Sheet 1)
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ARTEX C406-N Emergency Locator Transmitter (ELT) SARSAT Test Set-up Figure 602 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL • NOTE:
Message format (short). When ownership of an aircraft is transferred within the same country, the C406-N ELT must be registered with the applicable authority. When an aircraft with a C406-N ELT changes tail number or country registration, the ELT must have the new identification data entered. The ELT must be registered with the applicable authority.
G.
Restore Access (1) Close access panel 340A. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. End of task Task 25-60-00-723 5.
Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 Emergency Locator Transmitter (ELT) Functional Check A.
General (1) This task gives the procedures to do a functional check of the Dorne and Margolin, Pointer 30001, and Pointer 3000-11 Emergency Locator Transmitter (ELT).
B.
Special Tools (1) Amplitude Modulation (AM) Receiver
C.
Access (1) Remove the aft cabin partition or unzip the canvas wall to get access to the ELT. Refer to Rear Cargo Compartment Wall - Maintenance Practices.
D.
Do an inspection of the ELT, mounting tray, antenna, and the ELT battery for condition and correct installation. (1) Make sure that the ELT master switch, found on the forward end of the ELT, is set to the OFF position. (2) Remove the ELT from the mounting tray. Refer to Emergency Locator Transmitter System Maintenance Practices.
CAUTION: Do not use solvents to clean the ELT, mounting tray, or electrical contacts. Solvents used in these areas can cause damage to the ELT housing. (3) (4) (5)
Examine the ELT and the mounting tray for correct installation, cleanliness, cracks, or other damage. Examine the ELT battery for corrosion. Look at the battery expiration date. (a) Make sure that the battery life limit is not expired. (b) Make sure that the battery expiration date is shown correctly in the maintenance records. NOTE:
(c) (d)
The battery manufacturer puts a mark on the battery to show the battery life limit. When you install a new battery in an ELT, make sure that you make a record of the expiration date in the space given on the ELT switch nameplate on the side of unit, and on the instruction nameplate on the top of unit.
If it is necessary to replace the ELT battery, refer to Emergency Locator Transmitter System - Maintenance Practices. You must replace the ELT battery with a new battery if one or more of the conditions that follow occur: • The ELT battery is used in an emergency • Operation for an unknown amount of time • The ELT battery is used for more than one hour of cumulative time
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(6)
The voltage under load is less than 12.0 Vdc The replacement date shown on the battery label has expired or will expire before the next scheduled inspection. (e) Record the new battery expiration date in the maintenance log if you replaced it. Examine the ELT antenna for correct installation, cracks, or other damage.
E.
Do a G Switch Operational Check (Dorne and Margolin/Pointer 3000-1 Series). (1) While you hold the transmitter in one hand, sharply strike the end of the case in the direction of activation shown on the transmitter case. (a) Make sure that the G switch has been actuated. (2) Reset the G switch.
F.
Do a G Switch Operation Check (Pointer 3000-11 Series). (1) Hold transmitter firmly in one hand and make a throwing motion followed by a sudden reversal of the transmitter. (a) Make sure that the G switch has been actuated. (2) Reset the G switch.
G.
Do an Operational Check of the Radiated Signal with Local Monitoring. (1) Install the ELT in the airplane. Refer to Emergency Locator Transmitter System - Maintenance Practices. (2) Make sure that test is performed within five minutes before or after the hour. (3) Put a small, hand held AM radio tuned to any frequency, within six inches of the emergency locator transmitter antenna. NOTE: (4)
Use of the airplanes’s VLF receiver or the ADF will not do a sufficient power check of the radiated signal.
For Pointer 3000-1 ELT's, disconnect the remote connector from the ELT.
CAUTION: For Pointer 3000-1 ELT's, the remote connector must be disconnected from the ELT before you do maintenance. If the remote connector in not disconnected, it could cause the ELT’s internal fuse to blow. (5)
Put the master switch to the ON position to activate the emergency locator transmitter system. NOTE:
(6)
(7)
Activate the emergency locator transmitter system for no more than three sweeps of the audio signal. (a) Make sure that the signal has been detected on the AM radio. (b) If the ELT does not operate correctly during the functional check, remove the transmitter and return it to an authorized avionics repair shop for inspection and repair. Refer to Emergency Locator Transmitter System - Maintenance Practices. Restore the master switch to the AUTO position. NOTE:
(8) (9)
On Pointer 3000-11 series system, the transmitter can be activated from the cockpit by placing the remote mounted switch to ON position. Other transmitters must be activated from the tailcone mounting area.
On Pointer 3000-11 series system, momentarily place the remote mounted switch to the RESET position then release it. This will put the transmitter in the AUTO position.
For Pointer 3000-1 ELT's, connect the remote connector to the ELT. Make an entry in the Airplane Logbook to show that the test has been completed.
H.
Restore Access (1) Install aft cabin partition or zip the canvas wall. Refer to Rear Cargo Compartment Wall Maintenance Practices End of task
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MODEL 208 MAINTENANCE MANUAL Task 25-60-00-960 6.
Emergency Locator Transmitter Battery Discard A.
General (1) This section gives the information needed to complete the discard procedures for the emergency locator transmitter battery.
B.
Tools and Equipment (1) None
C.
Access (1) For Artex series ELT's, open access panel 340A on the right side of the vertical stabilizer to get access to the transmitter. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. (2) For Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 series ELT's, remove the aft cabin partition or unzip the canvas wall to get access to the transmitter. Refer to Rear Cargo Compartment Wall - Maintenance Practices.
D.
Discard of the Emergency Locator Transmitter Battery. (1) Remove emergency locator transmitter (ELT) from the airplane. (a) For the Artex C406-2 ELT, refer to ARTEX C406-2 Emergency Locator Transmitter System - Maintenance Practices. (b) For the Artex ME406 ELT, refer to Artex ME406 Emergency Locator Transmitter System Maintenance Practices. (c) For the Artex C406-N ELT, refer to ARTEX C406-N Emergency Locator Transmitter System - Maintenance Practices. (d) For Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 series ELT's, refer to Emergency Locator Transmitter System - Maintenance Practices. (2) Remove the battery pack from the ELT. (a) For Artex series ELT's, refer to the applicable Artex 406MHz Emergency Locator Transmitters Description, Operation, Installation and Maintenance Manual. Refer to the Introduction, List of Publications. (b) For Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 series ELT's refer to Emergency Locator Transmitter System - Maintenance Practices. (3) Install a new battery pack in the ELT. (a) For Artex series ELT's,refer to the applicable Artex 406MHz Emergency Locator Transmitters Description, Operation, Installation and Maintenance Manual. Refer to the Introduction, List of Publications. (b) For Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 series refer to Emergency Locator Transmitter System - Maintenance Practices. (4) Install the emergency locator transmitter in the airplane. (a) For the Artex C406-2 ELT, refer to ARTEX C406-2 Emergency Locator Transmitter System - Maintenance Practices. (b) For the Artex ME406 ELT, refer to Artex ME406 Emergency Locator Transmitter System Maintenance Practices. (c) For the Artex C406-N ELT, refer to ARTEX C406-N Emergency Locator Transmitter System - Maintenance Practices. (d) For Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 series ELT's, refer toEmergency Locator Transmitter System - Maintenance Practices. (5) Make an entry in the Airplane Logbook to show the battery pack replacement date and the next expiration date. The new battery expiration date can be put on a label on the face of the transmitter.
E.
Restore Access (1) For Artex series ELT's, close access panel 340A. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) For Dorne and Margolin, Pointer 3000-1, and Pointer 3000-11 series ELT's, install the aft cabin partition or zip the canvas wall. Refer to Rear Cargo Compartment Wall - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL EMERGENCY LOCATOR TRANSMITTER - DESCRIPTION AND OPERATION 1.
General A.
2.
3.
This section describes emergency locator transmitters used on Model 208 and 208B airplanes.
Description A.
Airplanes may be equipped with one of three different emergency locator transmitter systems. The Pointer 3000-11 is a replacement for both the Dorne and Margolin and Pointer 3000-1 systems, and meets TSO-C91A requirements. (1) Dorne and Margolin - Airplanes 20800001 Thru 20800127 and 208B0001 Thru 208B0078 were equipped with a Dorne and Margolin system. This system may be identified by a bright orange unit mounted in the tailcone. (2) Pointer Model 3000-1 - Airplanes 20800128 Thru 20800242 and 208B0079 Thru 208B0448 are equipped with a Pointer 3000-1 system. This unit may be identified as a black-on-grey unit, also mounted in the tailcone. (3) Pointer Model 3000-11 - Airplanes 20800243 and On and 208B0449 and On are equipped with a Pointer 3000-11 system.
B.
All transmitters are designed to provide a broadcast tone that is audio-modulated in a swept manner that is a distinct, easily recognizable distress signal for reception by search and rescue personnel and others monitoring the emergency frequencies. All units transmit an omni-directional signal on the international distress frequencies or 121.5 and 243.0 MHz simultaneously. General aviation and commercial airplanes, the FAA and CAP monitor 121.5 MHz, and 243.0 MHz is monitored by the military.
Operation A.
Dorne and Margolin system transmits after the unit has received a 5g (tolerances are +2g and -0g) impact force for a duration of 11 to 16 milliseconds. Power is supplied to the transmitter by an alkaline battery pack which enables transmitter to transmit on both frequencies at 75mw rated power output for 48 continuous hours in the temperature range of -4°F to +131°F. Transmitter exhibits line-of-sight transmission characteristics which correspond approximately to 100 miles at a search altitude of 10,000 feet. Alkaline battery pack has replacement date/and date of installation on top of transmitter.
B.
Pointer Model 3000-1 system transmits after unit has received a 5g (tolerances are +2g and -0g) impact force for a duration of 11 to 16 milliseconds. Power is supplied to transmitter by a battery pack consisting of four 1.4V magnesium "D" cell batteries in series. System transmits continuously on both distress frequencies simultaneously at 75mw rated power output between 7.5 hours at approximately -40°F and up to 150 hours at approximately +50°F. System will provide line-of-sight transmission up to 100 miles, depending on search aircraft altitude, weather and topography. Magnesium battery pack replacement date is marked in space on label at the end of the unit.
C.
Both the Dorne and Margolin and Pointer 3000-1 systems have a three-position switch on forward end of unit which controls operation. Placing switch in the ON position will energize the unit to start transmitting emergency signals. In OFF position, unit is inoperative. Placing switch in AUTO position will set unit to start transmitting emergency signals only after unit has received a 5g (tolerances are +2g and -0g) impact force for a duration of 11 to 16 milliseconds. Pointer 3000 Model also incorporates a transmitter annunciator light that illuminates red to indicate the transmitter is transmitting a distress signal. In addition, it also incorporates a G switch reset button that, when pushed in, will reset the inertia G switch to the OFF position.
D.
Pointer 3000-11 system is automatically activated by a deceleration sensing inertia switch. Inertia switch is designed to activate when unit senses longitudinal inertia forces as required in TSO-C91A. Power is supplied to the transmitter by a battery pack consisting of five 1.5 VDC alkaline “C” cell batteries in an impact resistant fabricated foam housing. Unit transmits continuously on both distress frequencies simultaneously. Alkaline battery pack replacement date is marked on battery pack and on label at end of unit.
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Pointer 3000-11 system incorporates a master ON-OFF-AUTO switch on unit, and a remote mounted ON-AUTO-RESET control switch. This switch is mounted to instrument panel and allows for remote checks of system without directly accessing the transmitter.
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MODEL 208 MAINTENANCE MANUAL EMERGENCY LOCATOR TRANSMITTER - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been prepared to aid maintenance technicians in system understanding. Refer to Figure 101.
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Emergency Locator Transmitter Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL EMERGENCY LOCATOR TRANSMITTER - MAINTENANCE PRACTICES 1.
General A.
2.
This section provides maintenance practices for the Dorne and Margolin, Pointer 3000-1 and Pointer 3000-11 series emergency locator transmitters.
Transmitter Removal/Installation A.
Remove Transmitter. (1) Gain access to transmitter by removing aft cabin partition or unzipping canvas wall. Refer to Rear Cargo Compartment Wall - Maintenance Practices. (2) Disconnect coaxial antenna cable from transmitter. (3) Remove transmitter from airplane according to transmitter type: (a) Dorne and Margolin (Refer to Figure 201) - Remove screws attaching transmitter to bracket and remove transmitter. (b) Pointer 3000-1 (Refer to Figure 202) - Loosen wing nut on transmitter mounting bracket and remove transmitter. (c) Pointer 3000-11 (Refer to Figure 202) - Remove remote connector from transmitter. Disengage attach strap from around transmitter and remove transmitter from airplane.
B.
Install Transmitter. (1) Install transmitter to airplane according to transmitter type:
CAUTION: Ensure that direction of flight arrow on transmitter is pointing toward nose of airplane. (a) (b) (c)
Dorne and Margolin (Refer to Figure 201) - Position transmitter to bracket and secure using screws. Pointer 3000-1 (Refer to Figure 202) - Position transmitter to bracket and secure using wing nut. Pointer 3000-11 (Refer to Figure 202) - Position transmitter to bracket and secure using attach strap. Reconnect remote connector to transmitter.
CAUTION: Coaxial antenna cable assembly contains a matching transformer which must be connected properly to obtain specified emergency locator transmitter performance. Marking sleeves (labels) on antenna cable ends specify correct hookup. (2) (3) 3.
Install coaxial antenna cable to transmitter. Perform an operational test of transmitter in accordance with Emergency Locator Transmitter Test.
Panel Mounted Rocker Switch Removal/Installation NOTE:
An instrument panel mounted rocker switch is used in conjunction with Pointer 3000-11 series emergency locator transmitter to provide remote testing capability of transmitter.
A.
Remove Switch (Refer to Figure 202 ). (1) Disengage BCN MONITOR circuit breaker on firewall mounted junction box. (2) Gain access to backside of remote mounted switch. Compress locking tabs on either side of switch and, at same time, pull switch aft and away from instrument panel. (3) Disconnect electrical connector from switch.
B.
Install Switch (Refer to Figure 202). (1) Reconnect electrical connector to remote mounted switch. (2) Grasp edges of remote mounted switch and insert into instrument panel cutout. Ensure locking tabs engage and that switch is secure. (3) Engage BCN MONITOR circuit breaker on firewall mounted junction box.
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Dome and Margolin Emergency Locator Transmitter Installation Figure 201 (Sheet 1)
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Dome and Margolin Emergency Locator Transmitter Installation Figure 201 (Sheet 2)
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Pointer 3000 Series Emergency Locator Transmitter Installation Figure 202 (Sheet 1)
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Pointer 3000 Series Emergency Locator Transmitter Installation Figure 202 (Sheet 2)
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Pointer 3000 Series Emergency Locator Transmitter Installation Figure 202 (Sheet 3)
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4.
Antenna Removal/Installation A.
Remove Antenna. (1) Disconnect coaxial antenna cable from base of antenna. (2) Remove antenna from airplane according to antenna type: (a) Dorne and Margolin (Refer to Figure 201) - On inside of airplane, remove nut and washer securing antenna base to fuselage. Withdraw antenna (with washer) from airplane fuselage hole. (b) Pointer 3000-1 and 3000-11 (Refer to Figure 202) - On outside of airplane, remove six screws attaching antenna base and cable to cabin top and antenna doubler. Remove antenna rod and hub assembly.
B.
Install Antenna. (1) Install antenna to fuselage according to antenna type:
CAUTION: Adhere rubber boot to antenna only using class VC adhesive. Do not apply adhesive to fuselage skin or damage to paint may result. (a)
(b)
Dorne and Margolin (Refer to Figure 201) - Position antenna (with washer and rubber boot) through hole on top of fuselage. From inside of fuselage, secure antenna to fuselage using washer and nut. Slide rubber boot down against fuselage and secure top of boot to antenna using Class VC adhesive. Refer to Adhesive and Solvent Bonding - Maintenance Practices. Pointer 3000-1 and 3000-11 (Refer to Figure 202) - From inside of airplane, position antenna base and cable to fuselage. Secure to fuselage using screws. Connect antenna rod and hub assembly to antenna base.
CAUTION: Coaxial antenna cable assembly contains a matching transformer which must be connected properly to obtain specified emergency locator transmitter performance. Marking sleeves (labels) on the antenna cable ens specify correct hookup. (2) 5.
Reconnect coaxial antenna cable to base of antenna.
Battery Pack Removal A.
Remove Dorne and Margolin Battery Pack (Refer to Figure 201). (1) Place transmitter switch in OFF position. (2) Remove transmitter from airplane. (3) Remove screws attaching cover to transmitter. (4) Disconnect battery pack electrical connector. (5) Remove battery pack from transmitter.
B.
Install Dorne and Margolin Battery Pack (Refer to Figure 201).
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WARNING: Dispose of unserviceable battery pack properly. incinerate or compact.
Do not
CAUTION: Battery packs are not interchangeable and transmitters can only use recommended battery pack, as operating life and/or signal strength of the transmitter will be seriously degraded. In some cases, changes in mechanical configuration may make transmitter prone to failure due to vibration and corrosion. (1) (2) (3)
Place battery pack in case. Connect electrical connector to battery pack. Attach cover to transmitter using screws.
WARNING: Ensure new battery pack expiration date is entered in airplane records. It is also recommended that expiration date be place in emergency locator transmitter owner’s manual for quick reference. (4) (5)
Stamp new replacement date on outside of transmitter. Date should be noted on switch nameplate, on side of unit and in instruction nameplate on top of unit. Install transmitter to airplane.
C.
Remove Pointer 3000-1 and 3000-11 Series Battery Packs (Refer to Figure 202). (1) Place transmitter switch in OFF position. (2) Remove transmitter from airplane. (3) Remove screws attaching base plate to transmitter. (4) Disconnect battery pack and transmitter electrical connectors. (5) Remove battery pack from transmitter.
D.
Install Pointer 3000-1 and 3000-11 Battery Packs (Refer to Figure 202).
WARNING: Dispose of unserviceable battery pack properly. incinerate or compact.
Do not
CAUTION: Battery packs are not interchangeable and transmitters can only use recommended battery pack, as operating life and/or signal strength of the transmitter will be seriously degraded. In some cases, changes in mechanical configuration may make transmitter prone to failure due to vibration and corrosion. (1) (2) (3)
Place battery pack in transmitter. Connect battery pack electrical connectors. Attach base plate and gasket to transmitter using screws.
WARNING: Ensure new battery pack expiration date is entered in airplane records. It is also recommended that expiration date be place in emergency locator transmitter owner’s manual for quick reference. Stamp new replacement date on outside of transmitter. The date should be noted on switch nameplate, on side of unit and in instruction nameplate on top of unit. Install transmitter to airplane. (a)
(4)
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6.
Emergency Locator Transmitter Test A.
Operational Test Of Radiated Signal With Control Tower Monitoring. (1) Request permission from control tower and/or flight service station to test emergency locator transmitter system.
CAUTION: Prior approval from control tower and/or flight service station must be obtained when checking/testing operation of the emergency locator transmitter system. If performing local monitoring test conduct test at five minutes before or after the hour. (2)
Perform an operational test for the Emergancy Locator Transmitter. If the ELT does not perform properly during test proceed to Step 2. If ELT performs properly during test, proceed to Step 5.
CAUTION: Remove remote connector from (Pointer 3000-11) elt before performing maintenance. Failure to remove the remote connector could cause inadvertent blowing of the ELT’s internal fuse. (3) (4) (5) (6) (7)
If the ELT fails to operate properly during the functional test, remove the transmitter and return it to any authorized avionics repair shop for inspection and repair. Install transmitter. Perform operational test for the Emergency Locator. Make an entry in the airplane logbook stating this test has been accomplished. Activate emergency locator transmitter system by placing master switch to ON position. NOTE:
(8) (9)
Contact control tower and/or flight service station to confirm proper emergency locator transmitter system operation. Restore master switch to AUTO position. NOTE:
B.
On Pointer 3000-11 series system, momentarily place remote mounted switch to RESET position and release. This will place transmitter in the AUTO position.
Operational Test Of Radiated Signal With Local Monitoring. (1) Verify that test is performed within five minutes before or after the hour. (2) Position a small, hand held AM radio tuned to any frequency, within six inches of the emergency locator transmitter antenna. NOTE: (3) (4)
Using airplanes’s VLF receiver or ADF will not properly check power of radiated signal.
Activate emergency locator transmitter system for no more than three sweeps of the audio signal. (a) Verify that signal has been detected on the AM radio. Restore master switch to AUTO position. NOTE:
C.
On Pointer 3000-11 series system, transmitter may be activated from cockpit by placing the remote mounted switch to ON position. Other transmitters must be activated from the tailcone mounting area.
On Pointer 3000-11 series system, momentarily place the remote mounted switch to RESET position and release. This will place the transmitter in the AUTO position.
G Switch Operational Check (Dorne and Margolin/Pointer 3000-1 Series). (1) Remove transmitter from airplane. (2) While holding transmitter in one hand, sharply strike end of case in the direction of activation indicated on transmitter case. (a) Verify that G switch has been actuated. (3) Reset the G switch.
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Reinstall transmitter in airplane.
G Switch Operation Check (Pointer 3000-11 Series). (1) Remove transmitter from airplane. (2) Hold transmitter firmly in one hand, and make a throwing motion followed by a sudden reversal of the transmitter. (a) Verify that G switch has been actuated. (3) Reset the G switch. (4) Reinstall transmitter in airplane.
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MODEL 208 MAINTENANCE MANUAL ARTEX C406-2 EMERGENCY LOCATOR TRANSMITTER SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
Description A.
3.
An Artex C406-2 Emergency Locator Transmitter (ELT) System is installed to help rescue teams Þnd the airplane in the event of a crash. It is made such that it can operate in a wide range of environmental conditions and is resistant to the forces caused by many types of accidents.
The Artex C406-2 Emergency Locator Transmitter (ELT) system has a transmitter, an ELT/NAV interface, mode S address box, warning buzzer, integral battery pack, an internal G-switch, an ELT antenna, a cockpit ELT control panel, the cable assembly, and an antenna coax cable. (1) The transmitter has an internal battery and internal G-switch. It is installed in a tray and will come on automatically if the G-switch is actuated or if the cockpit panel switch is actuated manually. When the airplane electrical system is on, the microprocessor in the transmitter uses power from the airplane's electrical system. Electrical power from the transmitter's internal alkaline battery pack is used for the system test sequence and will also keep the system on in the event of an emergency. (2) The Artex C406-2 system uses an ELT antenna that is installed on the top of the fuselage, at FS 292.44.00 and RBL 15.55 for the 208 and FS 340.44 and RBL 15.55 for the 208B. The antenna is connected to the transmitter with a coaxial cable. (3) A G-switch, installed in the transmitter, and a two-position cockpit panel switch (SZ09) on the right switch/meter panel are used to control the transmitter. The cockpit panel switch lets the ßight crew activate, reset or test the system. An ON/OFF toggle switch on the transmitter is set to the ON position for normal system operation, and to OFF during maintenance or service. (4) The Artex ELT ELT/NAV Interface is used to convert the longitude/latitude navigation information into a format that the ELT can recognize. The ELT/NAV unit actively updates and stores this information. In the event of a crash, the ELT will transmit the last known position information. The ELT/NAV interface is connected to the ELT and the navigation system with cable assemblies.
Operation A.
The Artex C406-2 Emergency Locator Transmitter (ELT) System can be activated automatically by the G-switch or manually by one of the two manual control switches. (1) The G-switch will operate and start the transmitter as a result of crash accelerations that are parallel to the longitudinal axis of the airplane in a forward direction. (2) A remote-mounted switch (SZ09) on the right switch/meter panel in the cockpit can be used to manually operate the transmitter when the switch is set to the ON position.
B.
When activated, the ELT transmits on emergency frequencies 121.50, 243.00 and 406 MHz, at the same time with a swept tone at three sweeps-per-second. (1) The 121.50 and 243.00 MHz frequencies are used to send a locator signal that can be followed by those that are receiving it. The 406 MHz frequency is used to activate a satellite tracking system. The Artex C406-2 system is connected to the navigational system of the airplane as well as the transponder system. When the ELT system is in operation, the location and the tail number of the airplane are transmitted on the 406 MHz frequency.
C.
The Artex C406-2 system also has a complete self-analysis program with test routines that are transmitted at reduced power over frequencies 121.50, 243.00 and 406. MHz. The test sequence examines the system microprocessor, antenna and transmitter. The test routine is started when the remote switch is set to the ON position for one second, then moved to the ARM position.
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MODEL 208 MAINTENANCE MANUAL ARTEX C406-2 EMERGENCY LOCATOR TRANSMITTER SYSTEM - MAINTENANCE PRACTICES 1.
2.
3.
General A.
This section contains information required to complete maintenance procedures on the Artex C406-2 Emergency Locator Transmitter (ELT) System.
B.
The Artex C406-2 ELT is found under an aft cargo floor panel that is marked by a location placard.
Emergency Locator Transmitter (ELT) Removal/Installation A.
Remove the Emergency Locator Transmitter (ELT) (Refer to Figure 201). (1) Make sure the battery switch is in the OFF position. (2) Disengage the ELT/NAV circuit breaker on the cockpit circuit breaker panel. (3) Open the aft baggage compartment access door. (4) Remove the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. (5) Make sure the ON/OFF switch on the ELT is in the OFF position. (6) Disconnect the coax connectors (PZ121 and PZ123) from the ELT. (7) Loosen the knurled nuts on the mounting tray end cap. (8) Pull the front cover away from the transmitter and mounting tray. (9) Remove the top cover from the transmitter. (10) Pull the mounting tray end cap away from the ELT to disconnect the connector (PZ117) from the ELT. (11) Lift up on the connector end of the ELT to remove it from the mounting tray. (12) Remove the screws that attach the mounting tray to the shelf assembly. (13) Remove the ELT system components from the airplane.
B.
Install the Emergency Locator Transmitter (ELT) (Refer to Figure 201). (1) Use the screws to attach the mounting tray to the support assembly. (2) Put the ELT in the mounting tray at an angle to engage the lock mechanism at the opposite end of the ELT. (3) Push the ELT down into the mounting tray until it is fully installed in the tray. (4) install the top cover onthe ELT transmitter (5) Put the mounting tray end cap onto the mounting tray and ELT to engage the electrical connector (PZ117) on the ELT. (6) Connect the coax connectors (PZ121 and PZ123) to the ELT. (7) Tighten the knurled nuts that hold the ELT in position in the mounting tray. (8) Make sure the ON/OFF switch on the ELT is in the ON position. (9) Engage the ELT/NAV circuit breaker on the cockpit circuit breaker panel. (10) Do a Transmitter Test of the ARTEX C406-2 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (11) Install the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. (12) Close the aft baggage compartment door.
ELT Antenna Removal/Installation A.
Remove the ELT Antenna (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6)
The ELT antenna is found on the top surface of the fuselage at FS 292.44 and RBL 15.55 for the Model 208. For the Model 208B, the antenna is at FS 340.44 and RBL 15.55.
Open the aft baggage compartment door. Remove the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. Make sure the ON/OFF switch on the ELT is in the OFF position. Remove the overhead panel directly under the ELT antenna. Put the master switch on the emergency locator transmitter in the OFF position. Get access to the ELT antenna. Remove the overhead panel directly under the antenna.
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Artex C406-2 Emergency Locator Transmitter System Installation Figure 201 (Sheet 1)
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Artex C406-2 Emergency Locator Transmitter System Installation Figure 201 (Sheet 2)
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Artex C406-2 Emergency Locator Transmitter System Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9) B.
Disconnect the coax connectors (PZ122 and PZ124) from the ELT antenna. Remove the screws that attach the ELT antenna to the top surface of the fuselage. Remove the antenna from the airplane.
Install the ELT Antenna (Refer to Figure 201). (1) Make sure that you remove all of the old sealant from the ELT antenna and from the airplane skin. Refer to Chapter 20, Fuel, Weather, Pressure and High Temperature Sealing - Maintenance Practices. (2) Connect the coax connectors (PZ122 and PZ124) to the ELT antenna. NOTE:
The coax connector with the 90-degree connector (PZ122) is to be attached to the forward connector.
(3)
Apply a chemically conductive chemical film treatment to the faying surfaces of the ELT antenna and the airplane structure to make sure that there is an electrical bond. Refer to Chapter 20, Electrical Bonding - Maintenance Practices. (4) Put the ELT antenna in position on the top surface of the fuselage. (5) Use the screws to attach the ELT antenna to the top surface of the fuselage. (6) Use Type I, Class B sealer to apply a fillet seal around the antenna where it touches the outside surface of the airplane. Refer to Chapter 20, Fuel, Pressure, Weather and High-Temperature Sealing - Maintenance Practices. (7) Put the master switch, on the emergency locator transmitter, in the ON position. (8) Do a Transmitter Test of the ARTEX C406-2 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (9) Install the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. (10) Install the overhead panel. (11) Close the aft baggage compartment door. 4.
5.
ELT/NAV Interface Unit Removal/Installation A.
Remove the ELT/NAV Interface Unit (Refer to Figure 201). (1) Open the aft baggage compartment access door. (2) Disengage the ELT/NAV on the cockpit circuit breaker panel. (3) Remove the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. (4) Disconnect the electrical connectors (PZ118 and PZ119) from the ELT/NAV interface unit. (5) Remove the screws and washers that attach the ELT/NAV interface unit to the shelf assembly. (6) Remove the ELT/NAV interface unit from the airplane.
B.
Install the ELT/NAV Interface Unit (Refer to Figure 201). (1) Put the ELT/NAV interface unit in position on the shelf assembly. (2) Use the screws and washers to attach the ELT/NAV interface unit to the support assembly. (3) Connect the electrical connectors (PZ118 and PZ119) to the ELT/NAV interface unit. (4) Engage the ELT/NAV on the cockpit circuit breaker panel. (5) Do a Transmitter Test of the ARTEX C406-2 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (6) Install the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. (7) Close the aft baggage compartment door.
Mode S Program Module Removal/Installation A.
Remove the Mode S Program Module (Refer to Figure 201). (1) Open the aft baggage compartment access door. (2) Remove the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. (3) Disengage the ELT/NAV circuit breaker on the cockpit circuit breaker panel. (4) Disconnect the electrical connector (PZ120) from the Mode S program module.
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MODEL 208 MAINTENANCE MANUAL (5) (6) B.
Remove the screws and washers that attach the Mode S program module to the shelf assembly. Remove the Mode S program module from the airplane.
Install the Mode S Program Module (Refer to Figure 201). (1) Put the Mode S program module in position on the shelf assembly. (2) Use the screws and washers to attach the Mode S program module to the support assembly. (3) Connect the electrical connector (PZ120) to the Mode S program module. (4) Engage the ELT/NAV on the cockpit circuit breaker panel. (5) Do a Transmitter Test of the ARTEX C406-2 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (6) Install the aft cargo floor panel between FS 264.00 and FS 307.91 for the 208 and FS 332.00 and FS 355.91 for the 208B. (7) Close the aft baggage compartment access door.
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MODEL 208 MAINTENANCE MANUAL ARTEX ME406 EMERGENCY LOCATOR TRANSMITTER SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
3.
An Artex ME406 Emergency Locator Transmitter (ELT) System is installed to help rescue teams Þnd the airplane in the event of a crash. It is made to operate in a wide range of environmental conditions and is resistant to the forces caused by many types of accidents.
Description A.
Artex ME406 ELT. (1) The Artex ME406 Emergency Locator Transmitter (ELT) system includes an ELT unit, an integral battery pack, warning buzzer, internal G-switch, antenna, remote switch, cable assembly, and antenna coaxial cable. The ELT unit transmits on 121.5 MHz and 406.028 MHz. (2) The battery pack has two D-size lithium cells mounted under a battery cover. The battery pack is replaced as necessary in the Þeld. (3) The ELT activates a buzzer that is installed near the ELT assembly. The buzzer makes a loud noise to let people know that the ELT is on. (4) The G-switch is installed in the ELT transmitter and is activated with a sudden reduction in forward speed.
B.
Artex ELT Antenna. (1) The ELT system uses an antenna to transmit the emergency locator signal. The ELT antenna is installed on top of the tailcone skin, forward of the vertical stabilizer at FS 311.45 and RBL 3.62 for the 208 and at FS 359.45 and RBL 3.62 for the 208B. The ELT antenna is connected with a coaxial cable to the ELT unit inside the dorsal.
C.
ELT Remote Switch. (1) The ELT remote switch is installed on the right panel. The ELT remote switch is a two-position rocker switch that can be set in the ARM or the ON positions.
Operation
CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the Þrst Þve minutes of each hour. If you must complete the functional test at a time other than the Þrst Þve minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than Þve seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz signal after it is activated for approximately 50 seconds. This signal is identiÞed as a distress signal. A.
Artex ME406 ELT. (1) When an accident occurs, the ELT will activate automatically and transmit a standard swept tone on the 121.5 MHz (emergency frequency). The 121.5 MHz transmission will continue until the ELT battery has expired. The 406.028 MHz transmitter is activated and will send a message to the satellite every 50 seconds for 440 milliseconds. The 406.028 MHz transmission will continue for 24 hours and then stop. During operation, the ELT will receive electrical power from the ELT battery pack only.
B.
ELT Remote Switch. (1) The ELT can also be activated manually in the cockpit with the ELT remote switch. To manually activate the ELT, put the ELT remote switch in the ON position. The red LED will come on when the remote switch is set in the ON position. The ELT remote switch can also be used to do a test of the ELT system (refer to Artex ME406 Emergency Locator Transmitter - Troubleshooting). During typical operation, the ELT remote switch will be in the ARM position.
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MODEL 208 MAINTENANCE MANUAL ARTEX ME406 EMERGENCY LOCATOR TRANSMITTER SYSTEM - TROUBLESHOOTING 1.
General A.
2.
Tools and Equipment A.
3.
This section contains the information that is needed to complete the self test for the ARTEX ME406 Emergency Locator Transmitter (ELT) system. The system transmits on two frequencies at the same time.
For information on tools and equipment, refer to Equipment and Furnishings - General.
ME406 Emergency Locator Transmitter (ELT) Self Test Preparation
CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the Þrst Þve minutes of each hour. If you must complete the functional test at a time other than the Þrst Þve minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular AC-91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than Þve seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz signal after it is activated for approximately 50 seconds. This signal is identiÞed as a distress signal. A.
Prepare the Airplane for the ME406 Emergency Locator Transmitter Troubleshooting. (1) Put the BATTERY switch in the ON position. (2) Examine the ELT battery to make sure that it is serviceable. (a) If the battery must be replaced, follow the manufacturer's instructions to replace it.
B.
Do a ELT 121.5 MHz Test. (1) Tune the receiver (usually the aircraft radio) to 121.5 MHz. (2) Turn the ELT instrument panel remote switch to the ON position and wait for 3 sweeps on the receiver which takes about 1 second. (3) Turn the remote switch back to the ARM (OFF) position immediately and the switch's LED and the buzzer will give 1 pulse. If more pulses are displayed, Þnd the problem from the list below: (a) One ßash - Indicates that the system is operational and that there were no error conditions found. (b) Three ßashes - Shows an open or short condition on the antenna output or cable. Use the list below to isolate and repair the problem: 1 Make sure the BNC cable is connected and in good condition. Do a continuity check of the center conductor and shield. Examine for a shorted cable. 2 Examine for a intermittent connection in the BNC cable. 3 Examine the antenna installation if this error code persists. You can examine it with a VSWR meter. Examine the antenna for opens, shorts, and a resistive ground plane connection. (c) Four ßashes - This shows a low power condition. This occurs if the output power is below approximately 33 dBm (2 watts) for the 406.028 MHz signal, or 17 dBm (50mW) for the 121.5 MHz signal. Also, this may indicate that the 406.028 MHz signal is off frequency. For this error code, the ELT must be sent back to ARTEX for repair or replacement. (d) Five ßashes - This shows that the ELT has not been programmed. However, this does not show erroneous or corrupted programmed data. (e) Six ßashes - This shows that the G-switch loop between pins 5 and 12 at the D-sub connector is not installed. The ELT will not activate during a crash. 1 Do a resistance test to make sure the harness D-sub jumper is installed. There must be less than 1 ohm of resistance between pins 5 and 12. (f) Seven ßashes - This shows that the ELT battery has too much accumulated operation time and you must replace it to meet FAA speciÞcations.
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Put the BATTERY switch in the OFF position.
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MODEL 208 MAINTENANCE MANUAL ARTEX ME406 EMERGENCY LOCATOR TRANSMITTER SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives maintenance practices for the emergency locator transmitter (ELT) system. Components in the ELT system include the ELT, antenna, remote switch, and buzzer.
Emergency Locator Transmitter (ELT) Removal/Installation A.
Remove the Emergency Locator Transmitter (ELT) (Refer to Figure 201). (1) Make sure the MASTER switch is in the OFF position. (2) Remove access panel 340A on the right side of the tail dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Keep the ON/ARM switch on the ELT in the ARM position.
CAUTION: Remove the electrical connector and the ELT is off. However, the ELT can be activated with the electrical connector removed if the switch on the front is moved to the ON position. Be careful not to move the switch to the ON position. (4)
Disconnect the BNC connector (PZ148) and the electrical connector (PZ153) from the ELT. NOTE:
(5) (6) (7) B.
The ELT is off when the electrical connector is removed from the ELT.
Open the latch on the ELT strap assembly and lift the hinged strap up and away from the ELT. Open the Velcro strap that holds the ELT to the mounting tray. Remove the ELT from the mounting tray.
Install the Emergency Locator Transmitter (ELT) (Refer to Figure 201). NOTE:
The ELT is off when the electrical connector is removed from the ELT.
CAUTION: Remove the electrical connector and the ELT is off. However, the ELT can be activated with the electrical connector removed if the switch on the front is moved to the ON position. Be careful not to move the switch to the ON position. (1) (2) (3) (4) (5) (6) (7) (8) 3.
Put the ELT in the mounting tray at the angle necessary to engage the lock mechanism at the opposite end of the ELT. Push the ELT down into the mounting tray until it is fully installed in the tray. Connect the Velcro strap that holds the ELT firmly to the mounting tray. Move the hinged strap down on the ELT and close the ELT strap assembly latch. Connect the BNC connector and the electrical connector to the ELT. Make sure the ON/ARM switch is in the ARM position. Do a Transmitter Test of the Artex ME406 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. Install access panel 340A on the right side of the tail dorsal assembly.
ELT Buzzer Removal/Installation A.
Remove the ELT Buzzer (Refer to Figure 201). (1) Remove electrical power from the airplane. (2) Remove access panel 340A on the right side of the tail dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Remove the screws from the terminals on the top of the buzzer. (4) Remove the electrical connectors from the terminals on the buzzer. (5) Remove the mounting nut on the bottom side of the bracket from the bottom of the buzzer. (6) Lift the buzzer from the bracket and remove the buzzer from the tailcone.
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MODEL 208 MAINTENANCE MANUAL
Artex ME406 Emergency Locator Transmitter (ELT) System Installation Figure 201 (Sheet 1)
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Artex ME406 Emergency Locator Transmitter (ELT) System Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL B.
4.
5.
Install the buzzer (Refer toFigure 201). (1) Carefully put the buzzer in position on the bracket. (2) Install the mounting nut on the bottom of the buzzer. (3) Put the electrical connectors on the terminals on the top of the buzzer. (4) Install screws to the terminals to hold the electrical connectors on the terminals. (5) Do a Transmitter Test of the Artex ME406 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (6) Install access panel 340A on the right side of the tail dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Remote Switch Removal/Installation A.
Remove the Remote Switch (Refer to Figure 201). (1) Remove electrical power from the aircraft. (2) Remove access panel 340A on the right side of the tail dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Disconnect the electrical connector (PZ153) from the ELT. (4) Remove the screws from the front of the switch. (5) Pull the remote switch from the panel to get to the electrical connector. (6) Disconnect the connector from the back of the switch.
B.
Install the Remote Switch. (1) Connect the electrical connector to the back of the switch. (2) Put the remote switch into the panel. (3) Install the screws on the front of the switch. (4) Connect the electrical connector to the ELT. (5) Do a Transmitter Test of the Artex ME406 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (6) Install access panel 340A on the right side of the tail dorsal assembly.
ELT Antenna Removal/Installation A.
Remove the ELT Antenna (Refer to Figure 201). NOTE: (1)
The ELT antenna is found on the top surface of the fuselage at FS 311.45 and RBL 3.62 for the Model 208. For the Model 208B, the antenna is at FS 359.45 and RBL 15.55.
Remove access panel 340A on the right side of the tail dorsal assembly. NOTE:
The ELT is off when the electrical connector is removed from the ELT.
CAUTION: Remove the electrical connector and the ELT is off. However, the ELT can be activated with the electrical connector removed if the switch on the front is moved to the ON position. Be careful not to move the switch to the ON position. (2) (3)
B.
Make sure the ELT ON/ARM switch is in the ARM position and disconnect the electrical connector (PZ153) from the ELT. Remove the ELT antenna. Refer to Figure 201. (a) Remove and keep the screws that attach the ELT antenna to the top surface of the fuselage. (b) Carefully lift the antenna from the fuselage surface. (c) Disconnect the BNC connector from the antenna.
Install the ELT Antenna (Refer to Figure 201). (1) Make sure that you remove all of the old sealant from the ELT antenna and the airplane skin. Refer to Chapter 20, Fuel, Weather, Pressure and High Temperature Sealing - Maintenance Practices. (2) Connect the BNC connector to the ELT antenna.
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6) (7) (8)
Apply a chemically conductive chemical film treatment to the faying surfaces of the ELT antenna and to the airplane structure to make sure that there is an electrical bond. Refer to Chapter 20, Electrical Bonding - Maintenance Practices. Put the ELT antenna in position on the top surface of the fuselage. Attach the ELT antenna to the top surface of the fuselage with the kept screws. (a) Torque the screws to a maximum of 20 inch-pounds. Use a Type I, Class B sealer to apply a fillet seal around the antenna where it touches the outside surface of the airplane. Refer to Chapter 20, Fuel, Pressure, Weather and High-Temperature Sealing - Maintenance Practices. With the ELT ON/ARM switch in the ARM position, connect the electrical connector to the ELT. Do a Transmitter Test of the Artex ME406 Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check.
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MODEL 208 MAINTENANCE MANUAL ARTEX C406-N EMERGENCY LOCATOR TRANSMITTER SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
3.
An ARTEX C406-N Emergency Locator Transmitter (ELT) System is installed to help rescue teams Þnd the airplane if there is a crash. It is made to operate in a wide range of environmental conditions and is resistant to the forces caused by many types of accidents.
Description A.
ARTEX C406-N ELT. (1) The ARTEX C406-N Emergency Locator Transmitter (ELT) system includes an ELT unit, integral battery pack, warning buzzer, internal G-switch, antenna, remote switch, cable assembly, antenna coaxial cable, and can include an optional programming adapter. The ELT unit transmits on 121.5/243.0 MHz and 406.028 MHz. (2) The battery pack has four cells mounted under a battery cover. The battery pack is replaced as necessary in the Þeld. (3) The ELT energizes a buzzer that is installed near the ELT assembly. The buzzer makes a noise to show that the ELT is on. (4) The G-switch is installed in the ELT transmitter and is started with a sudden reduction in forward speed.
B.
ARTEX ELT Antenna. (1) The ELT system uses an antenna to transmit the emergency locator signal. The ELT antenna is installed on top of the tailcone skin, forward of the vertical stabilizer at FS 311.45 and RBL 3.62 for the 208, and at FS 359.45 and RBL 3.62 for the 208B. The ELT antenna is connected with a coaxial cable to the ELT unit inside the dorsal.
C.
ELT Remote Switch. (1) The ELT remote switch is installed on the right panel. The ELT remote switch is a two-position rocker switch that you can set in the ARM or the ON positions.
Operation
CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the Þrst Þve minutes of each hour. If you must complete the functional test at a time other than the Þrst Þve minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular 91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than Þve seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz signal for 520 milliseconds approximately every 50 seconds. This transmission is an encoded digital message and is sent to a satellite as a distress signal. A.
ARTEX C406-N ELT. (1) When an accident occurs, the ELT will start automatically and transmit a standard swept tone on the 121.5 and 243.0 MHz (emergency frequencies). The 121.5/243.0 MHz transmission will continue until the ELT battery has expired, which will be more than 50 hours. The 406.028 MHz transmitter is activated and will send a message to the satellite every 50 seconds for 520 milliseconds. The 406.028 MHz transmission will continue for 24 hours, then stop. During operation, the ELT will receive electrical power from the ELT battery pack only.
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ELT Remote Switch. (1) You also can start the ELT manually in the cockpit with the ELT remote switch. To manually start the ELT, put the ELT remote switch in the ON position. The red LED will ßash twice each second when the remote switch is set in the ON position. You also can use the ELT remote switch to do a test of the ELT system (refer to ARTEX C406-N Emergency Locator Transmitter - Troubleshooting). During typical operation, the ELT remote switch will be in the ARM position.
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MODEL 208 MAINTENANCE MANUAL ARTEX C406-N EMERGENCY LOCATOR TRANSMITTER SYSTEM - TROUBLESHOOTING 1.
General A.
2.
Tools and Equipment A.
3.
This section contains the information that is needed to complete the self test for the ARTEX C406-N Emergency Locator Transmitter (ELT) system. The system transmits on three frequencies at the same time.
For information on tools and equipment, refer to Equipment and Furnishings - General.
C406-N Emergency Locator Transmitter (ELT) Self Test Preparation
CAUTION: Operate the Emergency Locator Transmitter (ELT) system only during the Þrst Þve minutes of each hour. If you must complete the functional test at a time other than the Þrst Þve minutes of the hour, you must do the test with a direct connection to the ELT and a 30 dB attenuator. Refer to the FAA Advisory Circular 91-44A. CAUTION: Do not operate the Emergency Locator Transmitter (ELT) for more than Þve seconds at a time. Do not operate the ELT again for 15 seconds. The ELT will transmit a 406.028 MHz signal for 520 milliseconds approximately every 50 seconds. This transmission is an encoded digital message and is sent to a satellite as a distress signal. A.
Prepare the Airplane for the C406-N Emergency Locator Transmitter Troubleshooting. (1) Put the BATTERY switch in the ON position. (2) Examine the ELT battery to make sure that it is serviceable. (a) If you must replace the battery, use the manufacturer's instructions. NOTE:
B.
If you use the navigation position function of the ELT, make sure that both the ELT and the airplane's navigational system are on at least 30 seconds before the ELT test.
Do the ELT Test. (1) Tune the receiver (usually the aircraft radio) to 121.5 MHz. (2) Turn the ELT instrument panel remote switch to the ON position and wait for 3 sweeps on the receiver which takes about 1 second. (3) Turn the remote switch back to the ARM (OFF) position and observe the LED activity at this time. NOTE:
(a) (b)
The microprocessor in the ELT makes a check of the G-switch (automatic activation switch) latching circuit, pins 12 and 13 on the 22-pin circular connector at the ELT; the 406.028 MHz transmitter for proper RF output; presence of valid navigation data (navigation system must be active) and a battery check. If the ELT operates correctly, the sequence that follows the entry to the "ARMED" (OFF) condition will result in the panel LED ON for approximately 1 second and then it will go off. If a problem is detected, the LED has a coded signal that follows the initial 1 second pulse. The coded signal and related problem are as follows (the LED will ßash in order of importance with approximately a 0.5 - to 1-second pause between each error code if there are multiple errors):
One ßash - Shows a G-switch loop open failure. Three ßashes - Shows a 406-028 MHz transmitter problem (i.e. bad or unconnected coax, antenna problem, or low power output).
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Five ßashes - This shows that there is no navigation data. This is possibly the result of incorrect wiring between the system interface connections, incorrect programming, or invalid navigation data (navigation system not powered up). NOTE:
(d)
(4)
This error is not present when the ELT is programmed with a short message (User Protocol).
Seven ßashes - This shows that the ELT battery has too much accumulated operation time and you must replace it to meet FAA speciÞcations.
NOTE:
There is a sequence to the problem-reporting, which is the same order as that listed above. That is, if the G-switch circuit has a failure, there will be a single ßash, then 3 ßashes if there was a transmitter problem and so on.
NOTE:
There is an error condition where the LED on the ELT and remote switch will ßash rapidly with a 2-ßash pulse made of a short and a long ßash. This occurs immediately after power is applied to the ELT, if either of the conditions that follow occur: 1. The ELT does not use the optional C406-N Programming Adapter (P/N 4535068) and pins 3 and 4 are not connected with a jumper (or connection is interrupted) at the mating circular connector. 2. The Programming Adapter is used and there is a communication problem or wiring error. NOTE: This error will continue for up to one minute even after aircraft power is removed due to the internal ELT power supply backup.
Do a monthly "self-test" of the ELT with the steps outlined in the section. NOTE:
The ELT is not to be tested more than once a month as excessive activations will decrease the battery life.
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MODEL 208 MAINTENANCE MANUAL ARTEX C406-N EMERGENCY LOCATOR TRANSMITTER SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives maintenance practices for the emergency locator transmitter (ELT) system. Components in the ELT system include the ELT unit, integral battery pack, warning buzzer, internal G-switch, antenna, remote switch, cable assembly, optional programming adapter, and antenna coaxial cable. The ELT unit transmits on 121.5 /243.0 MHz and 406.028 MHz
Emergency Locator Transmitter (ELT) Removal and/or Installation A.
Remove the Emergency Locator Transmitter (ELT) (Refer to Figure 201). (1) Make sure that the MASTER switch is in the OFF position. (2) Remove access panel 340A on the right side of the dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Keep the ON/OFF switch on the ELT in the OFF position.
CAUTION: The ELT is off when the electrical connector is removed from the ELT. However, you can activate the ELT with the electrical connector removed if you move the switch on the front of the ELT to the ON position. Be careful not to move the switch to the ON position. (4) (5) (6) (7) (8)
Loosen the thumbscrews on the end of the cap. Pull the end cap away from the ELT. Lift up the protective top Cover and push it away from the connector end of the ELT unit. Disconnect the BNC connector (PZ148) and the electrical connector (PZ146) from the ELT. Lift the ELT from the connector end to remove it from the mounting tray. NOTE:
B.
Careful use of a flat blade screwdriver as a lever makes this step easier.
Install the Emergency Locator Transmitter (ELT) (Refer to Figure 201).
CAUTION: The ELT is off when the electrical connector is removed from the ELT. However, you can activate the ELT with the electrical connector removed if you move the switch on the front of the ELT to the ON position. Be careful not to move the switch to the ON position. (1) (2) (3) (4) (5) (6)
Put the ELT in the mounting tray at an angle necessary to engage the lock-ears at the end opposite of the direction-of-flight arrow. Push the ELT down into the mounting tray until it is fully installed in the tray. Install the protective top cover on the ELT with the cover locking-slots over the locking-ears on the ELT. Push the cover toward the connector end of the ELT and down in position on the ELT. Insert the antenna BNC connector through the end cap access hole and connect it to the ELT unit. Put the 22-pin connector through the end cap and connect it to the ELT connector. Slide the end cap into position over the mounting tray and protective top cover and attach the end cap to the mounting tray with the two thumbscrews.
CAUTION: Do not tighten the thumbscrews more than 18 inch-pounds. (7) (8) (9)
Make sure that the ON/OFF switch is in the OFF position. Do a Transmitter Test of the ARTEX C406-N Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. Install access panel 340A on the right side of the dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation
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ARTEX C406-N Emergency Locator Transmitter (ELT) System Installation Figure 201 (Sheet 1)
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ARTEX C406-N Emergency Locator Transmitter (ELT) System Installation Figure 201 (Sheet 2)
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3.
4.
5.
ELT Buzzer Removal/Installation A.
Remove the ELT Buzzer (Refer to Figure 201). (1) Remove electrical power from the airplane. (2) Remove access panel 340A on the right side of the dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Remove the screws from the terminals on the top of the buzzer. (4) Remove the electrical connectors from the terminals on the buzzer. (5) Remove the mounting nut on the bottom side of the bracket from the bottom of the buzzer. (6) Lift the buzzer from the bracket and remove the buzzer from the tailcone.
B.
Install the buzzer (Refer to Figure 201). (1) Carefully put the buzzer in position on the bracket. (2) Install the mounting nut on the bottom of the buzzer. (3) Put the electrical connectors on the terminals on the top of the buzzer. (4) Install screws to the terminals to hold the electrical connectors on the terminals. (5) Do a Transmitter Test of the ARTEX C406-N Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (6) Install access panel 340A on the right side of the dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Remote Switch Removal/Installation A.
Remove the Remote Switch (Refer to Figure 201). (1) Remove electrical power from the aircraft. (2) Remove access panel 340A on the right side of the dorsal assembly. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Disconnect the electrical connector (PZ146) from the ELT. (4) Remove the screws from the front of the remote switch. (5) Pull the remote switch from the panel to get access to the electrical connector. (6) Disconnect the connector from the back of the switch.
B.
Install the Remote Switch. (1) Connect the electrical connector to the back of the switch. (2) Put the remote switch into the panel. (3) Install the screws on the front of the switch. (4) Connect the electrical connector to the ELT. (5) Do a Transmitter Test of the ARTEX C406-N Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check. (6) Install access panel 340A on the right side of the dorsal assembly.
ELT Antenna Removal/Installation A.
Remove the ELT Antenna (Refer to Figure 201). NOTE: (1)
The ELT antenna is on the top surface of the fuselage at FS 311.45 and RBL 3.62 for the Model 208. For the Model 208B, the antenna is at FS 359.45 and RBL 15.55.
Remove access panel 340A on the right side of the dorsal assembly.
CAUTION: The ELT is off when the electrical connector is removed from the ELT. However, you can activate the ELT with the electrical connector removed if you move the switch on the front of the ELT to the ON position. Be careful not to move the switch to the ON position. (2)
Make sure the ELT ON/OFF switch is in the OFF position and disconnect the electrical connector (PZ148) from the ELT.
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B.
Remove the ELT antenna. Refer to Figure 201. (a) Remove and keep the screws that attach the ELT antenna to the top surface of the fuselage. (b) Carefully lift the antenna from the fuselage surface. (c) Disconnect the BNC connector from the antenna.
Install the ELT Antenna (Refer to Figure 201). (1) Make sure that you remove all of the old sealant from the ELT antenna and the airplane skin. Refer to Chapter 20, Fuel, Weather, Pressure and High Temperature Sealing - Maintenance Practices. (2) Connect the BNC connector to the ELT antenna. (3) Apply a chemically conductive chemical film treatment to the faying surfaces of the ELT antenna and to the airplane structure to make sure that there is an electrical bond. Refer to Chapter 20, Electrical Bonding - Maintenance Practices. (4) Put the ELT antenna in position on the top surface of the fuselage. (5) Attach the ELT antenna to the top surface of the fuselage with the kept screws. (a) Torque the screws to a maximum of 20 inch-pounds. (6) Use a Type I, Class B sealer to apply a fillet seal around the antenna where it touches the outside surface of the airplane. Refer to Chapter 20, Fuel, Pressure, Weather and High-Temperature Sealing - Maintenance Practices. (7) With the ELT ON/OFF switch in the OFF position, connect the electrical connector to the ELT. (8) Do a Transmitter Test of the ARTEX C406-N Emergency Locator Transmitter (ELT) System. Refer to Emergency Equipment - Inspection/Check.
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MODEL 208 MAINTENANCE MANUAL SOUNDPROOFING - MAINTENANCE PRACTICES 1.
General A.
This section covers the soundproofing panels in the cabin area. The soundproofing material is a combination of fiberglass batting and damping foam panels. Specific percentages of soundproofing damping foam panels with self-adhesive backing are installed in designated areas. NOTE:
2.
Soundproofing damping foam panels do not overlap stringers, longerons, bulkheads or other raised surfaces.
B.
A layer of fiberglass batting is required in most of the cabin area and is cemented over the entire surface, including damping foam panels and all raised surfaces (stringers, longerons and bulkheads).
C.
The firewall, passenger door and lower cargo door soundproofing is 100 percent coverage, using a combination of a cushion (two layers of Therma-Sil with a Fiberfrax Durablanket filler) and a single layer Therma-Sil blanket attached with clips.
Soundproofing Damping Foam Panel Removal/Installation A.
Remove Soundproofing Damping Foam Panels (Refer to Figure 201). (1) Pull panel away from airplane surface. Damping foam will probably separate, leaving fragments stuck to airplane surface.
CAUTION: Do not allow solvent to touch window, painted trim, upholstery or carpet. (2) (3) B.
Loosen fragments by applying Toluene solvent (or equivalent). Apply two to five minutes for fragments and adhesive to soften. Using a non-metallic scraper, remove all fragments and residue.
Install Soundproofing Damping Foam Panels (Refer to Figure 201). NOTE:
Soundproofing panels cover either 80 percent or 100 percent of designated areas. Refer to Figure 201 for soundproofing coverage percentages explanation.
NOTE:
Soundproofing material shall be applied to the lower surface of floorboards, including all access panels (Model 208 only).
(1) (2) (3) (4)
Ensure surface to be covered is clean and smooth. Measure surface to be covered and make a paper template. Using paper template, cut out soundproofing panel. Remove protective cover sheet from adhesive surface of panel.
CAUTION: Care should be taken in positioning panels. Once panel adhesive contacts airplane surface, panel cannot be repositioned. (5) (6) 3.
Position panel to surface of airplane. Press entire panel surface against airplane surface.
Fiberglass Batting Removal/Installation A.
Remove Fiberglass Batting (Refer to Figure 201). NOTE: (1)
Remove only as much fiberglass batting as is necessary to perform repairs and maintenance.
Pull fiberglass batting away from structure. Batting separation will probably leave fragments stuck to structure.
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Soundproofing Installation Figure 201 (Sheet 1)
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Soundproofing Installation Figure 201 (Sheet 2)
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Soundproofing Installation Figure 201 (Sheet 3)
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Soundproofing Installation Figure 201 (Sheet 4)
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Soundproofing Installation Figure 201 (Sheet 5)
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Soundproofing Installation Figure 201 (Sheet 6)
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Soundproofing Installation Figure 201 (Sheet 7)
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CAUTION: Do not allow solvent to touch windows, painted trim, upholstery or carpet. (2) (3) B.
Loosen fragments by applying toluene, or its equivalent, and allowing two to five minutes to soften adhesive. Using a non-metallic scraper, remove all fragments and residue.
Install Fiberglass Batting (Refer to Figure 201). (1) Cut a piece of fiberglass batting sufficient to cover repair area. (2) Brush industrial cement (EC-1300L or equivalent) over entire repair area.
CAUTION: Care should be taken in positioning batting. Once batting has contacted cement, it may be difficult to reposition without damage to batting. (3) (4) 4.
Position batting to surface. Press entire surface of batting against airplane.
Therma-Sil Removal/Installation A.
Remove Therma-Sil Blanket and Cushion (Refer to Figure 201). (1) Remove clips securing Therma-Sil blanket and Fiberfrax Durablanket/Therma-Sil cushion. (2) Remove Therma- Sil blanket and Fiberfrax Durablanket/Therma-Sil cushion.
B.
Install Therma-Sil Blanket and Cushion (Refer to Figure 201). (1) Position Therma-Sil blanket and Fiberfrax Durablanket/Therma-Sil cushion to firewall. (2) Install clips securing Therma-Sil blanket and Fiberfrax Durablanket/Therma-Sil cushion to firewall.
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26 CHAPTER
FIRE PROTECTION
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LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
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26-10-00
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Aug 1/1995
26-10-00
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Aug 1/1995
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26-Title 26-List of Effective Pages 26-Record of Temporary Revisions 26-Table of Contents 26-List of Tasks
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS FIRE PROTECTION - GENERAL. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
26-00-00 Page 1 26-00-00 Page 1 26-00-00 Page 1
ENGINE FIRE DETECTION SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
26-10-00 Page 1 26-10-00 Page 1 26-10-00 Page 1
ENGINE FIRE DETECTION SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
26-10-00 Page 101 26-10-00 Page 101
ENGINE FIRE DETECTION SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Heat Detection Loop Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Box Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Alarm Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
26-10-00 Page 201 26-10-00 Page 201 26-10-00 Page 201 26-10-00 Page 201 26-10-00 Page 201
ENGINE FIRE DETECTION SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Fire Detection System General Visual Inspection . . . . . . . . . . . . . . . . . . . . . . .
26-10-00 Page 601 26-10-00 Page 601 26-10-00 Page 601
PORTABLE FIRE EXTINGUISHING - DESCRIPTION AND OPERATION . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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PORTABLE FIRE EXTINGUISHING - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Portable Fire Extinguisher Restoration (Internal Inspection) . . . . . . . . . . . . . . . . . . . . . Portable Fire Extinguisher Functional Check (Weight Check) . . . . . . . . . . . . . . . . . . . . Portable Fire Extinguisher Restoration (Hydrostatic Test) . . . . . . . . . . . . . . . . . . . . . . .
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26 - CONTENTS © Cessna Aircraft Company
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LIST OF TASKS 26-10-00-210
Engine Fire Detection System General Visual Inspection
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26-20-00-290
Portable Fire Extinguisher Restoration (Internal Inspection)
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26-20-00-720
Portable Fire Extinguisher Functional Check (Weight Check)
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26-20-00-780
Portable Fire Extinguisher Restoration (Hydrostatic Test)
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MODEL 208 MAINTENANCE MANUAL FIRE PROTECTION - GENERAL 1.
Scope A.
2.
This chapter provides maintenance of components which detect and alert flight crew to an engine fire.
Definition A.
This chapter is divided into sections to assist maintenance personnel in locating specific systems and information. The following is a brief description of each section. For locating information within the chapter, refer to the table of contents at the beginning of the chapter. (1) The section on fire detection provides information on the fire detection loop, fire detection control box, fire alarm, fire detector light and generator overheat detection. (2) The section on fire extinguishers provides description and operation information of hand fire extinguishers.
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MODEL 208 MAINTENANCE MANUAL ENGINE FIRE DETECTION SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
Fire protection consists of an engine fire detection system and a portable Halon 1211 fire extinguisher mounted in the pilot’s door. The fire detection system provides the means to detect a fire in the engine compartment and alert the crew by a visual audible indication.
Description and Operation A.
The fire detection system is installed in the engine compartment. The system consists of a flexible closed loop consisting of three section with high resistance at normal operating temperature, decreasing as temperature rises. The loop is connected to a control box which detects the change in resistance and triggers a warning light on the annunciator panel and an audible alarm when the temperature reaches a set point of 425°F on the first section, 625°F to 650°F on the second section and 450°F on the third section. The alarm also sounds through the pilot’s headphones.
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MODEL 208 MAINTENANCE MANUAL ENGINE FIRE DETECTION SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Fire Detection System Troubleshooting Chart Figure 101 (Sheet 1)
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Fire Detection System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL ENGINE FIRE DETECTION SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Maintenance practices for engine fire detection system consist of heat detection loop removal/ installation, control box removal/installation and alarm assembly removal/installation.
Heat Detection Loop Removal/Installation A.
Remove Heat Detection Loop (Refer to Figure 201 ). (1) Ensure BATTERY switch is OFF. (2) Remove left and right cowl doors. (3) Remove screws securing heat detection loop mounting flange to firewall. (4) Separate heat detection loop sections by loosening nuts. The loops with insulators snap out from clamps. (5) Disconnect electrical connectors from heat detection loop.
CAUTION: Extreme care should be exercised during maintenance not to twist, kink or dent the sensing loop cables. Minimum radius of bends may not be less than 0.50 inch. (6) (7) B.
Separate loop section by removing nuts. Remove loop sections from engine compartment. Remove and retain insulators from loop sections.
Install Heat Detection Loop (Refer to Figure 201). (1) Insert leads of loop through firewall.
CAUTION: Extreme care should be exercised during maintenance not to twist, kink or dent the sensing loop cables. Minimum radius of bends may not be less than 0.50 inch. (2) (3) (4) (5) (6) (7) 3.
4.
Attach heat detection loop mounting flange to firewall with screws. Ensure that there is a positive electrical contact between flanges and firewall. Attach each section of loop with brackets and nuts. Secure heat detection elements to engine mount assembly with clamps, and to engine flange brackets with screws using retained insulators. Connect electrical connectors to heat detection loop. Test system by activating TEST switch. Install left and right cowl doors.
Control Box Removal/Installation A.
Remove Control Box (Refer to Figure 201). (1) From inside cabin section under instrument panel, remove screws securing control box to firewall. (2) Disconnect control box electrical connector. (3) Remove control box from airplane.
B.
Install Control Box (Refer to Figure 201 ). (1) Install control box to firewall and secure with screws. (2) Connect electrical connector to control box.
Alarm Assembly Removal/Installation A.
Remove Alarm Assembly (Refer to Figure 201). (1) Remove screws securing alarm assembly to interior cabin top. (2) Disconnect electrical connector from alarm assembly. (3) Remove alarm assembly from airplane.
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Fire Detection System Figure 201 (Sheet 1)
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Fire Detection System Figure 201 (Sheet 2)
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Fire Detection System Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL B.
Install Alarm Assembly (Refer to Figure 201). (1) Install alarm assembly to interior cabin top and secure with screws. (2) Connect electrical connector to alarm assembly.
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MODEL 208 MAINTENANCE MANUAL ENGINE FIRE DETECTION SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the engine fire detection system in a serviceable condition.
Task 26-10-00-210 2.
Engine Fire Detection System General Visual Inspection A.
General (1) This task includes the steps necessary to do a general visual inspection of the engine fire detection system.
B.
Special Tools (1) None
C.
Access (1) Open the left and the right cowl doors. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices.
D.
Do a General Visual Inspection of the Engine Fire Detection Loop. (1) Examine all three fire detection loop sections for condition, kinks in the loop, and for evidence of damage. (2) Examine the loop fittings for loose connections and for cracking at the tube to connector solder joints. (3) Examine the loop section for proper routing, security of the clamps and clamp insulators. (4) Examine the clamps and clamp insulators for proper positioning on the tube. NOTE: (5)
E.
Examine the clamps and attachment hardware for condition and security.
Do a General Visual Inspection of the Engine Fire Detection Control Box. NOTE: (1) (2)
F.
The minimum loop bend radius is 0.50 inch (12.70 mm).
The fire detection control box is located on the aft side of the firewall.
Examine the fire detection control box for condition and security of installation. Examine the control box wire harness for condition, proper routing, and security of the wire harness connector.
Do a General Visual Inspection of the Engine Fire Detection Alarm Assembly. NOTE: (1) (2)
The fire detection alarm assembly is located on the cockpit ceiling.
Examine the fire detection alarm assembly for condition and security of installation. Examine the accessible wiring for condition, chafing, and security of the connectors.
G.
Restore Access (1) Close the left and the right cowl doors. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL PORTABLE FIRE EXTINGUISHING - DESCRIPTION AND OPERATION 1.
General A.
2.
Portable, hand operated Þre extinguishers are mounted in easily accessible locations for use in the event of a Þre. (Refer to Figure 1).
Description and Operation A.
One portable Þre extinguisher is mounted in the ßight compartment on the pilot’s door, or mounted between the pilot and copilot seats on the cargo barrier. The extinguishing agent is Halon 1211 and may be used on solid combustible, electrical or liquid Þres. Servicing of the extinguisher can be handled by most Þre equipment dealers. The Þre extinguishers are mounted within a quick release, clamp type bracket assembly. An optional additional portable Þre extinguisher can be installed at FS 257.0 between RBL 14.0 and RBL 23.5, installed on the seat tracks of the most aft right seat of the airplane forward of the main cabin door entry door. The Þre extinguisher and quick release bracket is identical to the one mounted in the ßight crew area.
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Portable Fire Extinguisher Installation Figure 1 (Sheet 1)
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Portable Fire Extinguisher Installation Figure 1 (Sheet 2)
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Portable Fire Extinguisher Installation Figure 1 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL PORTABLE FIRE EXTINGUISHING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the portable fire extinguishing system in a serviceable condition.
Task 26-20-00-290 2.
Portable Fire Extinguisher Restoration (Internal Inspection) NOTE:
One hand operated portable fire extinguisher is mounted in the flight compartment on the pilot’s door, or mounted between the pilot and the copilot seats on the cargo barrier.
A.
General (1) This task includes the steps necessary to do a restoration (internal inspection) of the portable fire extinguisher.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Portable Fire Extinguisher Restoration (Internal Inspection). (1) Remove the hand fire extinguisher from the quick release, clamp type bracket assembly. (2) Send the portable fire extinguisher to an approved fire extinguisher service facility for the internal inspection. (3) Install the hand fire extinguisher in the quick release, clamp type bracket assembly.
E.
Restore Access (1) None End of task Task 26-20-00-720 3.
Portable Fire Extinguisher Functional Check (Weight Check) NOTE:
One hand operated portable fire extinguisher is mounted in the flight compartment on the pilot’s door, or mounted between the pilot and the copilot seats on the cargo barrier.
A.
General (1) This task includes the steps necessary to do a functional check of the portable fire extinguisher.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Portable Fire Extinguisher Functional Check (Weight Check). (1) Remove the portable fire extinguisher from the quick release, clamp type bracket assembly. (2) Weigh the portable fire extinguisher. (3) Compare the weight of the portable fire extinguisher bottle to the weight shown on the NFPA or UL specification placard, attached to the portable fire extinguisher. NOTE:
It is possible for the pressure gage to show full pressure when the portable fire extinguisher is not fully charged.
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MODEL 208 MAINTENANCE MANUAL (4) (5)
Make sure that the weight of the portable fire extinguisher agrees with the weight shown on the NFPA or UL specification placard. If the portable fire extinguisher weight is less than the NFPA or UL specification placard weight, replace the portable fire extinguisher or send it to an approved fire extinguisher service facility.
E.
Restore Access (1) None End of task Task 26-20-00-780 4.
Portable Fire Extinguisher Restoration (Hydrostatic Test) NOTE:
One hand operated portable fire extinguisher is mounted in the flight compartment on the pilot’s door, or mounted between the pilot and the copilot seats on the cargo barrier.
A.
General (1) This task includes the steps necessary to do a restoration (hydrostatic test) of the portable fire extinguisher.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Portable Fire Extinguisher Restoration (Hydrostatic Test). (1) Remove the hand fire extinguisher from the quick release, clamp type bracket assembly. (2) Send the portable fire extinguisher to an approved fire extinguisher service facility for the hydrostatic test. (3) Install the hand fire extinguisher in the quick release, clamp type bracket assembly.
E.
Restore Access (1) None End of task
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27 CHAPTER
FLIGHT CONTROLS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT 27-Title 27-List of Effective Pages 27-Record of Temporary Revisions 27-Table of Contents 27-List of Tasks 27-00-00 27-00-00 27-10-00 27-10-00 27-10-00 27-10-00 27-10-01 27-10-02 27-10-02 27-10-02 27-10-02 27-20-00 27-20-00 27-20-00 27-20-01 27-20-02 27-20-02 27-20-02 27-20-03 27-30-00 27-30-00 27-30-00 27-30-01 27-30-01 27-30-02 27-30-02 27-30-02 27-30-02 27-30-03 27-30-03 27-30-03 27-30-03 27-31-00 27-31-00 27-31-00 27-50-00 27-50-00 27-50-00 27-50-01
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Pages 1-5 Pages 601-602 Page 1 Pages 101-103 Pages 501-516 Pages 601-604 Pages 201-220 Page 1 Pages 101-103 Pages 201-227 Pages 601-604 Page 1 Pages 101-103 Pages 601-603 Pages 201-210 Page 1 Pages 101-102 Pages 201-204 Pages 501-503 Page 1 Pages 101-102 Pages 601-603 Pages 201-211 Pages 501-518 Page 1 Pages 101-103 Pages 201-225 Pages 601-603 Page 1 Pages 101-102 Pages 201-206 Pages 501-507 Pages 101-103 Pages 201-209 Page 601 Page 1 Pages 101-105 Pages 601-606 Pages 201-230
Mar 1/2008 Jun 1/2011 Aug 1/1995 Aug 1/1995 Apr 1/2010 Jun 1/2011 Apr 1/2010 Jan 3/2005 Aug 1/1995 Jun 1/2011 Jun 1/2011 Aug 1/1995 Aug 1/1995 Jun 1/2011 Apr 1/2010 Aug 1/1995 Aug 1/1995 Apr 1/2010 Apr 1/2010 Aug 1/1995 Aug 1/1995 Jun 1/2011 Jul 1/2010 Mar 1/2008 Aug 1/1995 Aug 1/1995 Mar 1/2012 Mar 1/2012 Aug 1/1995 Aug 1/1995 Jul 1/2010 Jun 3/2002 Sep 4/2001 Jun 1/2011 Jun 1/2011 Aug 1/1995 Jun 1/1998 Jun 1/2011 Jan 3/2005
27 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL 27-50-01
Pages 501-505
Jun 1/2011
27-50-02
Pages 501-515
Jun 1/2011
27-70-00
Pages 201-204
Aug 1/1995
27-70-01
Pages 201-204
Sep 1/2000
27 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS FLIGHT CONTROLS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-00-00 27-00-00 27-00-00 27-00-00
Page 1 Page 1 Page 1 Page 2
FLIGHT CONTROLS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Controls Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-00-00 Page 601 27-00-00 Page 601
AILERON AND SPOILER SYSTEM - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-10-00 Page 1 27-10-00 Page 1 27-10-00 Page 1
AILERON AND SPOILER SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-10-00 Page 101 27-10-00 Page 101
AILERON AND SPOILER SYSTEM - ADJUSTMENT/TEST. . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Rigging (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Friction Band (Airplanes with 400B and 400B IFCS Autopilots Types AF550A and IF-550A Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Slip Clutch Adjustment, Values and Capstan (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) . . . . . . . . . . . . . . . . . . . . . . . . . KM 275 and KM 277 Slip Clutch Torque Adjustment (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed). . . . . . . . . . . . . Aileron Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Friction Band (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-10-00 Page 501 27-10-00 Page 501 27-10-00 Page 501 27-10-00 Page 506 27-10-00 Page 506 27-10-00 Page 506 27-10-00 Page 510 27-10-00 Page 516
AILERON AND SPOILER SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spoiler System Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron System Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-10-00 Page 601 27-10-00 Page 601 27-10-00 Page 601 27-10-00 Page 602
AILERONS AND CONTROL COLUMN - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Column Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Column Tube Assembly Bearings Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . Pulleys, Cables, Quadrants and Bell Cranks Removal/Installation . . . . . . . . . . . . . . . Rigging Aileron System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Friction Band Requirements. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-10-01 Page 201 27-10-01 Page 201 27-10-01 Page 201 27-10-01 Page 208 27-10-01 Page 208 27-10-01 Page 218 27-10-01 Page 220
AILERON TRIM SYSTEM - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-10-02 Page 1 27-10-02 Page 1
AILERON TRIM SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-10-02 Page 101 27-10-02 Page 101
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MODEL 208 MAINTENANCE MANUAL AILERON TRIM SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim System Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim System Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Actuator Disassembly (Airplanes with 2660044-1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Actuator Disassembly (Airplanes with 2661615-1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Actuator Disassembly (Airplanes with 2661615- 9 or 2661615-10 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection and Repair of Aileron Trim Tab Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2660044-1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2661615-1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2661615- 9 or 2661615-10 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection and Rigging Procedures - Aileron Trim Tab Actuator (Airplanes with 2660044-1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection and Rigging Procedures - Aileron Trim Tab Actuator (Airplanes with 2661615-1, 2661615-9, or 2661615-10 Trim Tab Actuator Installed) . . . . . . . . . . .
27-10-02 Page 201 27-10-02 Page 201 27-10-02 Page 201 27-10-02 Page 210
27-10-02 Page 226
AILERON TRIM SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab (Free Play) Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim System Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Actuator (2660044-1) Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Actuator (2661615-1, 2661615- 9, or 2661615-10) Lubrication. . .
27-10-02 Page 601 27-10-02 Page 601 27-10-02 Page 601 27-10-02 Page 601 27-10-02 Page 603 27-10-02 Page 604
RUDDER SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-00 Page 1 27-20-00 Page 1 27-20-00 Page 1
RUDDER SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-00 Page 101 27-20-00 Page 101
RUDDER SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder System Functional Check (Standard Rudder Installation) . . . . . . . . . . . . . . . . Rudder System Functional Check (Float Kit Installation) . . . . . . . . . . . . . . . . . . . . . . . . Rudder Bar Bearings and Rudder Pedals Lubrication. . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-00 Page 601 27-20-00 Page 601 27-20-00 Page 601 27-20-00 Page 602 27-20-00 Page 603
RUDDER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Pedals, Pulleys, and Cables Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . Rudder System Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-01 Page 201 27-20-01 Page 201 27-20-01 Page 201 27-20-01 Page 210
RUDDER TRIM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-02 Page 1 27-20-02 Page 1 27-20-02 Page 1
RUDDER TRIM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-02 Page 101 27-20-02 Page 101
RUDDER TRIM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim System Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim System Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-02 Page 201 27-20-02 Page 201 27-20-02 Page 201 27-20-02 Page 201
YAW DAMPER - ADJUSTMENT/TEST. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Yaw Damper Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-20-03 Page 501 27-20-03 Page 501
27-10-02 Page 211 27-10-02 Page 221 27-10-02 Page 221 27-10-02 Page 222 27-10-02 Page 222 27-10-02 Page 223 27-10-02 Page 224 27-10-02 Page 225
27 - CONTENTS © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ELEVATOR SYSTEM - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-00 Page 1 27-30-00 Page 1 27-30-00 Page 1
ELEVATOR SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-00 Page 101 27-30-00 Page 101
ELEVATOR SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator System Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-00 Page 601 27-30-00 Page 601 27-30-00 Page 601
ELEVATOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator System Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator System Rigging. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Friction Band Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-01 Page 201 27-30-01 Page 201 27-30-01 Page 201 27-30-01 Page 210 27-30-01 Page 211
ELEVATOR - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Rigging (Airplanes with 400B and 400B IFCS Autopilot Types AF-550A and IF-550A Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Friction Band (Airplanes with 400B and 400B IFCS Autopilot, Types AF550A and IF-550A Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Slip Clutch Adjustment, Values and Capstan (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) . . . . . . . . . . . . . . . . . . . . . . . . . KM 275 and KM 277 Slip Clutch Torque Adjustment (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed). . . . . . . . . . . . . Elevator Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Friction Band (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-01 Page 501 27-30-01 Page 501 27-30-01 Page 501 27-30-01 Page 508 27-30-01 Page 508 27-30-01 Page 512 27-30-01 Page 512 27-30-01 Page 517
ELEVATOR TRIM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-02 Page 1 27-30-02 Page 1 27-30-02 Page 1
ELEVATOR TRIM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-02 Page 101 27-30-02 Page 101
ELEVATOR TRIM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim System Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim System Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator Disassembly (Airplanes with 2660017-1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator Disassembly (Airplanes with 2661215-1 or 2661215-9 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator Inspection/Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator Lubrication and Assembly (Airplanes with 2660017- 1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator Lubrication and Assembly (Airplanes with 2661215-1 or 2661215-9 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator Inspection and Rigging (Airplanes with 2660017-1 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator Inspection and Rigging (Airplanes with 2661215-1 or 2661215-9 Trim Tab Actuator Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-02 Page 201 27-30-02 Page 201 27-30-02 Page 201 27-30-02 Page 210 27-30-02 Page 211 27-30-02 Page 211 27-30-02 Page 221 27-30-02 Page 221 27-30-02 Page 222 27-30-02 Page 223 27-30-02 Page 224
27 - CONTENTS © Cessna Aircraft Company
Page 3 of 5 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ELEVATOR TRIM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab (Free Play) Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator (2660017-1) Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Actuator (2661215-1 and 2661215-9) Lubrication . . . . . . . . . . . . .
27-30-02 Page 601 27-30-02 Page 601 27-30-02 Page 601 27-30-02 Page 601 27-30-02 Page 603
ELECTRIC ELEVATOR TRIM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-03 Page 1 27-30-03 Page 1 27-30-03 Page 1
ELECTRIC ELEVATOR TRIM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-03 Page 101 27-30-03 Page 101
ELECTRIC ELEVATOR TRIM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . Electric Elevator Trim System Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electric Elevator Trim System Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-03 Page 201 27-30-03 Page 201 27-30-03 Page 206
ELECTRIC ELEVATOR TRIM - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electric Trim Rigging (Airplanes with 400B and 400B IFCS Autopilots Types AF550A and IF-550A Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electric Elevator Trim Clutch Torque System Check (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) . . . . . . . . . . . . . . . . . . . . . . Electric Trim Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-30-03 Page 501 27-30-03 Page 501 27-30-03 Page 505 27-30-03 Page 506
STALL WARNING SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-31-00 Page 101 27-31-00 Page 101
STALL WARNING SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning Detector Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning Lift Transducer Removal/Installation For TKS Equipped Airplane . . . Stall Warning Thermostat Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning Horn Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning Horn Disconnect Switch Removal/Installation (Airplanes 20800316 and On and 208B0800 and On and Airplanes 20800001 thru 20800315 and 208B0001 thru 208B0799 incorporating CAB00-1). . . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning Flight Operational Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning Detector Adjustment (Airplanes 20800001 thru 20800056) . . . . . . . . . KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Audio Alert Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-31-00 Page 201 27-31-00 Page 201 27-31-00 Page 201 27-31-00 Page 201 27-31-00 Page 207 27-31-00 Page 207
STALL WARNING SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning System Operational Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-31-00 Page 601 27-31-00 Page 601 27-31-00 Page 601
FLAP SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-50-00 Page 1 27-50-00 Page 1 27-50-00 Page 1
FLAP SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-50-00 Page 101 27-50-00 Page 101
FLAP SYSTEM - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Actuator Mount Bracket Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Bellcrank Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap System Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Tracks and Rollers Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-50-00 Page 601 27-50-00 Page 601 27-50-00 Page 601 27-50-00 Page 601 27-50-00 Page 602 27-50-00 Page 603
27-31-00 Page 208 27-31-00 Page 208 27-31-00 Page 209 27-31-00 Page 209
27 - CONTENTS © Cessna Aircraft Company
Page 4 of 5 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FLAP SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Actuator Worm Gear Assembly Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . Flap Transmission Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Control Lever and Pointer Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Switch Actuator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Actuator Tube Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flaps Wing-to-Wing Interconnect Rod Assembly Removal/Installation. . . . . . . . . . . . Flap Interconnect Rods Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Connecting Rods Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Pushrods Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Cables and Pulleys Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Inboard Forward Bell Cranks Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . Flap Inboard Aft Bell Cranks Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Outboard Bell Cranks Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Actuator Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Switch Actuator Disassembly/Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Standby Flap Motor Switches Removal/Installation (Airplanes 20800001 thru 20800223 and 208B0001 thru 208B0326 not incorporating SK208-119A) . . . . . . Standby Flap Motor Switches Removal/Installation (Airplanes 20800224 and On, 208B0327 and On, 20800001 thru 20800223 and 208B0001 thru 208B0326 incorporating SK208-119A) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-50-01 Page 201 27-50-01 Page 201 27-50-01 Page 201 27-50-01 Page 201 27-50-01 Page 201 27-50-01 Page 201 27-50-01 Page 204 27-50-01 Page 204 27-50-01 Page 208 27-50-01 Page 210 27-50-01 Page 210 27-50-01 Page 213 27-50-01 Page 213 27-50-01 Page 216 27-50-01 Page 219 27-50-01 Page 221 27-50-01 Page 223 27-50-01 Page 223
FLAP SYSTEM - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Switch Actuator Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-50-01 Page 501 27-50-01 Page 501 27-50-01 Page 501
FLAP RIGGING GUIDE - ADJUSTMENT/TEST. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Component Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-50-02 Page 501 27-50-02 Page 501 27-50-02 Page 501 27-50-02 Page 504 27-50-02 Page 507
RUDDER GUSTLOCK - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Gustlock Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Gustlock Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Gustlock Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-70-00 Page 201 27-70-00 Page 201 27-70-00 Page 201 27-70-00 Page 201 27-70-00 Page 204 27-70-00 Page 204
RUDDER GUSTLOCK - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Gustlock Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Gustlock Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27-70-01 Page 201 27-70-01 Page 201 27-70-01 Page 201 27-70-01 Page 201 27-70-01 Page 201
27-50-01 Page 227 27-50-01 Page 230
27 - CONTENTS © Cessna Aircraft Company
Page 5 of 5 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 27-00-00-640
Flight Controls Lubrication
27-00-00 Page 601
27-10-00-720
Spoiler System Functional Check
27-10-00 Page 601
27-10-00-721
Aileron System Functional Check
27-10-00 Page 602
27-10-02-720
Aileron Trim Tab (Free Play) Functional Check
27-10-02 Page 601
27-10-02-640
Aileron Trim System Lubrication
27-10-02 Page 601
27-10-02-641
Aileron Trim Tab Actuator (2660044-1) Lubrication
27-10-02 Page 603
27-10-02-642
Aileron Trim Tab Actuator (2661615-1, 2661615- 9, or 2661615-10) Lubrication
27-10-02 Page 604
27-20-00-720
Rudder System Functional Check (Standard Rudder Installation)
27-20-00 Page 601
27-20-00-721
Rudder System Functional Check (Float Kit Installation)
27-20-00 Page 602
27-20-00-640
Rudder Bar Bearings and Rudder Pedals Lubrication
27-20-00 Page 603
27-30-00-720
Elevator System Functional Check
27-30-00 Page 601
27-30-02-720
Elevator Trim Tab (Free Play) Functional Check
27-30-02 Page 601
27-30-02-640
Elevator Trim Tab Actuator (2660017-1) Lubrication
27-30-02 Page 601
27-30-02-641
Elevator Trim Tab Actuator (2661215-1 and 2661215-9) Lubrication
27-30-02 Page 603
27-31-00-710
Stall Warning System Operational Check
27-31-00 Page 601
27-50-00-220
Flap Actuator Mount Bracket Detailed Inspection
27-50-00 Page 601
27-50-00-221
Flap Bellcrank Detailed Inspection
27-50-00 Page 601
27-50-00-720
Flap System Functional Check
27-50-00 Page 602
27-50-00-640
Flap Tracks and Rollers Lubrication
27-50-00 Page 603
27 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FLIGHT CONTROLS - GENERAL 1.
Scope A.
2.
This chapter provides maintenance of components which furnish a means of manually controlling the ßight attitude characteristics of the airplane, including ßaps and spoilers.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME Digital Inclinometer
NUMBER PRO 360
MANUFACTURER
USE
Kell-Strom Tool Co. Inc. 214 Church Street WethersÞeld, CT 06109
To measure angle of control surface deßection.
Rig Pin Tool
Locally fabricate (NOTE 1)
To secure control surfaces in a Þxed position.
Elevator Rigging Protractor
Locally fabricate (NOTE 1)
To measure angle of elevator.
Rudder Rigging Protractor
Locally fabricate (NOTE 1)
To measure angle of rudder.
Elevator Neutral Rigging Tool
SE1511
Cessna Aircraft Company Cessna Parts Distribution Department 701, CPD 2 5800 East Pawnee Road Wichita, KS 67218-5590 (P.O. Box 7704 Wichita, KS 67277-7704)
To rig elevator.
Tensiometer
T52002101
PaciÞc ScientiÞc Electro Kinetics Div. 402 E. Guitierrez St. Santa Barbara, CA 93102
To measure and obtain cable tension.
Hunter Spring Gage Type L-30
Ametek U.S. Gauge 900 Clymer Ave. Sellersville, PA 18960
To measure friction forces.
Dial Indicator (0 to 1.00 inch)
Commercially Available
To check trim tab free play.
Grease
B100-24
Commercial Aircraft Products Inc. 2633 West Pawnee Wichita, KS 67213
To lubricate ßap actuator.
Molykote Bonded Lubricant Spray
EC321R
Cessna Aircraft Company
To lubricate aileron trim system cables.
Light Consistency Silicone Grease
5565450-28
Cessna Aircraft Company
To lubricate aileron trim tab actuator components.
Iridate
Type 14-2 Powder
Witco Allied-Kelite Div. of Witco Chemical Corp. 400 Midland Avenue Highland Park, MI 48203
To Þnish elevator trim tab actuator housing.
27-00-00 © Cessna Aircraft Company
Page 1 Mar 1/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Non-Chromated Primer
Type I-P, E9191
Sterling Lacquer Mfg. 3150 Brannon Ave St. Louis, MO 63139
To prime elevator trim tab actuator housing.
Vivid Orange
SherwinWilliams Number 17H54958 335-566
Commercially available
To paint elevator trim tab actuator
Vesatal White Lacquer
SherwinWilliams Number 27H511
Commercially available
To paint elevator trim tab actuator
Locally fabricate
To check ßap switch actuation.
Flap Switch Actuator Rigging Fixture
NOTE 1: Refer to Figure 1 for fabrication details. 3.
DeÞnition A.
This chapter is divided into sections and subsections to assist maintenance personnel in locating speciÞc systems and information. The following is a brief description of each section. For locating information within the chapter, refer to the Table of Contents at the beginning of the chapter. (1) The aileron section provides information on control wheels, cables, linkage and aileron assemblies. (2) The rudder section provides information on rudder pedals, cables, linkage and rudder assembly. (3) The elevator section provides information on control column, cables, linkage and elevator assemblies. (4) The stall warning section provides information on mechanical and electrical components used in the stall warning system. (5) The ßap section provides information on the ßap actuator, cables, linkage, and the ßap assemblies. (6) The rudder gust lock section provides information on components used in the gust lock system.
27-00-00 © Cessna Aircraft Company
Page 2 Mar 1/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Rigging Tools Fabrication Figure 1 (Sheet 1)
27-00-00 © Cessna Aircraft Company
Page 3 Mar 1/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Rigging Tools Fabrication Figure 1 (Sheet 2)
27-00-00 © Cessna Aircraft Company
Page 4 Mar 1/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Rigging Tools Fabrication Figure 1 (Sheet 3)
27-00-00 © Cessna Aircraft Company
Page 5 Mar 1/2008
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FLIGHT CONTROLS - INSPECTION/CHECK Task 27-00-00-640 1.
Flight Controls Lubrication A.
General (1) This task provides the procedures to perform a Lubrication of the Flight Controls. NOTE:
B.
For more flight controls lubrications, Refer to Chapter 12, Flight Controls - Servicing.
Special Tools NOTE: (1) (2) (3) (4)
Equivalent tools and equipment can be used.
Oil - MIL-L-7870 Grease - MIL-G-21164 Grease - MIL-G-81322 Dry Solid Film Lubricant - MIL-L-23398.
C.
Access. (1) Remove floor panels 216AC and 216BC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Remove pedestal panels 226A, 226B, 226C, and 226D. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Remove vertical stabilizer panels 373AL and 374AR. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (4) Remove wing panels 503EB, 525AB, and 525CB left, and 603EB, 625AB, and 625CB right. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do a Lubrication of the Left and the Right Aileron and Spoiler Pushrods. (1) Disconnect the left and the right wing aileron and spoiler pushrods. Refer to Ailerons and Control Column - Maintenance Practices. (2) Examine for corrosion, condition, and pitting. (3) Lubricate by hand with MIL-G-21164 grease. (4) Connect the left and the right wing aileron and spoiler pushrods. Refer to Ailerons and Control Column - Maintenance Practices.
E.
Do a Lubrication of the Left and the Right Spoiler Hinges. (1) Wipe and clean the left and right spoiler hinge assemblies with a lint-free cloth. (2) Lubricate the left and the right spoiler hinge assemblies with dry solid film lubricant. (3) Wipe off excess lubricant.
F.
Do a lubrication of the Rudder Trim Control. (1) Turn the trim control wheel fully left or right. (2) Wipe the threads of the rudder trim shaft with a lint-free cloth. (3) Turn the trim control in the full opposite direction. (4) Wipe the threads of the rudder trim shaft with a lint-free cloth. (5) Lubricate the trim shaft, nut, and link with MIL-L-7870 oil. (6) Lubricate the trim wheel support bearing from the left side of pedestal with MIL-L-7870 oil. (7) Wipe off the excess oil.
G.
Do a Lubrication of the Elevator Trim Control. (1) Lubricate the elevator trim control wheel shaft and the support bearing with dry solid film lubricant. (2) Lubricate both sprocket shafts under the floor with dry solid film lubricant. (3) Wipe the cable chain with clean a lint-free cloth, but do not lubricate.
H.
Do a lubrication of the Elevator Trim Actuator Pushrods (Left and Right). (1) Lubricate the pushrods at the actuator and the trim tab horn with MIL-L-7870 oil.
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Do a lubrication of the Elevator Trim Cable Chains (Left and Right Stabilizer). (1) Wipe the left and the right chains with a clean lint-free cloth. NOTE: (2)
J.
Do not lubricate the chains unless you operate the airplane in a seacoast condition. Lubricant can cause dust and dirt to collect and cause the links to bind.
If you operate the airplane in seacoast conditions, apply a light coat of MIL-L-7870 oil to the chains for corrosion protection.
Do a lubrication of the Left and the Right Flap Outboard Bell Crank Bearings. NOTE:
(1) (2)
Airplanes 20800161 and On, 208B0190 and On, and airplanes that incorporate SNL8917 have sealed bearings installed and do not require lubrication. To identify these bell cranks, measure the bell crank mount tube outside diameter. The new bell crank mount tube outside diameter is 1.00 inch. The initial bell crank mount tube outside diameter is 0.687 and must be removed for lubrication.
Remove the flap bell cranks from the wings. Refer to Flap System - Maintenance Practices. Remove the bearings from the bell cranks. (a) Clean and examine the bearing for corrosion, condition, and pitting. NOTE:
(3) (4)
If bearing is found unserviceable, you can replace the bell crank with sealed bearings in accordance with SNL89-17.
Install the bearings in the bell cranks Install the bell cranks in the wing. Refer to Flap System - Maintenance Practices.
K.
Restore Access. (1) Install wing panels 503EB, 525AB, and 525CB left, and 603EB, 625AB, and 625CB right. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Install vertical stabilizer panels 373AL and 374AR. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Install pedestal panels 226A, 226B, 226C, and 226D. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (4) Install floor panels 216AC and 216BC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task
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MODEL 208 MAINTENANCE MANUAL AILERON AND SPOILER SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
This section describes those systems and components used to control the ailerons. Included are control columns, ailerons, spoilers and aileron trim systems.
Description and Operation A.
The aileron and spoiler system includes ailerons, right aileron servo operated trim tab, left aileron servo tab and left and right spoiler. A left and optional right control column contain a control wheel, control tube, bearings and a quadrant. An interconnect cable attaches left control wheel to right control wheel. the aileron control cable is divided into two loops. The low tension loop, located in the fuselage, is routed from left quadrant, under floorboards and up the left sidewall to a bellcrank in the cabin top. The loop runs from bellcrank across cabin top to right sidewall, down the sidewall, under floorboards and back to the left quadrant. The high tension loop is located in cabin top and left and right wings. It interconnects bellcrank in cabin top with left and right wing bellcranks. Aileron pushrods and spoiler pushrods connect wing bellcranks with ailerons and spoilers. The spoiler moves down three to four degrees during total aileron down travel. the spoiler moves up slightly during first five degrees of up aileron travel, then more proportionate during remainder of up aileron travel.
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MODEL 208 MAINTENANCE MANUAL AILERON AND SPOILER SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Aileron and Control System Troubleshooting Chart Figure 101 (Sheet 1)
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Aileron and Control System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL AILERON AND SPOILER SYSTEM - ADJUSTMENT/TEST 1.
General A.
2.
This section has procedures for the 400B Autopilot, 400B Integrated Flight Control System, KAP-150 Autopilot, KFC-150 and KFC-225 Flight Control System. The following procedures give instructions for aileron rigging and for the aileron friction band check.
Aileron Rigging (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) NOTE: A.
For the correct operation of the 400B Autopilot and 400B IFCS, the autopilot rigging must be examined in accordance with Chapter 5, Inspection Time Limits.
Rig the Ailerons (Refer to Figure 501). (1) Set the control wheels so the ailerons are in a neutral position. (2) Set a bar across the top of the control wheels and tape the bar to the control wheels. The bar connects them together and locks them in the neutral position. (3) Remove the pilot and copilot seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (4) Remove the floor covering and access plates to get access to the roll actuator and cable assemblies. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (5) Make sure the actuator support assembly is correctly riveted to the longeron at rib-line 19.00. (6) Make sure the electrical connector is correctly attached. (7) Make sure the actuator motor is attached to the actuator mount with four bolts and washers. NOTE: (8) (9)
The aileron actuator cable assembly is also utilized as the aileron interconnect cable (actuator/interconnect cable assembly).
Make sure the actuator mount is correctly attached to the actuator support assembly with four bolts and washers. Make sure the aileron actuator/interconnect cable assembly is correctly adjusted by the following inspection procedures. (a) Make sure the right hand clamp assembly is correctly set in position on the aileron cable assembly. It must be four inches outboard from the center of the actuator sprocket. (b) Make sure the upper and lower right hand clamp halves are attached together with two bolts, washers, and nuts. (c) Make sure the right hand aileron cable is routed through the small cable slot. (d) Make sure the actuator/interconnect cable terminal is routed through the large cable slot and set in position with two roll pins. (e) Make sure the chain portion of the actuator/interconnect cable assembly is closely centered on the actuator sprocket. (f) Make sure the chain guard posts have safety wire attached and are correctly set in position next to the actuator sprocket.
(g) (h) (i)
NOTE:
When the chain guard posts are loosened for removal, do not lose the lock washers.
NOTE:
To replace the actuator/interconnect cable assembly, you must first remove the safety wire attached to the two chain guard posts and remove the chain guard posts.
Make sure the chain part of the cable assembly is attached to the cable terminals with two chain link fasteners. Make sure the cable part of actuator/interconnect assembly is correctly routed through the pulley mount brackets and centered on the pulley. Make sure the two pins are installed correctly to the pulley mount brackets to safety the cable portion of the actuator/interconnect cable assembly.
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Aileron Rigging (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) Figure 501 (Sheet 1)
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Aileron Rigging (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) Figure 501 (Sheet 2)
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Aileron Rigging (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) Figure 501 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (j)
Make sure the two pulley mount brackets are correctly attached to the bulkhead at station 128.00. (k) Make sure the pulley is correctly attacked to the pulley mount brackets with a bolt, washer and nut. (l) Make sure the upper and lower left hand clamp halves are correctly attached together with two each bolts, washer and nut. (m) Make sure the left hand aileron cable is routed through the large cable slot. (n) Make sure the actuator interconnect cable is routed through the small cable slot. (10) Do an inspection of the cable tension on the actuator/interconnect cable assembly. (a) Make sure the aileron actuator/interconnect cable assembly is rigged correctly. (b) Attach the cable tensiometer to the actuator/interconnect cable. The tensiometer must show a value of 10 to 15 pounds. (c) If the cable tension is less than 10 pounds or more than 15 pounds, adjust the cables to the correct tension. 1 Loosen the bolts that attach the left hand clamp assembly. 2 To reduce the cable tension, move the left hand clamp inboard. To increase the cable tension, move the left hand clamp outboard. 3 When the correct tension is shown, tighten the bolts. (d) Remove the tensiometer from the actuator/interconnect cable. (e) Remove the bar attached to the control wheels. (11) To remove the gyro roll and pitch signals made by a non-erect gyro, the 400B autopilot has a GYRO switch located on the rear of the control head. Set the GYRO switch to the OUT position. The 400B lFCS does not have a gyro out switch. It will be necessary to hook up an outside vacuum source or operate the engine to erect the gyro. NOTE:
If an outside vacuum source is used, it must be calibrated in inches of mercury. The desired suction range necessary to erect the gyro is 4.6 to 5.4 In.Hg.
NOTE:
If the engine is operated to erect the gyro, the engine must be operated at 65 percent Ng to provide the amount of vacuum required and maintain the correct bus voltage.
(12) Turn the airplane battery switch and the 1 & 2 avionics power and autopilot (AP) switch to ON. (13) Use the autopilot's PULL-TURN knob (pull out on 400B autopilot), and visually check the operation of the aileron travel in each direction. NOTE:
After several seconds, the ailerons will move slowly in the correct direction. The aileron travel does not need to have full travel in each direction. The typical control wheel travel will be full travel.
(14) Do a check of the autopilot so that it may be overpowered using the control wheels at any time. This procedure must be kept to a minimum since the extended manual operation will overheat the actuator, causing a thermostatic switch to remove power from the actuator. (15) Turn OFF the autopilot (AP) switch, avionics power switches and the airplane battery switch. (16) Install the access plates and carpeting. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
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3.
Aileron Friction Band (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) A.
Do a check of the aileron friction band. NOTE:
(1)
(2) (3)
4.
The aileron friction band without an autopilot installed must be six pounds or less. With an autopilot installed, the aileron friction band must be eight pounds or less. The friction band must be measured and calculated each time the autopilot rigging is examined. Corrective action to decrease the aileron system friction band must be done if it is higher than the band range.
All friction measurements must be made with a load scale so that the force exerted to move the aileron is applied to and in place of the control wheel assembly. The load scale must be attached to control wheel assembly at the longest possible moment arm (inside the grip). The friction band requirements apply over the complete travel of the ailerons. (a) Turn the control wheel to approximately 30 degrees in a counter clockwise direction and attach the load scale. (b) Apply a force to turn the control wheel in a clockwise direction. Write down the scale value as the control wheel passes the aileron neutral position. Identify this force as F1. (c) Remove the load scale. (d) Turn the control wheel approximately 30 degrees in the clockwise direction. Attach the load scale and exert a force to rotate the control wheel in a counter clockwise direction. Write down the scale value as the control wheel passes the aileron neutral position. Identify this force as F2. (e) Remove the load scale. The aileron friction band is calculated by adding the values of F1 and F2. The friction band = F1 + F2. When the friction band is greater than the limits of 7 pounds, the following steps must be completed to reduce the system friction to an acceptable level. (a) Do a check of the aileron direct, aileron return, interconnect, and rudder cables for clearance. Remove all interference. (b) Decrease the tension on the aileron cable (minimum cabin loop tension is 15 pounds, minimum wing loop tension is 35 pounds) with ailerons in the neutral position. (c) Decrease the aileron actuator/interconnect cable tension to 10 to 15 pounds with the ailerons in the neutral position. (d) Do a check of the pulley alignment and adjust as necessary.
Slip Clutch Adjustment, Values and Capstan (Airplanes with KAP-150 Autopilot and KFC-150, KFC225 Flight Control System Installed) A.
Servo Slip Clutch Torque Settings (Refer to Figure 502). NOTE: (1)
The servo slip clutch torque settings are adjustable and must be set before servo installation.
Set the servo slip clutch torque to the appropriate value. NOTE:
5.
The fixtures and tools required to complete the adjustments are supplied with the KTS 150 test set and the KTS 158 test set.
KM 275 and KM 277 Slip Clutch Torque Adjustment (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) A.
Adjust the Slip Clutch Torque (Refer to Figure 502). (1) Refer to Slip Clutch Adjustment, Values and Capstan, to determine the servo mount part number and setting required for each axis of the airplane. (2) Remove the capstan guard from the KM 275 and KM 277 capstan plate.
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Clutch Adjustment Figure 502 (Sheet 1)
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Clutch Adjustment Figure 502 (Sheet 2)
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Clutch Adjustment Figure 502 (Sheet 3)
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Attach the KM 275 or KM 277 servo mount and servo motor to the KTS 158 test stand (0474238-01). (a) When you adjust a KM 275, place the adapter tool over the KM 275 capstan and insert the positioning pin (from the straight-up position) to attach the adapter tool. (b) When you adjust a KM 277, use the three sprocket pins (071-6065-00) to attach the adapter tool to the capstan. NOTE:
(4) (5) (6)
An alternative adjustment method for the KM 277 is to use the King gear adapter assembly (071-6018-06).
Insert a torque wrench (Snap-On TEP-6FUA or equivalent). Connect the servo motor to the appropriate KTS 158 Test Set connector and apply power to the servo motor. Do a test of the torque value. NOTE:
The desired torque value is the average of the maximum and minimum indications from the clockwise and counterclockwise rotations. The test must be repeated three times in each direction and then the average of the six values is used to determine the true torque value.
Use the appropriate switch on the KTS 158 Test Set and turn the servo motor in the clockwise direction. Write down the torque value shown on the wrench. (b) Use the appropriate switch on the KTS 158 Test Set and turn the servo motor in the counter clockwise direction. Write down the torque value shown on the wrench. (c) If the level measured is less than the desired value, rotate the clutch adjust nut clockwise. (d) If the level measured is more than the desired value, rotate the clutch adjust nut counterclockwise. After an adjustment, repeat the torque test. After wiring has been completed and the servos installed, make sure the rotation direction of the servo capstans is correct. (a)
(7) (8) 6.
Aileron Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) A.
Adjust the Ailerons (Refer to Figure 503). NOTE:
(1) (2) (3) (4)
To make sure of the correct operation of the KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System, the autopilot rigging must be checked in accordance with Chapter 5, Inspection Time Limits.
Set the control wheels so the ailerons are in a neutral position. Set a bar across the top of the control wheels and tape the bar to the control wheels. The bar connects them together and locks them in the neutral position. Remove the pilot and copilot seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. Remove the floor covering and access plates to get access to the roll servo and cable assemblies. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. NOTE:
(5) (6)
The KS 271A Roll Servo and KM 275 Servo Mount are located under the cabin floor on the right side of the airplane at FS 148.0.
Adjust the slip clutch of the KM 275 capstan. Refer to Slip Clutch Torque Adjustment. Make sure the KM 275 Servo Mount is correctly attached to the roll servo bracket.
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Aileron Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) Figure 503 (Sheet 1)
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Aileron Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) Figure 503 (Sheet 2)
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Aileron Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) Figure 503 (Sheet 3)
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Aileron Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) Figure 503 (Sheet 4)
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Make sure the servo mount and roll servo bracket are installed in the airplane under the right main aileron cable. NOTE:
The inboard servo bracket must be 1.88 inches aft of FS 143.0. The top of the KM 275 Capstan must be 0.25 to 0.50 inch below the right main aileron cable.
(8)
Make sure the roll bracket support is installed against the inboard side of the existing triangle box brace located to the right and the KM 275 capstan. The flange of the roll bracket support containing the two slots must be aligned over two nut plates existing in the triangle box brace. The longest flange of the roll bracket support must be against the forward side of the roll servo bracket. Make sure the roll bracket support is attached to the triangle box brace with two bolts and two washers. (9) Make sure the bridle cable is properly aligned in the bridle cable idler pulley at FS 131.0 and LBL 14.0 and the pulley correctly installed with one bolt, washer and nut. (10) With the ailerons in the neutral position, locate the KM 275 capstan ball straight up. (11) Make sure the roll bridle cable is wrapped around the KM 275 capstan. NOTE:
The short end of the roll bridle cable must be routed aft along the right main aileron cable. The long end of the roll bridle cable must be routed forward along the main aileron cable and around the idler pulley, then routed aft along the left main aileron cable.
(12) Make sure the ends of the roll bridle cable are correctly attached to the main aileron cables with four cable clamps, four bolts, eight washers and four nuts.
CAUTION: The roll bridle cable must not rub against any of the stringers and bulkheads. (13) Do an inspection of the roll bridle cable to make sure it passes through all the stringers and bulkheads correctly. Do an inspection of the cable for corrosion and worn areas on the cable. If necessary, increase the cable openings in the stringers and bulkheads to allow the roll bridle cable to travel without touch. (14) Install a tensiometer on the bridle cable. (a) Make sure the tensiometer on the cable shows 12.0 pounds, +5 or -5 pounds. (b) Make sure the torque on the four cable clamp nuts is 50 inch-pounds, +5 or -5 inch-pounds. (15) Do a check of the roll bridle cable route around the idler pulley at FS 131.0. Make sure the two cable guard pins for the main aileron cables give protection to the roll bridle cable idler pulley. (16) Remove the tensiometer. (17) Remove the tape and the bar attached to the control wheels. (18) Install the pilot and copilot seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (19) Install the floor cover and access plates. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
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7.
Friction Band (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) A.
Do a check on the aileron friction band. NOTE:
(1)
(2) (3)
The aileron friction band without an autopilot installed must be six pounds or less. With an autopilot installed, the aileron friction band must be eight pounds or less. The friction band must be measured and calculated each time the autopilot rigging is examined. Corrective action to decrease the aileron system friction band must be done if it is higher than the band range.
All friction measurements must be made with a load scale so that the force exerted to move the aileron is applied to and in place of the control wheel assembly. The load scale must be attached to the control wheel assembly at the longest possible moment arm (inside the grip). The friction band requirements apply over the complete travel of the ailerons. (a) With control wheel rotated to approximately 30 degrees in the counter clockwise direction, attach a load scale and exert a force to turn the wheel in a clockwise direction. Take the scale value as the control wheel passes the aileron neutral position. Identify this value as F1. (b) With the control wheel rotated to approximately 30 degrees in the clockwise direction, attach load scale and exert a force to turn the wheel in a counter clockwise direction. Take the scale value as the control wheel passes the aileron neutral position. Identify this value as F2. The aileron friction band is calculated by adding the measured values F1 and F2. Friction Band = F1 + F2. When the friction band is more than the limitations of six pounds or less without an autopilot, or eight pounds or less with an autopilot, do the steps below to decrease the system friction to an satisfactory level. (a) Do a check of the aileron direct cable, aileron return cable, interconnect cable, and rudder cable for clearance and remove any interference. (b) Decrease the tension on the aileron cable with the ailerons in the neutral position. The tension of the minimum cabin loop to 15 pounds. The tension of the minimum wing loop to 35 pounds. (c) Decrease the tension on the aileron actuator/interconnect cable with the ailerons in the neutral position. The acceptable range is 10 to 15 pounds. (d) Do a check and adjust pulley alignment as necessary.
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MODEL 208 MAINTENANCE MANUAL AILERON AND SPOILER SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the aileron and spoiler system in a serviceable condition.
Task 27-10-00-720 2.
Spoiler System Functional Check A.
General (1) This task gives the procedures to do a functional check of the spoiler system.
B.
Special Tools (1) Inclinometer (2) Cable Tensiometer
C.
Access (1) Remove the applicable wing panels to get access to spoiler components, Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Open (unzip) the fabric headliner (passenger) or remove the hard shelled headliner (cargo) to get access to the spoiler components. Refer to Chapter 25, Cabin Upholstery - Maintenance Practices.
D.
Do a functional check of the spoiler system. (1) Do a check of the cable movement for binding and full travel. (2) Examine the spoiler skins for loose rivets and cracks. (3) Examine the hinges for corrosion, condition, and cracks. (a) Examine the bearings and bonding jumpers for signs of damage or wear, unserviceable fasteners, and security of installation. (4) Examine the bolts and nuts at both ends of pushrods for correct cotter pin installation.
CAUTION: If the pushrod will not turn using hand force, remove the rod end attach bolt and examine for cause. Make sure that the rod ends are aligned to let the rod turn a small amount when installed (vertical plane of each rod end in-line with each other). (5) (6) E.
Examine the aileron/spoiler bellcrank tubes, bearings, pushrods, stop bolts, and brackets for corrosion, cracks, signs of damage, failed fasteners, security of installation, and correct safetying. Examine the attachment brackets on each spoiler for corrosion, condition, cracks, security, and correct attachment of the cable to the bracket.
Do a spoiler rigging check on the left and the right spoilers. (1) With the flaps at the full up position, slowly turn the control wheel and examine for a minimum of 0.010 inch (0.254 mm) to a maximum of 0.030 inch (0.762 mm) clearance between spoiler trailing edge and the top surface of the flap at the minimum position. This will occur before the aileron reaches the full down position. (2) With the aileron at the neutral position, install an inclinometer on left spoiler and adjust it to zero. (3) Install a rig pin in the upper quadrant and the lower quadrant. (4) With the ailerons held in the neutral position, and flaps fully retracted, make sure that the trailing edge of the spoiler is 0.55 inch +0.05 or -0.05 inch (13.97 mm +1.27 or -1.27 mm) above the surface of the flap at the outboard end of the spoiler. (5) Remove the rig pins from the upper quadrant and the lower quadrant. (6) Use the control wheel to raise the left spoiler to its full up position. (a) The inclinometer must read 40 +5 or -5 degrees. (7) With the aileron at the neutral position, install an inclinometer on the right spoiler and adjust it to zero.
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Use the control wheel to raise the right spoiler to its full up position. (a) The inclinometer must read 40 +5 or -5 degrees. NOTE:
If the system is found to be out of tolerance, perform all adjustments, Refer to Aileron and Spoiler - Maintenance Practice. Include aileron friction band check. Ensure all rigging pins are removed after this task is complete.
F.
Restore access. (1) Install the applicable panels and covers that were removed to get access to the spoiler components on both wings. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. (2) Close (zip) the fabric headliner (passenger) or install the hard shelled headliner (cargo). Refer to chapter 25, Cabin Upholstery - Maintenance Practices. End of task Task 27-10-00-721 3.
Aileron System Functional Check A.
General (1) This task gives the procedures to do a functional check of the aileron system.
B.
Special Tools (1) Inclinometer (2) Cable Tensiometer (3) Spring scale measuring from 0 to 20 pounds
C.
Access (1) Remove panels 212FR, 226B, 231BL, 231CL, 251CL, 251DL, 252BR, 252FR, 501BB, 501CB, 501DB, 501EB, 503AB, 503BB, 503CB, 503DB, 503EB, 601BB, 601CB, 601DB, 601EB, 603AB, 603BB, 603CB, 603DB, 603EB, and 651AB to get access to the aileron components. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Open (unzip) the fabric headliner (passenger) or remove the hard shelled headliner (cargo) to get access to the aileron components. Refer to Chapter 25, Cabin Upholstery - Maintenance Practices.
D.
Do a Functional Check of the Aileron System. (1) Examine the aileron control and the aileron trim tab system cable movement for binding and full travel. (2) Examine the aileron skins for corrosion, cracks, and loose rivets. (3) Examine the aileron trim tab skins for corrosion, cracks, and loose rivets. (4) Examine the aileron trim tab stop blocks (right wing) for corrosion, condition, and security of installation. (5) Examine the aileron trim tab control and indicator for corrosion, condition, and security of installation. (6) Examine the balance weights for looseness and the supporting structure for corrosion, cracks, and failed fasteners. (7) Examine the aileron hinges, hinge bolts, hinge bearings, attach fittings, horn, and bonding jumper for corrosion, cracks, signs of damage, wear, failed fasteners, security, and correct safeftying. (8) Examine the bellcracks in both wings and above headliner and the bearings, push rods, stop bolts, and brackets, for corrosion, cracks, signs of damage, failed fasteners, security of installation, and correct safetying. (9) Examine the aileron and aileron trim cable runs for correct routing, fraying, and twisting. (a) Make sure there is no interference with the adjacent structure, equipment, wiring, plumbing, and other controls. (10) Move a cloth along the full length of the cables to examine for broken wires. (a) If snags are found or you think that there are broken wires, refer to Chapter 20, Control Cable Wire Breakage and Corrosion Limitations - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL (11) Examine the turnbuckles for correct thread exposure. (a) Make sure that the turnbuckle locking clips are installed correctly. Refer to Chapter 20, Safetying - Maintenance Practices. (12) Examine the swage fittings reference marks for an indication of cable slippage inside of the fitting. (a) Examine the fittings for corrosion, distortion, cracks, and broken wires at the fittings. (13) Examine the pulleys, attach brackets, and guard pins for condition, wear, corrosion, and security. (a) You must turn the pulleys to make sure there freedom of movement and to make sure there is even wear of the pulleys. (b) If discrepancies are found with the brackets, examine the structure where the brackets are attached for hidden damage. (14) Examine the aileron trim tab actuators for corrosion, damage, and security. (15) Examine the aileron trim tab actuator mounting structures for corrosion, damage, cracks, and security of installation. (16) Examine the aileron trim tab actuator pushrods and attaching hardware for corrosion, condition, damage, wear, and security of installation. (17) Examine the chains for corrosion, tension, and correct alignment. (18) Examine the aileron trim control wheel bearings for wear. (19) Examine the control wheel for condition and security of installation. (20) Examine the control column for corrosion, signs of damage, failed fasteners, and security of installation. (21) Examine all welds in the column tube and the torque tube for corrosion and cracks. (22) Examine both torque tube support arms for corrosion, condition, and security of the attach bearings. (23) Examine the support arm attach structure for corrosion, condition, cracks, and correct safety of the attach bolts. (24) Examine the cable guards for corrosion, condition, and security on both column quadrants. (25) Examine for sufficient clearance of all components and structure at the full aft and the full forward positions. (26) Make sure that the chain is correctly centered and aligned on the sprocket. (a) The chain guard posts must be correctly installed and attached with safety wire. (27) Make sure that the chain is correctly attached to the cable assembly and turnbuckle terminal with the chain connecting links. E.
Examine the Cable Travel and Tensions. (1) Set the control wheels to put the ailerons in the neutral position. (a) Make sure that the ailerons are streamlined with the inboard trailing edge of the aileron aligned with the outboard trailing edge of the flap. (2) Attach an inclinometer on the left aileron and set it to zero degrees. (3) Examine the cable tensions and adjust if necessary. (a) For the aileron control cables, refer to Ailerons and Control Column - Maintenance Practices. (b) For the aileron trim cables, refer to Aileron Trim System - Maintenance Practices. (c) For airplanes equipped with 400B and 400B IFCS autopilot type AF-550A and IF-550A, refer to Aileron and Spoiler System - Adjustment/Test. (4) Operate the system through its full range of travel. (a) Make sure that all of the components that move do not hit, touch, or catch on structural components or other system components. (5) Turn the control wheel so that the stop bolt touches the right bellcrank. (a) Make sure that the inclinometer shows 25 +4 or -0 degrees up travel on the left aileron. (6) Turn the control wheel so that the stop bolt touches the left bellcrank. (a) Make sure that the inclinometer shows 16 +1 or -0 degrees down travel on the left aileron. (7) Remove the inclinometer from the left aileron. (8) Put the right aileron trim tab in the streamline position. (a) Refer to Aileron Trim System - Maintenance Practices if rigging is necessary. (9) Install an inclinometer on the right aileron trim tab and set it to zero degrees. (10) Put the right aileron trim tab in the full up position. (a) Make sure that the inclinometer shows 15 +2 or -2 degrees.
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MODEL 208 MAINTENANCE MANUAL (11) Put right aileron trim tab in the full down position. (a) Make sure that the inclinometer shows 15 +2 or -2 degrees. (12) Remove the inclinometer from right aileron trim tab. (13) Do a friction band check. Refer to Ailerons and Control Column - Maintenance Practices. F.
Restore Access (1) Install panels 212FR, 226B, 231BL, 231CL, 251CL, 251DL, 252BR, 252FR, 501BB, 501CB, 501DB, 501EB, 503AB, 503BB, 503CB, 503DB, 503EB, 601BB, 601CB, 601DB, 601EB, 603AB, 603BB, 603CB, 603DB, 603EB, and 651AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Close (zip) the fabric headliner (passenger) or install the hard shelled headliner (cargo). Refer to Chapter 25, Cabin Upholstery - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL AILERONS AND CONTROL COLUMN - MAINTENANCE PRACTICES 1.
General A.
2.
For lubrication requirements of the aileron and control system, refer to Chapter 12, Flight Controls Servicing.
Control Column Removal/Installation A.
Remove the Control Column (Refer to Figure 201). (1) Remove the carpet or vinyl cover to get access to the floorboard. Refer to Chapter 25, Floor Covering/Control Column Cover - Maintenance Practices. (2) Remove the screws from the control column covers, then disconnect the floor column covers from the column tubes. (3) Remove the screws from the scuff plates and disconnect the scuff plates from the floorboard. Remove the floor cover plates as required to get access to the torque tube assembly mounting bolts. (4) Remove the nuts, washers and screws. Remove the cover plate from the instrument panel. (5) Remove the nut, washer and bolt. Disconnect the control tube and control wheel from the shaft. Disconnect the housing cap and remove the control tube, cover plate and cable from the system. (6) Remove the safety wire or clip, and loosen the turnbuckles. Remove the safety wire and screws from the cable guard. Disconnect the cable guard from the column. (7) Remove the screws and spacers from the quadrant and release the cables. (8) Remove the nuts, washers and bolts. Disconnect the column from the torque tube. (9) Remove the screws from the inboard channels, outboard stiffeners and cap strips. Disconnect the inboard channels and outboard stiffeners from the longerons. Disconnect the cap strips from the supports. (10) Remove the nuts, washers and bolts from the cables. Disconnect the cables from the rudder arms. (11) Cut the safety-wire and remove the roll pin from the coupler. Disconnect the flex shaft from the coupler. (12) Remove the nuts, washer, bolts, bearing and bushing. Disconnect the link and the plates from the steering pushrod. (13) Remove the cotter pin, nut, washer and bolt. Disconnect the pushrod from the support arm. (14) Remove the cotter pin, nut, washer and bolt from the support arms. (15) Lift the torque tube assembly with the inboard and outboard cables attached from under the floorboard, and remove them from the system. NOTE:
B.
Remove the bearings from the support arms. Clean, dry with air, and lubricate the bearings with MlL-G- 81322 general purpose grease before installation of the torque tube assembly. Replace the bearings in the support arms.
Install Control Column (Refer to Figure 201). (1) Find the torque tube assembly with the inboard and outboard cables attached, under the floorboard mounting brackets. (2) Replace the bolts, washers, nuts and cotter pins. (3) Connect the pushrod to the support arm and replace the bolt, washer, nut and cotter pin. (4) Connect the link and the plates to the steering pushrod. Replace the bolts, bearing, washers, bushings and nuts. (5) Connect the flex shaft to the coupler, align the attaching holes and replace the roll pin. Use safety wire to connect the roll pin. (6) Connect the rudder cables to the rudder arms and replace the bolts, washers and nuts. (7) Attach the column tubes to the torque tube. Align the attaching holes and replace the bolts, washers and nuts. (8) Attach the control tubes to the shaft, replace the bolts, washers and nuts. (9) Connect the housing cap to the electrical plug, attach the cover plate to the instrument panel, and replace the screws, washers and nuts. (10) Connect the ball-ends of the right inboard and right outboard cables to the right quadrant and replace the screws and spacers. Use safety wire to connect the screws.
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Control Column Installation Figure 201 (Sheet 1)
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Control Column Installation Figure 201 (Sheet 2)
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Control Column Installation Figure 201 (Sheet 3)
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Control Column Installation Figure 201 (Sheet 4)
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Control Column Installation Figure 201 (Sheet 5)
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Control Column Installation Figure 201 (Sheet 6)
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MODEL 208 MAINTENANCE MANUAL (11) Connect the ball-ends of the left inboard and left outboard cables to the left quadrant and replace the screws and spacers. Use safety wire to connect the screws. (12) Use tape to attach a rod or bar across the top of the control wheels. (13) Tighten the turnbuckles until the cable tension is 30 pounds, +5 or -5 pounds (133.45 N, +22.24 or -22.24 N). Use safety wire or install clips in the turnbuckles. Remove the rod or bar from the top of the control wheels. NOTE:
All control surface cable tensions must be set at an ambient temperature of 70°F (39°C). Let the temperature of the airplane stabilize for four hours before setting the cable tension.
Attach the inboard channels and the outboard stiffeners to the Iongerons. Replace the screws. Attach the cap strips to the supports and replace the screws. Attach the scuff plates to the supports and replace the screws. Install the control column covers on the column tube assemblies and attach the control column covers to the scuff plates and instrument panels. Replace the screws. (18) Replace the access panels and carpet or vinyl floor cover.
(14) (15) (16) (17)
3.
4.
Column Tube Assembly Bearings Removal/Installation A.
Remove the Column Tube Assembly Bearings (Refer to Figure 201). (1) Cut the safety wire and remove the nut from the shaft. (2) Remove the cotter pin, nut, washer, and bolt. Disconnect the quadrant from the shaft. Disconnect the shaft from the column tube assembly. (3) Remove the seals and bearings from the column tube assembly.
B.
Install the Column Tube Assembly Bearings (Refer to Figure 201). (1) Clean the column tube assembly, bearings, seals and shaft and dry with air. (2) Replace the bearings and seals in the column tube assembly. Push the shaft through the seals, bearings, and column tube assembly. (3) Attach the quadrant to the shaft. Align the mounting holes and replace the bolt, washer, nut, and cotter pin. (4) Replace the nut on the shaft. Torque the nut at 5 to 10 inch pounds (0.56 to 1.13 N-m) and attach with safety wire.
Pulleys, Cables, Quadrants and Bell Cranks Removal/Installation A.
Remove the Pulleys, Cables, Quadrants and Bell Cranks (Refer to Figure 202). (1) Remove the carpet or vinyl floor covers, plywood floor covers in Model 208 and 208B airplanes, floorboard access covers, scuff plates, wing access covers and unzip the headliner to get access to the system. (2) Remove the screws, washers and nuts from the lower aileron quadrant. (3) Remove the cable guard from the support. (4) Cut the safety wire or remove the clips and loosen the turnbuckles on the direct and fuselage loop cables. (5) Cut the safety wire, remove the screws and spacers. Disconnect the ball-ends of the fuselage loop cables from the quadrant. (6) Remove the bolt, spacer, washer and nut and disconnect the pulleys from the support. (7) Remove the nuts, washers and bolts. Disconnect the pulleys from the supports. (8) (208 ) Remove the bolt and disconnect the pulley from the support. (9) (208B) Remove the bolt, washer and nut. Disconnect the pulley and cable guard from the bracket. (10) (208B) Remove the bolt, washer and nut. Disconnect the pulley from the bracket. (11) (208B) Remove the bolt, washer and nut. Disconnect the pulley from the bracket. (12) Remove the bolt and washer and disconnect the pulley and cable guard from the bearing. (13) Remove the cotter pins, washers and pins. Disconnect the ball-ends of the fuselage loop and direct cables from the lower quadrant. (14) Remove the nuts, washers and bolts. Disconnect the direct and carry-thru pulleys from the support.
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Aileron Systems Installation Figure 202 (Sheet 1)
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Aileron Systems Installation Figure 202 (Sheet 2)
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Aileron Systems Installation Figure 202 (Sheet 3)
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Aileron Systems Installation Figure 202 (Sheet 4)
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Aileron Systems Installation Figure 202 (Sheet 5)
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Aileron Systems Installation Figure 202 (Sheet 6)
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Aileron Systems Installation Figure 202 (Sheet 7)
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Aileron Systems Installation Figure 202 (Sheet 8)
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MODEL 208 MAINTENANCE MANUAL (15) Disconnect the cables from the upper quadrant and lower bell crank. (a) On Airplanes 20800001 thru 20800415 and 20800417 thru 20800420 and Airplanes 208B0001 thru 208B1215 and 208B1217 thru 208B1313 Incorporating CAB08-6 and Airplanes 20800416 and 20800421 and On and Airplanes 208B1216 and 208B1314 and On, remove the cotter pins, nuts (two each left and right), washers, lock washer and bolts from the upper quadrant. (b) On Airplanes 20800001 thru 20800415 and 20800417 thru 20800420 and Airplanes 208B0001 thru 208B1215 and 208B1217 thru 208B1313 Not Incorporating CAB08-6 remove the cotter pins, washers, pins and nuts. Disconnect the carry-thru cables from the upper quadrant. (c) Remove the cotter pins, nuts, washers and bolts from the lower bell crank. (d) Remove the terminal ends of the direct and carry-thru cables from the bell crank. NOTE:
The direct and carry-thru cables are routed to the bell crank in the opposite wing. Removal procedures for the cables, quadrants, pulleys, spoiler and aileron pushrods are identical for both wings.
(16) Remove the fuselage loop, direct and carry-thru cables from system. NOTE:
To make the replacement of cables easier, attach a length of wire to the cable to be removed from the system. After removing the cable, leave the wire in place, routed through the structure. To replace the cable, attach it to wire, and pull the cable into position with the wire.
(17) Remove the nut, washer and bolt. Disconnect the upper and lower quadrants from the support. (18) Disconnect the spacer, washers and the upper and lower quadrant bearings from the quadrant. NOTE:
Remove the upper and lower quadrant bearings from the quadrant. Clean and dry the bearings with air and lubricate by hand with MIL-PRF-81322 general purpose grease. Replace the bearings and stake each bearing a minimum of six places.
(19) Remove the cotter pins, nuts, washers and bolts. Disconnect the spoiler pushrod from the bell crank and the bracket. (20) Remove the washer, cotter pin, nut, washer and bolt. Disconnect the aileron pushrod from the bell crank and the support. (21) Remove the bolt and washer. Disconnect the bell crank from the support. B.
Install Pulleys, Cables and Bell Cranks (Refer to Figure 202). (1) Attach the bell crank to the support. Install the washer and bolt. (2) Attach the aileron pushrod to the bell crank and the support. Install the washer, bolts, nut and cotter pin. (3) Attach the spoiler pushrod to the bell crank and to the spoiler attach bracket. Install the bolts, washers, nuts and cotter pins. (4) Attach the spacer, washers, and the upper and lower quadrant to the support. Install the bolt, washer and nut. (a) Make sure that the minimum distance between the upper and lower quadrant is 0.040 inch (1.01600 mm). (5) Install the fuselage loop, carry-thru and direct cables in system. (6) On Airplanes 20800001 thru 20800415 and 20800417 thru 20800420 and Airplanes 208B0001 thru 208B1215 and 208B1217 thru 208B1313 Incorporating CAB08-6 and Airplanes 20800416 and 20800421 and On and Airplanes 208B1216 and 208B1314 and On do the steps that follow: (a) Install the fittings of the direct cables to the quadrant. (b) Install the fittings of the carry-thru cables to the quadrant. 1 If a square carry-thru cable fitting is installed, make sure that the carry thru cable fitting is turned so that it is flush with the surface of the upper quadrant before you tighten the nuts. 2 Torque the nut that is close to the carry thru cable fitting to 40 inch pounds, +5 or -5 inch pounds (4.5 N-m, +.564 or - .564 N-m).
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(c) (d) (e)
3 Install safety wire between the two nuts installed on the carry-thru cable. Install the cotter pins, nuts (two each left and right) , washers, lock washer and bolts from the upper quadrant. Install the cotter pins, nuts, washers and bolts from the lower bell crank. Install the fittings of the direct and carry-thru cables from the bell crank.
NOTE:
(7)
On Airplanes 20800001 thru 20800415 and 20800417 thru 20800420 and Airplanes 208B0001 thru 208B1215 and 208B1217 thru 208B1313 Not Incorporating CAB08-6 do the steps that follow: (a) Attach the carry-thru cables to the upper quadrant. Install the nuts, pins, washers and cotter pins. (b) Install the cotter pins, nuts, washers and bolts from the lower bell crank. (c) Install the fittings of the direct and carry-thru cables from the bell crank. NOTE:
(8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21) 5.
The direct and carry-thru cables are routed to the bell crank in the opposite wing. Installation procedures for the cables, quadrants, pulleys, spoiler and aileron pushrods are identical for both wings.
The direct and carry-thru cables are routed to the quadrant in the opposite wing. Installation procedures for cables, quadrant, pulleys, spoilers and aileron pushrods are the same for both wings.
Attach the ball-ends of the fuselage loop and direct cables to the lower quadrant. Install the pins, washers and cotter pins. Attach the direct and carry-thru pulleys to the support. Install the bolts, washers and nuts. Attach the pulley and cable guard to the bearing. Install the bolt and washer. (208) Attach the pulley to the support. Install the bolt and washer. (208B) Attach the pulley and cable guard to the bracket. Install the bolt, washer and nut. (208B) Attach the pulley to the bracket and install the bolt. (208B) Attach pulley to the bracket. Install the bolt, washer and nut. Attach the left and right pulleys to the supports. Install the bolts, washers, and nuts. Attach the left and right pulleys to the support (SD). Replace the bolt, spacer, washer, and nut (SC). Attach the ball-ends of the fuselage loop cables to the quadrant. Replace the spacers and bolts. Tighten and attach with safety wire or install clips on the fuselage loop and direct loop turnbuckles. Attach the cable guard to the support. Replace the screws, washers, and nuts. Do the rigging for the aileron system, refer to Chapter 27, Ailerons and Control ColumnMaintenance Practices. Close the headliner, replace the wing access covers, scuff plates, floorboard access covers and carpet or vinyl floor covers (plywood floor covers in Model 208 and 208B airplanes).
Rigging Aileron System A.
Rigging Procedures (Refer to Figure 202). (1) Remove the wing access plates and open the headliner to get access to the turnbuckles and bell cranks as needed. NOTE:
(2) (3) (4) (5)
All control surface cable tensions must be adjusted at an ambient temperature of 70°F (39°C). Let the temperature of the airplane stabilize for four hours before the cable tensions are set.
Remove the safety wire or clips from the fuselage loop and direct cable turnbuckles. If installed, remove the safety wire from between the nuts on the terminal ends of the carry-thru cables. Loosen the nuts on the terminal ends of the carry-thru cables. Release the tension on all cables.
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MODEL 208 MAINTENANCE MANUAL (6) (7) (8) (9) (10)
(11)
(12) (13) (14) (15)
Put a bar across the control wheels and look for level. If not level, refer to Control Column Removal/Installation for leveling control wheels. Put tape across the top of the control wheels to hold them in a neutral position. Install the rigging pins, in each of the bell cranks. Use tape to hold the rigging pins in place. Remove the bolts and washers to disconnect the aileron pushrod from the aileron. On Airplanes 20800001 thru 20800415 and 20800417 thru 20800420 and Airplanes 208B0001 thru 208B1215 and 208B1217 thru 208B1313 Incorporating CAB08-6 and Airplanes 20800416 and 20800421 and On and Airplanes 208B1216 and 208B1314 and On do the steps that follow: (a) If a square carry-thru cable fitting is installed, make sure that the carry thru cable fitting is turned so that it is flush with the surface of the upper quadrant before you tighten the nuts. (b) Tighten the nuts on the terminal ends of the carry-thru cables evenly to set the cable tension at 40 pounds, +5 or -5 pounds (177.93 N, +22.41 or -22.41 N). (c) Torque the nut that is close to the carry thru cable fitting to 40 inch pounds, +5 or -5 inch pounds (4.5 N-m, +.564 or - .564 N-m). (d) Install safety wire between the two nuts installed on the carry-thru cable. On Airplanes 20800001 thru 20800415 and 20800417 thru 20800420 and Airplanes 208B0001 thru 208B1215 and 208B1217 thru 208B1313 Not Incorporating CAB08-6 do the steps that follow: (a) Tighten the nuts on the terminal ends of the carry-thru cables evenly to set the cable tension at 40 pounds, +5 or -5 pounds (177.93 N, +22.41 or -22.41 N). (b) Tighten the turnbuckle on the direct cable to set the cable tension at 40 pounds, +5 or -5 pounds (177.93 N, +22.41 or -22.41 N). Use safety wire or install clips on the turnbuckle. Tighten the fuselage loop turnbuckles on the fuselage loop cables to set the cable tension at 20 pounds, +5 or -5 pounds (88.96 N, +22.41 or -22.41 N). Use safety wire or install clips on the turnbuckles. With the ailerons streamlined, (inboard trailing edge of aileron aligned with outboard trailing edge of flap), attach the aileron pushrods to the supports. Replace the washers and bolts. Remove the rigging pins from each of the bell cranks. Attach an inclinometer to the left aileron and set at zero degrees.
WARNING: If turning the control wheels counterclockwise does not put the left aileron in the raised position, the system is rigged backwards. The system must be correctly rigged. Check for crossed or wrapped cables. (16) Remove the bar from the control wheels and turn the control wheels counterclockwise. This will put the left aileron in a raised position. (17) Adjust the stop bolt so that it touches the right bell crank at 25 degrees (+4 or -0 degree tolerance) up travel on the left aileron. Tighten the locknut. (18) Turn the control wheels clockwise, and adjust the stop bolt so that it touches the left bell crank at 16 degrees (+1 or -0 degree tolerance) down travel on the left aileron. (19) Streamline the right aileron and attach an inclinometer; set at zero degrees. (20) Examine the travel on the right aileron. Set the locknuts and use safety wire on the pushrods. (21) When the ailerons have been rigged properly, put the rig pin into the upper quadrant and lower quadrant. (22) With the ailerons held in a neutral position and flaps completely retracted, make sure the trailing edge of the spoiler is 0.55 inch, +0.05 or -0.05 inch, (13.97 mm, +1.27 or -1.27 mm) above the surface of the flap at the outboard end of the spoiler. Adjust as required. (23) Remove the rig pin from the upper quadrant and lower quadrant. Turn the control wheels slowly back and forth from stop to stop. Adjust the spoiler pushrod as required to give a 0.01 to 0.03 inch (0.25 to 0.76 mm) clearance between the spoiler trailing edge and the top of the flap surface at the minimum clearance position. The total spoiler travel is 40 degrees up (+5 or -5 degrees), and 0 degrees down (+0 to -5 degrees). (24) Lock the adjusting nuts on the pushrod safety wire. (25) Replace the wing access plates and close the headliner.
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6.
Friction Band Requirements A.
Measuring Friction Band. NOTE:
(1) (2)
All friction measurements must be taken with load scale so that the force necessary to move the ailerons is applied tangentially to the direction of rotation of the control wheel. The load scale must be attached to the control wheel inside grip at the lowest possible moment arm. The friction band requirements apply over the complete travel range of the ailerons.
Rotate control wheel approximately 30 degrees counterclockwise from neutral position, attach load scale, rotate wheel clockwise, and check scale reading as wheel passes through neutral position. Make same check in opposite direction of control wheel rotation. NOTE:
(3) B.
The aileron friction band is calculated by adding scale readings from (1) to scale readings from (2), therefore (1) plus (2) equals friction band.
The maximum permitted friction band is six pounds or less without the autopilot installed, or eight pounds or less with the autopilot installed.
Adjusting Friction Band NOTE: (1) (2) (3)
When the friction band exceeds the limitations, the following steps shall be taken to reduce the system friction to an acceptable level:
Check fuselage loop and wing loop for clearance and eliminate all interference as indicated. Reduce aileron cable tension as required, 15 pounds (66.72 N), fuselage loop, 35 pounds (155.69 N), wing loop with ailerons in neutral position. Check and adjust pulley alignment as indicated.
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MODEL 208 MAINTENANCE MANUAL AILERON TRIM SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
The aileron trim system consists of the following components: (1) Trim Wheel - The trim wheel is located on the center pedestal and provides for manual input to the trim system. (2) Sheathed Cables - From the trim wheel, sheathed cables are routed through the windshield center post and out the right wing. (3) Trim Tab Actuator - The sheathed cables terminate in the wing and are connected to chains which wrap around trim tab actuators. These actuators are connected with integral pushrods to the aileron trim tab and serve to alter trim tab position.
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MODEL 208 MAINTENANCE MANUAL AILERON TRIM SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been included to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Aileron Trim System Troubleshooting Chart Figure 101 (Sheet 1)
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Aileron Trim System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL AILERON TRIM SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
This section includes information on removal, installation, and rigging of the aileron trim system. Also included are rebuilding procedures for aileron trim tab actuator. For lubrication requirements of the aileron trim system, refer to Chapter 12, Flight Controls Servicing.
Aileron Trim System Removal/Installation A.
Remove Aileron Trim System (Refer to Figure 201). (1) Remove the screws and washers from aft and upper pedestal covers and pins from the trim wheel. Remove the wheel and aft cover from the pedestal; the upper cover can be moved out of the way without removing it from the pedestal. (2) Remove the windshield center post cover strip, loosen the overhead fuel console, and remove the access panels on the bottom of the wings to gain access to the trim system and the left wing aileron servo tab. (3) Remove the connecting links from chain and disconnect the up cable and down cable from the chain. (4) Loosen the screws and unseat the chain guards, allowing them to clear the chain drive sprocket. Disconnect the chain from the chain drive sprocket. (5) Remove the lock nuts from the unions at the supports. NOTE:
The up cables and down cables are routed through the channel in the windshield center post. Access to the center post channel is gained by removing the cover strip. The cover strip snaps on and off the center post channel and no fasteners are necessary. The strip is fabricated from aluminum and must be removed and replaced carefully.
(6)
Push the spring-loaded knurled ends of the connectors toward the up cable and down cable far enough to remove the ball-ends of the up cable and down cable from the connectors. Leave the connectors attached to the up cable and down cable. (7) Remove the cables from the system with the unions attached to both ends of the cables. (8) Remove ties from the cable mounts and the cables at the right wing ribs. (9) Remove lock nuts from the unions and disconnect the unions from the support. (10) Remove screws and disconnect the stop blocks from the cables. (11) Remove the bolt and spacer from the sprocket guard. (12) Remove the connecting links and disconnect the cables from the chain. Remove the cables with the unions attached, from the right wing. NOTE:
To facilitate installation of the cables, attach a length of wire to the cables being removed from the airplane. Leave the wire in position, routed through the structure. Attach the cable to be installed to the wire and pull the cable in position with wire.
(13) Remove the cotter pins, nuts, washers, bolts, and bushings. Disconnect the pushrods from the actuator and right aileron trim tab. (14) Cut the safety wire and remove the bolts and washers from the actuator. Disconnect the actuator from the support. (15) Remove the cotter pins, nuts, washers, bolts, and bushings. Disconnect the pushrods from the left aileron servo tab and the bracket. NOTE:
There are no adjustments or rigging procedures required on the left aileron servo tab after the removal and installation of the pushrods and component parts.
(16) Cut safety wire, remove the bolts and washers. Disconnect the actuator from the support and set it on the bench. (17) Remove the connecting link from the chain and disconnect the chain from the primary sprockets. (18) Remove the groove pin and disconnect the secondary sprocket from the internal screw.
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Aileron Trim Installation Figure 201 (Sheet 1)
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Aileron Trim Installation Figure 201 (Sheet 2)
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Aileron Trim Installation Figure 201 (Sheet 3)
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Aileron Trim Installation Figure 201 (Sheet 4)
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Aileron Trim Installation Figure 201 (Sheet 5)
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Aileron Trim Installation Figure 201 (Sheet 6)
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Aileron Trim Installation Figure 201 (Sheet 7)
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Aileron Trim Installation Figure 201 (Sheet 8)
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MODEL 208 MAINTENANCE MANUAL B.
Install Aileron Trim System (Refer to Figure 201). (1) Attach the actuator to the outboard support using washers and bolts. Attach the actuator to the inboard support using washers and bolts. Use safety wire to attach the bolts. (2) Secure the right trim tab in a streamlined position. (3) Secure pushrods to the external screws. Replace the bolts, washers, nuts, and cotter pins. (4) Rotate the secondary sprocket until the pushrods are aligned with the trim tab attach bracket. Replace bushings, bolts, washers, nuts and cotter pins. Release the trim tab. (5) Attach the chain to the secondary sprocket with the ends equidistant from the center of the sprocket. (6) Replace the bolt and spacer on the sprocket guard. (7) Attach pushrods to the left aileron servo tab, replace the bushings, bolts, nuts, and cotter pins. (8) Replace the cables, with unions and connectors attached, in the right wing. NOTE:
Before installing any cables in the system, lubricate the inside of the cable housings with Dow Corning Molykote EC321R bonded lubricant spray.
NOTE:
Make sure the cables are correctly routed through the wing ribs, and unions are installed in the supports. Tighten the lock nuts securely on both sides of the supports.
(9) Find the cables on cable mounts in the wing ribs, and secure them with ties. (10) Attach the clevis ends of the cables to the chain and replace the connecting links. (11) Install stop blocks on the cables. Replace the screws, and the safety wire. NOTE:
For the location of the stop blocks, refer to Aileron Trim System Rigging.
(12) Replace the cables with unions connected to each end of the cable. NOTE:
Make sure the cables are correctly routed from the pedestal through the windshield center post and the support. Tighten the lock nuts securely on both sides of the supports.
(13) Attach the ball ends of the cables to the connectors. Attach the clevis ends of the cables to the chain. Replace the connecting links. NOTE:
If the chain has been removed from the sprocket, replace it with ends equidistant from center of the sprocket. Find the chain guards with dimples positioned in mating holes in the support. Tighten the screws.
(14) Replace the windshield cover strip on the center post. Tighten the overhead fuel console, and replace the access panels on the bottom of the wings. (15) Install the aft and upper pedestal covers and replace the mounting screws and washers. Install the trim wheel and replace the roll pins. 3.
Aileron Trim System Rigging A.
Rigging Procedures (Refer to Figure 201). (1) Secure the right aileron trim tab in the streamlined position. (2) Remove the mounting screws from the aft upper pedestal covers, roll pins, and trim wheel. (3) Remove the roll pin and aileron trim indicator wheel from the shaft, set the indicator in the neutral position. Replace the aileron trim indicator wheel and roll pin. (4) Loosen the screws in the stop blocks sufficiently to allow cables to move until either of the cables touches a union. (5) Release the right aileron trim tab and install an inclinometer set on zero degrees. (6) Turn the trim wheel counterclockwise until the inclinometer reads 15 degrees up, +2 or -2 degrees. (7) Move up the stop block to touch the union. Tighten the screw and safety wire. (8) Turn the trim wheel clockwise until the inclinometer reads 15 degrees down, +2 or -2 degrees. (9) Move down the stop block to touch the union. Tighten the screw and safety wire.
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MODEL 208 MAINTENANCE MANUAL (10) Remove the inclinometer from the right aileron trim tab. Remove the roll pin and disconnect the trim wheel from the shaft. (11) Install the aft and upper pedestal covers and replace the mounting screws. (12) Install the trim wheel on the shaft and replace the roll pins. (13) Rotate the trim wheel to stops in both directions and look at the clearance between the indicator and the aft cover. It can be necessary to hand form the indicator in order to clear the slot in the aft cover. (14) On Airplanes 2080001 thru 20800081, to set the cable tension at 3.0 pounds (13.34 N) maximum, adjust the locknut on the ends of the cable housings away from the bulkhead to increase tension, or toward the bulkhead to decrease tension. NOTE:
If adjustment of the cable housing does not cause tension in the cable, it is recommended that the 2660029-1 cable assembly be replaced with the 2660029-7 cable assembly.
(15) Beginning with Airplanes 20800082 and On, to set cable tension at 3.0 pounds (13.34 N) maximum, rotate the barrel in the required direction. Use safety wire to attach the barrel. NOTE:
If barrel rotation does not cause tension in the cables, adjust the locknut on the ends of the cable housings away from the bulkhead to increase tension or toward the bulkhead to decrease tension.
CAUTION: Make sure the system is rigged correctly. Turn the trim wheel counterclockwise to move the trim tab up. Maximum misalignment between the ailerons and the aileron trim tab trailing edges must not exceed 0.0 inch, +0.5 or -0.5 inch (0.0 mm, +12.7 or -12.7 mm). (16) Make sure the system is correctly rigged. (17) Do the Aileron Trim System Lubrication. Refer to Aileron Trim System - Inspection/Check, and to Chapter 5, Inspection Time Limits, for the lubrication intervals. 4.
Aileron Trim Tab Actuator Disassembly (Airplanes with 2660044-1 Trim Tab Actuator Installed) A.
Disassembly Procedures (Refer to Figure 202). (1) Cut the safety wire and remove the screws. Disconnect the chain guard from the actuator housing. Remove the plug button from the secondary sprocket. Disconnect the repair link and remove the chain from the secondary sprockets. (2) Remove the groove pins, and disconnect the primary sprocket and secondary sprockets from the internal screws. NOTE: (3)
Put index marks on the bearings and the actuator housing. Remove the groove pins from the actuator housing, and disconnect the bearings from the actuator housing. Remove and discard the O-rings from the bearings. NOTE:
(4) (5) (6)
If necessary, apply heat to the sprockets to loosen the Loctite seal between sprockets and the internal screws.
If the bearings are to be reused, they must be replaced in the same location and relative position from which they were removed.
Tap the ends of the external screws on a hard surface to remove the bearings and the races. Tap the internal screws at the location with groove holes to remove the bearings and races. Remove the internal screws from the actuator housing and disconnect the bearings, races, and external screws from the internal screws.
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 1)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 2)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 3)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 4)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 5)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 6)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 7)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 8)
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Aileron Trim Tab Actuator Disassembly/Assembly Figure 202 (Sheet 9)
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MODEL 208 MAINTENANCE MANUAL (7)
5.
Examine the condition of the bearings in the external screw and replace if necessary. NOTE:
Use an arbor press and mandrel.
NOTE:
Clean the actuator components with solvent and dry thoroughly. Do not clean the bearings, and do not allow cleaned parts to touch lint or dirt.
Aileron Trim Tab Actuator Disassembly (Airplanes with 2661615-1 Trim Tab Actuator Installed) A.
Disassembly Procedures (Refer to Figure 202). (1) Cut the safety wire and remove the screws. Disconnect the chain guard from the actuator housing. Remove the plug button from the secondary sprocket. Disconnect the repair link and remove the chain from the secondary sprockets. (2) Remove the groove pins, and disconnect the secondary sprockets and the primary sprocket from the internal screws. Remove the screw and end plate. NOTE: (3) (4)
Cut the safety wire from the screws and remove the screws and the end plate. Tap the ends of the external screws on a hard surface to remove the wipers and bearings. If the bearings are to be reused, keep the two halves together as a pair as they are removed. Unscrew the external screw from the internal screws. NOTE:
(5) (6) (7)
6.
If necessary, apply heat to the sprockets to loosen the Loctite seal between sprockets and the internal screws.
If the bearings are to be reused, they must be replaced in the same location and relative position from which they were removed.
Tap the internal screws at the location with groove holes to remove bearings. Remove the internal screws from the actuator housing. Examine the condition of the bearings in the external screw and replace if necessary. NOTE:
Use an arbor press and mandrel.
NOTE:
Clean the actuator components with solvent and dry thoroughly. Do not clean the bearings, and do not allow cleaned parts to touch lint or dirt.
Aileron Trim Tab Actuator Disassembly (Airplanes with 2661615- 9 or 2661615-10 Trim Tab Actuator Installed) A.
Disassembly Procedures (Refer to Figure 202). (1) Cut the safety wire and remove the two screws that attach the chain guard to the actuator. (2) Remove the chain guard. (3) Drive the groove pins from the secondary sprockets. (4) Remove the primary sprocket; plug button, secondary sprockets, and chain from the internal screws. NOTE: (5) (6) (7) (8) (9)
If necessary, apply heat to the sprockets to loosen the Loctite seal between sprockets and the internal screws.
Cut the safety wire from the center screw and remove the center screw that secures the end plate. Use a soft hammer to tap lightly against the externally threaded screws to remove the bearings. Remove the screws that secure the end plate. Use a soft hammer to tap lightly against the internally threaded screws to remove the split bearings. Remove the externally threaded screws and the internally threaded screws from the actuator housing.
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MODEL 208 MAINTENANCE MANUAL (10) Separate the split bearings and remove them from the externally threaded screws. NOTE: (11) (12) (13) (14) (15)
7.
Remove and discard the O-rings from the bearings. Remove the shank seal from the bearings. Remove the externally threaded screws from the internally threaded screws. Remove the wiper ring from the shaft of the externally threaded screw. Examine the condition of the bearings in the external screw and replace if necessary. NOTE:
Use an arbor press and mandrel.
NOTE:
Clean the actuator components with solvent and dry thoroughly. Do not clean the bearings, and do not allow cleaned parts to touch lint or dirt.
Inspection and Repair of Aileron Trim Tab Actuator A.
Inspection Criteria. NOTE: (1) (2) (3) (4) (5)
8.
If the bearings are to be reused, they must be put back into the same relative position from which they were removed.
Remove actuator from system. Clean, inspect, and lubricate detail parts. Replace any components that show damage or excessive wear. Refer to Chapter 5 for Time Limits.
Clean the detail parts with a solvent in a well ventilated area away from sparks or open flame. Avoid the inhalation of solvent vapors. Dry the parts with dry, compressed air, lint free cloth or lint free disposable tissue. Examine the parts visually, preferably under magnification. If any parts show wear or damage, do a dimensional examination and replace parts, if necessary. If the finish on the actuator housing or chain guard has worn away or bare metal is exposed, apply Iridite to the part. Apply two coats of nonchromated primer and repaint with lacquer. The finish must consist of vivid orange or white lacquer. For a list of iridite, primers and paints required for the actuator, refer to Chapter 27, Flight Controls - General.
Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2660044-1 Trim Tab Actuator Installed) A.
Lubrication/Assembly Procedures (Refer to Figure 202). NOTE: (1) (2) (3) (4) (5)
Lubricate each detail part of the actuator assembly before installation, using 5565450-28 light consistency silicone grease.
Install the new O-rings in bearings. Install the internal screws with the groove pin hole location up and the actuator housing in an upright position with the flat end face without screw hole location down. Replace the races and bearings. Find the bearings on the index marks and lightly tap or press them into the actuator housing until the groove pins can be installed through the actuator and bearings. Put the actuator housing in an upright position with the flat end face without screw hole location up. Install the races and bearings in the actuator housing. Make sure the bearings are positioned on index marks, and lightly tap or press them into the actuator housing until the groove pins can be installed through the housing and bearings. NOTE:
(6)
Steps (2) through (5) are applicable if existing bearings are used. If new bearings are required, steps (6) through (15) are applicable.
Heat-soak the new bearings in SAE 20-weight oil for 20 minutes at 140°F (60°C). Cool the bearings to ambient temperature before installation.
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9) (10) (11) (12) (13)
Install the internal screws in the actuator housing with the groove pin hole location up and the actuator housing in an upright position with the flat end face without screw hole location down. Install the races, bearings, and a 0.004 to 0.006 inch (0.102 to 0.152 mm) shim on the groove pin hole location of the internal screws. Replace the bearings, press or tap them lightly until the bearings are flush with the end of the actuator housing. Put the actuator housing in an upright position with the bearings on bottom. Install the races and the bearings in the actuator housing. Press or tap lightly until the bearings are flush with the end of the actuator housing. Put a clamp securely across the assembled bearings to prevent any linear movement of the internal screws. Drill 0.094 inch (2.40 mm) diameter holes (4 positions) through the existing 0.062 inch (1.60 mm) diameter holes in the actuator housing and through the bearings. NOTE:
For an oversize groove pin, order Part Number GP3H094X0625-14.
(14) Release the clamp and disassemble the actuator sufficiently to remove a 0.004 to 0.006 inch (0.102 to 0.152 mm) shim from the actuator housing. (15) Reassemble the actuator and install four 0.094 inch (2.40 mm) groove pins through the actuator housing and bearings. (16) Replace the external screws in the actuator housing at the flat end face without screw hole location. NOTE:
After the external screw threads touch the internal screw threads, make sure they are not cross-threaded, and turn the external screws all the way in. Engagement should be smooth with no tight spots. If the threads drag or tight spots are found, disassemble the actuator and replace the internal and external screws.
(17) Apply No. 609 Loctite to the mating surfaces of the primary sprocket, secondary sprockets, and internal screws. Install the sprockets and replace the groove pins and plug button. (18) Proceed to Inspection and Rigging Procedures - Aileron Trim Tab Actuator. (Airplanes with 2660044-1 Trim Tab Actuator Installed), and complete the assembly of the aileron trim tab actuator. 9.
Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2661615-1 Trim Tab Actuator Installed) A.
Lubrication/Assembly Procedures (Refer to Figure 202). (1) Before assembly, apply a 0.118 inch (3.00 mm) coating of 5565450-28 light consistency silicone grease on the threads of the internal screws, outer surface of internal screws, and on the outside diameter of the external screws. (2) Install the internal screws with the groove pin hole location up and the actuator housing in the upright position with the flat end face without screw hole location down. (3) Install the bearings on the internal screws, then install the end plate on the actuator housing with the screw. NOTE: (4) (5)
Lubricate the threads of the external screw. Heat the wipers to make them more pliable. NOTE:
(6)
After the installation of the end plate, the outer race of the bearing must not move in the actuator housing.
Be very careful not to damage the wipers, when installing them over the external screw threads.
Position the end plate and wipers over the external screws. Inspect the wipers to make sure the threads did not damage the wipers during installation. Make sure the flat side of the wiper is installed to face the bearing.
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9)
Connect the external screws to the internal screws. Fill the bearing halves with 5565450-28 grease. Position the bearing halves around the external screws and press the bearings into the actuator housing. (10) Install the wipers in the actuator housing, then install the end plate on the actuator housing using the screws. NOTE:
After the external screw threads touch the internal screw threads, make sure they are not cross-threaded and turn the external screws all the way in. Engagement should be smooth with no tight spots. If the threads drag or tight spots are found , disassemble actuator and replace the internal and external screws.
(11) Safety wire the screws. (12) Apply No. 609 Loctite to mating surfaces of the secondary sprockets and the primary sprocket and install them on the internal screws using the groove pins. (13) Apply common RTV sealant to the plug button and the secondary sprocket and install the plug button in the secondary sprocket. (14) Work the screws all the way in and out 2 to 3 times. Wipe the excess grease from both ends after each cycle. NOTE:
There must be no end play between the bearings, inner race, internal screws and sprockets when the groove pins are installed.
NOTE:
After the assembly, the maximum longitudinal movement of the external screws and the actuator housing is not to exceed 0.007 inch (0.177 mm).
(15) Install the chain guard on the actuator housing using the screws. Apply No. 609 Loctite to the screws before installation. Proceed to Inspection and Rigging Procedures - Aileron Trim Tab Actuator (Airplanes with 2661615-1 and 2661615-9 Trim Tab Actuator Installed). 10.
Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2661615- 9 or 2661615-10 Trim Tab Actuator Installed) A.
Lubrication/Assembly Procedures (Refer to Figure 202). (1) Before assembly, apply an 0.118 inch (3.00 mm) coating of 5565450-28 light consistency silicone grease on the threads of the internal screws, outer surface of internal screws, and on the outside diameter of the external screws. (2) Install the internal screws with the groove pin hole location up and the actuator housing in the upright position with the wiper end location down. (3) Install the bearings on the internal screws, then install the end plate on the actuator housing with the screw. NOTE: (4) (5)
Lubricate the threads of the external screw. Apply heat to the wipers to make them more pliable if necessary. NOTE:
(6)
Be careful not to damage the wipers, when installing them over the external screw threads.
Position the end plate and wipers over the external screws. NOTE:
(7) (8)
After the installation of the end plate, the outer race of the bearing must not move in the actuator housing.
Inspect the wipers to make sure the threads did not damage the wipers during installation. Make sure the flat side of the wiper is installed towards the bearing.
Position the new O-rings onto the external screws with 5565450-28 grease. Turn the external screws into the internal screws.
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MODEL 208 MAINTENANCE MANUAL (9)
Position the bearing halves over the O-rings around the external shafts and press the bearings into the actuator housing. NOTE:
Seal the shank of the bearing halves with Pro-Seal 870, Type X, Class B sealant, (U544044).
(10) Install the wipers in the actuator housing. (11) Install the end plate on the actuator housing with the screws. NOTE:
After the external screw threads touch the internal screw threads, make sure they are not cross-threaded and turn the external screws in until seated. Engagement should be smooth with no tight spots. If the threads drag or tight spots are found , disassemble the actuator and replace the internal and external screws.
(12) Safety wire the screws. (13) Apply No. 609 Loctite to mating surfaces of the secondary sprockets and the primary sprocket and install them on the internal screws using the groove pins. (14) Apply common RTV sealant to the plug button and the secondary sprocket and install the plug button in the secondary sprocket. (15) Operate the actuators through two to three complete cycles of travel. Wipe the excess grease from both ends after each cycle. NOTE:
There must be no end play between the bearings, inner race, internal screws and sprockets when the groove pins are installed.
NOTE:
After the assembly, the maximum longitudinal movement of the external screws and the actuator housing is not to exceed 0.007 inch (1.77 mm).
(16) Apply Loctite 609 to the screws, then install the chain guard on the actuator housing. Proceed to Inspection and Rigging Procedures - Aileron Trim Tab Actuator (Airplanes with 2661615-1 and 2661615-9 Trim Tab Actuator Installed). (17) Complete the Inspection and Rigging Procedures. Refer to Inspection and Rigging Procedures Aileron Trim Tab Actuator (Airplanes with 2661615-1 and 5661615-9 Trim Tab Actuator Installed). 11.
Inspection and Rigging Procedures - Aileron Trim Tab Actuator (Airplanes with 2660044-1 Trim Tab Actuator Installed) A.
Inspection/Rigging Procedures (Refer to Figure 202). (1) After assembling the detailed parts, turn the primary sprocket clockwise, then counterclockwise far enough to get approximately 0.75 inch (19.05 mm) linear movement of the external screws in each direction. Movement must be smooth in each direction with no torque change in one direction or the other. NOTE: (2)
The bearings in the external screws must be aligned within 0.010 inch (0.254 mm) before installation of the actuator in the system. NOTE:
(3) (4) (5)
Starting torque of the primary sprocket must not exceed 4.76 inch-pounds (0.54 N.m) at the ambient temperature of 65°F (18°C).
A surface plate or table, two threaded rods or bolts (10-24 NC 3A thread), V-blocks, an angle block, clamps, height gage and a dial indicator (or equivalent precision measuring equipment) are necessary to perform this procedure.
Attach the bolts or threaded rods to both sides of the actuator housing at the location that contains bearings with O-ring. Bolts or rods must be tightened. Mount the unit in V-blocks in a vertical position. Turn one of the external screws in the required direction to allow the installation of the No. 11 drill rod (0.191 inch (4.85 mm)) through both of the bearings.
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MODEL 208 MAINTENANCE MANUAL (6) (7)
Examine the dimension from the top of the bolts or rods at the location that contains bearings with O-ring to the top of the No. 11 drill rod outside of each bearing. Remove the No. 11 drill rod and turn one of the screws in the required direction. Replace the No. 11 drill rod through the bearings and examine the alignment. Continue to turn the screws as required to align the bearings within 0.010 inch (0.254 mm). NOTE:
(8) (9) 12.
If the bearings cannot be aligned to 0.010 inch (0.254 mm) with the chain removed, turn one of the secondary sprockets one or two teeth in the desired direction. Secondary sprockets have two sets of mounting holes positioned 75 degrees apart. It can be necessary to move the sprocket from one set to the other. If it is determined that the excessive free play of the aileron trim tab is caused by the actuator, internal screws and external screws must be replaced along with any detail part worn beyond dimensional tolerance. However, if special optical inspection equipment is available and it is verified that threads on the internal screw and external screw are not worn beyond dimensional tolerance, the screws can be installed again in the assembly.
Install the chain on the secondary sprockets and replace the connector link. Install the chain guard on the actuator housing and replace the screws. Safety wire the screws.
Inspection and Rigging Procedures - Aileron Trim Tab Actuator (Airplanes with 2661615-1, 26616159, or 2661615-10 Trim Tab Actuator Installed) A.
Inspection/Rigging Procedures (Refer to Figure 202). (1) After assembling the detailed parts, turn the primary sprocket clockwise, then counterclockwise far enough to get approximately 0.75 inch (19.05 mm) linear movement of the external screws in each direction. Movement must be smooth in each direction with no torque change in either direction. NOTE: (2)
Align the bearings in the external screws to within 0.010 inch (0.254 mm) before installation of the actuator in the system. NOTE:
(3) (4) (5) (6)
The starting torque of the primary sprocket must not be more than three inch-pounds at the ambient temperature of 65°F (18°C).
A surface plate or table, two threaded rods or bolts (10-24 NC 3A thread), V-blocks, an angle block, clamps, height gage and a dial indicator (or equivalent precision measuring equipment) are necessary to do this procedure.
Attach the bolts or threaded rods to both sides of the actuator housing at the location that contains bearings with O-ring. The bolts or rods should be tightened. Mount the unit in the V-blocks in a vertical position. Turn one of the external screws in the required direction to allow installation of the No. 11 drill rod (0.191 inch (4.85 mm) diameter) through both of the bearings. Measure the dimension from the top of the bolts or rods at the location that contains bearings with O-ring to the top of the No. 11 drill rod outside of each bearing.
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(8) (9)
Remove the No. 11 drill rod and turn one of the screws in the required direction. Replace the No. 11 drill rod through the bearings and examine the alignment. Continue to turn the screws as required to align the bearings within 0.010 inch (0.254 mm). NOTE:
If the bearings cannot be aligned to 0.010 inch (0.254 mm) with the chain removed, turn one of the secondary sprockets one or two teeth in the desired direction. The secondary sprockets have two sets of mounting holes positioned 75 degrees apart. It can be necessary to move the sprocket from one set to the other.
NOTE:
If it is determined that the excessive free play of the aileron trim tab is caused by the actuator, internal screws and external screws must be replaced along with any detail part worn beyond the dimensional tolerance. However, if special optical inspection equipment is available and it is verified that the threads on the internal screw and external screw are not worn beyond the dimensional tolerance, the screws can be installed again in the assembly.
Install the chain on the secondary sprockets and replace the connector link. Install the chain guard on the actuator housing and install the screws. Apply No. 609 Loctite to the screws before installation.
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MODEL 208 MAINTENANCE MANUAL AILERON TRIM SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the aileron trim system in a serviceable condition.
Task 27-10-02-720 2.
Aileron Trim Tab (Free Play) Functional Check A.
General (1) This task gives the procedures to do an aileron trim tab (free play) functional check.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Aileron Trim Tab (Free Play) Functional Check (Refer to Figure 601). (1) Put the ailerons and the trim tab in the neutral position and secure them from movement. (2) Determine maximum allowable free play, measuring chord length at the extreme inboard end of the trim tab then multiply chord length by 0.025 to get the maximum allowable free play. (3) Use fingertip pressure and move the trim tab trailing edge up and down to examine free play. NOTE: (4) (5)
Measure free play at the same point on the trim tab that the chord length was measured. Total free play must not exceed the maximum allowable.
If the trim tab free play is less than the maximum allowable, no additional inspection is required. If the trim tab free play is more than the maximum allowable, the following items must be examined: (a) Look for loose fasteners on the trim tab doubler. (b) Examine the hinge, hinge pin, and fasteners on the trim tab doubler. (c) Examine both ends of the push-pull rods and fasteners for wear and loose component parts. (d) If corrosion, worn parts, or loose fasteners are found, replace the fasteners and install new parts in system. (e) Do a second free play inspection. 1 If the free play is still excessive, remove the aileron trim tab actuator from the airplane and set it on a bench. Refer to Aileron Trim System - Maintenance Practices. 2 Disassemble the actuator and examine the detail parts for corrosion and excessive wear. Refer to Aileron Trim System - Maintenance Practices. 3 If corrosion or worn parts are found, replace the parts and reassemble the actuator. (f) Install the actuator in the airplane . Refer to Aileron Trim System - Maintenance Practices. (g) Do the free play inspection again.
E.
Restore Access (1) None End of task Task 27-10-02-640 3.
Aileron Trim System Lubrication A.
General (1) This task gives the procedures to do the aileron trim system lubrication.
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Aileron Trim Tab (Free Play) Functional Check Figure 601 (Sheet 1)
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Special Tools (1) Dow Corning Molykote EC321R Bonded Lubrication Spray
C.
Access (1) Remove the applicable wing panels to get to the aileron trim control cables. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do the Aileron Trim System Lubrication (Refer to Figure 201 found in Aileron Trim System Maintenance Practices). (1) Move the aileron trim cables to the right until they stop. (2) Apply Dow Corning Molykote EC321R lubrication spray on a clean dry cloth until it is damp. NOTE:
This cloth is used to lubricate the aileron trim cables and to help keep the lubrication mist from a spray bottle off of the wing.
(a)
(3)
Rub the cloth with the lubrication along the exposed aileron trim cables between the cable ends and the cable housing. 1 Make sure that all exposed sides of the cables are coated with the lubrication. (b) Make sure that you apply the Dow Corning Molykote EC321R lubrication where the cable enters the cable housing opening. Move the aileron trim cables to the left until they stop and rub the cloth with the lubrication along the areas that were not initially lubricated.
E.
Restore Access (1) Install the wing access panels. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. End of task Task 27-10-02-641 4.
Aileron Trim Tab Actuator (2660044-1) Lubrication A.
General (1) This task gives the procedures to do the aileron trim tab actuator (2660044-1) lubrication.
B.
Special Tools (1) Grease
C.
Access (1) None
D.
Do the Aileron Trim Tab Actuator (2660044-1) Lubrication (Refer to Figure 202 found in Aileron Trim System - Maintenance Practices). (1) Remove the aileron trim tab actuator from the airplane and put it on a bench. Refer to Aileron Trim System - Maintenance Practices. (2) Disassemble the aileron trim tab actuator. Refer to Aileron Trim System - Maintenance Practices. (3) Do the Inspection and Repair of Aileron Trim Tab Actuator. Refer to Aileron Trim System Maintenance Practices. (4) Do the lubrication and the assembly steps found in Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2660044-1 Trim Tab Actuator Installed). Refer to Aileron Trim System - Maintenance Practices. (5) Install the aileron trim tab actuator in the airplane. Refer to Aileron Trim System - Maintenance Practices.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL Task 27-10-02-642 5.
Aileron Trim Tab Actuator (2661615-1, 2661615- 9, or 2661615-10) Lubrication A.
General (1) This task gives the procedures to do the aileron trim tab actuator (2661615-1, 2661615- 9, or 2661615-10) lubrication.
B.
Special Tools (1) Grease
C.
Access (1) None
D.
Do the Aileron Trim Tab Actuator (2661615-1, 2661615- 9, or 2661615-10) Lubrication (Refer to Figure 202 found in Aileron Trim System - Maintenance Practices). (1) Remove the aileron trim tab actuator from the airplane and put it on a bench. Refer to Aileron Trim System - Maintenance Practices. (2) Disassemble the aileron trim tab actuator. Refer to Aileron Trim System - Maintenance Practices. (3) Do the Inspection and Repair of Aileron Trim Tab Actuator. Refer to Aileron Trim System Maintenance Practices. (4) For aileron trim tab actuator (2661615-1), do the lubrication and the assembly steps found in Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 2661615-1 Trim Tab Actuator Installed). Refer to Aileron Trim System - Maintenance Practices. (5) For aileron trim tab actuators (2661615- 9 or 2661615-10), do the lubrication and the assembly steps found in Lubrication and Assembly of Aileron Trim Tab Actuator (Airplanes with 26616159 or 2661615-10 Trim Tab Actuator Installed). Refer to Aileron Trim System - Maintenance Practices. (6) Install the aileron trim tab actuator in the airplane. Refer to Aileron Trim System - Maintenance Practices.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL RUDDER SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
The rudder is a movable airfoil hinged to the aft spar of the vertical stabilizer. The rudder produces airplane yaw.
Description and Operation A.
The rudder system includes pilot’s and copilot’s rudder pedals attached to torque tubes. Control cables are attached to arms on torque tubes and routed under floorboard through a series of pulleys to a rudder bellcrank in stinger of airplane. Rudder is controlled by depressing left or right rudder pedal corresponding to desired direction of deflection.
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MODEL 208 MAINTENANCE MANUAL RUDDER SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Rudder System Troubleshooting Chart Figure 101 (Sheet 1)
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Rudder System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL RUDDER SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the rudder system in a serviceable condition.
Task 27-20-00-720 2.
Rudder System Functional Check (Standard Rudder Installation) A.
General (1) This task gives the procedures to do a functional check of the rudder system.
B.
Special Tools (1) Cable Tensiometer (2) Rudder Travel Protractor (3) Nose Wheel Turning Bar
C.
Access (1) Remove the applicable floor panels to get access to the rudder control system. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Remove the tail stinger from the airplane to get access to the rudder stop bolts. Refer to Tail Stinger - Maintenance Practices.
D.
Do a Functional Check of the Rudder System (Standard Rudder Installation). (1) Examine all cable runs for correct routing, fraying, and twisting. (a) Look for interference with the adjacent structure, equipment, wiring, plumbing, and other controls. (2) Examine the cable movement for binding and full range of travel. (3) Examine the swage fitting reference marks for signs of cable slippage inside of the fitting. (a) Examine the fitting for corrosion, distortion, cracks, and broken wires at the fitting. (4) Examine the turnbuckles for correct thread engagement. (a) Make sure that the turnbuckle locking clips are installed. Refer to Chapter 20, Safetying Maintenance Practices. (5) Move a cloth along the full length of the cable to examine for broken wires. (a) If snags are found or you think that there are broken wires, Refer to Chapter 20, Control Cable and Corrosion Limitations - Maintenance Practices. (6) Examine the cable attach holes in the rudder torque tube arm for excessive wear. (7) Examine the rudder torque tube, bellcrank, and lower hinge area for corrosion and condition. (8) Examine the rudder stop bolts for condition, corrosion, and security. (9) Examine the rudder hinge, hinge bearing, rudder pedals, and bonding jumper, for correct installation, corrosion, signs of damage, and unserviceable fasteners. (10) Examine the rudder skins for cracks, loose rivets, and corrosion. (11) Examine the balance weight for looseness and the supporting structure for damage.
E.
Examine the Travel and Cable Tensions. (1) Examine the cable tension in the tailcone area. NOTE:
(2) (3) (4) (5)
Cable tensions must be measured at least one foot from any pulley or cable turnbuckle.
(a) The tension must be 30 pounds + 5 or -5 pounds (133.45 N + 22.24 or - 22.24 N). If necessary, do the Rudder System Rigging. Refer to Rudder - Maintenance Practices. Install the rudder travel protector. Put the rudder trim system in the neutral position. Operate the system through its full range of travel. (a) Make sure that all of the components that move do not hit, touch, or catch on structural components or other system components.
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With the nose wheel turning bar, turn the nose wheel left until the rudder stop block contacts the bolt. (a) The rudder travel on the protractor must be 25 +2 or -2 degrees. With the nose wheel turning bar, turn the nose wheel right until the rudder stop block contacts the bolt. (a) The rudder travel on the protractor must be 25 +2 or -2 degrees. Turn the nose wheel to center and make sure that the rudder pedals and the rudder are centered. (a) If the rudder pedals and the rudder are not centered, make sure that the nose gear steering rigging is correct. Refer to Chapter 32, Nose Gear Steering - Maintenance Practices. Remove the rudder travel protractor.
F.
Restore access. (1) Install the tail stinger. Refer to Tail Stinger - Maintenance Practices. (2) Install the floor panels. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task Task 27-20-00-721 3.
Rudder System Functional Check (Float Kit Installation) A.
General (1) This task gives the procedures to do a functional check of the rudder system.
B.
Special Tools (1) Cable Tensiometer (2) Rudder Travel Protractor
C.
Access (1) Remove the applicable floor panels to get access to the rudder control system. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Remove the tail stinger from the airplane to get access to the rudder stop bolts. Refer to Tail Stinger - Maintenance Practices.
D.
Do a Functional Check of the Rudder System (Standard Rudder Installation). (1) Examine all cable runs for correct routing, fraying, and twisting. (a) Look for interference with the adjacent structure, equipment, wiring, plumbing, and other controls. (2) Examine the cable movement for binding and full range of travel. (3) Examine the swage fitting reference marks for signs of cable slippage inside of the fitting. (a) Examine the fitting for corrosion, distortion, cracks, and broken wires at the fitting. (4) Examine the turnbuckles for correct thread engagement. (a) Make sure that the turnbuckle locking clips are installed. Refer to Chapter 20, Safetying Maintenance Practices. (5) Move a cloth along the full length of the cable to examine for broken wires. (a) If snags are found or you think that there are broken wires, Refer to Chapter 20, Control Cable and Corrosion Limitations - Maintenance Practices. (6) Examine the cable attach holes in the rudder torque tube arm for excessive wear. (7) Examine the rudder torque tube, bellcrank, and lower hinge area for corrosion and condition. (8) Examine the rudder stop bolts for condition, corrosion, and security. (9) Examine the rudder hinge, hinge bearing, rudder pedals, and bonding jumper, for correct installation, corrosion, signs of damage, and unserviceable fasteners. (10) Examine the rudder skins for cracks, loose rivets, and corrosion. (11) Examine the balance weight for looseness and the supporting structure for damage.
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Examine the Travel and Cable Tensions. (1) Examine the cable tension in the tailcone area. NOTE:
(2) (3) (4) (5) (6) (7) (8)
Cable tensions must be measured at least one foot from any pulley or cable turnbuckle.
(a) The tension must be 30 pounds + 5 or -5 pounds (133.45 N + 22.24 or - 22.24 N). If necessary, do the Rudder System Rigging. Refer to Rudder - Maintenance Practices. Install the rudder travel protector. Put the rudder trim system in the neutral position. Operate the system through its full range of travel. (a) Make sure that all of the components that move do not hit, touch, or catch on structural components or other system components. Move the rudder left until the rudder stop block contacts the stop bolt. (a) The rudder travel on the protractor must be from 22 to 25 degrees. Move the rudder right until the rudder stop block contacts the stop bolt. (a) The rudder travel on the protractor must be from 22 to 25 degrees. Remove the rudder travel protractor.
F.
Restore access. (1) Install the tail stinger. Refer to Tail Stinger - Maintenance Practices. (2) Install the floor panels. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task Task 27-20-00-640 4.
Rudder Bar Bearings and Rudder Pedals Lubrication A.
General (1) This task gives the procedures to do a lubrication of the rudder bar bearings and rudder pedals.
B.
Special Tools (1) MIL-L-7870 or equivalent
C.
Access. (1) Remove or loosen top bearing blocks one at a time to get sufficient access to friction surface.
D.
Do a Lubrication of the Rudder Bar Attach Bearings and Rudder Pedals. (1) Clean and lubricate rudder bar attach bearings and all accessible component pivot points on the rudder bar.
E.
Restore Access. (1) Install or tighten the bearing blocks. End of task
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MODEL 208 MAINTENANCE MANUAL RUDDER - MAINTENANCE PRACTICES 1.
General A.
2.
Rudder maintenance practices consist of rudder pedals, pulleys and cables removal/installation, and rigging rudder system.
Rudder Pedals, Pulleys, and Cables Removal/Installation A.
Remove Rudder Pedals, Pulleys, and Cables (Refer to Figure 201). (1) Remove carpet or vinyl cover, plywood floor covers (118), floorboard access covers, tailcone upholstery panel, and tailcone stinger to gain access to system. (2) Remove cotter pins (75). nuts (74), washers (73), and bolts (72) from clevis ends of cables (70) and (71). Detach cables from cable arms (49). (3) Remove nuts (23) and (23A), washers (22) and (22A), bolts (21) and (21A), bearing (36), bushing (36A), washers (37), safety wire (35), and roll pin (25). Detach steering pushrod (20) from link (34), and flex shaft (26) from coupler (24). (4) Remove nuts (12), washers (11), bolts (10), cotter pins (8), spacers (7), and pins (6). Detach pedals (1), pedal arm assemblies (9). and links (5) from link assemblies (42) and forward and aft torque tubes (58) and (59). (5) Loosen nuts (47), detach hoses (62) and (63). Remove fittings (46) from right and left brake cylinders (14) and (15). Plug cylinder ports. (6) Remove cotter pins (19), washers (18), and pins (16) and (17). Detach left and right brake cylinders (14) and (15) from pedals (2) and bearings (54). (7) Remove nuts (57), washers (56), and bolts (55) from bearings (54) and spacers (38). Detach remaining rudder controls from floorboard of airplane and place on bench. (8) Remove cotter pins (8), spacer (7), and pins (6). Detach links (13) from pedals (2) and link assemblies (42) and (61). (9) Remove roll pins (4) and pins (3). Disconnect rudder pedals (2) from forward and aft torque tubes (58) and (59). (10) Remove nuts (45), washers (44), and bolts (43). Detach link assemblies (42) and (61) from forward and aft brake interconnect tubes (40) and (41). (11) Remove forward and aft interconnect tubes from forward and aft torque tubes (58) and (59). (12) Remove nuts (80), washers (79), and bolts (78). Detach pulleys (76) and spacers (77) from supports (81). (13) Remove nuts (85), washers (84), and bolts (83). Detach pulleys (82) from supports (86). (14) Airplanes 20800001 Thru 20800185 and 208B0001 Thru 208B0214 except airplanes incorporating SK208- 76, remove nuts (94), washers (93), and bolts (92). Detach fairleads (90) from supports (91). (15) Remove bolts (97) and washers (98). Detach pulleys (95) from supports (96). (16) Remove nuts (103), washers (102), and bolts (101). Detach pulleys (100) from supports (104). (17) Remove safety wire or clips and disconnect turnbuckles (106) and (116) cables (71), (107), (70), and (108). (18) Remove nuts (112), washers (111), spacers (110), and bolts (109). (19) Remove cables (70), (71), (107), and (108) from system. NOTE:
B.
To ease replacement of cables, attach a length of wire to cable to be removed from system. After removing cable, leave wire in place, routed through structure. When replacing cable, attach it to wire and pull cable into position with wire.
Rudder Pedals, Pulleys, and Cables Installation (Refer to Figure 201). NOTE:
(1)
Before assembling rudders, check condition of bushings in each end of forward and aft torque tubes (58) and (59). If bushings are scored or excessively worn, they should be replaced.
Install forward and aft brake interconnect tubes (40) and (41) in forward and aft torque tubes (58) and (59).
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Rudder System Installation Figure 201 (Sheet 1)
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Rudder System Installation Figure 201 (Sheet 2)
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Rudder System Installation Figure 201 (Sheet 3)
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Rudder System Installation Figure 201 (Sheet 4)
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Rudder System Installation Figure 201 (Sheet 5)
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Rudder System Installation Figure 201 (Sheet 6)
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Rudder System Installation Figure 201 (Sheet 7)
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MODEL 208 MAINTENANCE MANUAL (2)
Attach link assemblies (42) and (61) to forward and aft brake interconnect tubes (40) and (41). Replace bolts (43), washers (44), and nuts (45). NOTE:
(3)
(4) (5) (6)
Rig torque tubes as noted: (a) Align 0.093 inch diameter holes in rigging arms (50) and install 3/32 x 3/4 inch cotter pin (51). (b) Place lower halves of bearings (52) and (54) on a flat surface. Install forward and aft torque tubes (58) and (59) in bearings (52) and (54). (c) Check that gear teeth (60) are engaged and tape forward (58) and aft (59) torque tubes securely adjacent to teeth (60). Attach links (13) to link assemblies (42) and (61). Replace pins (6), spacers (7), and cotter pins (8). Attach pedals (2) to forward and aft torque tubes (58) and (59) and links (13). Replace pins (3), roll pins (4), pins (6) and cotter pins (8). Attach upper halves to lower halves of bearings (52) and (54). NOTE:
(7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21) (22) (23) (24) (25)
Forward and aft torque tubes (58) and (59) must be rigged before installing rudder controls in airplane.
To facilitate installation of rudder control in airplane, install one bolt (55) in each bearing (52) and (54) and hand-tighten nut (57). Move partially assembled rudder control from bench and install in airplane.
Remove hand-tightened nuts and install remaining bolts (55) and spacers (38) to attaching holes in floorboard. Replace washers (56) and nuts (57). Remove tape from forward and aft torque tubes (58) and (59), and cotter pin (51) from rigging arms (50). Attach pedals (1), pedal arm assemblies (9), and links (5) to forward and aft torque tubes (58) and (59). Replace bolts (10), washers (11), and nuts (12). Attach links (5) to link assemblies (42). Replace pins (6) and cotter pins (8). Connect right and left brake cylinders (14) and (15) to pedals (2) and bearings (54). Replace pins (16) and (17), washers (18), and cotter pins (19). Remove plugs from ports of right and left brake cylinders (14) and (15). Attach fittings (46), 0-rings (48), and nuts (47) to right and left brake cylinders (15). Tighten nuts (47). Attach hoses (62) and (63) to fittings (46). Add brake fluid and bleed system. Refer to Chapter 32 for brake system maintenance. Replace cables (70), (71), (107), and (108) in system. Attach clevis ends of cables (107) and (108) to rudder bellcrank (105). Replace bolts (109), bushings (110), washers (111), and nuts (112). Attach pulleys (100) and cables (70) and (71), to supports (104). Replace bolts (101), washers (102), and nuts (103). Airplanes 20800001 Thru 20800185 and 208B0001 Thru 208B0214 except airplanes incorporating SK208-76, attach fairleads (90) to supports (91); replace bolts (92), washers (93), and nuts (94). Attach pulleys (95) and cables (70) and (71), to supports (96). Replace bolts (97) and washers (98). Attach pulleys (82) and cables (70) and (71), to supports (86). Replace bolts (83), washers (84), and nuts (85). Attach pulleys (76) and cables (70) and (71), to supports (81). Replace bolt (78), spacer (77), washer (79), and nut (80). Attach clevis ends of cables (70) and (71) to cable arms (49). Replace bolts (72), washers (73), nuts (74), and cotter pins (75). Attach cables (71) and (107) to turnbuckle (106). Attach cables (70) and (108) to turnbuckle (116). Tighten turnbuckles (106) and (116) evenly; set cable tension at 25 to 35 pounds. Safety wire or install clips on turnbuckles. Check that threads in shaft (39) are centered in link (34) and nut (39A).
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MODEL 208 MAINTENANCE MANUAL (26) Attach flex shaft (26) to coupler (24). Align attaching holes and replace roll pin (25) and safety wire. (27) Connect steering pushrod (20) to link (34). Replace bearing (36), bushings (36A), bolts (21) and (21A), washers (22), (22A), (37), and nuts (23) and (23A). NOTE:
Make sure the rudder trim system is correctly rigged.
WARNING: Be sure rudder moves in correct direction when rudder pedals are actuated. Check for crossed or wrapped cables. (28) Replace carpet or vinyl cover, floorboard access covers, plywood floor covers (118), tailcone upholstery panel, and tailcone stinger. 3.
Rudder System Rigging A.
Rig Rudder System (Refer to Figure 201). NOTE: (1) (2) (3) (4)
All control surface cable tensions should be rigged at an ambient temperature of 70°F. Allow temperature to stabilize for a period of four hours before setting cable tensions.
Remove carpet or vinyl floor cover, plywood floor covers (118), floorboard access covers, upholstery panel at tailcone entrance, and tailcone stinger to gain access to system. Remove nut (23), washer (22), and bolt (21). Detach pushrod (20) from link (24). Move pushrod away from path of link. Remove safety wire and roll pin (25). Disconnect flex shaft (26) from coupler (24). Remove safety wire or clips and loosen turnbuckles (106) and (116). NOTE:
(5) (6) (7) (8)
Hold rudder rigging protractor against vertical stabilizer with dial adjacent to trailing edge of rudder 1.56 inches (39.62 mm) above rudder skin trim (measure along fin skin trim) with aft end at lower end of rudder trailing edge.
Clamp a block across face of rudder pedals (1) and (2). Adjust turnbuckles (106) and (116) to place trailing edge in line with zero mark on rudder travel protractor. Set cable tension at 25 to 35 pounds; safety wire or install clips on turnbuckles. Remove block from rudder pedals (1) and (2). Loosen locknuts (115) and set left and right stop bolts (113) and (114) to allow maximum rudder travel in either direction of 23 to 27 degrees on landplanes regardless of which rudder is installed. Tighten locknuts (115). NOTE:
Amphibian or floatplanes equipped with Wipline Model 8000 floats and floatplane rudder: set rudder travel at 23 to 25 degrees left and right.
(9) Remove rudder rigging protractor from vertical stabilizer. (10) Check that threads on shaft (39) are centered in link (34) and nut (39A). Adjust if required. (11) With rudder streamlined, install flex shaft (26) on coupler (24) and align attaching holes. Replace roll pin (25) and safety wire. (12) Attach steering pushrod (20) to link (34). Replace bolt (21), washer (22), and nut (23). (13) Replace tailcone stinger, tailcone upholstery panel, floorboard access covers, plywood floor covers (118), and carpet or vinyl floor cover. NOTE:
If pointer (28) fails to center when rudder is streamlined, loosen screw (29) until pointer (28) can be moved to neutral position. Tighten screw after setting pointer.
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MODEL 208 MAINTENANCE MANUAL RUDDER TRIM - DESCRIPTION AND OPERATION 1.
General A.
2.
Rudder trim is provided to ease required pedal force while airplane is operating.
Description and Operation A.
The rudder trim system consists of a trim wheel on the control pedestal, a flexible shaft attached to the forward rudder torque tube and the nose gear steering pushrod. A trim indicator is attached to the trim wheel.
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MODEL 208 MAINTENANCE MANUAL RUDDER TRIM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101
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Rudder Trim System Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL RUDDER TRIM - MAINTENANCE PRACTICES 1.
General A.
2.
Rudder system maintenance practices consist of rudder trim system removal/installation and rudder trim system rigging.
Rudder Trim System Removal/Installation A.
Remove Rudder Trim System (Refer to Figure 201). (1) Remove screws (3) and washers and loosen upper cover (2). (2) Remove roll pins (7) and disconnect aileron trim wheel (6) from shaft (27). (3) Remove screws (5) and disconnect aft cover (4) from pedestal (1). NOTE:
Remove firewall shutoff valve control mounting brackets, and disconnect control (8) from firewall shutoff valve (31). Leave control (8) attached to aft cover (4) and move cover away from control pedestal (1).
(4)
Remove screw (12) and spacer (13) from bracket (28). Disconnect pointer (11) and remove from pedestal (1). (5) Remove roll pin (10) and disconnect rudder trim wheel (9) from flexible shaft (14). (6) Remove clamp (15) from control pedestal (1). Remove roll pins (17) from coupling (16). Disconnect flexible shaft (14) from control pedestal bracket and coupling (16). (7) Disconnect coupling (16) from shaft (21). (8) Remove snap rings (20) and (19) from shaft (21) and nut (18). (9) Remove nuts (25) and (25A), washers (24) and (24A), bolts (23) and (23A). Disconnect link (22) from steering pushrod (26). (10) Disconnect link (22) from shaft (21). (11) Disconnect shaft (21) from nut (18). Remove snap rings (19) from nut (18). Disconnect nut (18) from steering arms (30). B.
3.
Install Rudder Trim System (Refer to Figure 201). (1) Replace nut (18) to steering arms (30). Replace snap rings (19). Connect shaft (21) to nut (18). (2) Connect link (22) to shaft (21). (3) Connect link (22) to steering pushrod (26). Replace bolts (23) and (23A), washers (24) and (24A), and nuts (25) and (25A). (4) Replace snap rings (20) on shaft (21). (5) Connect coupling (16) to shaft (21). Replace roll pin (17). Safety-wire roll pin. (6) Connect flexible shaft (14) to control pedestal bracket and coupling (16). Replace roll pin (17) and clamp (15). Safety-wire roll pin and clamp. (7) Connect rudder trim wheel (9) to flexible shaft (14). Replace roll pin (10). (8) Connect pointer (11) to bracket (28). Replace spacer (13) and screw (12). (9) Connect fuel shutoff cable (8) to fuel shutoff valve. Replace fuel shutoff mounting brackets. (10) Connect aft cover (4) to pedestal (1). Replace screws (5) and washers. (11) Connect aileron trim wheel (6) to shaft (27). Replace roll pins (7) (12) Replace screws (3) and washers in upper cover (2).
Rudder Trim System Rigging A.
Rig (1) (2) (3) (4) (5) (6) (7)
Rudder Trim System (Refer to Figure 201). Remove nut (25), washer (24), and bolt (23). Disconnect link (22) from steering pushrod (26). Secure rudder in streamlined position. Loosen screw (12) and locate pointer (11) in neutral position. Tighten screw (12). Remove roll pin (17). Disconnect coupling (16) from shaft (21). Rotate shaft (21) until threads extend equidistant on each side of nut (18). Connect shaft (21) to coupling (16). Replace roll pin (17). Safety-wire roll pin. Rotate forward end of shaft (21) up and thread on link (22) until 0.50 inch of thread is showing on shaft (21). (8) Connect link (22) to steering pushrod (26). Replace bolt (23), washer (24), and nut (25).
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Rudder Trim System Installation Figure 201 (Sheet 1)
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Rudder Trim System Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (9)
Release rudder from streamlined position. NOTE:
After rigging rudder trim system, check nose gear steering rigging instructions. (Refer to Chapter 32, Nose Landing Gear - Maintenance Practices ).
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MODEL 208 MAINTENANCE MANUAL YAW DAMPER - ADJUSTMENT/TEST 1.
Yaw Damper Rigging A.
Adjust the Yaw Damper (Refer to Figure 501). (1) Make sure the rudder control system has been adjusted correctly. Refer to Chapter 27, Rudder - Maintenance Practices. (2) Adjust the slip clutch yaw damper servo to a setting of 50 +5 or -5 inch pounds. Refer to the King drawing number 155-9347-00 (or latest revision) for the clutch setting procedures. (3) Connect the yaw damper servo cables to the rudder cables with the clamp, bolts, washers and nuts. (a) Torque the cable clamp nuts each to 50 +5 or -5 inch pounds. (4) Set the rudder control cables in the center of the travel position and the ball position of the yaw damper servo capstan straight forward. (a) Wrap the bridle cable around the capstan. (b) Use the turnbuckle and apply tension to the bridle cable to 20 +5 or -5 inch pounds. (c) Use a lock clip and lock the turnbuckles. Refer to Chapter 20, Safetying - Maintenance Practices. (d) Install the capstan cable guard with one leg aft. (5) The KRG331 yaw rate gyro must be installed so that it is level +2 or -2 degrees in all directions when the airplane is in level flight. NOTE:
The inverter, yaw computer and yaw rate gyro are located below the copilot's seat.
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Yaw Damper Rigging Figure 501 (Sheet 1)
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Yaw Damper Rigging Figure 501 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL ELEVATOR SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
Model 208 elevator system consists of components used to control the elevator up or down.
Description and Operation A.
The elevator is controlled by a control column support that is attached to a pushrod and attached to a bellcrank. Left and right bellcrank arms are equipped with links attached to cables. Left and right cables are routed under the floorboard to turnbuckles in the tailcone. A second set of cables, connected to turnbuckles, are routed to a bellcrank in the tailcone. A pushrod attaches bellcrank to an elevator torque tube. The remainder of system consists of pulleys, cables, supports, and attaching hardware.
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MODEL 208 MAINTENANCE MANUAL ELEVATOR SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Elevator System Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ELEVATOR SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the elevator system in a serviceable condition.
Task 27-30-00-720 2.
Elevator System Functional Check A.
General (1) This task gives the procedures to do a functional check of the elevator system.
B.
Special Tools (1) Inclinometer (2) Cable Tensiometer (3) Elevator Neutral Rigging Tool (4) Elevator Rigging Protractor (5) Spring Scale (0 to 20 Pounds) (6) External Electrical Power Unit
C.
Access (1) Remove the applicable floor panels to get access to the elevator control system. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Remove vertical stabilizer panel 320A to get access to the elevator control system. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do a Functional Check of the Elevator System. (1) Do a check of the cable movement for binding and full travel. (2) Examine the elevator skins for cracks and loose rivets. (3) Examine the elevator hinges, hinge bolts, hinge bearings, torque tube, horn, attach fittings, and bonding jumper for corrosion, cracks, signs of damage, wear, unserviceable fasteners, security, and correct safeftying. (4) Examine the balance weights and the support structure for corrosion, looseness, cracks, and damage. (5) Examine the outboard tips for cracks in the rib flange. (6) Examine the bellcracks, bearings, push rods, stop bolts, and brackets, for corrosion, cracks, signs of damage, failed fasteners, security of installation, and correct safetying. (7) Examine the elevator and elevator trim cable runs for correct routing, fraying, and twists. (a) Make sure there is no interference with the adjacent structure, equipment, wiring, plumbing, and other controls. (8) Move a cloth along the full length of the cables to examine for broken wires. (a) If snags are found or you think that there are broken wires, refer to Chapter 20, Control Cable Wire Breakage and Corrosion Limitations - Maintenance Practices. (9) Examine the turnbuckles for correct thread exposure. (a) Make sure that the turnbuckle locking clips are installed correctly. Refer to Chapter 20, Safetying - Maintenance Practices. (10) Examine the swage fittings reference marks for an indication of cable slippage inside of the fitting. (a) Examine the fittings for corrosion, distortion, cracks, and broken wires at the fittings. (11) Examine the pulleys, attach brackets, and guard pins for condition, wear, corrosion, and security. (a) You must turn the pulleys to make sure there freedom of movement and to make sure there is even wear of the pulleys. (b) If discrepancies are found with the brackets, examine the structure where the brackets are attached for hidden damage. (12) Examine the elevator trim tab actuators for corrosion, damage, and security. (13) Examine the elevator trim tab actuator mounting structure for corrosion, damage, cracks, and security of installation at the horizontal stabilizer rear spar.
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MODEL 208 MAINTENANCE MANUAL (14) Examine the elevator trim tab actuator pushrods and attaching hardware for corrosion, condition, damage, wear and security of installation. (15) Examine the elevator trim actuator motor for corrosion, security of installation, and signs of overheating. (16) Examine the chain for corrosion, tension, and correct alignment (17) Examine the control column for corrosion, signs of damage, unserviceable fasteners, and security of installation. (18) Examine the column lock for correct operation. (19) Examine all welds in the column tube and the torque tube for corrosion and cracks. (20) Examine both torque tube support arms for corrosion, condition, and security of the attach bearings. (21) Examine the support arm attach structure for condition, cracks, and correct safety of the attach bolts. (22) Examine the cable guards for corrosion, condition, and security on both column quadrants. (23) Examine for sufficient clearance of all components and structure at the full aft and full forward positions. (24) Make sure that the elevator, elevator actuator, and elevator mount are correctly attached together with four bolts and washers. (25) Make sure that the actuator mount is correctly attached to the mounting bracket with four bolts. (26) Make sure that the lower cable assembly is correctly attached to the aft part of the bellcrank assembly attach bracket with one bolt, nut, washers, and cotter pin. (27) Make sure the chain is correctly centered and aligned on the sprocket. (a) The chain guard posts must be correctly installed and attached with safety wire. (28) Make sure that the chain is correctly attached to the cable assembly and the turnbuckle terminal with the chain connecting links. E.
Examine the Cable Travel and Tensions. (1) Set the control wheels to put the elevators in the neutral position. (2) Make sure that the left elevator is at the streamlined position (3) Attach an inclinometer on the left elevator's trailing edge and set it to zero degrees.
CAUTION: Do not attempt to align the horn (balance weight portion) on the elevator to the stabilizer. CAUTION: Make sure that the support stand is under the tail to prevent the tail cone from dropping while working in the tail cone. (4)
(5) (6) (7) (8) (9) F.
Examine the cable tensions and adjust if necessary. (a) For the elevator control cables, refer to Elevator - Maintenance Practices. (b) For the elevator trim cables, refer to Elevator Trim System - Maintenance Practices. (c) For airplanes equipped with 400B and 400B IFCS autopilot type AF-550A and IF-550A, refer to Elevator - Adjustment/Test. Operate the system through its full range of travel. (a) Make sure that all of the components that move do not hit, touch, or catch on structural components or other system components. Move the elevator to contact the down stop bolt. (a) Make sure that the inclinometer shows 20 +2 or -2 degrees. Move the elevator to contact the up stop bolt. (a) Make sure that the inclinometer shows 25 +2 or -2 degrees. Remove the inclinometer from the left elevator trailing edge. Do an electric elevator trim clutch torque system check, (Refer to Electric Elevator Trim Adjustment/Test).
Do an Electric Elevator Trim System Operational Check (1) Connect external electrical power to the airplane. (2) Set the External Power Switch to BUS. (3) Set the Battery switch to ON.
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MODEL 208 MAINTENANCE MANUAL (4)
(5) (6) (7)
Do a check to make sure that the left and the right elevator trim switch halves operate correctly. (a) Move the right switch half forward to the DN position momentarily, then release it to the center position. 1 Make sure that the elevator trim wheel does not move. (b) Move the right switch half aft to the UP position momentarily, then release it to the center position. 1 Make sure that the elevator trim wheel does not move. (c) Move the left switch half forward to the DN position momentarily, then release it to the center position. 1 Make sure that the elevator trim wheel does not move. (d) Move the left switch half aft to the UP position momentarily, then release it to the center position. 1 Make sure that the elevator trim wheel does not move. (e) Move and hold the left and the right switch halves forward to the DN position and do the following before the elevators reach the full down position. 1 Make sure that the elevator travel direction is correct. 2 Push and release the A/P Trim Disconnect push button. a Make sure that the elevator trim wheel stops does not move after the A/P Trim Disconnect push button is pushed and released. (f) Move and hold the left and the right switch halves aft to the UP position and do the following before the elevators reach the full up position. 1 Make sure that the elevator travel direction is correct. 2 Push and release the A/P Trim Disconnect push button. a Make sure that the elevator trim wheel stops does not move after the A/P Trim Disconnect push button is pushed and released. (g) Release the left and the right switch halves to the center OFF position. (h) Operate the system through the full range of travel and examine for binding, jerking movements, and sluggish operation. (i) Examine the operating time for the full range of motion. 1 Airplanes equipped with King KFC-150 or -250 autopilot must complete the full range of travel from 26 to 38 seconds. 2 Airplanes equipped with King KFC-225 autopilot must complete the full range of travel from 16 to 24 seconds. Set the Battery switch to OFF. Set the External Power Switch to OFF. Disconnect the external electrical power unit from the airplane.
G.
Restore Access (1) Install vertical stabilizer panel 320A. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Install the applicable floor panels that were removed to get access to the elevator control system. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task
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MODEL 208 MAINTENANCE MANUAL ELEVATOR - MAINTENANCE PRACTICES 1.
General A.
2.
This section includes maintenance practices for the removal/installation of the elevator system and the elevator system rigging procedures.
Elevator System Removal/Installation A.
Remove the Elevator System (Refer to Figure 201). (1) Remove the carpet or vinyl cover, plywood floor covers, floorboard access covers, and upholstery panel to get access to the tailcone. (2) Remove the cotter pins, nuts, washers and bolts, then disconnect the pushrod from the support arm and the bell crank. (3) Remove the bolts, washers and nuts, then remove the upper and lower links from the forward bell crank and cables. (4) Remove the bolt and washers, then disconnect the forward bell crank from the support. (5) Remove the nuts, washers and bolts, then remove the pulleys from the supports. (6) Remove the nuts, washers and bolts, then disconnect the fairleads from the supports (Airplanes 20800001 thru 20800185 and 208B0001 thru 208B0214 not incorporating SK208-76). NOTE:
Airplanes 20800186 and On and 208B0215 and On, and airplanes incorporating SK208-76 do not have the fairleads installed.
(7) (8) (9)
Remove the nuts, washers and bolts, then disconnect the pulleys from the supports. Remove the safety wire or clips and disconnect the turnbuckles from the cables. Remove the nuts, washers and bolts, then remove the right and left links from the cables and the aft bell crank. (10) Remove the cotter pins, nuts, washers and bolts, then disconnect the pushrod from the bell crank and the elevator pushrod arm. (11) Remove the cotter pins, nut, washer, races, bearings, spacer and bolt, then disconnect the aft bell crank from the support. NOTE:
To make the removal and installation of the cables easy, attach a length of wire on the opposite end of the removal end of the cable. When the cable is removed, leave the wire in position to route the cable through the airplane structure. Pull the replacement cable into the correct location with the wire.
(12) Remove the cables from the airplane. B.
Install the Elevator System (Refer to Figure 201). (1) Install the cables in the airplane. (2) Attach the aft bell crank to the support, then install the spacer, bolt, races, bearings, washer, nut and cotter pins. (3) Attach the pushrod to the aft bell crank and the elevator pushrod arm, then install the bolts, washers, nuts and cotter pins. (a) Install the pushrod so that the placard is on the lower forward end to facilitate inspections. (4) Attach the forward bell crank to the support, the install the bolts and washers. (5) Connect the pushrod to the support arm and the forward bell crank, then install the bolts, washers, nuts and cotter pins. (6) Connect then links to the forward bell crank and cables, then install the bolts, washers and nuts. (7) Attach the pulleys to the support, then install the bolts, washers and nuts. (8) Connect the fairleads to the support, then install the bolts, washers and nuts (Airplanes 20800001 thru 20800185 and 208B0001 thru 208B0214 not incorporating SK208-76). NOTE: (9)
Airplanes 20800186 and On and 208B0215 and On, and airplanes incorporating SK208-76 do not have the fairleads installed.
Attach the pulleys to the supports, then install the bolts, washers and nuts.
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Elevator System Installation Figure 201 (Sheet 1)
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Elevator System Installation Figure 201 (Sheet 2)
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Elevator System Installation Figure 201 (Sheet 3)
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Elevator System Installation Figure 201 (Sheet 4)
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Elevator System Installation Figure 201 (Sheet 5)
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Elevator System Installation Figure 201 (Sheet 6)
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Elevator System Installation Figure 201 (Sheet 7)
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Elevator System Installation Figure 201 (Sheet 8)
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MODEL 208 MAINTENANCE MANUAL (10) Attach the links to the forward bell crank and cables, then install the bolts, washers and nuts. (11) Connect the cables to the turnbuckles that connect the up and down cables to the forward and aft bell cranks. (12) Tighten the turnbuckles gradually to set the cable tension at 55 to 65 pounds, then safety the cables with wire or install clips on the turnbuckles.
WARNING: Make sure that elevators move in correct direction while actuating control column forward and aft. Also, check for crossed or wrapped cables. (13) Adjust the elevator system rigging. Refer to Elevator System Rigging. (14) Install the upholstery panel at the entrance to the tailcone, floorboard access covers, plywood floor covers, carpet and the vinyl covers. 3.
Elevator System Rigging A.
Rig the Elevator System (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7) (8)
All control surface cable tensions should be rigged at an ambient temperature of 70°F. When the temperature has been stable for four hours, then you can set the cable tensions.
Remove the control lock and replace it with an elevator neutral rigging tool. Remove the upholstery panel from the entrance to the tailcone. Remove the safety wire, then loosen the turnbuckles that connect the up and down cables to the bell crank. Use the elevator rigging protractor to set the elevator in the neutral position, then tighten the turnbuckles sufficiently to attach the elevator in the neutral position. Attach the inclinometer to the elevator, then set the elevator in the zero position. Tighten the turnbuckles until the elevator cannot be moved from the zero setting. Remove the access covers 212FR and 211EL (Airplanes 20800130 and On and 208B0087 and On). Refer to Chapter 6, Access Plates and Panels Identification - Maintenance Practices. Remove the elevator neutral rigging tool, then set the elevator to 18 to 22 degrees in the DOWN position on the elevator rigging protractor. NOTE:
(9) (10) (11) (12) (13) (14) (15) (16)
You can adjust the DOWN stop bolt if necessary.
For the Model 208 airplanes without the TKS anti-ice system installed and all Model 208B airplanes, set the elevator from 23 to 27 degrees on the UP mark on the elevator rigging protractor, then adjust the UP stop bolt if necessary. For the Model 208 airplanes with the TKS anti-ice system installed, set the elevator to 17 to 19 degrees on the UP mark on the elevator rigging protractor, then adjust the UP stop bolt if necessary. For Airplanes 20800130 and On and 208B0087 and On, remove or install stop blocks as necessary for full normal up travel and to get the same dimensions that are shown in Detail G. Attach the tensiometer to the elevator cables, then do a check of the elevator tension settings in at more than one location. Set the cable tension to 55 to 65 pounds, then safety the turnbuckles with wire. Remove the tensiometer from the elevator cable, then replace the upholstery panel at the entrance to the tailcone. Install the access covers if they were removed. Refer to Chapter 6, Access Plates and Panels Identification - Maintenance Practices. Do a check of the friction band, then replace the control lock after you complete the check of the friction band. Refer to Friction Band Check.
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4.
Friction Band Check A.
Do a Check of the Friction Band NOTE:
(1) (2)
All friction band measurements must be done with the load applied to a load scale parallel to the direction of movement of the control column. The control column movement must be slow and stable over the full range of travel.
With the control tube one inch forward of the neutral position, make a load scale reading as the column crosses the neutral position. (a) Identify this reading as F1. With control tube one inch aft of neutral position, make a load scale reading in the opposite direction. (a) Identify this reading as F2. NOTE:
(3) (4)
If during step 2 the control tube moves forward by itself after being released from the starting point, then subtract the F2 reading from the F1 reading.
The sum of the F1 reading plus the sum of the F2 reading equals the friction band. The friction band should be set at 15 pounds maximum without the autopilot installation, and set at 20 pounds maximum with the autopilot installation. NOTE: (a) (b) (c) (d)
If the friction band is more than the acceptable limit, do the checks that follow:
Do a check of the elevator cable tension to make sure it is set it at 55 pounds. Do a check of the alignment of the pulleys and cables. Do a check for binding of the pulleys and cables. Make sure there is no interference with pulleys and cables and the airplane structure or adjacent cables.
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MODEL 208 MAINTENANCE MANUAL ELEVATOR - ADJUSTMENT/TEST 1.
2.
General A.
This section has procedures for the following airplanes. (1) Airplanes with the 400B, 400B IFCS Autopilot Types AF-550A, IF-550A installed. (2) Airplanes with the KAP-150 Autopilot, KFC-150, KFC-225 Flight Control System installed.
B.
The following procedures give instructions for the elevator rigging and friction band for these airplanes.
Elevator Rigging (Airplanes with 400B and 400B IFCS Autopilot Types AF-550A and IF-550A Installed) A.
Adjust the elevator rigging (Refer to Figure 501 and 502). (1) Set the control wheels in the neutral position of elevator. (2) Attach a neutral rigging tool to the control column. If necessary, make a neutral rigging tool for the control column. (3) Use an inclinometer on the left elevator's trailing edge and adjust inclinometer to show zero degrees. The left elevator must be in the streamlined position.
CAUTION: Do not attempt to align the horn (balance weight portion) on the elevator to the stabilizer. (4)
Remove the aft baggage partition to get access to the elevator actuator, chain, and cable assemblies.
CAUTION: Make sure a support stand is under the tail to prevent the tail cone from dropping while working inside the tail cone. (5)
Make sure the elevator actuator cable assemblies are correctly rigged. (a) Make sure the elevator actuator and the elevator actuator mount are correctly attached together with four bolts and washers. (b) Make sure the actuator mount is correctly attached to the mounting bracket with four bolts. (c) Make sure the mounting bracket assembly is correctly attached to the bulkhead station 308.00. (d) Make sure the lower cable assembly is correctly attached to the aft portion of the bell crank assembly, attach bracket with one bolt, washers, nut, and cotter pin. (e) Make sure the chain is correctly centered and aligned on the actuator sprocket. The chain guard posts must be correctly installed and attached with safety wire. (f) Make sure the chain is correctly attached to the cable assembly and turnbuckle terminal with the chain connecting links. (g) Make sure the turnbuckle is attached with safety wire. Refer to Chapter 20, Safetying Maintenance Practices. NOTE:
If the chain is to be replaced, you must Þrst remove the safety wire attached to the two chain guard posts. Then remove the chain guard posts.
CAUTION: Do not lose the lock washers when you loosen the chain guard posts for removal. (h) (6)
(i) Do (a) (b)
Make sure the upper cable fork is correctly attached to the upper forward part of the bell crank attach bracket with a bolt, washer, nut , and cotter pin. Make sure the chain and terminal are correctly attached together with a chain connect link. a check of the elevator actuator cable tension. Make sure the pitch actuator is rigged. Attach a cable tensiometer to the elevator actuator cable. The tensiometer must indicate 10 to 15 pounds.
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Elevator Rigging Figure 501 (Sheet 1)
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Elevator Rigging Figure 501 (Sheet 2)
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Elevator Rigging Figure 501 (Sheet 3)
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Elevator Rigging Figure 501 (Sheet 4)
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Inclinometer Figure 502 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (c)
(7) (8) (9) (10)
(11) (12) (13) (14) (15)
(16)
If the cable tension is less than 10 pounds or more than 15 pounds, adjust the cable to bring it to the desired tension. 1 Remove the safety wire from the turnbuckle. 2 If cable tension is too high, loosen the turnbuckle. If cable tension is too low, tighten the turnbuckle. 3 Attach safety wire to the turnbuckle. Refer to Chapter 20, Safetying - Maintenance Practices. 4 Remove the cable tensiometer from the actuator cable. Check the inclinometer on the left elevator trailing edge to make sure it shows 0 degrees. Remove the inclinometer if value is 0 degrees. If the value is not 0 degrees, repeat the elevator rigging procedure. Remove the rigging tool attached to the control wheels. The 400B autopilot has a GYRO switch located on the rear of the control head to eliminate the gyro roll and pitch signals generated by a non-erected gyro. (a) Set the GYRO switch to the OUT position. The 400B IFCS does not have a gyro out switch. It will be necessary to hook up an outside vacuum source or operate the engine to erect the gyro. (a) If an outside vacuum source is used, it must be calibrated in inches of mercury. The desired suction range required to erect the gyro is 4.6 to 5.4 inches of mercury. (b) If the engine is operated to erect the gyro, the engine must be run at 65 degrees No. This will provide the desired amount of vacuum and maintain correct bus voltage. Turn to ON, the airplane battery switch, 1 & 2 avionics power switches, and the autopilot (A) switch. Allow the pitch wheel to run to electrical neutral. Pull the control wheel back until the control forces are neutralized. Hold the control wheel to prevent full travel and rotate the PITCH command wheel in the NOSEDOWN direction. The control wheel must move forward. Turn the autopilot (A) switch to OFF and then back to ON. Allow the PITCH command wheel to neutralize. Pull the control wheel back until there is a slight tendency for the control wheel to travel forward. Rotate the PITCH command wheel in the NOSE-UP direction. The control wheel must be prevented to move forward by the actuator force. Push control wheel forward 2 to 3 inches to conÞrm that actuator force is being applied. Check that the autopilot can be overpowered using control wheels at any time. This practice must be kept to a minimum because wear of the slip-clutch will result from extended periods of manual overpower. Extended periods of manual overpower of the actuator will cause a thermostatic switch to remove power from the actuator. After approximately 10 minutes, the switch will automatically reset to close the autopilot interlock circuit. Power can then be reapplied to the actuator by engaging the AP/ON-OFF switch. NOTE:
Because of ground loads applied to the elevators, elevator travel from stop to stop cannot be obtained in most instances while you conduct the ground tests. The service technician must not assume the elevator actuator is not completely operational. Flight tests in accordance with the appropriate 400B Service/Parts manual must be completed after rigging procedures are completed to fully determine if the system is completely operational. Make sure the Gyro switch is set in the IN position before ßight tests are initiated.
(17) Turn the autopilot (AP) switch, avionics power switches, and the airplane battery switch to OFF. (18) Install the aft baggage partition.
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3.
Elevator Friction Band (Airplanes with 400B and 400B IFCS Autopilot, Types AF-550A and IF-550A Installed) A.
Do a check of the elevator friction band. NOTE:
(1)
(2)
(3)
4.
The elevator friction band without an autopilot installed must be 15 pounds or less. The maximum allowable friction band is 20 pounds with an autopilot installed. You must check the friction band every time the autopilot rigging is checked. The friction band must be measured and calculated. Correct action to reduce the elevator system friction band must be completed.
All friction band measurements must be made with the load scale so the force exerted to move the elevator is applied in a direction parallel to the control column. The movement of the control wheel must be slow and steady. The friction band requirements apply to all of the elevator travel range. The production inspection must be made at or in 1.00 inch of either side of the position occupied by the control wheel when the elevator is set in the neutral position. (a) Put the elevators in a down position (the control wheel one inch forward of the position occupied when elevator is in neutral position). 1 Move the control wheel aft by exerting force described in Step 3.A.(1). 2 Do a load scale check as the control wheel passes the elevator neutral position. Identify the load scale check as F1. (b) Put the elevators in an up position (the control wheel assembly one inch aft of the position occupied when elevator is in neutral position). 1 Move the control wheel assembly forward by exerting force described in Step 3.A.(1). 2 Do a load scale check as the control wheel passes the elevator neutral position. Identify the load as F2. The elevator friction band must be calculated from the measurements taken from the previous steps. (a) The friction band is the sum of forces F1 and F2 where a force must be applied in a forward direction toward the panel to return the elevators to a down position. The friction band = F1 + F2. (b) The friction band is the difference between two forces where a restraining force must be applied in a aft direction away from the panel to prevent the elevator from abruptly returning to a down position. The friction band = F1 - F2. When the friction band exceeds the limits, the following instructions must be completed in the order shown to reduce friction to the system. (a) Decrease the tension to the elevator cable to 15 pounds with the elevator set in the neutral position. (b) Do a check for binding and correct alignment of the pulleys. Correct alignment if necessary. (c) Make sure the hinges are clean and lubricated.
Slip Clutch Adjustment, Values and Capstan (Airplanes with KAP-150 Autopilot and KFC-150, KFC225 Flight Control System Installed) A.
Servo Slip Clutch Torque Settings (Refer toFigure 503). NOTE: (1)
The servo slip clutch torque settings are adjustable and must be set before servo installation.
Set the servo slip clutch torque to the appropriate value. NOTE:
The Þxtures and tools required to complete the adjustments are supplied with the KTS 150 test set and the KTS 158 test set.
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Clutch Torque Adjustment Figure 503 (Sheet 1)
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Clutch Torque Adjustment Figure 503 (Sheet 2)
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Clutch Torque Adjustment Figure 503 (Sheet 3)
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5.
KM 275 and KM 277 Slip Clutch Torque Adjustment (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) A.
Adjust the Slip Clutch Torque (Refer to Figure 503). (1) Refer to Slip Clutch Adjustment, Values and Capstan to determine the servo mount part number and setting required for each axis of the airplane. (2) Remove the capstan guard from the KM 275 and KM 277 Capstan plate. (3) Attach the KM 275 or KM 277 servo mount and servo motor to the KTS 158 test stand (0474238-01). (a) When you adjust a KM 275, place the adapter tool over the KM 275 capstan and insert the positioning pin (from the straight-up position) to attach the adapter tool. (b) When you adjust a KM 277, use the three sprocket pins (071-6065-00) to attach the adapter tool to the capstan. NOTE: (4) (5) (6)
An alternative adjustment method for the KM 277 is to use the King gear adapter assembly (071-6018-06).
Insert a torque wrench (Snap-On TEP-6FUA or equivalent). Connect the servo motor to the appropriate KTS 158 Test Set connector and apply power to the servo motor. Do a test of the torque value. NOTE:
The desired torque value is the average of the maximum and minimum indications from the clockwise and counter clockwise rotations. The test must be repeated three times in each direction and then the average of the six values is used to determine the true torque value.
(a)
(7) (8) 6.
Use the appropriate switch on the KTS 158 Test Set and turn the servo motor in the clockwise direction. Write down the torque value shown on the wrench. (b) Use the appropriate switch on the KTS 158 Test Set and turn the servo motor in the counter clockwise direction. Write down the torque value shown on the wrench. (c) If the level measured falls below the desired value, rotate the clutch adjust nut clockwise. (d) If the level measured falls above the desired value, rotate the clutch adjust nut counterclockwise. After an adjustment, repeat the torque test. After wiring has been completed and the servos installed, make sure the rotation direction of the servo capstans is correct.
Elevator Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) A.
Adjust the elevator rigging (Refer to Figure 504). (1) Set the control wheels in the neutral position of elevator. (2) Attach to the control column a neutral rigging tool. If necessary, make a neutral rigging tool for the control column. (3) Install an inclinometer on the left elevator's trailing edge and adjust the inclinometer to read zero degrees while the left elevator is set in the neutral position.
CAUTION: Do not align the horn (balance weight portion) on the elevator to the stabilizer. (4) (5) (6)
Set in position a support stand under the tail skid to prevent the tailcone from dropping while working inside the tailcone. Remove the aft baggage partition to get access to the elevator, servo mount and bridle cable assembly. Adjust the elevator servo slip clutch values and capstan. Refer to Slip Clutch Adjustment, Values and Capstan and Slip Clutch Torque Adjustment.
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Elevator Rigging Figure 504 (Sheet 1)
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Elevator Rigging Figure 504 (Sheet 2)
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Elevator Rigging Figure 504 (Sheet 3)
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Elevator Rigging Figure 504 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (7)
The KS 270A Pitch Servo, KM 275 Servo Mount and the ßap sense circuit breakers are installed for the KAP-150 Autopilot System and KFC-150, KFC-225 Flight Control System. The KS 270A and KM 275 Servo Mount are located just aft of FS 308.0 in the center of the airplane. (8) Make sure the pitch servo bracket is correctly attached to the servo gusset. (9) Make sure the pitch KM 275 servo mount is correctly attached to the pitch servo bracket and servo gusset. (10) Do an inspection to make sure the bridle cable is correctly attached to the top and bottom attach brackets. (11) Place the elevator in the full nose down position. (a) Do a check to make sure the KM 275 capstan ball location is rotated approximately 30 degrees aft from the straight up position. (b) Make sure the pitch bridle cable is wrapped on the KM 275 Capstan. (c) Make sure the turnbuckle end of the pitch bridle cable is routed from the top of the KM 275 Capstan to the top pitch link of the bell crank. (d) Make sure the opposite end of the pitch bridle cable is routed from the bottom of the KM 275 Capstan to the bottom pitch link of the bell crank. (e) Make sure the attach ends of the pitch bridle cable are correctly attached to the bell crank pitch links. (f) Install the Tensiometer on the pitch bridle cable and adjust the tension on the turnbuckle of the pitch bridle cable to 20 +5 or -5 pounds. (g) Remove the Tensiometer and make sure the turnbuckle clips are installed. (12) Check the pitch bridle cable for fraying and corrosion. 7.
Elevator Friction Band (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) A.
Do a check of the elevator friction band. NOTE:
(1)
(2)
Without an autopilot installed the elevator friction band is 15 pounds or less. With an autopilot installed, the elevator friction band is 20 pounds or less. Measurement of the elevator friction band is necessary when you examine the autopilot rigging. An adjustment to decrease the aileron system friction band is necessary if it is higher than the band range speciÞed.
All friction band measurements must be made with the load scale. The load scale must be set in a position so that the force exerted to move the elevator is applied in a direction parallel to the control column. The movement of the control wheel assembly must be slow and steady. The friction band requirements apply over the complete elevator travel range. The inspection must be made at or within one inch from either side of the position occupied by the control wheel assembly when the elevator is in neutral position. (a) Start with the elevators in a down position (control wheel assembly is one inch forward of the position occupied when elevator is in neutral position). Move the control wheel assembly aft by exerting force as instructed in Step 7.A. (1). 1 Write down the load scale value as the wheel assembly passes the elevator neutral position. Identify the value as F1. (b) Start with the elevators in an up position (control wheel assembly is one inch aft of the position occupied when elevator is in neutral position). Move the control wheel assembly forward by exerting force as instructed in Step 7.A. (1). 1 Write down the load scale value as the wheel assembly passes the elevator neutral position. Identify the value as F2. The elevator friction band must be calculated from the values given. (a) In a system where a force must be exerted in a direction toward the panel (forward) to return the elevators to a down position, the friction band is the sum of forces F1 and F2. The friction band = F1 + F2. (b) In a system where a restraining force must be exerted in a direction away from the panel (aft) to prevent the elevator from abruptly returning to a down position, the friction band is the difference of the two forces. The friction band = F1 - F2.
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MODEL 208 MAINTENANCE MANUAL (3)
When the friction band exceeds the limits, the following steps must be completed in the order shown to reduce system friction: (a) Reduce the tension on the elevator cable to 15 pounds with the elevator in the neutral position. (b) Do a check for binding and alignment of pulleys. Correct as necessary. (c) Make sure the hinges are clean and lubricated.
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MODEL 208 MAINTENANCE MANUAL ELEVATOR TRIM - DESCRIPTION AND OPERATION 1.
General A.
2.
Elevator trim is provided to ease required control force while airplane is operating at different flight attitudes.
Description and Operation A.
An elevator trim tab is located on trailing edges of the right and left elevator. The tabs are manually actuated by rotating a trim wheel, located on left side of control pedestal. The system consists of following components: trim control wheel attached to a sprocket, a roller chain attached to up and down cables. The up and down cables are routed under floorboard, through pulleys, and through tailcone of airplane to roller chains attached to sprockets on left and right trim tab actuators located in horizontal stabilizers. Two pushrods connect left and right actuators to trim tabs.
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MODEL 208 MAINTENANCE MANUAL ELEVATOR TRIM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Elevator Trim System Troubleshooting Chart Figure 101 (Sheet 1)
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Elevator Trim System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL ELEVATOR TRIM - MAINTENANCE PRACTICES 1.
General A.
2.
Elevator trim maintenance practices consist of elevator trim system removal/installation, elevator trim system rigging, elevator trim tab free play inspection, elevator trim tab actuator disassembly, elevator trim tab actuator inspection and repair, elevator trim tab actuator lubrication and assembly and elevator trim tab actuator inspection and rigging.
Elevator Trim System Removal/Installation A.
Remove Elevator Trim System (Refer to Figure 201). (1) Remove carpet or vinyl cover, plywood ßoor covers, ßoorboard and horizontal stabilizer access plates or covers, and upholstery panel at entrance to tailcone. (2) Remove connector links; detach chain from cables. (3) Remove nut and washer . Detach trim wheel from shaft. (4) Detach chain from sprockets and remove sprocket from shaft. (5) Remove screw from pedestal. Detach pointer, spacer, and bushing. (6) Remove nuts, washers, bolts, and bushings; detach sprockets from support. (7) Remove bolt and washer. Detach pulley from support. (8) Remove nuts, washers, and bolts. Detach pulleys from supports. (9) Airplanes 20800001 Thru 20800185 and 208B0001 Thru 208B0214 except airplanes incorporating SK208-76, remove nuts, washers, and bolts, detach fairleads from supports. (10) Remove nut, washer, and bolt. Detach pulleys from support. (11) Remove nuts, washers, and bolts. Detach stop blocks from cables. (12) Remove safety wire or clips from turnbuckles. (13) Detach turnbuckle from cables. (14) Detach turnbuckle from cables. (15) Remove bolts and washer. Detach pulleys from supports. (16) Remove connector links from chain. Detach chain from cables. NOTE:
To ease removal and installation of cables, attach a length of wire opposite removal end of cable. When cable is removed, leave wire in place, routed through structure. Pull replacement cable into correct location with wire.
(17) (18) (19) (20) (21)
Remove cables from system. Remove safety wire and bolts. Detach sprocket guards from actuators. Remove cotter pins, nuts, bushings, and bolts. Detach pushrods from actuators and horns. Remove chains from actuator sprockets. Remove safety wire and bolts. Detach actuators from supports, and remove through stabilizer access port. (22) Remove plug buttons and groove pins. Detach actuator sprockets from actuators. B.
Install Elevator Trim System (Refer to Figure 201). (1) Attach actuator sprockets to actuators. Replace groove pins and plug buttons. (2) Replace actuators in stabilizers. Attach actuators to supports. Replace bolts and safety wire. (3) Secure trim tabs in streamlined position. Attach pushrods to horn and actuator. Replace bushings, bolts, nuts, and cotter pins. Release trim tabs. NOTE: (4)
Torque nuts to 10 inch-pounds, then overtorque until Þrst cotter pin slots line up with holes in bolts.
Attach chains on sprockets with connector link ends of chains equidistant from centerline of outboard sprockets.
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Elevator Trim System Installation Figure 201 (Sheet 1)
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Elevator Trim System Installation Figure 201 (Sheet 2)
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Elevator Trim System Installation Figure 201 (Sheet 3)
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Elevator Trim System Installation Figure 201 (Sheet 4)
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Elevator Trim System Installation Figure 201 (Sheet 5)
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Elevator Trim System Installation Figure 201 (Sheet 6)
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Elevator Trim System Installation Figure 201 (Sheet 7)
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Elevator Trim System Installation Figure 201 (Sheet 8)
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WARNING: Pitch of sprockets must be synchronized with pitch of chain. Do not place unequal load on pushrods. Sprockets are provided with two sets of mounting holes and may have to be removed and replaced in a different set of holes to synchronize with pitch of chain. (5) (6) (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21) (22) (23) (24) 3.
Attach sprocket guards to actuators. Replace bolts. Safety wire bolts. Replace cables in system. Attach cable to chains. Replace connector links. Attach cables and to chains. Replace connector links. Attach pulleys to support. Replace washers and bolts. Attach bushing, spacer, and pointer to pedestal. Replace screw. Attach sprocket to shaft. Attach sprockets to support. Replace bushings, bolts, washers, and nuts. Attach chain to sprockets and with connector link ends of chain centered. Attach trim wheel to shaft. Replace washer and nut. Attach chain to cables. Replace connector links. Attach pulleys to support). Replace washer and bolt. Attach pulleys to supports. Replace bolts, washers, and nuts. Airplanes 20800001 Thru 20800185 and 208B0001 Thru 208B0214 except airplanes incorporating SK208-76, attach fairleads to supports. Replace bolts, washers, and nuts. Attach pulleys to support. Replace bolt, washer, and nut. Attach stop blocks to cables. Replace bolts, washers, and nuts. Refer to Elevator Trim System Rigging. Attach turnbuckle to cables. Attach turnbuckle to cables. With trim tabs streamlined, tighten turnbuckles evenly until cable tension checks 20 pounds, +5 or -5 pounds. Safety wire or install clips on turnbuckles. Replace tailcone upholstery panel, ßoorboard access covers and vinyl cover or speed tape access covers, and plywood ßoor covers.
Elevator Trim System Rigging A.
Rig Elevator Trim System (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13)
All control surface cable tensions should be rigged at an ambient temperature of 70°F. Allow temperature to stabilize for a period of four hours before setting cable tension.
Remove upholstery panel at entrance of tailcone. With elevator pinned at neutral, set the elevator trim tab at 0 degrees. Secure trim tabs in streamlined position (faired with the elevator), and attach an inclinometer on left tab. Set at zero degrees. Cut safety wire or remove clips and loosen turnbuckles. Check that chains are centered on sprockets. Check pointer. If it does not indicate neutral trim, loosen nut and washer and disengage trim wheel from shaft far enough to set pointer on neutral. Tighten nut on washer and trim wheel. Tighten turn buckles evenly and set cable tension at 20 pounds, +5 or -5 pounds. Safety wire or install clips on turnbuckles. Measure 28 inches aft from bulkhead and set stop block on cable. Replace bolt, washer, and nut. Release trim tabs and rotate trim wheel forward until inclinometer checks 15 degrees, +2 or -2 degrees. Attach stop block to cable in contact with stop block. Replace bolt, washer, and nut. Rotate trim wheel aft until inclinometer checks 15 degrees, +2 or -2 degrees. Attach stop block to cable in contact with stop block. Replace bolt, washer and nut.
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MODEL 208 MAINTENANCE MANUAL (14) Check pointer through 15 degrees in each direction of travel. If required, bend pointer slightly to clear pedestal cover. (15) Verify correct tab movement in response to trim wheel movement. (16) Remove inclinometer from elevator trim tab and replace upholstery panel at entrance to tailcone. 4.
Elevator Trim Tab Actuator Disassembly (Airplanes with 2660017-1 Trim Tab Actuator Installed) A.
Disassemble Elevator Trim Tab Actuator (Refer to Figure 202). (1) Cut safety wire and remove bolts; remove chain guard from actuator housing. Remove plug buttons from secondary sprockets. Detach repair link and remove chain from sprockets. (2) Remove groove pins and sprockets from internal screws. NOTE: (3)
Place index marks on bearings and actuator housing. Remove groove pins and from actuator housing and remove bearings and from actuator housing. Remove and discard O-rings from bearings. NOTE:
(4) (5) 5.
It may be necessary to apply heat to sprocket to loosen Loctite seal between sprockets and internal screws.
If bearings are to be reused, they must be replaced in the same location and relative position from which they were removed.
Tap ends of external screws on table top to remove bearings, races, bearings, and washers. Discard washers. Tap internal screws at ends “B" and remove internal screws from actuator housing and separate bearings, races, washers and external screws from internal screws. Discard washers.
Elevator Trim Tab Actuator Disassembly (Airplanes with 2661215-1 or 2661215-9 Trim Tab Actuator Installed) A.
Disassemble Trim Tab Actuator (Refer to Figure 202). (1) Remove the safety wire from the chain guard bolts. (2) Remove the bolts, chain guard and spacers from actuator housing. (3) Remove plug buttons from sprockets. (4) Detach repair link and remove chain from sprockets. (5) Remove groove pins from the sprockets. (6) Remove the sprockets from the internal screws. NOTE: (7) (8)
It may be necessary to apply heat to sprocket to loosen Loctite seal between sprockets and internal screws.
Remove screw and end plate. Remove the countersunk screws. NOTE:
The screws that are countersunk are installed with Loctite. These may be difÞcult to remove. If normal loosening techniques will not free the screws, applying some heat may help. Do not force the screws loose, the screw head may break. Instead, use a drill and easy-out to remove the screws. This will avoid damage to the end plate.
(9) Remove screw and slide end plate up external screws. (10) For actuator 2661215-1, tap the ends of the external screws on table top to remove wipers and bearings. NOTE:
If bearings are to be reused, keep the two halves together as a pair as they are removed. They must be replaced in the same location and relative position from which they were removed.
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Elevator Trim Tab Actuator Figure 202 (Sheet 1)
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Elevator Trim Tab Actuator Figure 202 (Sheet 2)
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Elevator Trim Tab Actuator Figure 202 (Sheet 3)
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Elevator Trim Tab Actuator Figure 202 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL
Elevator Trim Tab Actuator Figure 202 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL
Elevator Trim Tab Actuator Figure 202 (Sheet 6)
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MODEL 208 MAINTENANCE MANUAL
Elevator Trim Tab Actuator Figure 202 (Sheet 7)
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MODEL 208 MAINTENANCE MANUAL
Elevator Trim Tab Actuator Figure 202 (Sheet 8)
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MODEL 208 MAINTENANCE MANUAL
Elevator Trim Tab Actuator Figure 202 (Sheet 9)
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MODEL 208 MAINTENANCE MANUAL (11) For actuator 2661215-9, tap the ends of the internal screws on table top to remove wipers and bearings. NOTE:
The bearings are shank sealed on the actuator 2661215-9.
NOTE:
If bearings are to be reused, keep the two halves together as a pair as they are removed. They must be replaced in the same location and relative position from which they were removed.
(12) Unscrew the external screws from the internal screws. (a) For actuator 2661215-9, remove the O-ring from the external screw. (13) Tap internal screws at the sprocket end to remove the bearings. (a) Check condition of bearings; replace if required, utilizing an arbor press and mandrel. (14) Remove internal screws from actuator housing. (15) Clean actuator components and dry thoroughly. DO NOT clean bearing. DO NOT allow cleaned parts to contact lint or dirt. 6.
Elevator Trim Tab Actuator Inspection/Repair A.
Inspect/Repair Elevator Trim Tab Actuator (Refer to Figure 202). NOTE: (1) (2) (3) (4) (5)
7.
Remove actuator from system. Clean, inspect, and lubricate detail parts. Replace any components that show damage or excessive wear. Refer to Chapter 5 for Time Limits.
Clean detail parts with solvent in a well ventilated area away from sparks or open ßame. Avoid inhalation of solvent vapors. Dry parts with dry compressed air, lint free cloth, or lint free disposable tissue. Check parts visually, preferable under magniÞcation. If any parts show wear or damage, perform a dimensional check and replace parts, if necessary. If Þnish on actuator housing or chain guard has worn away or bare metal is exposed, apply Iridite 14-2, followed by two coats of epoxy primer. The Þnish shall consist of vivid orange or white lacquer.
Elevator Trim Tab Actuator Lubrication and Assembly (Airplanes with 2660017- 1 Trim Tab Actuator Installed) A.
Lubricate and Assemble Trim Tab Actuator (Refer to Figure 202). NOTE: (1) (2) (3) (4) (5)
Lubricate each detail part of actuator assembly before installation with 5565450-28 light consistency silicone grease, which may be purchased from Cessna Parts Distribution.
Install new O-rings in bearings. Install internal screws with ends "B" up and actuator housing in upright position with end "C", down. Install races, washers, bearings, washers, and bearings. Locate bearings on index marks and lightly tap or press them into actuator housing until groove pins can be installed through actuator and bearings. Place actuator housing in upright position with end “C” up. Install races, washers, bearings and bearings in actuator housing. Ensure bearings are located on index marks, and lightly tap or press them into actuator housing until groove pins can be installed through housing and bearings. NOTE:
(6) (7)
Steps through are applicable if existing bearings and are utilized. If new bearings are required, steps through are applicable.
Heat-soak new bearings in SAE 20-weight oil for 20 minutes at 140°F. Cool bearings to ambient temperature before installation. Install internal screws in actuator housing with ends "B" up and actuator housing in upright position with end "C" down.
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MODEL 208 MAINTENANCE MANUAL (8) (9) (10) (11) (12) (13) (14) (15) (16)
Install races, bearings and a 0.004 to 0.006 inch shim on ends "B" of internal screws. Replace bearings, press or tap lightly until bearings are ßush with end of actuator housing. Place actuator housing in upright position with bearings on bottom. Install races, bearings and bearings in actuator housing. Press or tap lightly until bearings are ßush with end of actuator housing. Place a clamp securely across assembled bearings to prevent any linear movement of internal screws. Drill 0.094 inch diameter holes (4 places) through existing 0.062 inch diameter holes in actuator housing and through bearings. Release clamp and remove bearings and 0.004 to 0.006 inch shim from actuator housing. Replace bearings and install new 0.094 inch groove pins (4 each) through actuator housing and bearings. Replace external screws in actuator housing at end “C”. NOTE:
After external screw threads contact internal screw threads, ensure they are not cross-threaded, and turn external screws approximately 25 turns clockwise after threads engage. Engagement should be smooth with no tight spots. If threads drag or tight spots are observed, disassemble actuator and replace internal and external screws.
(17) Apply No. 609 Loctite to mating surfaces of sprockets and internal screws. Install sprockets and replace groove pins and plug button. 8.
Elevator Trim Tab Actuator Lubrication and Assembly (Airplanes with 2661215-1 or 2661215-9 Trim Tab Actuator Installed) A.
Lubricate and Assemble Elevator Trim Tab Actuator (Refer to Figure 202). (1) Prior to assembly, apply a 1/8 inch coating of 5565450-28 light consistency silicone grease, which may be purchased from Cessna Parts Distribution, on the threads of internal screws, the outer surface of internal screws, and the outside diameter of external screws. (2) Install bearings into the actuator housing. (3) Put the internal screws into the actuator housing. (4) Install the end plate on the actuator housing with screw. NOTE: (5) (6)
Lubricate threads of the external screw. Heat the wipers to make more pliable. NOTE:
(7)
After installation of end plate, outer race of bearing must not move in actuator housing.
Be very careful not to damage the wipers when installing them over threads of the external screw.
Position end plate and wipers over external screws. (a) Inspect wipers to ensure threads do not damage wipers during installation. NOTE:
(8) (9)
Make sure the ßat side of wiper is seated into bearing recess.
For actuator 2661215-9, install the O-ring on the external screws. Screw the external screws into the internal screws. NOTE:
After external screw threads contact internal screw threads, ensure they are not cross-threaded, and turn external screws all the way in. Engagement should be smooth with no tight spots. If threads drag or tight spots are observed, disassemble actuator and replace internal screws and external screws.
(10) For actuator 2661215-1, position the bearing halves around the external screws and press the bearings into the actuator housing.
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MODEL 208 MAINTENANCE MANUAL (11) For actuator 2661215-9, position the bearing halves around the external screws. NOTE:
Make sure that the O-ring is in the groove of both bearing halves.
(a)
(12) (13) (14) (15) (16) (17)
Shank seal the bearing halves with Type X Class B sealant and press them into the actuator housing. Push the wipers into position in the bearing halves. Install the screws that attach the end plate on to actuator housing. (a) Safety wire screw . Apply No. 609 Loctite to mating surfaces of sprockets and internal screws. Install the sprockets on to the internal screws . (a) Install the groove pins. Apply common RTV sealant to plug buttons and sprockets. (a) Install the plug buttons in sprockets. Work the screws all the way in and out 2 to 3 times, wipe excess grease from both ends after each cycle. NOTE:
There must be no end play between bearing inner race, internal screw and sprocket when groove pins are installed.
NOTE:
After assembly, maximum longitudinal movement of external screws and actuator housing is not to exceed 0.007 inch.
(18) Install chain guard and spacers on to the actuator housing using bolts. (a) Safety wire the bolts. 9.
Elevator Trim Tab Actuator Inspection and Rigging (Airplanes with 2660017-1 Trim Tab Actuator Installed) A.
Inspect and Rig Elevator Trim Tab Actuator (Refer to Figure 202). (1) After assembling detail parts, rotate sprockets clockwise, then counterclockwise far enough to obtain approximately 0.75 inch linear movement of external screws in each direction. Movement should be smooth in each direction with no perceived torque change in either direction. (2) Bearings in external screws must be aligned within 0.010 inch before installing actuator in system. NOTE:
(3) (4) (5) (6)
A surface plate or table, four threaded rods or bolts (2 10-24 NC 3A thread and 2 1/4-20 UNC 2B thread), V-blocks, angle block, clamps and height gage and dial indicator or equivalent precision measuring equipment are required to perform preceding check.
Attach bolts or threaded rods to both sides of actuator housing at points “D”. Bolts or rods should be tightened. Mount unit in V-blocks in vertical position. Rotate either external screw in required direction to allow installation of No. 11 drill rod (0.191 inch diameter) through both bearings. Check dimension from top of bolts or rods at points “D” to top of No. 11 drill rod outside of each bearing.
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MODEL 208 MAINTENANCE MANUAL (7)
(8) (9) 10.
Remove No. 11 drill rod and rotate either screw in required direction. Replace No. 11 drill rod through bearings and check alignment. Continue rotating screw(s) as required to align bearings within 0.010 inch. NOTE:
If bearings cannot be aligned to 0.010 inch with chain removed, rotate either sprocket one or two teeth in desired direction. Sprockets have two sets of mounting holes located 75 degrees apart. It may be necessary to move sprocket from one set to the other.
NOTE:
If it is determined elevator trim tab excessive free play is caused by actuator, internal screws and external screws must be replaced along with any detail part worn beyond dimensional tolerance. However, if special optical inspection equipment is available and it is veriÞed threads on internal screws and external screws are not worn beyond dimensional tolerance, screws may be reinstalled in assembly.
Install chain on sprockets and replace connector link. Install chain guard on actuator housing and replace screws. Safety wire screws.
Elevator Trim Tab Actuator Inspection and Rigging (Airplanes with 2661215-1 or 2661215-9 Trim Tab Actuator Installed) A.
Rig and Inspect Elevator Trim Tab Actuator (Refer to Figure 202). (1) After assembling detail parts, rotate sprockets clockwise, then counterclockwise far enough to obtain approximately 0.75 inch linear movement of external screws in each direction. Movement should be smooth in each direction with no perceived torque change in either direction. NOTE: (2)
Bearings in external screws must be aligned within 0.010 inch before installing actuator in system. NOTE:
(3) (4) (5) (6) (7)
Starting torque of primary sprocket shall not exceed 3 inch-pounds at ambient temperature of 65°F.
A surface plate or table, four threaded rods or bolts (2 10-24 NC 3A thread and 2 1/4-20 UNC 2B thread), V-blocks, angle block, clamps and height gage and dial indicator or equivalent precision measuring equipment are required to perform preceding check.
Attach bolts or threaded rods to both sides of actuator housing at points “D”. Bolts or rods should be tightened. Mount unit in V-blocks in vertical position. Rotate either external screw in required direction to allow installation of No. 11 drill rod (0.191 inch) through both bearings. Check dimension from top of bolts or rods at points "D" to top of No. 11 drill rod outside of each bearing . Remove No. 11 drill rod and rotate either screw in required direction. Replace No. 11 drill rod through bearings and check alignment. Continue rotating screw(s) as required to align bearings within 0.010 inch. NOTE:
If bearings cannot be aligned to 0.010 inch with chain removed, rotate either sprocket one or two teeth in desired direction. Sprockets have two sets of mounting holes located 75 degrees apart. It may be necessary to move sprocket from one set to the other.
NOTE:
If it is determined elevator trim tab excessive free play is caused by actuator, internal screws and external screws must be replaced along with any detail part worn beyond dimensional tolerance. However, if special optical inspection equipment is available and it is veriÞed threads on internal screws and external screws are not worn beyond dimensional tolerance, screws may be reinstalled in assembly.
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MODEL 208 MAINTENANCE MANUAL (8) (9)
Install chain on sprockets and replace connector link. Install chain guard and spacers on actuator housing and replace bolts. Safety wire bolts.
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MODEL 208 MAINTENANCE MANUAL ELEVATOR TRIM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the elevator trim system in a serviceable condition.
Task 27-30-02-720 2.
Elevator Trim Tab (Free Play) Functional Check A.
General (1) This task gives the procedures to do a elevator trim tab (free play) functional check.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Elevator Trim Tab (Free Play) Functional Check (Refer to Figure 601). (1) Put the elevator and trim tab in the neutral position and secure from movement. (2) Determine the maximum allowable free play, measuring chord length at the extreme inboard end of the trim tab then multiply the chord length by 0.025 to get the maximum allowable free play. (3) Use Þngertip pressure and move the trim tab trailing edge up and down to examine free play. NOTE: (4) (5)
Measure free play at the same point on the trim tab that the chord length was measured. Total free play must not exceed the maximum allowable.
If the trim tab free play is less than the maximum allowable, no additional inspection is required. If the trim tab free play is more than the maximum allowable, the following items must be examined: (a) Look for loose fasteners on the trim tab doubler. (b) Examine the hinge, hinge pin, and fasteners on the trim tab doubler. (c) Examine both ends of the push-pull rods and fasteners for wear and loose component parts. (d) If corrosion, worn parts, or loose fasteners are found, replace the fasteners and install new parts in system. (e) Do a second free play inspection. If the free play is still excessive, remove the elevator trim tab actuator from the 1 airplane and set it on a bench. Refer to Elevator Trim - Maintenance Practices. Disassemble the actuator and examine the detail parts for corrosion and excessive 2 wear. Refer to Elevator Trim - Maintenance Practices. If corrosion or worn parts are found, replace the parts and reassemble the actuator. 3 (f) Install the actuator in the airplane . Refer to Elevator Trim - Maintenance Practices. (g) Do the free play inspection again.
E.
Restore Access (1) None End of task Task 27-30-02-640 3.
Elevator Trim Tab Actuator (2660017-1) Lubrication A.
General (1) This task gives the procedures to do the elevator trim tab actuator (2660017-1) lubrication.
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MODEL 208 MAINTENANCE MANUAL
Elevator Trim Tab (Free Play) Functional Check Figure 601 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
Special Tools (1) Grease
C.
Access (1) None
D.
Do the Elevator Trim Tab Actuator (2660017-1) Lubrication (Refer to Figure 202 found in Elevator Trim - Maintenance Practices). (1) Remove the elevator trim tab actuator from the airplane and put it on a bench. Refer to Elevator Trim - Maintenance Practices. (2) Disassemble the elevator trim tab actuator. Refer to Elevator Trim - Maintenance Practices. (3) Do the Elevator Trim Tab Actuator Inspection/Repair. Refer to Elevator Trim - Maintenance Practices. (4) Do the lubrication and the assembly steps found in Elevator Trim Tab Actuator Lubrication and Assembly (Airplanes with 2660017- 1 Trim Tab Actuator Installed). Refer to Elevator Trim Maintenance Practices. (5) Install the elevator trim tab actuator in the airplane. Refer to Elevator Trim - Maintenance Practices.
E.
Restore Access (1) None End of task Task 27-30-02-641 4.
Elevator Trim Tab Actuator (2661215-1 and 2661215-9) Lubrication A.
General (1) This task gives the procedures to do the elevator trim tab actuator (2661215-1 and 2661215-9) lubrication.
B.
Special Tools (1) Grease
C.
Access (1) None
D.
Do the Elevator Trim Tab Actuator (2661215-1 and 2661215-9) Lubrication (Refer to Figure 202 found in Elevator Trim - Maintenance Practices). (1) Remove the elevator trim tab actuator from the airplane and put it on a bench. Refer to Elevator Trim - Maintenance Practices. (2) Disassemble the elevator trim tab actuator. Refer to Elevator Trim - Maintenance Practices. (3) Do the Elevator Trim Tab Actuator Inspection/Repair. Refer to Elevator Trim - Maintenance Practices. (4) Do the lubrication and the assembly steps found in Elevator Trim Tab Actuator Lubrication and Assembly (Airplanes with 2661215-1 or 2661215-9 Trim Tab Actuator Installed). Refer to Elevator Trim - Maintenance Practices. (5) Install the elevator trim tab actuator in the airplane. Refer to Elevator Trim - Maintenance Practices.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL ELECTRIC ELEVATOR TRIM - DESCRIPTION AND OPERATION 1.
General A.
2.
Elevator trim on the Model 208 may be accomplished by a remote mounted electric motor that drives the elevator trim tab actuator.
Description and Operation A.
Electric elevator trim system consists of two trim switches, and a disconnect switch located on pilot's control wheel, circuit breaker, electric trim actuator, and associated wiring. The electric trim actuator and clutch are located on a support on the control pedestal. A chain attaches the motor and clutch to elevator trim wheel. The electric trim clutch is disengaged from trim motor at all times except when motor is operating and, except when trim motor is operating, has no effect on manual trim operation. However, the electric elevator trim system is inoperative, whenever the autopilot system is turned on. If an unwanted power signal activates the trim disconnect switch, the clutch will disengage. In operation, one of the two trim switches provides power to run the trim motor up or down. The other trim switch provides power to the trim clutch. The clutch is grounded through the normally closed trim disconnect relay. If a fault occurs to cause power to keep the clutch on, the trim disconnect switch momentarily connects the clutch relay to ground. Then the clutch relay latches on until power is removed from the clutch.
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MODEL 208 MAINTENANCE MANUAL ELECTRIC ELEVATOR TRIM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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MODEL 208 MAINTENANCE MANUAL
Electric Elevator Trim System Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ELECTRIC ELEVATOR TRIM - MAINTENANCE PRACTICES 1.
Electric Elevator Trim System Removal/Installation A.
Remove Elevator Electric Trim System (Refer to Figure 201 ). (1) Remove the screws and the washers and detach the right cover plate from the control pedestal. (2) Remove the nut and the washer. Detach the elevator trim wheel from the shaft. (3) If installed, remove the chain cover by removing the screws. (4) Remove the screws and detach the left cover plate from the control pedestal. (5) Loosen the bolts and move the support up to relieve the tension on the chain. (6) Remove the connecting link from the chain and detach the chain from the sprockets. (7) Disconnect the plug from the connector. (8) Remove the bolts and the washers. Detach the electric trim actuator and the actuator mounting bracket from the support. (9) Remove the bolts and the washers. Detach the support from the control pedestal. (10) Attach the right cover plate to the control pedestal and replace the screws and the washers. (11) Attach the left cover plate to the control pedestal and replace the screws. (12) Attach the elevator trim wheel to the shaft and replace the washer and the nut.
B.
Install Elevator Electric Trim System (Refer to Figure 201 ). (1) Remove the screws and the washers and disconnect the right cover plate from the control pedestal. (2) Remove the nut and the washer and disconnect the elevator trim wheel from the shaft. (3) Attach the electric trim actuator to the actuator mounting bracket and install the bolts. (a) Tighten the bolts per torque sequence diagram and torque to 15-20 inch- pounds. (4) Attach the support to the control pedestal and install the washers. Install the bolts and handtighten. (5) Attach the plug to the connector. (6) Attach the chain to the upper and the lower sprocket and install the connector link on the chain. (7) Adjust the support to set the required tension of the chain. Tighten the bolts. (8) Attach the left cover plate to the control pedestal and replace the screws. (9) If the chain cover was removed, install it with the screws. (10) Attach the elevator trim wheel to the shaft and replace the washer and the nut. (11) Attach the cover to the control pedestal and replace the screws and the washers.
C.
Electric Trim Clutch Torque Check (Refer to Figure 201 ). (1) Remove the screws and the washers and detach the right cover plate from the control pedestal. (2) Remove the nut and the washer. Detach the elevator trim wheel from the shaft. (3) If installed, remove the chain cover by removing the screws. (4) Remove the screws and detach the left cover plate from the control pedestal. (5) Loosen the bolts and move the support up to relieve the tension on the chain. (6) Remove the connecting link from the chain and detach the chain from the sprockets. (7) Disconnect the plug from the connector. (8) Remove the bolts and the washers. Detach the electric trim actuator and the actuator mounting bracket from the support, and move the actuator and the bracket from the airplane and put it on a bench. (9) Remove the bolts and detach the electric trim actuator from the actuator mounting bracket. (10) Remove the front seal retainer and detach the front seal from the sprocket housing. (11) Lock the shaft to prevent rotation and install a 1 5/16-inch socket over the sprocket housing. Attach a torque wrench to the socket. (12) Turn the sprocket housing slowly clockwise or counterclockwise and examine the torque. (13) The torque must be 30 inch-pounds, +5 or -5 inch-pounds. (14) Adjust the electric trim clutch torque as follows: (a) Loosen the locking screw.
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MODEL 208 MAINTENANCE MANUAL
Electric Elevator Trim System Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Electric Elevator Trim System Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Electric Elevator Trim System Installation Figure 201 (Sheet 3)
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Electric Elevator Trim System Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (b)
Turn the adjusting nut clockwise to increase the torque or counterclockwise to decrease the torque. NOTE:
When readjusting the clutch torque, turn the adjusting nut clockwise until the torque reaches two to four inch-pounds under the required setting. Then tighten the locking screw to 40-45 inch-pounds. If a closer setting is required, the adjusting nut may be turned without loosening the screw.
(15) Attach the front seal to the sprocket housing and replace the front seal retainer. (16) Attach the electric trim actuator to the actuator mounting bracket and replace the bolts. Follow the torque sequence diagram and torque the bolts to 15-20 inch-pounds. (17) Attach the electric trim actuator and the actuator mounting bracket to the support. Replace the washers and the bolts. (18) Connect the plug to the connector. (19) Attach the chain to the sprockets and replace the connector link. (20) Adjust the support to set the necessary chain tension and tighten the bolts. (21) Attach the left cover plate to the control pedestal and replace the screws. (22) If the chain cover was removed, install it with the screws. (23) Attach the elevator trim wheel to the shaft and replace the washer and the nut. (24) Attach the right cover plate to the control pedestal and replace the screws and the washers. 2.
Electric Elevator Trim System Operational Check A.
Check Individual Segments of Electric Trim. (1) Push forward to DN position momentarily, release to center position. Pull aft to UP position momentarily, release to center position. OBSERVE N0 MOVEMENT of elevator trim wheel as individual trim switch segments are cycled.
B.
Check Both Segments of Electric Trim. (1) Push forward and HOLD. During nose down cycle, DEPRESS and RELEASE A/P Trim Disconnect pushbutton. OBSERVE movement of elevator trim wheel in proper direction before A/P Trim Disconnect pushbutton is depressed and released. OBSERVE NO MOVEMENT after A/P Trim pushbutton is depressed and released. (2) Repeat step (2), except pull aft and HOLD. NOTE: (3) (4)
To reactivate system, release both segments to CENTER OFF position.
Operate system through full range of travel and check for binding, jerky movements and sluggish operation. Check operating time for full range of motion. (a) Airplanes equiped with King KFC-150 or -250 autopilot shall demonstrate full range of motion within 26 to 38 seconds. (b) Airplanes equiped with King KFC-225 autopilot shall demonstrate full range of motion within 16 to 24 seconds.
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MODEL 208 MAINTENANCE MANUAL ELECTRIC ELEVATOR TRIM - ADJUSTMENT/TEST 1.
Electric Trim Rigging (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) A.
Adjust the Electric Trim (Refer to Figure 501). (1) Set the control wheels in the neutral position of elevator. (2) Use a control column neutral rigging tool attach the control wheels in the neutral position. (3) Make sure the actuator motor is attached to the actuator mount with four bolts, washers and safety wire. (4) Make sure the actuator mount is attached to the support with four bolts and washers. (5) Make sure the support is attached to the brackets which attach to the pedestal using two each bolts and washers. (6) Make sure that chain is correctly aligned on the actuator motor sprocket and sprocket. (7) Make sure that chain is installed with a connecting link. (8) Make sure the two chain guard posts are correctly installed and have safety wire attached to the actuator assembly. (9) Make sure the electrical connector is attached. (10) Adjust the tension of the chain assembly. (a) Apply five pounds of scale tension to the chain the same distances from each of the sprockets. (b) While you apply the five pounds of pressure, examine the chain to have a total horizontal travel of 0.40 inch. (c) If the chain tension is too high or too low, loosen the bolts and adjust he two support brackets to allow 0.40 inch chain deflection. (d) Tighten the bolts. (e) Remove the scale from chain. (11) If the chain has to be removed for replacement or maintenance, you must remove the two chain guard posts. When the chain guard posts are removed, do not lose the lock washers. (12) To correctly adjust the electric trim system, you must make sure that you have a continuous 28.8 volts DC applied to the electronics side of the airplane's bus bar. This can be accomplished in one of the following methods. (a) Use the standard airplane starting procedures for the engine and operate the engine at 52 degrees Ng to maintain the normal operating aircraft voltage (28.8 VDC). (b) With the battery switch and avionics power switches set in the OFF position, connect a well regulated and filtered external power supply directly to the battery side of the battery contactor. Adjust the power supply for 28.8 volts DC and then turn to ON the battery switch and avionics power switches to supply power to the system. (13) To remove the gyro roll and pitch signals generated by a non-erected gyro, the 400B autopilot has a GYRO switch located on the rear of the control head. Set the GYRO switch to the OUT position. (14) If an outside vacuum source is used, it must be calibrated in inches of mercury and the suction range required to erect the gyro is 4.6 to 5.4 inches of mercury. (15) If the airplane engine is operated to erect the gyro, the engine must operate at 65 degrees Ng to provide the amount of vacuum and maintain correct bus voltage. (16) The 400B lFCS does not have a gyro out switch. You must hook up an outside vacuum source or operate the airplane engine to erect the gyros. (17) Remove the neutral rigging tool. (18) Set a piece of tape or a mark on the top of the airplane's ELEVATOR TRIM command wheel so that a full revolution of the ELEVATOR TRIM command wheel will be observed and timed with a stop watch. (19) With 28.8 volts DC applied to the electronics bus, place AP/ON autopilot (ON/OFF) switch to the ON position. Allow the autopilot to sync out. (Autopilot PITCH wheel stops running.) (20) Monitor and time one complete rotation of the airplane's ELEVATOR TRIM command wheel by placing pitch command wheel in the UP position. Make sure that you get a time of 30 +3 or -3 seconds for one complete rotation of the airplane's ELEVATOR TRIM command wheel. (a) Remove the cover assembly and sidewall to get access to the computer amplifier.
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MODEL 208 MAINTENANCE MANUAL
Electric Trim Rigging Figure 501 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Electric Trim Rigging Figure 501 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Electric Trim Rigging Figure 501 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (b) (c) (d)
Remove the center plug button from the top of the computer amplifier to get access to the R46 on computer module A6. Engage the autopilot and apply a NOSE-DOWN pitch command via the PITCH wheel, sufficient to cause trim to run. Adjust the R46 (counter clockwise - slows trim) on computer module A6 to get a rotation of the airplane's ELEVATOR TRIM wheel. One rotation in 30 +3 or -3 seconds in the NOSEDOWN direction and one rotation in 30 +3 or -3 seconds in the NOSE-UP position must be completed. NOTE:
With autopilot disengaged, the electric trim actuator operates on a bus voltage of 28.8 volts DC in both the NOSE-UP and NOSE-DOWN directions. The average time for ELEVATOR TRIM command wheel to make three complete rotations is 23.5 +2 or -2 seconds.
(e) Do a check of the system friction and voltage if time limit is exceeded. (21) Reverse the procedure and make sure that you are getting a reading of 30 +3 or -3 seconds for one full rotation of the airplane's ELEVATOR TRIM command wheel in the NOSE-DOWN position. If the rate of the travel for one full rotation does not agree with the prior travel time limits, then use the following procedures to get the desired rate of pitch trim tab travel. (a) Turn AP/ON autopilot (ON-OFF) switch, avionics power switches, and airplane battery switch to the OFF positions. (b) Replace the access plates removed. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (c) Remove the external power source if installed. (d) Remove the outside vacuum source if installed or reset the GYRO switch to the IN position. 2.
Electric Elevator Trim Clutch Torque System Check (Airplanes with 400B and 400B IFCS Autopilots Types AF-550A and IF-550A Installed) A.
Do a Check of the Trim Clutch Torque. (1) Remove the left and right cover plates from the pedestal. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Remove the nut and washer. Remove the elevator trim wheel from the shaft. (3) Loosen the bolts and move the supports up to relieve the tension on the chain. (4) Remove the connecting link from the chain and detach the chain from the sprockets. (5) Disconnect the plug from the connector. (6) Remove the bolts and washers. Remove the electric trim actuator and actuator mount bracket from the support. Move the actuator and bracket from the airplane to a table. (7) Remove the bolts and remove the electric trim actuator from the actuator mount bracket. (8) Remove the front seal retainer and the front seal from the sprocket housing. (9) Lock the shaft to prevent rotation and install 1 5/16 inch socket over sprocket housing. Attach a torque wrench to socket. (10) Rotate the sprocket housing slowly clockwise or counter clockwise and check the torque value. (11) The torque value must check 30 +5 or -5 inch-pounds. (12) Adjust the electric trim clutch torque as follows: (a) Loosen the lock screw. (b) Turn the adjusting nut clockwise to increase the torque or counter-clockwise to reduce the torque. NOTE:
If a closer setting is required, the adjusting nut may be turned without loosening the locking screw.
When you adjust the clutch torque, rotate the adjusting nut clockwise until the torque shows 2.0 to 4.0 inch-pounds by the required setting. Tighten the lock screw 40-45 inch-pounds. 2 (13) Install the front seal to the sprocket housing. (14) Install the front seal retainer. 1
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MODEL 208 MAINTENANCE MANUAL (15) Attach the electric trim actuator to the actuator mount bracket and install the bolts. Follow the torque sequence diagram and torque the bolts to 15-20 inch-pounds. (16) Attach the electric trim actuator and the actuator mount bracket to the support. Install the washers and bolts. (17) Connect the plug to the connector. (18) Attach the chain to the sprockets and install the connector link. (19) Adjust the support to set the required tension of the chain and tighten the bolt. (20) Attach the elevator trim wheel to the shaft and install the washer and nut. (21) Attach the left and right cover plate to the control pedestal and install the screws. 3.
Electric Trim Rigging (Airplanes with KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Installed) A.
Adjust the Electric Trim Rigging (Refer to Figure 502). (1) Set the control wheels in the neutral position of the elevator. (2) Lock the control wheels in the neutral position by using a control column neutral rigging tool. (3) Adjust the pitch trim servo slip clutch values of the KM 277 Capstan. Refer to Slip Clutch Torque Adjustment. (4) The KS 272A Trim Servo and KM 277 Servo Mount are installed with the electric pitch trim control in the KAP-150 Autopilot System and KFC-150, KFC-225 Flight Control System. The KS 272A and KM 277 units are installed inside the engine control pedestal just aft of FS 114.4. (5) Access the pitch trim servo installation and remove the five bolts that attach the chain guard to the pedestal. (6) Make sure the trim servo bracket is correctly installed inside the engine control pedestal. Inspect to make sure the trim servo bracket is attached at the forward end using two bolts and two washers. The two bolts fit the holes that exist in the engine control pedestal. (7) Do a check to make sure the aft end of the trim servo bracket is correctly attached using two bolts and two washers. The two bolts fit the holes (existing slots) in the engine control pedestal. Make sure the trim servo bracket is adjusted to the top of the slots if adjustment is necessary. (8) Make sure the two nut plates are correctly attached to the back of the KS 272A servo with four screws. (9) Do a check to make sure the aft trim bracket is correctly attached to the trim servo bracket with one bolt, two washers and one nut. (10) Do a check to make sure the forward trim bracket is correctly attached to the trim servo bracket with two bolts, four washers and two nuts. (11) Make sure the KM 277 servo mount is correctly attached to the aft trim bracket with two bolts, two washers and two nuts. (12) Make sure the KS 272A Pitch trim servo is correctly installed with two bolts. (13) Do a check to make sure the pitch trim chain is routed under the KM 277 Capstan and over the pilot's trim wheel shaft sprocket. Make sure the two ends of the trim chain are properly connected using the connecting link. (14) Adjust the tension of trim chain. (a) Apply five pounds of scale tension to the trim chain as close to equal distance from the trim wheel sprocket and KM 277 Capstan. (b) While applying five pounds of pressure, monitor the trim chain has a total horizontal travel of 0.25 inch maximum. (c) If chain tension is too high or too low, loosen the two bolts and adjust the trim servo bracket, up or down as required to allow 0.25 inch chain deflection. (d) Tighten the two bolts that attach the trim servo bracket. (e) Remove the scale from trim chain. (15) Remove the neutral rigging tool. (16) Make sure the servo's electrical connector is attached. (17) Install the chain guard with the five bolts.
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Electric Trim Rigging Figure 502 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL STALL WARNING SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101 and Figure 102.
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Stall Warning System Troubleshooting Chart Figure 101 (Sheet 1)
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Stall Warning Horn Disconnect Switch Troubleshooting Chart Figure 102 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL STALL WARNING SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
The following data contains instructions for the removal and installation of the detector, thermostat, horn, and the horn disconnect switch. This section also includes the stall warning operational check, detector adjustment and the stall warning heat check.
Stall Warning Detector Removal/Installation A.
Remove the Stall Warning Detector (Refer to Figure 201).
CAUTION: Do a careful check of the detector before you remove it from airplane. It is possible for the detector to be too hot to handle without hand protection. NOTE:
(1) (2) (3) (4) (5) (6) (7)
B.
3.
(Airplanes 20800316 and On and 208B0800 and On and Airplanes 20800001 thru 20800315 and 208B0001 thru 208B0799 incorporating CAB00-1) To preclude or disable nuisance stall warnings during ground operation, push the control yoke forward to the stop position. This will engage the disconnect switch for the ground stall warning horn.
Make sure the airplane master switches are set in the OFF position. Remove the screws and access plate from lower surface of wing. Remove screws attaching detector to airplane. Mark exact fore and aft location of vane on wing to ensure correct installation. Identify and disconnect electrical wires from detector. Remove detector from wing. Check the airplane's logbook for the detector tip gram setting. NOTE:
Airplane's with a TKS system installed from the factory have the tip gram setting written in the airplane's logbook.
NOTE:
The service facility must use tool T2680003-10 for detector installations on aircraft with and without anti-ice systems.
Install Stall Warning Detector (Refer to Figure 201). (1) Position detector up to wing. (2) Connect electrical wires to detector. (3) Insert detector into leading edge. (4) Secure detector to airplane using screws. (5) Attach access plate to lower wing skin using screws. (6) Perform an operational check of stall warning system. Refer to Stall Warning Operational Check.
Stall Warning Lift Transducer Removal/Installation For TKS Equipped Airplane A.
Remove the Stall Warning Lift Transducer (Refer to Figure 202).
CAUTION: Do a careful check of the lift transducer before you remove it from airplane. It is possible for the lift transducer to be too hot to handle without hand protection. (1) (2) (3) (4)
Make sure the airplane master switches are set in the OFF position. Remove and keep the mounting screws and access plate 503AB from the lower surface of the wing adjacent to the lift transducer. Mark the exact fore and aft location of vane on the wing for correct installation. Remove the sealant between the perimeter of the lift transducer and the wing porous panel.
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Stall Warning System Installation Figure 201 (Sheet 1)
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Stall Warning System Installation Figure 201 (Sheet 2)
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Stall Warning System Installation Figure 201 (Sheet 3)
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Stall Warning System Installation Figure 202 (Sheet 1)
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Stall Warning System Installation Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (5)
(6) (7) (8)
B.
4.
5.
Remove and keep the screws and washers that attach the lift transducer to the wing. Discard the nuts. (a) Make a note of the location of the shim washers that are between the lift transducer and the wing leading edge. Identify and disconnect the electrical wires from lift transducer. Remove the lift transducer from wing. Check the airplane's logbook for the lift transducer tip gram setting. NOTE:
Airplanes with a TKS system installed from the factory have the tip gram setting written in the airplane logbook.
NOTE:
When you order a new lift transducer from Cessna you must supply the tip gram setting that is in the airplane's logbook. If no tip gram setting is in the airplane's logbook, you must use a service facility to install the lift transducer.
Install the Stall Warning Lift Transducer (Refer to Figure 202). (1) Discard the gasket that comes with a new lift transducer. (2) Connect the electrical wires to the lift transducer. (3) Install the NAS1149FN816P Washer(s) and/or NAS1149FN832P Washer(s) to make the forward edge of the lift transducer flush or inset up to 0.010 inch with the outer surface of the porous panel. (a) Make sure to install these washers in their original location and add or delete washers until the mounting plate is flush with the porous panel. (4) Put the lift transducer in the left wing porous panel cutout. (5) Make sure that the lift transducer vane is at the mark that you made on the wing during the removal procedure. (6) Put the upper edge of the mounting plate 0.00 to 0.13 inches from the porous panel and each side of the mounting plate approximately a 0.04 inch distance from the porous panel. (a) Make sure the space between the edges of the mounting plate and the porous panel is even and parallel. (7) Install the lift transducer to the wing with the kept screws, washers, and new MS21044N08 nuts. (8) Apply a fillet seal around the mounting plate of the lift transducer with U060031 Sealant. (9) Attach the access plate to the lower wing skin with the kept screws. (10) Do an operational check of the stall warning system. Refer to Stall Warning Operational Check.
Stall Warning Thermostat Removal/Installation A.
Remove Stall Warning Thermostat (Refer to Figure 201). (1) Ensure airplane electrical power is OFF. (2) Remove screws and access plate from lower surface of wing. (3) Remove screws attaching thermostat to mounting bracket. (4) Identify, disconnect electrical wires and remove thermostat.
B.
Install Stall Warning Thermostat (Refer to Figure 201). (1) Position thermostat up to wing and connect electrical wires. (2) Position thermostat on mounting bracket and install screws. (3) Install access plate to lower surface of wing using screws.
Stall Warning Horn Removal/Installation A.
Remove Stall Warning Horn (Refer to Figure 201). (1) Remove screws, cover, and trim panel from headliner above pilot. (2) Remove screw securing mounting plate. (3) Pull aft end of mounting plate down. (4) Identify and disconnect horn electrical wires. (5) Remove nuts, washers, and stall warning horn.
B.
Install Stall Warning Horn (Refer to Figure 201). (1) Position stall warning horn to mounting plate. (2) Install washers and nuts.
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6) 6.
7.
Connect electrical wires. Push mounting plate up and secure with screw. Position cover and trim panel to mounting plate. Install screws to secure cover and trim panel.
Stall Warning Horn Disconnect Switch Removal/Installation (Airplanes 20800316 and On and 208B0800 and On and Airplanes 20800001 thru 20800315 and 208B0001 thru 208B0799 incorporating CAB00-1) A.
Remove Stall Warning Horn Disconnect Switch (Refer to Figure 201). (1) Remove electrical power from airplane. (2) Remove cabin floorboard panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Identify and disconnect electrical wires from the switch assembly. (4) Loosen screws and slide mounting bracket fully forward. (5) Loosen screws securing switch assembly to mounting bracket. (6) Remove switch assembly from airplane.
B.
Install Stall Warning Horn Disconnect Switch (Refer to Figure 201). (1) Position switch assembly on mounting bracket. (2) Install screws to loosely secure switch assembly to mounting bracket. (3) Slide mounting bracket fully forward and tighten aft screw. (4) Rotate switch assembly up until cam lobe makes contact with lower surface of bell crank. (5) Rotate switch assembly up 0.062 inch past actuation point and tighten screws to secure to mounting bracket. (6) Loosen aft screw and slide mounting bracket fully aft. (7) With control column in full forward position, slide mounting bracket forward until switch assembly actuates. (8) Slide mounting bracket 0.13 inch forward, past actuation point, and secure by tightening screws. (9) Connect electrical wires to the switch assembly. (10) Install cabin floorboard panel 232AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (11) Restore electrical power to airplane.
Stall Warning Flight Operational Check A.
Perform a Flight Operational Check (Refer to Figure 201). (1) Put the airplane in straight flight and trim it. (2) Set the engine power to flight idle (approximately 65% Ng) while slowing to approximately 1.5 times the stall speed. NOTE: (3) (4)
(5)
The stall warning system will be checked at idle power.
Reduce the airspeed with the elevator control until it is approximately 1.1 times the stall speed. Move the elevator control to slowly increase pitch attitude so that the airspeed decreases at a rate of no more than 1 knot per second. Do this until there is an unstoppable pitch down of the airplane, or the elevator control reaches the stop. NOTE:
An approach rate of 1 knot per second is a much slower entry rate then is used on a normal training stall.
NOTE:
Up to the time the airplane pitches, it must be possible to produce and correct both roll and yaw by normal use of the controls. During recovery, with normal use of the controls, it must be possible to prevent: more than 15 degrees roll, more than 15 degrees yaw, and more than 30 degrees pitch below level flight.
Record airplane speed that stall warning signal sounds.
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Record the airplane speed that the stall occurred. NOTE:
(7) (8) 8.
9.
The stall is the indicated airspeed at which the airplane pitches down with the nose up elevator being held by the pilot (usually around 10-12 degrees nose up), or it is the minimum speed observed by the pilot with the elevator on the nose up stop if no pitch down has occurred. The maximum time on the aft stop for elevator-limited stalls must be no more than 2 seconds.
Make sure that you hear the stall warning signal at 6 to 14 KIAS with the flaps fully up before the stall. Make sure that you hear the stall warning signal at 8 to 18 KIAS with the flaps fully extended before the stall.
Stall Warning Detector Adjustment (Airplanes 20800001 thru 20800056) A.
Adjust the stall warning vane to give a stall warning signal at 6 to 14 KIAS with the flaps fully up before the stall. Refer to Flight Operational Check.
B.
Adjust the stall warning vane to give a stall warning signal at 8 to 18 KIAS with the flaps fully extended before the stall. Refer to Flight Operational Check.
C.
Adjust the stall warning vane to a higher position to set the stall warning horn to a faster airspeed, or adjust the stall warning vane to a lower vane position to set the stall warning horn to a slower airspeed.
KAP-150 Autopilot and KFC-150, KFC-225 Flight Control System Audio Alert Removal/Installation A.
Remove the Autopilot Flight Control Audio Alert. NOTE: (1) (2) (3)
B.
The Sonalert is located in the pilot's overhead speaker area just forward of FS 166.45.
To gain access to the Sonalert, remove the long narrow speaker cover just right of the pilot's overhead radio speaker. The existing airplane radio alerter bracket is hinged at the forward end, loosen the alerter bracket at the aft end and swing the alerter bracket down. Disconnect the wiring and unscrew the threaded collar on the horn for removal.
Install the Autopilot Flight Control Audio Alert. NOTE: (1) (2) (3)
The Sonalert must be installed with the KAP-150 Autopilot System and the KFC-150, KFC225 Flight Control System.
Connect the wiring and attach the threaded collar on the horn. Swing the alerter bracket up and tighten the bracket at the aft end. Install the cover to the pilot's overhead speaker.
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MODEL 208 MAINTENANCE MANUAL STALL WARNING SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the stall warning system in a serviceable condition.
Task 27-31-00-710 2.
Stall Warning System Operational Check A.
General (1) This task gives the procedures to do a stall warning system operational check.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Stall Warning System Operational Check.
CAUTION: Make sure that the lift transducer is not hot before you do a check. If the thermostat does not operate correctly, the stall vane can overheat. (1) (2) (3) (4)
Make sure that the STALL HEAT switch on the DEICE/ANTI-ICE switch panel in the cockpit is at the OFF position. Apply electrical power to the airplane. Make sure that the STALL WARN circuit breaker is engaged. Move the stall warning vane to the up position and note the stall warning horn audible warning signal. NOTE:
(5) (6) (7) (8)
The airplane includes a stall warning ground disconnect switch. The elevator must be off of the forward stop before the stall warning horn will come on.
(a) If stall warning horn does not to come on, refer to Stall Warning System - Troubleshooting. Set the STALL HEAT switch on the DEICE/ANTI-ICE switch panel in the cockpit to the ON position for 30 seconds, then to the OFF position. Examine the stall warning vane to make sure that it is warm. If you do not feel heat, refer to the applicable troubleshooting chart. Refer to Stall Warning System - Troubleshooting. Remove electrical power from the airplane.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL FLAP SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
The flap system consists of mechanical and electrical components. The flap control lever on the control pedestal provides input to the flap switch actuator, which controls the primary flap motor. The flap actuator assembly drives a bellcrank in the root of the right wing. The flaps are connected to the right inboard forward bellcrank through a series of pushrods, connecting rods, interconnecting rods, and bellcranks. The standby system consists of an independent switch and motor.
Description and Operation A.
The flap control lever controls the flap switch actuator which allows the pilot to select any flap position between 0 and 30 degrees, with detents at UP, 10, 20, and FULL down settings. Any difference between the selected and actual flap position closes one of two micro switches located on the flap switch actuator. The closed micro switch actuates a relay which applies power to turn the primary motor in the proper direction. The motor turns the actuator drive screw, moving the stop nut attached to a tube connected to the right inboard forward bellcrank. As the tube moves, it rotates the right inboard forward bellcrank. The left inboard forward bellcrank is rotated by the right inboard forward bellcrank through the wing-to-wing interconnect rod. The inboard forward bellcranks rotate corresponding inboard aft bellcranks through interconnect rods. Connecting rods from the inboard aft bellcranks rotate the outboard bellcranks. Pushrods connect the flaps to the inboard and outboard bellcranks near the inboard and center flap tracks, respectively. Outboard flap travel is assisted by a cable attached to the inboard aft bellcrank. When the flap position matches the selected position, the micro switch in the flap switch actuator is opened, opening the corresponding relay, stopping motor and flap movement. A follow up cable provides flap position indication through a pointer at the control pedestal. The flap system is equipped with a standby motor which may be utilized in the event of primary system failure. The standby system is controlled by two toggle switches mounted in the overhead console. Before using the standby UP/DOWN switch, the NORMAL/STBY selector switch must be positioned to STBY. If the standby UP/DOWN switch is used to run the flaps with the NORMAL/STBY selector switch in the NORMAL position, the primary and standby motors will run in opposition to each other. Also, the standby system bypasses the limit function of the flap switch actuator, so extreme care must be exercised to prevent running the flaps past their up and down stops.
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MODEL 208 MAINTENANCE MANUAL FLAP SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Flap System Troubleshooting Chart Figure 101 (Sheet 1)
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Flap System Troubleshooting Chart Figure 101 (Sheet 2)
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Flap System Troubleshooting Chart Figure 101 (Sheet 3)
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Flap System Troubleshooting Chart Figure 101 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL FLAP SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the flap system in a serviceable condition.
Task 27-50-00-220 2.
Flap Actuator Mount Bracket Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the flap actuator mount bracket.
B.
Special Tools (1) None
C.
Access (1) Open (unzip) the fabric headliner (passenger) or remove the hard shelled headliner (cargo) to get access to the flap actuator mount bracket. Refer to Chapter 25, Cabin Upholstery - Maintenance Practices.
D.
Do a Detailed Inspection of the Flap Actuator Mount Brackets.
WARNING: If cracks are found in the support structure, reinforce or replace the structure as necessary. Stop drilling cracks is not sufficient; more reinforcement is necessary. (1)
Examine the flap actuator support structure for corrosion, cracks, deformation, or other signs of damage.
E.
Restore Access. (1) Close (zip) the fabric headliner (passenger) or install the hard shelled headliner (cargo). Refer to Chapter 25, Cabin Upholstery - Maintenance Practices. End of task Task 27-50-00-221 3.
Flap Bellcrank Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the flap bellcranks.
B.
Special Tools (1) None
C.
Access. (1) Remove the necessary panels and covers to get access to the flap components on both wings. Refer to Flap Rigging Guide - Adjustment/Test, Figure 501.
D.
Do a detailed inspection of the flap bellcranks. (1) Clean and lubricate the flap bellcranks, interconnect rods, and pushrods. Refer to Chapter 12, Flight Controls - Servicing.
CAUTION: Bellcrank supports are subject to high loads. Close inspection of the supports and the adjacent structure is mandatory. (2)
Examine the flap bellcranks, bellcrank tubes, bearing, and bushings for corrosion, cracks, condition, deformations, signs of damage, and security of installation.
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MODEL 208 MAINTENANCE MANUAL (3) (4)
Examine the bellcrank supports for corrosion, cracks, condition, buckling, and security of installation. Examine the tube ends for interference with the adjacent structure.
E.
Restore Access. (1) Install the applicable panels and covers that were removed to get access to the flap components on both wings. Refer to Flap Rigging Guide - Adjustment/Test, Figure 501. End of task Task 27-50-00-720 4.
Flap System Functional Check A.
General (1) This task gives the procedures to do a functional check of the flap system.
B.
Special Tools (1) Cable Tensiometer (2) Inclinometer (3) External Electrical Power Unit (4) Torque Wrench
C.
Access (1) Remove the necessary panels and covers to get access to the flap components on both wings. Refer to Flap Rigging Guide - Adjustment/Test, Figure 501. (2) Open (unzip) the fabric headliner (passenger) or remove the hard shelled headliner (cargo) to get access to the flap actuator and the wing-to-wing interconnect rod. Refer to Chapter 25, Cabin Upholstery - Maintenance Practices.
D.
Complete a Functional Check of the Flap System. (1) Examine the flap control lever and pointer for security of installation, travel, and signs of damage. (2) Examine the flaps for loose rivets, cracks, condition, and security of installation. (3) Examine the flap cable runs for interference with the structure, correct routing, frozen pulleys, fraying, twisting, and corrosion. (a) Look for interference with the adjacent structure, equipment, wiring, plumbing, and other controls. (4) Move a cloth along the full length of the flap cables to examine for broken wires. (a) If snags are found or you think that there are broken wires, Refer to Chapter 20, Control Cable and Corrosion Limitations - Maintenance Practices. (5) Examine the pulleys, attach brackets, and guard pins for condition, wear, corrosion, and security. (6) Turn the pulleys with your hand to make sure that there is freedom of movement, and to keep even wear on the pulleys. (7) Examine the cable attachment brackets on each flap for condition, corrosion, security, and correct attachment of the cable to the bracket. (8) Examine the motors and the transmission for condition, wear, corrosion, and security. (9) Do the Flap Component Inspection. Refer to Flap Rigging Guide - Adjustment/Test.
E.
Do a Travel and Cable Tension Check. (1) To examine the flap travel and cable tensions, do the Operational Check of the flaps. Refer to Flap Rigging Guide - Adjustment/Test.
F.
Do a Standby Flap Motor Operational Check (Refer to Figure 601).
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CAUTION: You must set the NORMAL/STBY switch to STBY before you operate the standby UP/DOWN switch. Since the standby flap system bypasses the limit function of the flap switch actuator, you must stop the operation of the standby UP/DOWN switch before the flaps reach their limits. This will help prevent overloading and damage to the flap system. (1) (2)
For Airplanes 20800224 and On and 208B0327 and On, and airplanes that incorporate SK208119A, break the frangble copper wire on the UP/DOWN switch guard and the NORMAL/STBY switch guard. Set the battery switch to ON.
WARNING: Before you move the flaps, make sure that the area around the flaps is clear. This will prevent injuries to personnel and damage to the equipment and the flaps. (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17)
Use the flap control lever in the control pedestal to move the flaps to the 10 degree position. Open the NORMAL/STBY switch guard. Set the NORMAL STBY switch to STBY. Move the flaps to the 20 degree position with the standby UP/DOWN switch. Move the flaps to the 10 degree position with the standby UP/DOWN switch. Close the NORMAL/STBY switch guard to set the NORMAL/STBY switch to NORMAL. Move the flaps to the UP position with the flap control lever. Set the battery switch to OFF. For Airplanes 20800224 and On and 208B0327 and On, and airplanes that incorporate SK208119A, use frangible copper wire to safety the NORMAL/STBY switch guard and the UP/DOWN switch guard in the closed position. Make sure that all rod end inspection holes are covered. Make sure that the rod ends are positioned so maximum rotational freedom is available to each rod (so rod housings are perpendicular to attaching bolts). Remove the inclinometers from left and the right flaps. Make sure that the necessary flap system components are secure, torqued, and safety wired. Put the External Power Switch to OFF. Remove external electrical power from the airplane.
G.
Restore Access. (1) Close (zip) the fabric headliner (passenger) or install the hard shelled headliner (cargo). Refer to Chapter 25, Cabin Upholstery - Maintenance Practices. (2) Install the applicable panels and covers that were removed to get access to the flap components on both wings. Refer to Flap Rigging Guide - Adjustment/Test, Figure 501. End of task Task 27-50-00-640 5.
Flap Tracks and Rollers Lubrication A.
General (1) This task provides the procedures to perform a lubrication of the flap tracks and rollers.
B.
Tools and Equipment (1) External Electrical Power Unit, 28 VDC. (2) Dry Solid Film Lubricant (MIL-L-23398D)
C.
Access (1) None
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Standby Flap Motor Switches Figure 601 (Sheet 1)
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Standby Flap Motor Switches Figure 601 (Sheet 2)
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Complete a Lubrication of the Flap Tracks and Rollers. (1) Connect the external electrical power unit to the airplane. (2) Set the External Power Switch to the BUS position. (3) Set the Battery Switch to the ON position. (4) Extend / retract the flaps as necessary to get access to the tacks and rollers. (5) Wipe the flap tracks and the rollers clean and examine for corrosion. (6) Lubricate the flap tracks and the rollers with dry solid film lubricant (MIL-L-23398D). (7) Wipe off unwanted spray. (8) Fully retract the flaps. (9) Set the Battery Switch to the OFF position. (10) Set the External Power Switch to the OFF position. (11) Remove the external electrical power unit from the airplane.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL FLAP SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Tools and Equipment A.
3.
4.
The flap system maintenance practices section has removal and installation procedures for the flap system components.
For a list of tools and equipment, refer to Flight Controls - General.
Flap Actuator Worm Gear Assembly Removal/Installation A.
Remove the Flap Actuator Worm Gear Assembly (Refer to Figure 201). (1) Remove the flap actuator assembly. Refer to Flap Actuator Removal/Installation. (2) Remove the safety wire from the standby motor. (3) Remove the standby motor hex nuts, the standby motor, and the standby motor coupling. (4) Turn the adapter counterclockwise and remove it. (5) Remove the primary motor hex nuts, the primary motor, and the primary motor coupling. (6) Remove the worm gear assembly from the transmission assembly.
B.
Install the Flap Actuator Worm Gear Assembly (Refer to Figure 201). (1) Pack the flap actuator worm gear with B100-24 Grease. Refer to Flight Controls - General. (2) Install the worm gear assembly in the transmission assembly. (3) Make sure the slot in the worm gear assembly aligns with the primary motor coupling. (4) Apply silicone sealer to the base of the threads on the adapter. (5) Turn the adapter clockwise to tighten it on the transmission assembly. (6) Put a new standby motor coupling on the standby motor shaft. (7) Put the standby motor shaft into the adapter. (8) Align the slot in the sleeve with the standby motor coupling by turning the standby motor coupling. Refer to the Model 208 Series Illustrated Parts Catalog for the standby motor coupling part number. (9) Install the primary and secondary motor hex nuts. (10) Tighten the primary and secondary motor hex nuts. (11) Install safety wire on the standby motor. Refer to Chapter 20, Safetying - Maintenance Practices. (12) Install the flap actuator assembly. Refer to Flap Actuator Removal/Installation.
Flap Transmission Removal/Installation A.
Remove the Flap Transmission (Refer to Figure 201). NOTE:
(1) (2) (3) B.
5.
If the standby flap motor operates, but the flaps do not move, it is possible that the flap transmission and the primary and standby motor couplings need to be replaced. Refer to Chapter 27, Flap System Troubleshooting.
Remove the flap actuator assembly. Refer to Flap Actuator Removal/Installation. Remove the bolts that attach the transmission to the actuator assembly. Remove the transmission assembly from the actuator assembly.
Install the Flap Transmission (Refer to Figure 201). (1) Install the bolts that attach the transmission to the actuator assembly. (2) Install the flap actuator assembly. Refer to Flap Actuator Removal/Installation.
Flap Control Lever and Pointer Removal/Installation A.
Remove the Flap Control Lever and the Pointer (Refer to Figure 202). (1) Remove the screws from the cover on top of the control pedestal to get access to the flap control lever and the pointer. (2) Remove the cotter pin, washer, and pin from the flap control lever. (3) Disconnect the flap control cable from the flap control lever. (4) Remove the nut from the pointer.
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Flap Actuator Worm Gear Assembly Figure 201 (Sheet 1)
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Flap Control Lever and Pointer Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7) (8) B.
6.
7.
Remove the follow-up cable from the clamp bolt. Remove the thin washers, clamp bolt, and spring from the pointer. Remove the knob from the control pedestal. Remove the spring, flap control lever, spacer, pointer, and step spacer from the control pedestal.
Install the Flap Control Lever and the Pointer (Refer to Figure 202). (1) Start the knob through the hole on the right side of the control pedestal. (2) Install the spring, flap control lever, spacer, pointer, and step spacer as the knob is pushed through the control pedestal. (3) Tighten the knob with your hand when the threads in the knob contact the threads in the mating part. (4) Connect the flap control cable to the flap control lever. (5) Install the pin, washer, and cotter pin to the flap control lever. (6) Connect the spring to the pointer. (7) Install the clamp bolt through the pointer. (8) Install the thin washers and the follow-up cable through the clamp bolt. (9) Install the nut on the clamp bolt. (10) Put the cover on the control pedestal and install the screws.
Flap Switch Actuator Removal/Installation A.
Remove the Flap Switch Actuator (Refer to Figure 203). (1) Remove the headliner to get access to the flap switch actuator. (2) Remove the cotter pin, washer, and pin from the flap control arm. (3) Disconnect the flap control cable from the flap control arm. (4) Remove the cotter pin, washer, and pin from the follow-up arm. (5) Disconnect the follow-up cable from the follow-up arm. (6) Disconnect the follow-up barrel assembly from the follow-up arm and the stud. (a) Remove the nut and bolt that attaches the follow-up barrel assembly to the follow-up arm. (b) Remove the cotter pin and nut that attaches the follow-up barrel assembly to the stud. (7) Remove the bolts and washers from the support assembly. (8) Disconnect the flap switch actuator from the support assembly. (9) Disconnect and tag the electrical leads to the down and up switch.
B.
Install the Flap Switch Actuator (Refer to Figure 203). (1) Connect the electrical leads to the up and down switch. (2) Put the flap switch actuator in the support assembly. (3) Install the washers and bolts in the support assembly. (4) Connect the follow-up barrel assembly to the stud and the follow-up arm. (a) Install the cotter pin and nut that attaches the follow-up barrel assembly to the stud. (b) Install the nut and bolt that attaches the follow-up barrel assembly to the follow-up arm. (5) Connect the follow-up cable to the follow-up arm. (6) Install the pin, washer, and cotter pin in the follow-up arm. (7) Connect the flap control cable to the flap control arm. (8) Install the pin, washer, and cotter pin in the flap control arm. (9) Make sure all the components are secure. (10) Install the headliner.
Flap Actuator Tube Removal/Installation A.
Remove the Flap Actuator Tube (Refer to Figure 204 and Figure 214). (1) Remove the panel on the bottom of the right wing to get access to the right inboard forward bell crank. (2) Remove the headliner to get access to the flap actuator assembly. (3) Remove the cotter pin and nut from the stud. (4) Disconnect the follow-up barrel assembly from the stud. (5) Remove the setscrew from the actuator tube. (6) Remove the drivescrew stop nut from the actuator tube. (7) Remove the nut, washer, and bolt from the actuator tube.
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Flap Switch Actuator Figure 203 (Sheet 1)
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Flap Actuator Tube Figure 204 (Sheet 1)
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Flap Actuator Tube Figure 204 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (8) Disconnect the actuator tube from the right inboard forward bell crank. (9) Remove the bolts and tie wrap from the right forward wing rib seal assembly. (10) Remove the actuator tube. B.
Install the Flap Actuator Tube (Refer to Figure 204 and Figure 214). (1) Put the actuator tube between the right inboard forward bell crank and the flap actuator assembly. (2) Connect the actuator tube to the right inboard forward bell crank. (3) Install the bolt, washer, and nut in the actuator tube. (4) Make sure the seal assembly is correctly put on the actuator tube. (5) Turn the drivescrew stop nut into the actuator tube. (6) Install the setscrew wet with Loctite 242 or equivalent MlL-S-22473, Grade B Adhesive. (7) Torque the setscrew to 40 inch-pounds. (8) Connect the right forward wing rib seal assembly to the right wing rib and install the bolts and tie wrap. NOTE: (9) (10) (11) (12)
8.
The right forward wing rib seal assembly must be extended with the flaps in the down position.
Connect the follow-up barrel assembly to the stud. Install the nut and cotter pin on the stud. Install the headliner. Install the panels on the bottom of both wings.
Flaps Wing-to-Wing Interconnect Rod Assembly Removal/Installation A.
Remove the Flaps Wing-to-Wing Interconnect Rod Assembly (Refer to Figure 205 and Figure 214). (1) Remove the panels on the bottom of both wings to get access to the left and right inboard forward bell cranks. (2) Remove the headliner to get access to the short and long wing-to-wing interconnect rod assemblies. (3) Remove the nut, washer, and bolt from the left inboard forward bell crank. (4) Disconnect the short wing-to-wing interconnect rod assembly from the left inboard forward bell crank. (5) Remove the nut, washer, and bolt from the right inboard forward bell crank. (6) Disconnect the long wing-to-wing interconnect rod assembly from the right inboard forward bell crank. (7) Remove the bolts and tie wrap from the right aft wing rib seal assembly. (8) Remove the bolts and tie wrap from the left wing rib seal assembly. (9) Loosen the left threaded jam nut and the right threaded jam nut. (10) Remove the barrel from the short wing-to-wing interconnect rod assembly. (11) Remove the short wing-to-wing interconnect rod assembly. (12) Put the long wing-to-wing interconnect rod assembly as far left as possible. (13) Remove the nuts, washers, bolts, and wear plates from the wear plate support. (14) Disconnect the aft end of the wear plate support and remove the barrel from the long wing-towing interconnect rod assembly. (15) Remove the long wing-to-wing interconnect rod assembly.
B.
Install the Flaps Wing-to-Wing Interconnect Rod Assembly (Refer to Figure 205 and Figure 214). (1) Put the long wing-to-wing interconnect rod assembly between the wear plate support and above the right side of cabin. (2) Put the short wing-to-wing interconnect rod assembly above the left side of the cabin. (3) Put the left wing rib seal assembly on the short wing-to-wing interconnect rod assembly. (4) Install the barrel between the short wing-to-wing interconnect rod assembly and the long interconnect rod assembly. (5) Adjust the wing-to-wing interconnect rod assembly length to 73.37 inches and tighten the left and the right threaded jam nuts. (6) Put the right aft wing rib seal assembly on the long wing-to-wing interconnect rod assembly.
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Flaps Wing-to-Wing Interconnect Rod Assembly Figure 205 (Sheet 1)
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Connect the right aft wing rib seal assembly to the right wing rib and install the bolts and tie wrap. NOTE:
(8)
The right aft wing rib seal assembly must be extended with the flaps in the up position.
Connect the left wing rib seal assembly to the left wing rib and install the bolts and the tie wrap. NOTE:
The left wing rib seal assembly must be extended with the flaps in the down position.
(9) Use rivets to install the aft end of the wear plate support. (10) Use the bolts, washers, and nuts to attach the wear plates to the wear plate support. (11) Use a bolt, washer, and nut to connect the short wing-to-wing interconnect rod assembly to the left inboard forward bell crank. (12) Use a bolt, washer, and nut to connect the long wing-to-wing interconnect rod assembly to the right inboard forward bell crank. (13) After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide. (14) Install the headliner. (15) Install the panels on the bottom of both wings. 9.
10.
Flap Interconnect Rods Removal/Installation A.
Remove the Flap Interconnect Rods (Refer to Figure 206 and Figure 214). (1) Remove the wing panels and the covers to get access to the left and the right inboard forward and aft bell cranks. (2) Remove the nut, washer, and bolt from the left inboard forward bell crank. (3) Remove the nut, washer, and bolt from the inboard aft bell crank. (4) Disconnect the left interconnect rod assembly from the left inboard forward bell crank and the inboard aft bell crank.
B.
Install the Flap Interconnect Rods (Refer to Figure 206 and Figure 214). (1) Connect the left interconnect rod assembly to the left inboard forward bell crank and the inboard aft bell crank. (2) Install the bolt, washer, and nut in the inboard aft bell crank. (3) Install the bolt, washer, and nut in the left inboard forward bell crank. (4) After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide. (5) Install the wing panels and the covers.
Flap Connecting Rods Removal/Installation A.
Remove the Flap Connecting Rods (Refer to Figure 207 and Figure 214). (1) Remove the wing panels and the covers necessary to get access to the left and the right connecting rods and the corresponding bell cranks. (2) Remove the nut, washer, and bolt from the inboard aft bell crank. (3) Disconnect the connecting rod assembly from the inboard aft bell crank. (4) Remove the nut, washer, and bolt from the outboard bell crank. (5) Disconnect the connecting rod assembly from the outboard bell crank. (6) Remove the connecting rod assembly through the access hole at Wing Station 53.00.
B.
Install the Flap Connecting Rods (Refer to Figure 207 and Figure 214). (1) Install the connecting rod assembly through the access hole at Wing Station 53.00. (2) Connect the connecting rod assembly to the outboard bell crank. (3) Install the bolt, washer, and nut in the outboard bell crank. (4) Connect the connecting rod assembly to the inboard aft bell crank. (5) Install the bolt, washer, and nut in the inboard aft bell crank. (6) After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide. (7) Install the wing panels and the covers.
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Flap Interconnect Rods Figure 206 (Sheet 1)
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Flap Connecting Rods Figure 207 (Sheet 1)
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11.
Flap Pushrods Removal/Installation A.
Remove the Flap Pushrods (Refer to Figure 208 and Figure 214). (1) Remove the wing panels and the covers to get access to the left and the right inboard and outboard pushrods. (2) Remove the nut, washers, and bolt from the inboard aft bell crank. (3) Remove the nut, washers, and bolt from the flap bracket. (4) Disconnect the inboard pushrod assembly from the inboard aft bell crank and the flap bracket. (5) Remove the nut, washers, bolt from the outboard bell crank. (6) Remove the nut, washers, and bolt from the flap bracket. (7) Disconnect the outboard pushrod assembly from the outboard bell crank and the flap bracket.
B.
Install the Flap Pushrods (Refer to Figure 204, Figure 205, Figure 206, Figure 207, Figure 208, and Figure 214). NOTE:
Airplanes 20800001 thru 20800126 and 208B0001 thru 208B0042 not incorporating CAB88-13 have a nonadjustable right wing-to-wing interconnect rod assembly. The rod is 73.37 inches long.
NOTE:
The following installation instructions are for the left inboard and outboard pushrods. The right inboard and outboard pushrods are similar.
(1)
To provide sufficient adjustment to the rig system, adjust the pushrods, connecting rods, and interconnecting rods to the following nominal lengths: NOTE:
(2) (3) (4) (5) (6) (7) (8) (9) 12.
The rod lengths are measured from center-to-center of the rod ends.
(a) Inboard pushrod assemblies - 8.57 inches. (b) Outboard pushrod assemblies - 7.62 inches. (c) Connecting rod assemblies - 73.82 inches. (d) Left interconnect rod assembly - 11.04 inches. (e) Right interconnect rod assembly - 14.20 inches. Connect the outboard pushrod assembly to the outboard bell crank and the flap bracket. Install the bolt, washers, and nut in the outboard bell crank. Install the bolt, washers, and nut in the flap bracket. Connect the inboard pushrod assembly to the inboard aft bell crank and the flap bracket. Install the bolt, washers, and nut in the inboard aft bell crank. Install the bolt, washers, and nut in the flap bracket. After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide. Install the wing panels and the covers.
Flap Cables and Pulleys Removal/Installation A.
Remove the Flap Cables and Pulleys (Refer to Figure 209 and Figure 214). (1) Remove the wing panels and the covers to get access to the left and the right flap cables and pulleys. (2) Remove the cotter pin, washer, and pin from the inboard aft bell crank. (3) Disconnect the turnbuckle from the inboard aft bell crank. (4) Remove the nut, washer, and bolt from the support. (5) Disconnect the pulley from the support. (6) Remove the cotter pin, washer, and pin from the flap bracket. (7) Disconnect the cable from the flap bracket. (a) Attach a wire to the removed end of the cable. (b) Remove the cable from the system. (c) Let the wire stay in position and do not change the routing of the wire through the structure.
B.
Install the Flap Cables and Pulleys (Refer to Figure 209 and Figure 214). (1) Attach a wire to the cable.
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Flap Pushrods Figure 208 (Sheet 1)
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Flap Cables and Pulleys Figure 209 (Sheet 1)
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Use the attached wire to pull the cable through the system between the inboard aft bell crank and the flap bracket. (3) Connect the turnbuckle to the inboard aft bell crank. (4) Install the pin, washer, and cotter pin in the inboard aft bell crank. (5) Put the cable around the pulley. (6) Install the pulley in the support. (7) Install the bolt, washer, and nut in the support. (8) Connect the cable to the flap bracket. (9) Install the pin, washer, and cotter pin in the flap bracket. (10) After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide. (11) Install the wing panels and the covers.
13.
Flap Inboard Forward Bell Cranks Removal/Installation A.
Remove the Flap Inboard Forward Bell Cranks (Refer to Figure 210 and Figure 214). (1) Remove the panels on the bottom of both wings to get access to the left and the right inboard forward bell cranks. (2) Remove the nut, washer, and bolt that connect the short wing-to-wing interconnect rod assembly to the left inboard forward bell crank. (3) Disconnect the short wing-to-wing interconnect rod assembly from the left inboard forward bell crank. (4) Remove the nut, washer, and bolt that connect the left interconnect rod assembly to the left inboard forward bell crank. (5) Disconnect the left interconnect rod assembly from the left inboard forward bell crank. (6) Remove the nut, washer, bolt, and special washers from the left inboard forward bell crank. (7) Disconnect the left inboard forward bell crank from the supports. (8) Remove the nut, washer, and bolt that connect the actuator tube to the right inboard forward bell crank. (9) Disconnect the actuator tube from the right inboard forward bell crank. (10) Remove the nut, washer, and bolt that connect the long wing-to-wing interconnect rod assembly to the right inboard forward bell crank. (11) Disconnect the long wing-to-wing interconnect rod assembly from the right inboard forward bell crank. (12) Disconnect the interconnect rod from the right inboard forward bell crank. (13) Remove the bolt and special brass washers (if necessary), washers, nut, and cotter pin (if necessary) from the right inboard forward bell crank. NOTE:
The installation of the right inboard forward bell crank changed on Airplanes 20800350 and 20800362 and On, 208B0931, 208B0947, 208B0972, 208B0973, 208B0976, 208B0979, 208B0988, 208B0989, and 208B0991 and On, and airplanes incorporating SK208-148A (CAB02-12, Revision 1). Three washers are used between the lower support and the nut and optional washers are permitted between the supports.
(14) Disconnect the right inboard forward bell crank from the upper and the lower supports. B.
Install the Flap Inboard Forward Bell Cranks (Refer to Figure 210 and Figure 214). (1) Put the left inboard forward bell crank in the supports. (2) Install the bolt, special washers, washer, and nut in the left inboard forward bell crank. NOTE: (3) (4) (5) (6)
The special washers are installed above and below the left inboard forward bell crank between the left inboard forward bell crank and the supports.
Connect the short wing-to-wing interconnect rod assembly to the left inboard forward bell crank. Install the bolt, washer, and nut that connect the short wing-to-wing interconnect rod assembly to the left inboard forward bell crank. Connect the left interconnect rod assembly to the left inboard forward bell crank. Install the bolt, washer, and nut that connect the left interconnect rod assembly to the left inboard forward bell crank.
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Flap Inboard Forward Bell Cranks Figure 210 (Sheet 1)
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Flap Inboard Forward Bell Cranks Figure 210 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (7) (8)
Connect the right inboard forward bell crank to the upper and the lower supports. Install the bolt, special washers (if necessary), washers, nut, and cotter pin (if necessary) in the right inboard forward bell crank. NOTE:
The special washers, if necessary, are installed above and below the right inboard forward bell crank between the right inboard forward bell crank and the supports.
NOTE:
The installation of the right inboard forward bell crank changed on Airplanes 20800350 and 20800362 and On, 208B0931, 208B0947, 208B0972, 208B0973, 208B0976, 208B0979, 208B0988, 208B0989, and 208B0991 and On, and airplanes incorporating SK208-148A (CAB02-12, Revision 1). The right inboard forward bell crank is installed with three washers between the lower support and the nut. Optional washers are permitted between the supports. A maximum of three optional thin washers or a combination of one optional thick and one thin washer is permitted between the supports.
(9) Connect the actuator tube to the right inboard forward bell crank. (10) Install the bolt, washer, and nut that connect the actuator tube to the right inboard forward bell crank. (11) Connect the long wing-to-wing interconnect rod assembly to the right inboard forward bell crank. (12) Install the bolt, washer, and nut that connect the long wing-to-wing interconnect rod assembly to the right inboard forward bell crank. (13) Connect the interconnect rod to the right inboard forward bell crank. (14) After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide. (15) Install the panels on the bottom of both wings. 14.
Flap Inboard Aft Bell Cranks Removal/Installation A.
Remove the Flap Inboard Aft Bell Cranks (Refer to Figure 211 and Figure 214). (1) Remove the wing panels and the covers to get access to the left and the right inboard aft bell cranks. (2) Remove the nut, washer, and bolt that connect the left interconnect rod assembly to the inboard aft bell crank. (3) Disconnect the left interconnect rod assembly from the inboard aft bell crank. (4) Remove the nut, washer, and bolt that connect the connecting rod assembly to the inboard aft bell crank. (5) Disconnect the connecting rod assembly from the inboard aft bell crank. (6) Remove the nut, washers, and bolt that connect the inboard pushrod assembly to the inboard aft bell crank. (7) Disconnect the inboard pushrod assembly from the inboard aft bell crank. (8) Remove the cotter pin, washer, and pin from the inboard aft bell crank. (9) Disconnect the turnbuckle from the inboard aft bell crank. (10) Remove the nut, washer, special washers, and rod that connect the inboard aft bell crank to the supports. (11) Disconnect the inboard aft bell crank from the supports.
B.
Install the Flap Inboard Aft Bell Cranks (Refer to Figure 211 and Figure 214). (1) Put the inboard aft bell crank in the supports. (2) Apply Loctite 242, or equivalent MIL-S-22473, Grade B Adhesive to the threads of the rod. (3) Install the rod, special washers, washer, and nut that connect the inboard aft bell crank to the supports. NOTE: (4) (5) (6)
The special washers are installed above and below the inboard aft bell crank between the inboard aft bell crank and the supports.
Connect the turnbuckle to the inboard aft bell crank. Install the pin, washer, and cotter pin in the inboard aft bell crank. Connect the inboard pushrod assembly to the inboard aft bell crank.
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Flap Inboard Aft Bell Cranks Figure 211 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9) (10) (11) (12) (13) 15.
Install the bolt, washer, and nut that connect the inboard pushrod assembly to the inboard aft bell crank. Connect the connecting rod assembly to the inboard aft bell crank. Install the bolt, washer, and nut that connect the connecting rod assembly to the inboard aft bell crank. Connect the left interconnect rod assembly to the inboard aft bell crank. Install the bolt, washer, and nut that connect the left interconnect rod assembly to the inboard aft bell crank. After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide. Install the wing panels and the covers.
Flap Outboard Bell Cranks Removal/Installation A.
Remove the Flap Outboard Bell Cranks (Refer to Figure 212 and Figure 214). NOTE: (1) (2) (3) (4) (5) (6)
The following removal instructions are for the left outboard bell crank. The right outboard bell crank is similar.
Remove the wing panels and the covers to get access to the left and the right outboard bell cranks. Remove the nut, washer, and bolt that connect the connecting rod assembly to the outboard bell crank. Disconnect the connecting rod assembly from the outboard bell crank. Remove the nut, washers, and bolt that connect the outboard pushrod assembly from the outboard bell crank. Disconnect the outboard pushrod assembly from the outboard bell crank. Remove the bolt, washer(s), and special washers (if applicable) from the outboard bell crank. NOTE:
(7) B.
The outboard bell cranks with part numbers 2622091-1 and 2622091-2 are installed with a washer under the head of the bolt and a bushing with two special washers. The special washers are installed one above and one below the bushing. The other outboard bell cranks are installed with a washer under the head of the bolt and a maximum of three optional thin washers between the supports.
Disconnect the outboard bell crank and the bushing (if applicable) from the supports.
Install the Flap Outboard Bell Cranks (Refer to Figure 212 and Figure 214). NOTE: (1)
The following installation instructions are for the left outboard bell crank. The right outboard bell crank is similar.
Put the outboard bell crank and the bushing (if applicable) in the supports. NOTE:
(2) (3) (4) (5) (6) (7) (8)
The outboard bell cranks with part numbers 2622091-1 and 2622091-2 are installed with a washer under the head of the bolt and a bushing with two special washers. The special washers are installed one above and one below the bushing. The other outboard bell cranks are installed with a washer under the head of the bolt and a maximum of three optional thin washers between the supports.
Install the special washers (if applicable), washer(s), and bolt in the outboard bell crank. Connect the outboard pushrod assembly to the outboard bell crank. Install the bolt, washers, and nut that connect the outboard pushrod assembly to the outboard bell crank. Connect the connecting rod assembly to the outboard bell crank. Install the bolt, washer, and nut that connect the connecting rod assembly to the outboard bell crank. After the installation of all the flap components, adjust the flaps. Refer to Flap Rigging Guide Adjustment/Test. Install the wing panels and the covers.
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Flap Outboard Bell Cranks Figure 212 (Sheet 1)
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16.
Flap Actuator Removal/Installation A.
Remove the Flap Actuator (Refer to Figure 213). (1) Remove the headliner to get access to the flap actuator assembly. (2) Disconnect the electrical connectors from the receptacles. (3) Remove the setscrew from the actuator tube. (4) Remove the drivescrew stop nut from the actuator tube. (5) Remove the nuts, washers, special washers, spacer, and bolt from the transmission assembly. (6) Remove the actuator assembly from the support. (7) Examine the leather washer for excessive wear. If the leather washer is worn, remove the leather washer from the transmission assembly at the drivescrew and discard it.
B.
Install the Flap Actuator (Refer to Figure 213). (1) If necessary, use EC1300L Adhesive to install the leather washer to the transmission assembly at the drivescrew. Refer to Chapter 20, Adhesive and Solvent Bonding - Maintenance Practices. (2) Position the actuator assembly in the support. (3) Install the spacer in the transmission assembly. (4) Install the special washers, washers, bolt, and nut in the transmission assembly. NOTE:
One of the special washers in the transmission assembly is installed above the spacer. The other special washer in the transmission assembly is installed below the spacer between the transmission assembly and the support. The washers in the transmission assembly are installed above and below the special washers.
(5) (6)
Turn the drivescrew stop nut into the actuator tube. Install the setscrew wet with Loctite 242 or equivalent MlL-S-22473, Grade B Adhesive. Refer to Chapter 20, Adhesive and Solvent Bonding - Maintenance Practices. (7) Torque the setscrew to 40 inch-pounds. (8) Connect the receptacles to the electrical connectors. (9) Adjust the flap system. Refer to Flap Rigging Guide. (10) Install the headliner. 17.
Flap Switch Actuator Disassembly/Assembly A.
Disassemble the Flap Switch Actuator (Refer to Figure 215). (1) Remove the flap switch actuator. Refer to Flap Switch Actuator Removal/Installation. (2) Remove the roll pin from the flap control arm. (3) Disconnect the flap control arm, stop assembly, and spacer from the switch actuator body. (4) Remove the spring and ball detent from the flap control arm. (5) Loosen the lock nuts on the down and up switch. (6) Remove the down and up switch from the switch actuator body. (7) Remove the cover from the switch actuator body. (8) Remove the snap ring above the short shaft. (9) Disconnect the short shaft, bearing above the short shaft, and spring washer above the short shaft. (10) Remove the snap ring below the long shaft. (11) Disconnect the long shaft, bearing below the long shaft, spring washer below the long shaft, and striker. (12) Loosen the lock nut on the up stop bolt. (13) Remove the up stop bolt from the stop assembly. (14) Loosen the lock nut on the full-down stop bolt. (15) Remove the full-down stop bolt from the stop assembly. (a) Examine the detents on the stop assembly. If the detents are worn or groves are apparent between the detents, you must replace the stop assembly. (16) Remove the roll pins from the gears on the short and long staff. (17) Disconnect the gear from the short shaft. (18) Disconnect the gear from the long shaft. (19) Remove the cotter pin, washers, spring washer, and gear from the striker.
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Flap Actuator Figure 213 (Sheet 1)
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Flap System Access Figure 214 (Sheet 1)
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Flap Switch Actuator Figure 215 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
Assemble the Flap Switch Actuator (Refer to Figure 215). (1) Install the gear, washers, and spring washer on the striker. (2) Install the cotter pin in the striker. (3) Install the gear on the short shaft. (4) Install the gear on the long shaft. (5) Install new roll pins in the gears on the short and long staff. (6) Install the up stop bolt in the stop assembly. (7) Install the full-down stop bolt in the stop assembly. (8) Tighten the lock nuts on the up stop bolt and the full-down stop bolt. (9) Put the striker in the switch actuator body. NOTE:
The long shaft must be put through the attaching hole in the striker before you assemble the spring washer below the long shaft, the bearing below the long shaft, and the snap ring below the long shaft.
(10) Install the long shaft through the attaching hole in the striker. (11) Assemble the spring washer below the long shaft and the bearing below the long shaft to the switch actuator body. (12) Install the snap ring below the long shaft. (13) Connect the short shaft, spring washer above the short shaft, and bearing above the short shaft to the switch actuator body. (14) Install the snap ring above the short shaft. (15) Install the cover on the switch actuator body. (16) Install the down switch in the switch actuator body. (17) Install the up switch in the switch actuator body. (18) Tighten the lock nuts on the down and up switch. (19) Connect the stop assembly and spacer to the switch actuator body. (20) Install the bolts and washers through the stop assembly and the switch actuator body. (21) Install the spring and detent ball in the flap control arm. (22) Connect the flap control arm to the long shaft. (23) Install a new roll pin in the flap control arm. (24) Adjust and install the flap switch actuator. Refer to Flap Switch Actuator Removal/Installation. (25) Adjust the flap system. Refer to Flap Rigging Guide. 18.
Standby Flap Motor Switches Removal/Installation (Airplanes 20800001 thru 20800223 and 208B0001 thru 208B0326 not incorporating SK208-119A) NOTE:
If the standby flap motor operates, but the flaps do not move, it is possible that the flap transmission and the primary and standby motor couplings need to be replaced. Refer to Chapter 27, Flap System Troubleshooting.
A.
Remove the Standby Flap Motor Switches (Refer to Figure 216). (1) Remove the Wemacs and the knobs. (2) Remove the retainers from the light assemblies. (3) Remove the trim panel. (4) Pull the headliner loose to get access to the overhead panel mounting screws. (5) Remove the screws and lower the overhead panel. (6) Open the NORMAL/STBY switch guard. (7) Remove the jam nut and special washer from the NORMAL/STBY switch. (8) Disconnect the NORMAL/STBY switch guard from the NORMAL/STBY switch. (9) Remove the jam nut and special washer from the standby UP/DOWN switch. (10) Remove the NORMAL/STBY switch and the standby UP/DOWN switch from the overhead panel. (11) Disconnect and tag the electrical leads to the NORMAL/STBY switch and the standby UP/DOWN switch.
B.
Install the Standby Flap Motor Switches (Refer to Figure 216). (1) Connect the electrical leads to the NORMAL/STBY switch and the standby UP/DOWN switch. (2) Install the NORMAL/STBY switch and the standby UP/DOWN switch in the overhead panel. (3) Install the special washer and jam nut on the standby UP/DOWN switch.
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Standby Flap Motor Switches Figure 216 (Sheet 1)
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Standby Flap Motor Switches Figure 216 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (4) (5) (6) (7) (8) (9) (10) 19.
Connect the NORMAL/STBY switch guard to the NORMAL/STBY switch. Install the special washer and jam nut on the NORMAL/STBY switch. Close the NORMAL/STBY switch guard. Put the overhead panel in position and install the screws. Attach the headliner. Put the trim panel in position and install the knobs and the Wemacs. Install the retainers on the light assemblies.
Standby Flap Motor Switches Removal/Installation (Airplanes 20800224 and On, 208B0327 and On, 20800001 thru 20800223 and 208B0001 thru 208B0326 incorporating SK208-119A) A.
Remove the Standby Flap Motor Switches (Refer to Figure 216). (1) Remove the Wemacs and the knobs. (2) Remove the retainers from the light assemblies. (3) Remove the trim panel. (4) Pull the headliner loose to get access to the overhead panel mounting screws. (5) Remove the screws and lower the overhead panel. (6) Break the frangible wire on the NORM/STBY switch guard. (7) Open the NORM/STBY switch guard. (8) Remove the jam nut and special washer from the NORM/STBY switch. (9) Disconnect the NORM/STBY switch guard from the NORM/STBY switch. (10) Break the frangible wire on the standby UP/DOWN switch guard. (11) Open the standby UP/DOWN switch guard. (12) Remove the jam nut and special washer from the standby UP/DOWN switch. (13) Disconnect the standby UP/DOWN switch guard from the standby UP/DOWN switch. (14) Remove the NORM/STBY switch and the standby UP/DOWN switch from the overhead panel. (15) Disconnect and tag the electrical leads to the NORM/STBY switch and the standby UP/DOWN switch.
B.
Install the Standby Flap Motor Switches (Refer to Figure 216). (1) Connect the electrical leads to the NORM/STBY switch and the standby UP/DOWN switch. (2) Install the NORM/STBY switch and the standby UP/DOWN switch in the overhead panel. (3) Connect the standby UP/DOWN switch guard to the standby UP/DOWN switch. (4) Install the special washer and jam nut on the standby UP/DOWN switch. (5) Connect the NORM/STBY switch guard to the NORM/STBY switch. (6) Install the special washer and jam nut on the NORM/STBY switch. (7) Close the NORM/STBY switch guard. (8) Put the overhead panel in position and install the screws. (9) Attach the headliner. (10) Put the trim panel in position and install the knobs and the Wemacs. (11) Install the retainers on the light assemblies. (12) Use frangible copper wire to safety the NORM/STBY switch guard and the standby UP/DOWN switch guard in the closed position .
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MODEL 208 MAINTENANCE MANUAL FLAP SYSTEM - ADJUSTMENT/TEST 1.
General A.
2.
This section has procedures for adjusting the flap switch actuator.
Flap Switch Actuator Adjustment A.
Adjust the Flap Switch Actuator (Refer to Figure 501 and Figure 502). (1) Fabricate a flap switch actuator rigging fixture. Refer to Figure 501. (a) Make the fixture from 0.125-inch (3.175 mm) aluminum as shown in Figure 501. (b) Cut the material to the size shown in Figure 501. (2) Remove the flap switch actuator from the system. Refer to Flap Switch Actuator Removal/ Installation. (3) Temporarily install the bolts in the switch actuator body to keep the stop assembly stable during the adjustment. (4) Loosen the lock nuts on the down and up switch. (5) Remove the down and the up switch from the switch actuator body. (6) Remove the cover from the switch actuator body. (7) Remove the snap ring and follow-up arm assembly from the flap switch actuator. (8) Put the rigging fixture in a vice with the notched end up. (9) Put the flap switch actuator in the rigging fixture so that the sides of the switch actuator are flush with the rigging fixture mating surfaces. (10) Adjust the flap control arm so that the hole (B) is aligned with the A radial on side 2 of the rigging fixture. The up stop bolt may be adjusted to hold the flap control arm in the necessary position. NOTE:
To help adjust the flap control arm so that the hole (B) is aligned with the A radial on side 2 of the rigging fixture, a short #10 diameter screw can be put in hole (B) and used as a pointer.
(11) Adjust the striker so the gear points toward and is aligned with the center of the spacer. (12) Put the follow-up arm assembly in the flap switch actuator so that the hole (C) is aligned with the A radial on side 1 of the rigging fixture. (13) Install the snap ring on the flap switch actuator. NOTE:
To help adjust the flap control arm so that the hole (C) is aligned with the A radial on side 1 of the rigging fixture, a short #10 diameter screw can be put in hole (C) and used as a pointer.
(14) Hold the flap control arm and the follow-up arm stable. (15) Put the cover on the switch actuator body. (16) Install the up and the down switch on the switch actuator body. Turn each switch clockwise until a click is heard. This is the approximate position for the switch. (17) Place the flap switch actuator in the rigging fixture. (18) Make sure that the flap control arm hole (B) is still aligned with the A radial on side 2 of the rigging fixture. (19) Turn the follow-up arm counterclockwise, away from the A radial on side 1 of the rigging fixture. (20) Attach the ohmmeter leads to the normally open and the common terminals of the up switch. NOTE:
The up switch must be set in the limits shown on Figure 502.
(21) Adjust the up switch so the normally open terminal opens as near as possible to the A radial when the follow-up arm hole (C) is turned toward the A radial in a clockwise direction. (22) Tighten the lock nut on the up switch. (23) Make sure the up switch is adjusted correctly. NOTE:
The follow-up arm positions for the switch to open and to close are different due to the flap switch actuator backlash. It is very important to do the instructions that follow to adjust the flap switch actuator accurately.
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Flap Switch Actuator Rigging Fixture Figure 501 (Sheet 1)
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Flap Switch Actuator Rigging Fixture Figure 501 (Sheet 2)
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Flap Switch Positions Figure 502 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (24) Put a mark on side 1 of the rigging fixture to show the position of the follow-up arm hole (C) where the up switch opens. (25) Turn the follow-up arm clockwise away from the A radial on side 1 of the rigging fixture. (26) Attach the ohmmeter leads to the normally open and the common terminals of the down switch. NOTE:
The down switch must be set in the limits shown on Figure 502.
(27) Adjust the down switch so the normally open terminal opens as near as possible to the mark made in step (24) when the follow-up arm is turned in a counterclockwise direction. (28) Tighten the lock nut on the down switch. (29) Make sure the down switch is adjusted correctly. NOTE:
Both switches can be turned for the same distance to stay in the limits shown on Figure 502.
(30) To make sure there is sufficient deadband between the switches, measure the travel, at point F, where both switches are open when you turn the follow-up arm past the set point in a clockwise or a counterclockwise direction. (31) If there is less than 1/8 inch of travel measured in step, slightly adjust either switch out to get sufficient deadband between the switches. (32) Install the flap switch actuator assembly. Refer to Flap Switch Actuator Removal/Installation. (33) Operate the flaps through one cycle to make sure the switch actuator is adjusted correctly. NOTE:
If the flap motor oscillates, there is not sufficient deadband between the switches.
(34) If there is not sufficient deadband between the switches, slightly adjust the inboard switch out. NOTE:
After the adjustments the switch position must be in the ranges shown on Figure 502.
(35) Do the operational check in the Flap Rigging Guide.
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MODEL 208 MAINTENANCE MANUAL FLAP RIGGING GUIDE - ADJUSTMENT/TEST 1.
General A.
This Rigging Guide consists of the following steps: FLAP COMPONENT INSPECTIONDetermines if any parts need repair or replacing prior to rigging. OPERATIONAL CHECKDetermines if flaps need rigging and provides criteria for system acceptance. FLAP RIGGINGDetailed procedure to adjust flaps to meet OPERATIONAL CHECK requirements.
WARNING: Follow appropriate safety precautions when operating and/ or adjusting flap system components.
2.
NOTE:
Two people are required to properly and efficiently perform these procedures.
NOTE:
The rigging procedures in this guide are for flap systems with adjustable interconnect rods on both sides. (Refer to Figure 502 for location of interconnect rods). 20800001 through 20800126 and 208B0001 through 208B0042 that have not complied with CAB88-13 have a non-adjustable interconnect rod between the right inboard forward and aft bellcranks. Compliance with CAB88-13 or installation of 2660020-4 interconnect rod assembly is required to perform the following rigging procedures.
Flap Component Inspection NOTE:
This inspection is necessary to find parts that are unserviceable before you start the flap rigging.
CAUTION: Repair or replace damaged parts before proceeding to operational check. A.
Exterior Inspection (1) Remove panels and covers to gain access to flap components in both wings. (Refer to Figure 501)
Wing Access Panels (Left Shown) Figure 501
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MODEL 208 MAINTENANCE MANUAL (2)
Inspect bell cranks and pulleys (Refer to Figure 502) for worn bearings or deformation. Check bellcrank mounting bolts for proper torque of 35 inch pounds.
Left Wing and Flap Components Figure 502 (3) (4) (5) (6) (7)
Inspect pushrods, interconnect rods, connecting rods, and rod ends for binding, wear, deformation, and other indications of distress or malfunction. (Refer to Figure 502) Inspect flap cables for fraying, frozen pulleys, corrosion, and interference with structure. Inspect wire bundle tie-wraps located below the flap cable. The tie-wrap "tails" should be in the 12 o'clock position - rotate tie wraps as required. Inspect flap tracks and rollers for corrosion, frozen bearings, flat spots, wear, deformation, or other indications of distress or malfunction. Hold flap to prevent it from falling to the end of its track. Remove bolts that attach the left and right interconnect rods to their respective forward inboard bellcranks (Refer to Figure 503). Manually move one flap at a time through its full range of travel to verify free movement.
Left Inboard Forward Bellcrank and Interconnect Rod Figure 503 (8) B.
Reconnect interconnect rods to left and right forward inboard bellcranks.
Interior Inspection (1) Unzip or remove headliner as required to gain access to flap actuator and wing-to-wing interconnect rod.
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WARNING: If cracks are found in support structure, reinforce or replace structure as required. Stop drilling cracks alone is insufficient; additional reinforcement is required. (2)
Inspect flap actuator support structure for cracks, deformation, or other signs of distress. (Refer to Figure 504)
Flap Actuator Support Structure Figure 504 (3)
Check flap actuator drivescrew threads for cleanliness and lubrication. If required, clean and lubricate with #10 weight non-detergent oil. (Refer to Figure 505)
Flap Switch Actuator Figure 505 (4) (5)
Inspect flap switch actuator assembly, and associated linkage for indications of binding, wear, interference, or other signs of distress. (Refer to Figure 505) Inspect the wing-to-wing interconnect rod and wear plates for binding or deformation. (Refer to Figure 506)
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Wing-to-Wing Interconnect Rod Figure 506 (6) 3.
Inspect the left and right fuselage pushrod seal assemblies for condition and security. (Refer to Figure 506)
Operational Check NOTE:
Perform the following OPERATIONAL CHECK completely and note discrepancies. If criteria in any of the OPERATIONAL CHECK steps are not met, perform the complete FLAP RIGGING procedure.
WARNING: Before you move the flaps, make sure that the area around the flaps is clear. This will prevent injuries to personnel and damage to the equipment and the flaps. A.
Position flap lever to UP (1) Check and note cable tension of both flaps. With flaps UP the tension must be 35 pounds +5 or -5 pounds. (Refer to Figure 501 for tensiometer access locations). (2) Lower flaps in small increments to 10 degrees while monitoring cable tension. Minimum cable tension is 10 pounds between UP and 10 degrees.
B.
Return flaps to UP position. (1) Check both flaps at each flap track position (inboard, center, outboard) for fore/aft movement by grasping flap with one hand while holding flap track with other hand. Push flap forward. Free play in mechanical linkage will allow slight forward movement if flap rollers are not contacting the end of each track. Movement must be negligible, indicating each flap roller is contacting the end of its track. NOTE: (2)
Slight up/down movement is acceptable due to roller/flap track clearance.
Loosen nuts on the interconnect rod attach bolts at the inboard forward bellcranks. (Refer to Figure 507 shows left, right typical). Bolts must move up and down by hand, indicating no preload. Retorque after check.
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Left Inboard Forward Bellcrank and Interconnect Figure 507 (3)
(4)
Lower flaps to 10 degrees, then raise to UP, while observing flap actuator support structure for deflection caused by preload. Also, listen and watch for indications of the drivescrew stopnut bottoming out at the end of the drivescrew, indicated by an abrupt stop. Transmission support structure must not deflect, and drivescrew stopnut must not bottom out at the end of the drivescrew in the UP position. (Refer to Figure 504 and Figure 505). With the flap control lever in the UP position, the flap control arm must contact the Up Stopbolt (Refer to Figure 508) and the flap pointer must point to the UP indication on the pedestal cover.
Flap Control Arm Against Stopbolt Figure 508 (5)
Extend the flaps to FULL DOWN and check clearance between the leather washer against the flap actuator transmission and the drivescrew stop nut. Clearance must be a minimum of 0.06 inches. (Refer to Figure 509)
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Flap Actuator Shown in Full Down Position Figure 509 (6) C.
Operate flaps through full range of travel and observe for erratic motion, binding, and interference.
Return flaps to the UP position. (1) Attach an inclinometer to each flap on the trailing edge rib, W.S. 68.00, located approximately thirty-four inches from the inboard edge of the flap. Set inclinometers to 0 degrees. (a) Record inclinometer reading for each flap at the following positions on Table 501. 1 Lower flap control lever to 10 degrees. 2 Lower flap control lever to 20 degrees. 3 Lower flap control lever to FULL DOWN. Raise flap control lever to 20 degrees. 4 5 Raise flap control lever to 10 degrees. 6 Raise flap control lever to UP.
Table 501. Flap Extension and Retraction Tolerances Flap Handle Position
Inclinometer Reading
Required Flap Position Left
0°
0°
10° extending
10° +1 or -2°
20° extending
20° +2 or -2°
30° extending
30° +1 or -2°
20° retracting
20° +2 or -2°
10° retracting
10° +1 or -2°
0° retracting
0°
Right
Flap positions must be within tolerances, symmetrical within 1/2 degree in all positions and within one degree at corresponding extending and retraction positions. D.
If the criteria in any of the previous steps was not met proceed with FLAP RIGGING. If all criteria were met, secure the flap system as follows: (1) Verify all rod end inspection holes are covered. Check that rod ends are clocked so maximum rotational freedom is available to each rod (so rod end housings are perpendicular to attaching bolts).
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) 4.
Remove inclinometers from the left and right flaps. Verify flap system components are secure, torqued, and safety wired as required. Install panels, covers and close headliner. Operate flaps through their full travel range. Check for adequate clearances to panels and covers. Verify smooth operation with no interference.
Flap Rigging NOTE:
If any of the following steps cannot be achieved without exposing rod end inspection holes, or exceeding bellcrank throw limits, set all rods to the nominal lengths specified in Flap System Maintenance Practices, Install Pushrods, then repeat the procedure.
A.
Flap switch actuator re-rigging (1) If flap travel failed Operational Check, Flap extension and retraction tolerances, remove, rig, and reinstall the flap switch actuator assembly in accordance with the instructions in Flap system Maintenance Practices, Flap Switch Actuator Disassembly/Assembly.
B.
Interconnect rod bolt removal (1) Lower flaps to 10 degrees. (2) Perform this step for both flaps. (a) Hold flap to prevent it from falling to the end of its track. (b) Remove interconnect rod attach bolt from the inboard forward bellcrank. (Refer to Figure 510). (c) Position interconnect rod to allow movement of flap and forward bellcrank without interference.
Left Interconnect Rod (Right Typical) Figure 510 C.
Adjustment of pushrods, and/or connecting rods to ensure full roller travel in all flap tracks.
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Roller Travel Checks Figure 511
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MODEL 208 MAINTENANCE MANUAL (1) (2)
D.
Perform the following procedure for both flaps. Adjust cable tension to obtain 35 pounds, + or - 5 pounds in the UP position. (a) Check flap at each track position (inboard, center, outboard) for fore/aft movement by grasping flap with one hand while holding flap track with other hand. Push flap forward. Free play in mechanical linkage will allow slight forward movement if flap rollers are not contacting the end of each track. Movement must be negligible, indicating each flap roller is contacting the end of its track. 1 If inboard, and center or outboard flap rollers contact the forward ends of the track go on to checks at full down position and check for CONDITION "C" or "D". 2 If Condition "A" or "B" exists, adjust as follows: a CONDITION "A" - Release cable tension, and shorten inboard pushrod and/ or lengthen outboard pushrod until inboard, and center or outboard flap rollers contact the forward ends of the track. b CONDITION "B" - Release cable tension, and lengthen inboard pushrod and/ or shorten outboard pushrod until inboard, and center or outboard flap rollers contact the forward ends of the track. (b) Move flap to FULL DOWN position. 1 If both inboard, and center rollers contact the end of their tracks at the same time, go on to Flap actuator adjustment for maximum travel. 2 If CONDITION "C" or "D" exists with flap in FULL DOWN position, adjust connecting rod (Refer to Figure 502) as follows: a CONDITION "C" - Release cable tension, and lengthen connecting rod as required to reclock the outboard bellcrank which increases outboard pushrod travel. Repeat checks and adjustments at UP and full down until both inboard, and center rollers contact the end of their tracks at the same time. b CONDITION "D" - Release cable tension, and shorten connecting rod as required to reclock the outboard bellcrank which decreases outboard pushrod travel. Repeat checks and adjustments at UP and full down until both inboard, and center rollers contact the end of their tracks at the same time.
Flap actuator adjustment for maximum travel. (1) Disconnect flap control cable from flap control arm. (Refer to Figure 512)
Flap Control Arm Against Up Stopbolt Figure 512 NOTE:
Locknuts on Up and Full Down stopbolts must be tightened after each adjustment. (Refer to Figure 511)
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Move flap control arm to the UP position. Flap control arm must be contacting Up stopbolt. (Refer to Figure 512) NOTE:
(3)
The flap control arm must be moved off, then back in contact with the Up stopbolt after each adjustment.
Loosen locknut, then turn in Up stopbolt (Refer to Figure 512) in small increments until drivescrew stopnut contacts the end of the drivescrew (Refer to Figure 513). A distinctive "thump" will be heard when the stopnut contacts the end of the drivescrew. (Refer to Figure 513).
Flap Actuator Shown in UP Position Figure 513 (4) (5)
Mark the drivescrew near the transmission to indicate rotation of drivescrew. (Refer to Figure 513) Back out Up stopbolt (Refer to Figure 514) in small increments until the drivescrew stopnut is 1 turn from the end of the drivescrew. NOTE:
Each full turn of the Up stopbolt is approximately 1 drivescrew turn.
Flap Control Arm Against Up Stopbolt Figure 514 (6)
Tighten UP stopbolt locknut then recheck.
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MODEL 208 MAINTENANCE MANUAL E.
Adjustment and re-connection of interconnect rods to ensure flaps are fully up in tracks without preload. NOTE: (1) (2)
When installing interconnect rod attach bolts at bellcrank, do not tighten nuts at this time.
Verify drivescrew stopnut is positioned in UP position previously set (full up; one turn from end). Repeat this step for both flaps. (a) Place and hold flap in the UP position. (b) Adjust and reconnect interconnect rod so the following conditions are met. 1 Flaps are UP in tracks by procedure in OPERATIONAL CHECK. 2 With interconnect rod reconnected to forward bellcrank, attach bolt must move up/ down by hand, indicating no preload. (Refer to Figure 515) NOTE:
If adjustment limits of the left interconnect rod are reached, the left forward bellcrank can be reclocked by adjusting the wing-to-wing interconnect rod barrel located on left side of cabin overhead.
Left Inboard Forward Bellcrank (Right Typical) Figure 515 3 F.
Move flap control arm to 10 degree detent, then back to UP. Recheck flaps and readjust interconnect rods if necessary.
Adjustment of 10 degree position using the follow-up barrel assembly. NOTE:
Anytime the follow-up barrel assembly (Refer to Figure 516) is adjusted, both the UP and Full Down stopbolts must be readjusted.
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Flap Actuator Shown in UP Position Figure 516 (1) (2)
Verify flaps are in the UP position, then adjust inclinometers to 0 degrees. Move flap control arm to the 10 degree detent. Lengthen or shorten follow-up barrel assembly (Refer to Figure 516) as required to obtain 10 degree position. NOTE:
(3) (4) G.
If 10 degree position cannot be obtained with follow-up barrel assembly adjustment, remove and rig flap switch actuator assembly in accordance with the instructions in Flap system - Maintenance Practices, Flap Switch Actuator Disassembly/Assembly. Repeat this procedure before proceeding to the final adjustment of UP and down stopbolts.
Move flap control arm to the 20 degree detent (second detent aft of the UP detent). Tighten locknuts and recheck for 10 degrees.
Final adjustment of the Up and the Full Down stopbolts. NOTE:
If follow-up barrel assembly was shortened in the previous step, back out the UP stopbolt (Refer to Figure 517) until the end of the bolt is flush with the nutplate.
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Flap Control Arm Against UP Stopbolt Figure 517
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Move flap control arm to UP position. Adjust Up stopbolt in small increments until flaps are full up in tracks and interconnect rod attach bolts move up/down by hand. NOTE:
(3) (4)
The flap control arm must be moved off, then back in contact with the Up stopbolt after each adjustment.
Back out the Full Down stopbolt until the end of the bolt is flush with the nutplate. Move flap control arm to contact Full Down stopbolt. (Refer to Figure 517) Adjust Full Down stopbolt in small increments to obtain a clearance of at least 0.06 inch between the leather washer next to the flap actuator transmission and the drivescrew stopnut. (Refer to Figure 518)
Flap Actuator Shown Full Down Figure 518 NOTE:
H.
Adjustment to obtain symmetrical flaps. (1) Move flap control arm to UP and verify inclinometers are set to 0 degrees. (2) Move flap control arm to FULL DOWN position and note inclinometer readings on both flaps. If the difference is 1/2 degree or less, proceed to Adjustment of Flap Pointer. If the difference is greater than 1/2 degree, accomplish the following procedures. NOTE:
(3) (4) I.
If the upper rollers contact the end of their tracks, or if flap position exceeds 31 degrees, back out Full Down stopbolt to turn off flap motor before either condition occurs.
Decision to lengthen, or shorten pushrods depends on available adjustment and difference from nominal lengths listed in in Flap System - Maintenance Practices, Install Pushrods. Only one side should have to be adjusted.
Remove interconnect rod attach bolt at the inboard forward bellcrank (Refer to Figure 514) on chosen side; release cable tension, and lengthen the pushrods on a flap that is short of travel; or, shorten the pushrods on a flap that has excess travel. Re-rig flaps omitting Flap Actuator Adjustment for Maximum Travel.
Adjustment of flap pointer. (1) Move flap control arm to 10 degree detent and note position of pointer. (2) Move flap control arm to 20 degree detent and note position of pointer. (3) Adjust flap follow-up cable housing at the flap switch actuator support assembly (Refer to Figure 519) as required to position flap pointer as close as possible to 10 and 20 degree marks on the pedestal cover.
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Flap Switch Actuator Support Assembly Figure 519 J.
Adjustment and re-connection of flap control cable. (1) Move flap control arm to the UP position. The flap control arm must be contacting the UP stopbolt. (2) Place flap control lever (located at center pedestal) in the UP position. (3) Adjust the flap control cable clevis end (Refer to Figure 519) to obtain a slight spring-back of the flap control lever in the UP position with the cable reconnected to the flap control arm. If adjustment limits of the clevis end are reached, the flap control cable housing at the pedestal can also be adjusted. NOTE:
K.
If spring-back is excessive, flaps may move up past the point set in (7) b. when flap control lever is moved to the forward end of the slot. Move flap control lever to the forward end of slot to verify this condition does not exist.
Perform previous OPERATIONAL CHECK.
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MODEL 208 MAINTENANCE MANUAL RUDDER GUSTLOCK - MAINTENANCE PRACTICES 1.
General A.
2.
This section covers description and operation, removal/installation and rigging procedures for the rudder gustlock installed on Airplane 20800030 Thru 20800236 and 208B0001 Thru 208B0381.
Description and Operation A.
A rudder gustlock may be installed on the airplane to lock the rudder in neutral position, thus preventing damage to the system by buffeting winds when the airplane is parked. The gustlock is operated by a red T-handle. A placard designating "UP" position of the handle is located on the end of the handle. The T-handle is located below the instrument panel, adjacent to the upper right corner of the pedestal.
B.
The rudder gustlock may be engaged by placing the T-handle in the UP position and pulling it aft until tension on the rudder cables prevents engagement of the handle with the "next tighter" locking tooth. The rudder gustlock may be released by the following methods: (1) gasp the T-handle and rotate it in either direction from its vertical "locked" position until the lock spring disengages from the locking teeth on the handle and allows rudder cable tension and the return spring to pull the T-handle forward to the released position; (2) move the fuel condition lever from CUTOFF to LOW IDLE position. NOTE:
3.
The rudder gustlock T-handle will not stay in locked position unless fuel condition lever is placed in CUTOFF position.
Rudder Gustlock Removal/Installation A.
Remove Rudder Gustlock (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6)
B.
Mounting bracket (7), support (10), reinforcement (40), doubler and grommet (37), bracket (47), and clamps (41) may be left in airplane unless they are excessively worn or damaged.
Remove pins (23) and (29), washers (24) and (31), and cotter pins (25) and (30). Disconnect terminals (35) from housing (22) and lock (43). Remove bolts (12) and (38), nuts (33), and firewall and rudder clamps from control (34). Disconnect control (34) and clamps (32) and (39) from system. Remove pin (23), washer (24), and cotter pin (25) from lock (43) and terminal (35). Disconnect terminal (35) and spring (45) from system. Remove bolts (42), nuts (48), and washers (49) from lock (43) and retainer (44). Disconnect lock (43) and retainer (44) from system. Remove screws (11), pin (23), bolt (18), nut (26), washer (24), and cotter pin (25) from guide (28), slide (27) and housing (22), disconnect T-handle (17), housing (22), spacers (19), spring (20), bushing (21), slide (27), and guide (28) from system. Remove pins (3), (23), and (51), washers (4), (24), and (52), and cotter pins (5), (25), and (53) from fuel condition lever (1), pushrod (2), and mounting bracket (7). Disconnect pushrod (2) and bellcrank (6) from system.
Install Rudder Gustlock (Refer to Figure 201). (1) Connect support (10) to pedestal and bulkhead at FS 100.00. Install rivets, screws (13), washers (14), and nuts (15). Tighten screws. (2) Rivet mounting brackets (7) to support (10). (3) Connect T-handle (17) to support (10) and housing (22). Attach slide (27), guide (28), spacers (19), spring (20), and bushing (21) to housing (22). (4) Connect housing (22) and guide (28) to support (10). Install screws (11), bolt (18), and nuts (26). Tighten bolt and screws. (5) Connect bellcrank (6) to mounting bracket (7) and slide (27). Install pins (23), washers (24), and cotter pins (25). NOTE:
Before proceeding to the next step, clamp a block across the face of the rudder pedals to secure rudder in neutral position.
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Rudder Gustlock Installation Figure 201 (Sheet 1)
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Rudder Gustlock Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (6)
Connect clamps (41) to rudder cables (46). Install and tighten bolts in clamps. NOTE:
(7) (8) (9) (10) (11) (12) (13) (14) (15)
Before tightening bolts in clamps (41), locate aft sides of clamps 5.50 inches from aft face of bulkhead (36).
Remove block from rudder pedals. Locate and rivet doubler and grommet (37) to bulkhead (36). Install lock (43) in doubler and grommet (37). Connect lock (43) and retainer (44) to rudder cables (46) and install bolts (42), washers (49), and nuts (48). Tighten bolts. Locate and rivet bracket (47) to bulkhead. Connect spring (45) to bracket (47) and terminal (35). Locate and rivet reinforcement (40) to aft face of bulkhead (36). Locate control (34) and attach clamps (32) and (39) to support (10) and reinforcement (40). Install and tighten bolts (12) and (38) and nuts (33). Tighten bolts. NOTE:
Before tightening bolts (38), locate aft end of control (34) 3.20 inches from aft face of bulkhead (36). Clamps are furnished to secure control (34) to firewall and rudder bracket assembly.
(16) Connect terminals (35) to housing (22) and lock (48). Install pins (23) and (29), washers (31) and (24), and cotter pins (30) and (25). (17) Connect terminal (35) to lock (43) and install pin (23), washer (24), and cotter pin (25). NOTE:
Ensure T-handle (17) is in the unlocked position. Dimensions from center of lock (43) to aft face of bulkhead (36) should check 8.00 inches.
(18) Connect pushrod (2) to fuel condition lever (1) and bellcrank (6). Install pins (3) and (51), washers (4) and (52). and cotter pins (5) and (53). 4.
Rudder Gustlock Rigging A.
Rigging Procedures. (1) Remove pin (3), washer (4), and cotter pin (5). Detach pushrod (2) from bellcrank (6). (2) Engage rudder gustlock by pulling T-handle (17) aft to the "tightest" locking tooth with decal (16) pointing up. (3) Place fuel condition lever (1) in CUTOFF position. (4) Loosen locknut (54) and turn clevis (55) in required direction to align holes in clevis with hole in bellcrank. (5) Install pin (3) in clevis (55) and bellcrank (6). (6) Move fuel condition lever (1) from CUTOFF to LOW IDLE position. NOTE:
5.
Rudder gustlock must release at or before fuel condition lever reaches LOW IDLE position. If gustlock does not release, adjust clevis (55) in required direction until gustlock releases. Cycle fuel condition control lever through its full travel and check that bellcrank (6) does not bottom out on either end of slots (8) and (9). If bellcrank bottoms out on either end, remove bellcrank and lengthen slots.
Rudder Gustlock Inspection A.
Ensure that clamps (41) are securely attached to rudder cables (46). Clamps must not slip on rudder cables.
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MODEL 208 MAINTENANCE MANUAL RUDDER GUSTLOCK - MAINTENANCE PRACTICES 1.
General A.
2.
3.
4.
This section covers description and operation, removal/installation and rigging procedures for the rudder gustlock installed on Airplanes 20800237 and On and 208B0382 and On.
Description and Operation A.
A rudder gustlock may be installed on the airplane to lock the rudder in neutral position, thus preventing damage to the system by buffeting winds when the airplane is parked. The gustlock is engaged by a handle on the left side of the tailcone stinger.
B.
The rudder gustlock may be engaged with the elevator in the full down position, then placing the gustlock handle, on the left side of the stinger, in the lock position. The gustlock is disengaged automatically any time the elevator is moved up to neutral position.
Rudder Gustlock Removal/Installation A.
Remove Rudder Gustlock (Refer to Figure 201). (1) Remove screw (10) from handle (11), and remove handle. (2) Remove screws (12), and cover plate (8). (3) Remove screws securing stinger, retain screws for reinstallation of stinger. (4) Remove nuts (7), washers (6), washers (4), bolts (3), bolts (1), washers (2), and remove lock assembly (5). (5) Remove pins (21) and remove hinge (22). (6) Remove nut (20), washer (19), and rod-release (18) from pin- eye (15). (7) Remove cotter pin (27), nut (26), washers (24), pin-eye (15), washers (14), and bolt (13). (8) Remove cotter pin (33), nut (34), washers (35), rod end (37) and bolt (40). (9) Remove cotter pin (47), nut (48), washer (49), bolt (52), and washer (51) and remove tube assembly (46). (10) Remove cotter pin (32), nut (31), washer (30), wedges (29). bushing (28), washer (25), bellcrank (36), bushing (23), bellcrank (39), washer (41), bushing (42), wedges (43), washer (44), and bolt (45).
B.
Install Rudder Gustlock (Refer to Figure 201). (1) Position tube assembly (46) on push rod (50) and install washer (51), bolt (52), washer (49), nut (48), and new cotter pin (47). (2) Install bellcranks (36) and (39) using bolt (45), washer (44), wedges (43), bushing (42), washer (41), bushing (23), washer (25), bushing (28), wedges (29), washer (30), nut (31), and new cotter pin (32). (3) Connect rod end (37) between bellcranks (36) and (39) using bolt (40), washers (38), washers (35), nut (34), and new cotter pin (33). (4) Connect pin-eye (15) between bellcranks (36) and (39) using bolt (13), washers (14), washers (24), nut (26), and new cotter pin (27). (5) Position rod-release (18) on pin-eye (15) and install hinge (22) using pins (21). (6) Install washer (19) and nut (20) on pin-eye (15). Do not tighten at this time. (7) Install lock assembly (5) using bolts (3), washers (4), washers (6), nuts (7), washers (2), and bolts (1). (8) Temporarily install handle (11) on lock assembly (5) using screw (10), and rig gustlock.
Rudder Gustlock Rigging A.
Rigging Procedures. (1) Remove screw (10), handle (11), screws (12), and plate (8). (2) Remove screws securing stinger and remove stinger. (3) With lock assembly (5) in the lock position and elevators in the down position, use nuts (16) and (20) to adjust rod-release so that it trips the lock assembly (5) when elevators are raised to the neutral position. (4) Tighten nuts (16) and (20).
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Rudder Gustlock Installation Figure 201 (Sheet 1)
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Rudder Gustlock Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (5) (6)
Install stinger and plate (8) using screws (12). Install handle (11) using screw (10).
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28 CHAPTER
FUEL
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MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
28-00-00
Pages 1-2
Sep 4/2001
28-01-00
Pages 1-3
Aug 1/1995
28-01-00
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Aug 1/1995
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Sep 1/2000
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Aug 1/1995
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Apr 1/2010
28-10-01
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Mar 1/2008
28-10-01
Pages 601-603
Jun 1/2011
28-10-02
Pages 201-205
Sep 2/2002
28-10-03
Pages 201-207
Jun 3/2002
28-10-03
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Jun 1/2011
28-21-00
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Jun 1/2011
28-21-00
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Jun 1/2011
28-22-00
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Mar 3/1997
28-23-00
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Aug 1/1995
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Jun 1/2011
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Aug 1/1995
28-Title 28-List of Effective Pages 28-Record of Temporary Revisions 28-Table of Contents 28-List of Tasks
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS FUEL - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-00-00 28-00-00 28-00-00 28-00-00
Page 1 Page 1 Page 1 Page 2
FUEL SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-01-00 Page 1 28-01-00 Page 1 28-01-00 Page 1
FUEL SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-01-00 Page 101 28-01-00 Page 101
FUEL SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Defueling. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Bay Purging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-01-00 Page 201 28-01-00 Page 201 28-01-00 Page 201 28-01-00 Page 201 28-01-00 Page 201
FUEL STORAGE - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-10-00 Page 1 28-10-00 Page 1 28-10-00 Page 1
FUEL TANKS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Drain Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Drain Valve Poppet Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Float-type Fuel Quantity Transmitters and Low Fuel Level Switches Removal/ Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CAN Bus Fuel Level Sensor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filler Adapter Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filler Cap Leak Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Reservoir Tank and Wing Fuel Bay Primer Adhesion Inspection and Repair (Airplanes 20800001 Thru 20800108 and 208B0001 Thru 208B0008) . . . . . . . . .
28-10-01 Page 201 28-10-01 Page 201 28-10-01 Page 201 28-10-01 Page 201 28-10-01 Page 201 28-10-01 Page 209 28-10-01 Page 211 28-10-01 Page 211 28-10-01 Page 212
FUEL TANKS - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Quantity System Calibration and Check Setup (Airplanes with CAN bus type fuel level sensors).. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Quantity System Calibration (Airplanes with CAN bus type fuel level sensors). Fuel Quantity System Check (Airplanes with CAN bus type fuel level sensors). . . . Fuel Calibration Data Load. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-10-01 Page 501 28-10-01 Page 501 28-10-01 Page 502 28-10-01 Page 503 28-10-01 Page 505
FUEL TANKS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Filler Assembly Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Storage System Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-10-01 Page 601 28-10-01 Page 601 28-10-01 Page 601 28-10-01 Page 601
FUEL TANK SEALING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Integral Fuel Bay Sealant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mixing and Applying Sealant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cure Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Leak Classes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sealing Fuel Leaks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Sealing During Structural Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Integral Fuel Bay Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-10-02 Page 201 28-10-02 Page 201 28-10-02 Page 201 28-10-02 Page 201 28-10-02 Page 201 28-10-02 Page 202 28-10-02 Page 202 28-10-02 Page 204 28-10-02 Page 205
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MODEL 208 MAINTENANCE MANUAL FUEL VENTILATION SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vent Lines and Valve Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Union, Tubes and Cross Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Seals and Reservoir Vent Lines Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Vent Line Float Valve Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-10-03 Page 201 28-10-03 Page 201 28-10-03 Page 201 28-10-03 Page 206 28-10-03 Page 206 28-10-03 Page 206
FUEL VENTILATION SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Vent Line Float Valve Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-10-03 Page 601 28-10-03 Page 601 28-10-03 Page 601
FUEL LINES, VALVES AND FILTERS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Screen and Fuel Shutoff Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lines and Filter Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lines and Filters Removal/Installation (208B Airplanes). . . . . . . . . . . . . . . . . . . . . . . . . Firewall Mounted Fuel Filter Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Filter Disassembly/Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-21-00 Page 201 28-21-00 Page 201 28-21-00 Page 201 28-21-00 Page 201 28-21-00 Page 211 28-21-00 Page 213 28-21-00 Page 213
FUEL LINES, VALVES AND FILTERS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Firewall and Wing Fuel Shutoff Valves Operational Check . . . . . . . . . . . . . . . . . . . . . . Firewall Fuel Shutoff Valve Control Operational Check. . . . . . . . . . . . . . . . . . . . . . . . . .
28-21-00 Page 601 28-21-00 Page 601 28-21-00 Page 601 28-21-00 Page 602
FUEL RESERVOIR - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reservoir Components Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-22-00 Page 201 28-22-00 Page 201 28-22-00 Page 201
FUEL SELECTOR CONTROLS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Selector Controls Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-23-00 Page 201 28-23-00 Page 201 28-23-00 Page 201
FUEL SELECTOR CONTROLS - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Shutoff Valve Linkage Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-23-00 Page 601 28-23-00 Page 601 28-23-00 Page 601
INDICATING - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-40-00 Page 1 28-40-00 Page 1 28-40-00 Page 1
FUEL QUANTITY INDICATING SYSTEMS - MAINTENANCE PRACTICES . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Quantity Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Fuel Level Switch Removal/Installation.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reservoir Low Fuel Level Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . Wing and Reservoir Low Fuel Level Switches Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Quantity Indication System Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-41-00 Page 201 28-41-00 Page 201 28-41-00 Page 201 28-41-00 Page 201 28-41-00 Page 201 28-41-00 Page 204 28-41-00 Page 204
FUEL QUANTITY INDICATING SYSTEMS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Reservoir Warning System Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Quantity and Low Fuel Warning Systems Functional Check . . . . . . . . . . . . . . . .
28-41-00 Page 601 28-41-00 Page 601 28-41-00 Page 601 28-41-00 Page 602
FUEL SELECTORS OFF WARNING SYSTEM - MAINTENANCE PRACTICES . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Warning Horns Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Relay Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Off Warning Switches Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Warning Switches Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Selector Off Warning System Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28-42-00 Page 201 28-42-00 Page 201 28-42-00 Page 201 28-42-00 Page 201 28-42-00 Page 201 28-42-00 Page 207 28-42-00 Page 207 28-42-00 Page 207
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 28-10-01-220
Fuel Filler Assembly Detailed Inspection
28-10-01 Page 601
28-10-01-221
Fuel Storage System Detailed Inspection
28-10-01 Page 601
28-10-03-710
Fuel Vent Line Float Valve Operational Check
28-10-03 Page 601
28-21-00-710
Firewall and Wing Fuel Shutoff Valves Operational Check
28-21-00 Page 601
28-21-00-711
Firewall Fuel Shutoff Valve Control Operational Check
28-21-00 Page 602
28-23-00-640
Wing Shutoff Valve Linkage Lubrication
28-23-00 Page 601
28-41-00-710
Fuel Reservoir Warning System Operational Check
28-41-00 Page 601
28-41-00-720
Fuel Quantity and Low Fuel Warning Systems Functional Check
28-41-00 Page 602
28 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
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MODEL 208 MAINTENANCE MANUAL FUEL - GENERAL 1.
Scope A.
2.
This chapter provides information on systems and components associated with fuel storage, fuel distribution, refueling and fuel quantity indicating.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Sealant Type I
CS-3204 Class A-1/2 Class A-2
Flame Master Chem Seal Div. 11120 Sherman Way Sun Valley, CA 91352
To seal fuel tank area.
Sealant Type I
P/S 890 Class A-2
PRC-DeSoto International 5454 SanFernando Rd. Glendale, CA 91203
To seal fuel tank area.
Sealant Type I
PR-1440 Class A-1/2 Class A-2 Class B-2
PRC-DeSoto International
To seal fuel tank area.
Sealant Type I
P/S 890 Class B-1/2 Class B-2
PRC-DeSoto International
To seal fuel tank area.
Sealant Type I
PR-1440 Class B-1/2 Class B-2
PRC-DeSoto International
To seal fuel tank area.
Sealant Type I
PR-1826 Class B
PRC-DeSoto International
To seal fuel tank area.
Sealant Type I
CS-3204 Class B-1/2 Class B-2
Flame Master, Chem Seal Div.
To seal fuel tank area.
Sealant Type VIII
PR-1428 Class B-1/2 Class B-2
PRC-DeSoto International
To seal fuel tank access panels.
Sealant Type VIII
FR-1081 Class B-1/2 Class B-2
Fiber Resin Corp. 170 W. Providencia Ave. Burbank, CA 91502
To seal fuel tank access panels.
Water Manometer (Range of not more than 70 inches of water.)
Commercially Available
To check wing pressure.
Pressure Regulator
Commercially Available
To regulate input pressure.
Fahrenheit Thermometer
Commercially Available
To monitor test area temperature.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Leak Detector
Eldorado LD-4
Eldorado Chemical Co. Inc. 14350 Lookout Road P. O. Drawer 34837 San Antonio, TX 78265-4837
To locate source of leak.
Primer
1200
Dow-Corning Corp. South Saginaw Rd. Midland, MI 48640
To prime mating surfaces of tube covers.
Thermal Coating
TBS-758
General Electric Co.. Waterford, NY 12188
To seal mating surfaces of tube covers.
Mold Release
MS122
Cessna Aircraft Company Cessna Parts Distribution Department 701, CPD 2 5800 East Pawnee Road Wichita, KS 67218-5590
Use on access panels prior to applying sealant.
Sealant
EC776P
Cessna Aircraft Company
To seal inside of fuel reservoir and inside of wing fuel bays.
Cleaner
Methyl n-Propyl Ketone
Commercially Available
To clean surfaces prior to sealing.
Scotchbrite Pad
N/A
Commercially Available
To remove loose primer.
Gasket
2616017-27
Cessna Aircraft Company
For access cover on fuel reservoir.
Gasket
2696001-1 (Used with 1613-00 Pump) 2616024-2 (Used with 2C6-8 Pump)
Cessna Aircraft Company
For auxiliary fuel pump.
3.
Definition A.
This chapter is divided into sections and subsections to assist maintenance personnel in locating specific systems and information. For locating information within the chapter, refer to the Table of Contents at the beginning of the chapter.
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MODEL 208 MAINTENANCE MANUAL FUEL SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
The fuel system consists of two vented, integral fuel tanks (one in each wing), fuel reservoir, two selector valves, fuel strainer, electrically operated auxiliary fuel pump, and ejector boost pump (both pumps are submerged in the fuel reservoir). Refer to Figure 1 for a fuel system schematic.
Description and Operation A.
Fuel flows from the tanks through two fuel tank shut-off valves in each tank. The valves are mechanically controlled by two fuel selectors in the overhead console. The selectors are labeled: LEFT TANK ON, OFF and RIGHT TANK ON, OFF. Left or right fuel tank, or both at the same time may be selected. In normal operation both tanks should be selected, however, in level cruise flight, fuel may be supplied from either left or right tank.
B.
Fuel flows by gravity from each tank to the fuel reservoir, located under the floorboard at the lowest point in the fuel system. The ejector boost pump and auxiliary electric fuel pump are located in the fuel reservoir. The fuel pumps are submerged in fuel, which helps prevent pump cavitation. Fuel is pumped from the reservoir through the fuel manifold into the engine by the ejector boost pump or auxiliary fuel pump. The ejector boost pump is driven by motive fuel flow from the engine fuel control unit. If the ejector boost pump should fail, the auxiliary electric fuel pump will automatically energize and supply fuel to the engine. The auxiliary electric fuel pump is also utilized to supply fuel to the engine during starts.
C.
A firewall shutoff valve is provided to isolate the fuel supply from engine compartment in case of fire. The manually operated shutoff valve control knob is located on the control pedestal.
D.
The fuel filter is located on the firewall downstream of the shutoff valve. A red warning button is located on top of the filter. If the button is "popped up", the filter screen is clogged, and fuel is forced to bypass the filter. Do not fly the airplane until after the source of fuel contamination is discovered and eliminated. The fuel is routed from the filter through the fuel heater and engine-driven fuel pump. The engine-driven pump delivers fuel, under pressure to the engine fuel filter. The engine fuel filter routes the fuel to a flow divider which distributes the fuel to 14 fuel nozzles in the combustion chamber of the engine.
E.
Excessive fuel, accumulated by engine shutdown, drains into a fire proof EPA fuel reservoir can on engine side of firewall. The reservoir should be drained daily, otherwise, it could overflow on the ground.
F.
The fuel system is vented by two vent valves attached to lines running from each tank outboard to the wing tips, then aft to wing trailing edge. Fuel reservoir is vented to each fuel tank. Complete blockage of the vent system will result in zero fuel flow to engine.
G.
The auxiliary fuel pump switch is located on left sidewall switch and circuit breaker panel, and labeled as follows: FUEL BOOST OFF, NORM and ON. When the switch is in NORM position, auxiliary fuel pump is armed and will operate anytime fuel manifold pressure drops below 4.75 psi. Place the switch in ON position for engine starting and continuous operation of auxiliary fuel pump.
H.
Fuel quantity is measured by four fuel quantity transmitters in each tank (one inboard, one center inboard, one center outboard, and one outboard) and indicated by two electrically operated fuel quantity indicators located on the instrument panel. The indicators are calibrated in pounds and gallons. An empty fuel tank is indicated by a red line and the letter E.
I.
Amber and red fuel system condition warning lights are located on the annunciator panel. Two amber lights (one for each fuel tank) are labeled, LEFT FUEL LOW and RIGHT FUEL LOW. Each light will come on when the fuel level of its respective tank reaches 25 gallons, +5 or -5 gallons. A red light, labeled RESERVOIR FUEL LOW will come on when fuel level in the reservoir drops below 2.15 to 1.95 gallons. An amber light, labeled FUEL PRESSURE LOW will come on when the fuel manifold pressure drops below 4.75 psi.
28-01-00 © Cessna Aircraft Company
Page 1 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Fuel System Schematic Figure 1 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL J.
The fuel system is equipped with drain valves for the purpose of examining fuel and removing contamination from system. The fuel should be drained and examined before the first flight of the day, and after each refueling. An EPA approved fuel reservoir can is mounted on the front LH side of the firewall to catch residual fuel after engine shutdown. Before the first flight of the day, utilize the drain valve on the bottom of the cowling to empty the EPA fuel reservoir can into a suitable container. NOTE:
K.
Do not drain fuel on asphalt or concrete surfaces.
An additional fuel selectors off warning system is incorporated to alert pilot if one or both of the fuel tank selectors are left in the OFF position inadvertently. The system includes redundant warning horns, a red annunciator light labeled FUEL SELECT OFF, actuation switches, and miscellaneous electrical hardware. The dual aural warning system is powered through the START CONT circuit breaker with a non pullable FUEL SEL WARN circuit breaker installed in series to protect the integrity of the start system. The annunciator is powered from the ANNUN PANEL circuit breaker.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FUEL SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to assist the maintenance technician in system understanding. Refer to Figure 101.
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Page 101 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Engine Fuel System Troubleshooting Chart Figure 101 (Sheet 1)
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Page 102 Aug 1/1995
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MODEL 208 MAINTENANCE MANUAL
Engine Fuel System Troubleshooting Chart Figure 101 (Sheet 2)
28-01-00 © Cessna Aircraft Company
Page 103 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Engine Fuel System Troubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL
Engine Fuel System Troubleshooting Chart Figure 101 (Sheet 4)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FUEL SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Fuel system maintenance practices consist of defueling and fuel bay purging.
Precautions A.
Always ground airplane before performing any maintenance on fuel system.
B.
Control drainage of residual fuel from disconnected fuel lines and hoses.
C.
Ensure battery switch is turned OFF when performing maintenance on fuel system, unless otherwise specified. NOTE:
3.
Plugs or caps should be installed on lines, hoses, and fittings to prevent thread damage and contamination).
D.
Use thread compound, Antisieze, Graphite Petrolatum (MlL-T-5544) or equivalent to lubricate, and/or seal leaking connections. Apply sparingly to male fittings, keeping compound off first two threads. DO NOT allow compound to enter fuel system.
E.
Before repairing fuel leaks or performing any maintenance on integral fuel tanks, it may be necessary to defuel airplane.
Defueling NOTE:
Drain valve poppet O-ring replacement can be performed, without defueling the airplane. Refer to Chapter 28, Fuel Tanks - Maintenance Practices.
WARNING: During all fuel system servicing procedures, fire fighting equipment must be available. Two ground wires from TLE-down rings on the airplane to approved ground stakes shall be used to prevent accidental dlsconnectlon of one ground wire. Make sure battery switch is turned off, unless otherwise specified. A.
4.
Defuel airplane. (1) Ground airplane to a suitable ground or stake. (2) Ensure battery switch is turned OFF. (3) Turn fuel selector valves OFF. (4) Remove filler cap(s) from tank(s) to be defueled; insert defueling nozzle. (5) Remove as much fuel as possible with defueling nozzle. (6) Remove drain valves from bottom of fuel tank and drain remaining fuel.
Fuel Bay Purging
WARNING: Purge fuel tank(s) with argon or carbon dioxide prior to repairing leaks to minimize possibility of explosion. Use a portable vapor detector to determine when it is safe to repair fuel tank(s). A.
Purge Fuel Bays. (1) Ground airplane to suitable ground or stake. (2) Ensure battery is disconnected from electrical system. (3) Drain all fuel from tank(s). (4) Remove access door and place inert gas supply hose in fuel tank. (5) Allow gas to flow into tank until fuel vapor can not be detected. Non-sparking tools shall be used to make repairs (air motors, plastic hammers and scrapers, etc.)
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MODEL 208 MAINTENANCE MANUAL FUEL STORAGE - DESCRIPTION AND OPERATION 1.
General A.
2.
This section covers that part of the system which stores fuel. Fuel tanks, tank sealing, vent system and filler caps are included.
Description and Operation A.
An integral fuel tank is located in each wing. The wet-wing cavity starts at WS 53.00, extends outboard to WS 214.30, and fore and aft between the front and rear wing spars. Airplanes 2080001 Thru 20800130 and 208B0001 Thru 208B0089 total capacity of each tank was 167.5 U.S. gallons, with 165 U.S. gallons usable. Airplanes 20800131 and On and 208B0090 and On and airplanes incorporating SK208-52, external wing tank sumps have been installed and the total capacity of each tank is 167.8 U.S. gallons, usable fuel remains at 165 U.S. gallons. The tank consists of upper and lower skins, with bonded stringers, ribs, front and rear spars and access panels on top and bottom skins. The tanks contain fuel drains, strainers, fuel quantity transmitters, and low fuel level switches. A filler cap is located adjacent to WS 214.30; an inboard filler cap may also be located near WS 64. Clean caps regularly with Stoddard solvent to ensure proper sealing.
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MODEL 208 MAINTENANCE MANUAL FUEL TANKS - MAINTENANCE PRACTICES 1.
General A.
2.
Fuel tank maintenance practices consist of fuel tanks component removal/installation and fuel tank leak check.
Fuel Drain Valve Removal/Installation A.
Remove Fuel Drain Valve (Refer to Figure 201). (1) Defuel and purge fuel tank. Refer to Fuel System - Maintenance Practices. (2) Remove drain valve (33) and O-ring (32) from nut (31). (3) Remove lower wing access covers (3) adjacent to WS 53.00. (4) Drill out rivets attaching retainer (30) to skin (34) and remove retainer (30) and nut (31) from fuel tank. (5) Clean sealant from mating surfaces of retainer (30) and skin (34). NOTE:
B.
3.
If nut (31) has not been damaged, or if no fuel leaks are noted in this area, removal of retainer (30), nut (31), and access covers (3) is not necessary.
Install Fuel Drain Valve (Refer to Figure 201). (1) Replace nut (31), attach and seal retainer (30) to skin (34). (2) Replace O-ring (32) on drain valve (33). Install and tighten drain valve (33). (3) Replace lower wing access covers (3) at WS 53.00.
Fuel Drain Valve Poppet Removal/Installation A.
Remove Fuel Drain Valve Poppet O-Ring (Refer to Figure 201). NOTE: (1) (2)
B.
Drain valve poppet O-ring replacement can be performed, without defueling the airplane.
With cross point screwdriver, turn fuel drain valve (33) poppet clockwise, until poppet drops down, exposing poppet O-ring. Using a pointed tool, remove O-ring from poppet. Discard O-ring.
Install Fuel Drain Valve Poppet O-Ring (Refer to Figure 201). (1) Position new O-ring over exposed end of fuel drain valve (33) poppet. Ensure O-ring is properly seated in groove. (2) With cross point screwdriver, push poppet upward and turn poppet counterclockwise until poppet locks in place, flush with bottom of drain valve.
CAUTION: Continuing to turn drain valve poppet counterclockwise past the locked position will result in the fuel drain valve (33) being locked in the open position. 4.
Float-type Fuel Quantity Transmitters and Low Fuel Level Switches Removal/Installation A.
Remove Fuel Quantity Transmitters and Low Fuel Level Switches (Refer to Figure 201). (1) Defuel and purge fuel tank. (2) Remove lower wing access covers (3) adjacent to WS 53.00, WS 101.00, WS 145.00 and WS 214.30. (3) Detach connector (10), disconnect and identify wires at inboard, center inboard, center outboard, and outboard fuel transmitters (5), (23), (24), and (24A). (4) Remove screws (22) from cover (23), and detach cover fuel bulkhead (4). (5) Remove nuts (9) and (11), and detach switch (6) from bracket (7). (6) Remove screws (20), lockwashers (19), washers (18), cover plate (17) from bulkheads (4) and (24), and wing spar (25).
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Fuel Bay Area Figure 201 (Sheet 1)
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Fuel Bay Area Figure 201 (Sheet 2)
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Fuel Bay Area Figure 201 (Sheet 3)
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Fuel Bay Area Figure 201 (Sheet 4)
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Fuel Bay Area Figure 201 (Sheet 5)
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Fuel Bay Area Figure 201 (Sheet 6)
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Fuel Bay Area Figure 201 (Sheet 7)
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MODEL 208 MAINTENANCE MANUAL (7)
Detach fuel transmitters (5), (23), (24A), and gaskets (16) from fuel bulkheads (4) and (24), and wing spar (25). Discard gaskets (16). NOTE:
(8) B.
Remove fuel transmitters (5), (23), (24), and (24A) carefully from bulkheads (4) and (24), wing spar (25), and nut rings (8) to prevent bending float arms (21).
Clean sealant from mating surfaces of fuel bulkheads (4) and (24), wing spar (25), covers (17) and (23), connector (10), and nuts (11) and (9).
Install Fuel Quantity Transmitters and Low Fuel Level Switches (Refer to Figure 201). (1) Attach switch (6) to bracket (7) using nut (9). Connect connector (10) to wiring, and tighten nut (11). NOTE: (2) (3)
Fillet seal top of nut (9) common to threads of switch (6) with PR-1428B 1/2 access cover sealant.
Attach cover (23) to inboard fuel bulkhead (4), install screws (22). Apply sealant to wires, connector (10), nuts (11 and 9), and cover (23). Seal cover on sides and bottom edge, adjacent to fuel bulkhead (4).
CAUTION: After sealing sides and bottom edge of cover (23) to fuel bulkhead (4), verify hole in bottom and holes in sides of cover are not plugged. (4)
Replace fuel transmitters (5), (23), (24) and (24A) in bulkheads (4) and (24), and wing spar (25). Install gaskets (16), cover plates (17), washers (18), lock washers (19) and screws (20). Tighten and torque screws. NOTE:
(5) (6) (7) 5.
Be careful when replacing fuel transmitters (5), (23), (24) and (24A), to keep from bending float arms (21) as they are inserted through nut rings (8).
Check fuel tanks for cleanliness, replace and seal access covers at WS 53.00, WS 101.00, WS 145.00 and WS 214.00. Calibrate fuel system. Refuel airplane and perform visual inspection for fuel leaks.
CAN Bus Fuel Level Sensor Removal/Installation A.
Remove the Sensor (Refer to Figure 202.) (1) Make sure the electrical power is OFF. (2) Defuel the airplane. Refer to Chapter 12, Fuel - Servicing. (3) Remove access panels 523BB and 511AB (left wing) or 623BB and 611AB (right wing). Refer to Chapter 6, Access Plates and Panels Identification. (4) Disconnect the electrical connector from the fuel level sensor. (5) Disconnect the fuel level sensor ground braid. (6) Remove the screws that attach the fuel level sensor to the bracket.
CAUTION: Do not bend the tube. A bent tube will give incorrect operation. (7) B.
Carefully remove the fuel level sensor from the fuel bay
Install the Sensor (Refer to Figure 202.) NOTE: (1)
If the fuel level sensors are replaced, the system must be calibrated. Refer to Fuel Quantity Calibration and Check.
Install the new gaskets on the fuel level sensor.
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Can Bus Fuel Level Sensor Figure 202 (Sheet 1)
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CAUTION: Do not bend the tube. A bent tube will give incorrect operation. (2) (3) (4) (5) (6) (7) 6.
Carefully install the fuel level sensor in the fuel bay. Attach the fuel level sensor with screws and torque the screws to 30 inch-pounds. Connect the electrical connector to fuel level sensor. Connect the fuel level sensor ground braid. Install access panels 523BB and 511AB (left wing) or 623BB and 611AB (right wing). Do a check of the fuel quantity system. Refer to Fuel Quantity Calibration and Check.
Filler Adapter Removal/Installation A.
Remove Filler Adapter (Refer to Figure 201). (1) Defuel and purge fuel tank. (2) Remove cap (35) and detach lanyard (36) from adapter (38). Separate cap (35) and lanyard (36) from adapter (3) Remove screws (39), detach adapter (38) from doubler (41) (4) Clean sealant from mating surfaces of adapter (38) and doubler (41).
B.
Install Filler Adapter (Refer to Figure 201). (1) Seal adapter (38) and doubler (41). (2) Replace screws (39). NOTE: (3)
7.
Removal and installation procedures are identical for inboard and outboard filler adapters.
Attach lanyard (36) to adapter (38).
Filler Cap Leak Test A.
Filler Cap for Leaks (Refer to Figure 201). Fill each tank with approved fuel. Place fuel selectors in OFF position. Plug outboard end of one vent line and both 0.040 inch diameter holes located six inches from outboard ends of both vent lines. (4) Connect a rubber hose and tee into the unplugged vent line. (5) Attach a pressure measuring instrument into hose tee; water manometer, manifold pressure gage, or airspeed indicator.
Test (1) (2) (3)
NOTE:
Pressure must not exceed 0.7 PSI (0.7 PSI equals 20 inches of water on a water manometer; 1.43 inches Hg on manifold pressure gage) or 174 knots on airspeed indicator.
WARNING: Never apply regulated or unregulated air from an air compressor to fuel vent. Major structural damage to the fuel tank may occur if more than 0.7 PSI is applied. (6) (7) (8)
Blow into the open end of hose until pressure reaches 0.7 PSI. It may take several breaths to reach 0.7 PSI. ALWAYS blow into hose, NEVER inhale fuel or vapor. Pinch or clamp hose to maintain pressure in fuel tanks. Apply soap solution to filler caps and inspect for bubbles. If either cap leaks, RELEASE ALL pressure from system before removing defective cap.
CAUTION: Never attempt to remove cap with pressure in system. (9)
Remove leaking cap, replace O-ring. NOTE:
If leak is observed in recessed area around stem, cap must be disassembled and stem O-ring must be replaced.
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MODEL 208 MAINTENANCE MANUAL (10) Clean cap and adapter with Stoddard solvent and perform leak test. NOTE: 8.
Minor leaks may be stopped by removing cotter pin and turning nut on cap lock tab clockwise to apply more tension on O-ring (37).
Fuel Reservoir Tank and Wing Fuel Bay Primer Adhesion Inspection and Repair (Airplanes 20800001 Thru 20800108 and 208B0001 Thru 208B0008) NOTE:
The following procedures provide instructions for inspection of fuel for paint primer particles and repair of fuel reservoir and/or wing fuel bays if evidence of paint primer particles are found in the fuel.
A.
Inspect Fuel (Refer to Figure 203). (1) Drain a pint of fuel from fuel filter (1), thru drain valve (2), into a clear container. Check this fuel sample for any evidence of paint primer particles. Repeat this same procedure, draining at least a quart of fuel from the reservoir tank thru drain valve (4) or drain line (6). (2) If no paint primer particles are evident, no further action is required. Make an entry in the Airplane Logbook stating the inspection has been performed. (3) If there is any evidence of paint primer particles in the fuel, within 25 hours of operation, the reservoir tank must be inspected for loose primer and repaired. If there is evidence of paint primer particles in the fuel and the primer in the reservoir is not loose, wing fuel tanks must be inspected.
B.
Fuel Reservoir Tank Primer Adhesion Inspection and Repair (Refer to Figure 203, Detail C and D). (1) Turn main fuel selector valves to OFF position. (2) Drain fuel from reservoir thru drain valve (4). Airplanes with cargo pod, drain fuel from the reservoir drain valve control (7). (3) Airplanes without cargo pod, remove access panel (3). (4) Airplanes with cargo pod, gain access thru cargo pod door and remove fuel drain line cover (8), lines (6) and drain control (7). Remove fuselage access cover (5). (5) Remove bolts (14) and washers (13); then remove reservoir access plate (15). Discard gasket (16). (6) Clean and purge reservoir before starting inspection. (7) Inspect interior of reservoir for paint primer which has peeled, blistered or separated from the walls. (8) If no primer separation can be found, proceed to Step (17). (9) If loose primer areas are found, disconnect line (12) from adapter (11) and from check valve (18). Disconnect fitting from ejector pump. Remove nut (29), O-ring (30) and ejector pump (10); discard O-ring (30). Disconnect drain line (22) from adapter (21); then remove adapter. Loosen nut (25) and remove adapter (24), nut (25), and O-ring (26). Remove bolts (27), washers (28); disconnect electrical lead (20) and remove auxiliary fuel pump (17); discard gasket (19). Cap check valve (18). (10) Remove any loose primer using a ScotchBrite pad. Using the pad, feather primer edges and rough up surrounding primer surface to improve adhesion of rubber sealing solution. (11) Ensure tank is clean of all particles using clean lint free rags or cheese cloth slightly dampened with Methyl n-Propyl Ketone. (12) Allow tank to completely dry. (13) Brush a coat of EC776 rubber sealant solution over prepared area. Tank temperature must be 60 degrees Fahrenheit or higher. To accelerate curing, ventilate tank with forced warm dry air until solution dries to touch. (14) Apply a second coat of EC776. (15) Inspect treated areas for signs of loose or wrinkled paint primer where sealant solution has been applied. If this condition exists, the area must be cleaned with an acrylic scraper and wiped clean with Methyl n-Propyl Ketone. Repeat Steps (11) thru (14). Sealer must cure for 48 hours before tank is refueled. (16) Position ejector pump (10) in reservoir, and install new O-ring (30) and nut (29). Connect fitting to ejector pump. Uncap check valve (18) and connect line (12) to check valve and ejector pump. Position auxiliary fuel pump (17) in reservoir, using new gasket (19) and install washers (28)
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Fuel Reservoir and Wing Fuel Bay Inspection and Repair Figure 203 (Sheet 1)
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Fuel Reservoir and Wing Fuel Bay Inspection and Repair Figure 203 (Sheet 2)
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Fuel Reservoir and Wing Fuel Bay Inspection and Repair Figure 203 (Sheet 3)
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Fuel Reservoir and Wing Fuel Bay Inspection and Repair Figure 203 (Sheet 4)
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(17) (18) (19) (20)
(21) (22) (23) C.
and bolts (27). Connect adapter (24) to auxiliary fuel pump, using nut (25) and new O-ring (26). Connect auxiliary fuel pump electrical lead (20). Install adapter (21) and connect drain line (22) to adapter. Lightly coat new gasket (16) on both sides with PR1428 sealer. Reinstall access plate (15), gasket (16), washers (13) and bolts (14). Do NOT refuel tank for 48 hours. Turn main fuel selector valves to ON position. Check for leaks around access cover (15). Airplanes without cargo pod, reinstall access panel (3) below reservoir. Airplanes with the cargo pod installed, clean off all excess sealant on access cover (5) with Methyl n-Propyl Ketone or equivalent and let dry. Use MS122 mold release and reseal with PR1428 sealant. Install and secure access cover (5). Install drain lines (6) and drain control (7). Clean and reseal drain line cover (8) with PR-1440 sealant and install drain line cover. If loose primer was found and repairs were made, the airframe and engine fuel pump filters must be cleaned after 25 hours of operation to ensure removal of any lint and/or fiber which may have remained. If no loose primer was found in reservoir inspection, proceed to Wing Fuel Tank Primer Adhesion Inspection and Repair. Make entry in Airplane Logbook stating compliance with the inspection, and method of compliance.
Wing Fuel Tank Primer Adhesion Inspection and Repair (Refer to Figure 203 ). (1) Drain all fuel from wing fuel bays. (2) Remove all wing fuel bay access plates (31).
WARNING: To avoid harmful fumes, ensure area is properly ventilated. (3) (4) (5) (6) (7) (8) (9)
Inspect all interior surfaces of each bay for any paint primer which has peeled, blistered or separated from interior surfaces. If no loose primer is found, proceed to Step 6.C.(14). Dry all fuel from wing fuel bays with clean lint free rags or cheese cloth. Remove loose primer from affected areas using a ScotchBrite pad and a minimal amount of Methyl n-Propyl Ketone. Use Scotch Brite pad to feather primer edges and rough up surrounding primer surface a minimum of 3 inches to improve adhesion of rubber solution. Ensure wing fuel bays are clean and free of all particles, using clean lint free rags or cheese cloth slightly dampened with Methyl n-Propyl Ketone. Allow wing tank to completely dry. NOTE:
Ensure rubber solution does not obstruct or plug fuel drain passages.
(10) Brush a coat of EC776 or PR1005L synthetic rubber solution over prepared area. Wing temperature must be at least 60 degrees Fahrenheit or higher. To accelerate curing, ventilate with forced warm air until solution dries to touch. (11) Apply a second coat of EC776 or PR1005L. (12) Inspect treated areas for signs of loose or wrinkled primer where rubber solution has been applied. If this condition exists, the area must be cleaned with an acrylic scraper and wiped clean with Methyl n-Propyl Ketone. Repeat Steps 6.C.(8) thru 6.C.(11). (13) Inspect all fuel drain passages to ensure they are not plugged or obstructed. (14) Seal and install wing access plates (31). Sealer must cure for 48 hours before fuel bays are refueled. (15) If loose primer was found and repairs were made, airframe and engine fuel pump filters must be cleaned after 25 hours of operation to ensure removal of any lint and/or fiber which may have remained. (16) Make an entry in the Airplane Logbook stating compliance of this inspection and method of compliance.
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MODEL 208 MAINTENANCE MANUAL FUEL TANKS - ADJUSTMENT/TEST 1.
Fuel Quantity System Calibration and Check Setup (Airplanes with CAN bus type fuel level sensors). NOTE:
All G1000 aircraft must have software version 0767.00 or later. The software version is shown on the upper right corner of the MFD on the Þrst page shown after the MFD is powered on in normal operation.
NOTE:
If the fuel quantity indicator on the Garmin G1000 system has a red X on it during normal operation, examine the fuel quantity sensors and wiring and refer to the Garmin G1000 Line Maintenance Manual for more Garmin system troubleshooting. If the values given on the PFD are not the same as the values given in the calibration procedure, refer to the Garmin G1000 Line Maintenance Manual for troubleshooting.
A.
Do a Fuel Quantity Calibration and Check Setup. (1) Make the airplane level. (a) Make the wings level to 0.0 degrees, +0.25 or -0.25 degree. Refer to Chapter 8, Leveling - Maintenance Practices. (b) Make the airplane level to 1.50 degrees, +0.25 or -0.25 degrees nose up position. Refer to Chapter 8, Leveling - Maintenance Practices. (2) Put the selector valves in the ON position. (3) Defuel the airplane. Refer to Chapter 12, Fuel - Servicing. (a) Drain the fuel tanks until the two tanks are empty. (4) Put the fuel selector valves in the OFF position. (5) Add unusable fuel to each fuel tank. Refer to the Pilot's Operating Handbook for the unusable fuel quantity. (6) Put the BATTERY switch and the AVIONICS 1 and AVIONICS 2 switches to the ON position to start the G1000 system in normal mode. (7) Make sure the fuel quantity indications do not show red Xs. (8) Disengage the RIGHT FUEL QTY and LEFT FUEL QTY circuit breakers and make sure that red Xs are shown for the fuel quantity. (9) Engage the RIGHT FUEL QTY and LEFT FUEL QTY circuit breakers and make sure no red Xs are shown for the fuel quantity. (10) Make sure L-R FUEL LOW annunciation is seen. NOTE:
The annunciation will be seen when the fuel quantity indicates less than 25 gallons in each tank.
(11) Disengage the MFD and both PFD circuit breakers. (12) Engage the PFD 2 circuit breaker while the ENT button is pushed on PFD 2. (13) Release the ENT button after the words INITIALIZING SYSTEM show on PFD 2. NOTE:
PFD 2 is now in the conÞguration mode.
(14) Engage the MFD circuit breaker while the ENT button is pushed on the MFD. (15) Release the ENT button after the words INITIALIZING SYSTEM show on the MFD.
CAUTION: Before you do the calibration procedure, you must turn on the G1000 system and let it become stable for a minimum of three minutes. NOTE:
The MFD is now in the conÞguration mode.
(16) Use the FMS outer knob to go to the GRS page group on the MFD. (17) Use the FMS inner knob to go to the GRS/GMU CALIBRATION page on the MFD. (18) Engage the PFD 1 circuit breaker while the ENT button is pushed on PFD 1.
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MODEL 208 MAINTENANCE MANUAL (19) Release the ENT button after the words INITIALIZING SYSTEM show on PFD 1. NOTE:
PFD 1 is now in the conÞguration mode.
(20) Use the Flight Management System (FMS) outer knob to go to the CAL page group. (21) Use the FMS inner knob to go to the FUEL TANK CALIBRATION page. (22) Do the steps that follow: (a) If only a fuel quantity system check is needed, do a Fuel Quantity System Check (Airplanes with CAN bus type fuel level sensors). (b) If a fuel quantity system calibration is needed, do a Fuel Quantity System Calibration (Airplanes with CAN bus type fuel level sensors). 2.
Fuel Quantity System Calibration (Airplanes with CAN bus type fuel level sensors). A.
Do the Fuel Quantity System Calibration. (1) If not completed before, do the Fuel Quantity System Calibration and Check Setup (Airplanes with CAN bus type fuel level sensors). (2) Push the softkeys on the FUEL CALIBRATION page of the PFD, in the sequence that follows, to enter the password. (a) Push Softkey 12 (far right softkey). (b) Push Softkey 11. (c) Push Softkey 10. (d) Push Softkey 9. (3) Push the SCALE softkey to select the scale function.
(4) (5) (6) (7) (8) (9)
NOTE:
The SCALE softkey will be gray with black letters when it is selected.
NOTE:
The scale function needs to be selected when overwriting the 0.00 LBS calibration points for both the left and right tank.
Push the TNK SEL softkey to highlight the CURRENT TANK Þeld. Turn the inner FMS knob to select LEFT. Push the ENT button to select the tank. Make sure that the airplane is level at 1.50 degrees, +0.25 or -0.25 degrees nose up and 0.0 degrees wings level attitude. After 90 seconds, make sure that the CALIBRATED TOTAL value shown for the LEFT tank is stable. Push the EMPTY softkey and push the enter (ENT) button to overwrite the 0.00 LB calibration point in the CALIBRATION TABLE. (a) Push the ENT button again to make sure the calibration is complete. NOTE:
(10) (11) (12) (13) (14)
There will be several calibration points in 67.10LB increments in the CALIBRATION TABLE. If the SCALE function operates correctly, small changes to the calibration points can occur when the EMPTY and ENT buttons are pushed.
Push the TNK SEL softkey to highlight the CURRENT TANK Þeld. Turn the inner FMS knob to select RIGHT. Push the ENT button to select the tank. Make sure that the CALIBRATED TOTAL value shown for the RIGHT tank is stable. Push the EMPTY softkey and push the ENT button to overwrite the 0.00 LB calibration point in the CALIBRATION TABLE. (a) Push the ENT button again to make sure the calibration is complete. NOTE:
There will be several calibration points in 67.10LB increments in the CALIBRATION TABLE. If the SCALE function operates correctly, small changes to the calibration points can occur when the EMPTY and ENT buttons are pushed.
(15) Do the Fuel Quantity System Check (Airplanes with CAN bus type fuel level sensors).
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3.
Fuel Quantity System Check (Airplanes with CAN bus type fuel level sensors). A.
Do the Fuel Quantity System Check. (1) If not completed before, do the Fuel Quantity System Calibration and Check Setup (Airplanes with CAN bus type fuel level sensors). (2) Make sure that the airplane is level at 1.50 degrees, +0.25 or -0.25 degrees nose up and 0.0 degrees wings level attitude. (3) Make sure that the left and right fuel quantity pointers are on the red line on the MFD on the GRS group GRS/GMU CALIBRATION page. (4) Make sure the CALIBRATED VALUE on the CALIBRATION TABLE on PFD 1 agrees with the values in Table 501, Calibration Table. (a) On PFD 1 on the FUEL TANK CALIBRATION page, Þnd the CALIBRATED VALUE next to 0.00LB. Find that number on Table 501 in a "0.00LB Calibrated Value" column. (b) On Table 501, Þnd the value in the 67.10LB Calibrated Value (+0.3 or -0.3) column next to the chosen 0.00LB Calibrated Value. (c) Compare the 67.10LB Calibrated Value on Table 501 to the 67.10LB CALIBRATED VALUE on PFD 1. NOTE: (d)
For example, if the 0.00LB CALIBRATED VALUE is 5.00 then the 67.10 CALIBRATED VALUE should be 25.55, +0.3 or -0.3.
If the 67.10LB CALIBRATED VALUE does not agree with Table 501, +0.3 or -0.3, do a Fuel Calibration Data Load and a Fuel Quantity System Calibration (Airplanes with CAN bus type fuel level sensors).
Table 501. Calibration Table
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
0.00
21.08
0.10
21.17
0.20
21.26
0.30
21.35
0.40
21.44
0.50
21.53
0.60
21.62
0.70
21.71
0.80
21.80
0.90
21.88
1.00
21.97
1.10
22.06
1.20
22.15
1.30
22.24
1.40
22.33
1.50
22.42
1.60
22.51
1.70
22.60
1.80
22.69
1.90
22.78
2.00
22.87
2.10
22.96
2.20
23.05
2.30
23.14
2.40
23.23
2.50
23.32
2.60
23.41
2.70
23.50
2.80
23.58
2.90
23.67
3.00
23.76
3.10
23.85
3.20
23.94
3.30
24.03
3.40
24.12
3.50
24.21
3.60
24.30
3.70
24.39
3.80
24.48
3.90
24.57
4.00
24.66
4.10
24.75
4.20
24.84
4.30
24.93
4.40
25.02
4.50
25.11
4.60
25.19
4.70
25.28
4.80
25.37
4.90
25.46
5.00
25.55
5.10
25.64
5.20
25.73
5.30
25.82
5.40
25.91
5.50
26.00
5.60
26.09
5.70
26.18
5.80
26.27
5.90
26.36
6.00
26.45
6.10
26.54
6.20
26.63
6.30
26.72
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MODEL 208 MAINTENANCE MANUAL Table 501. Calibration Table (continued)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
26.81
6.50
26.89
6.60
26.98
6.70
27.07
6.80
27.16
6.90
27.25
7.00
27.34
7.10
27.43
7.20
27.52
7.30
27.61
7.40
27.70
7.50
27.79
7.60
27.88
7.70
27.97
7.80
28.06
7.90
28.15
8.00
28.24
8.10
28.33
8.20
28.42
8.30
28.50
8.40
28.59
8.50
28.68
8.60
28.77
8.70
28.86
8.80
28.95
8.90
29.04
9.00
29.13
9.10
29.22
9.20
29.31
9.30
29.40
9.40
29.49
9.50
29.58
9.60
29.67
9.70
29.76
9.80
29.85
9.90
29.94
10.00
30.03
10.10
30.12
10.20
30.20
10.30
30.29
10.40
30.38
10.50
30.47
10.60
30.56
10.70
30.65
10.80
30.74
10.90
30.83
11.00
30.92
11.10
31.01
11.20
31.10
11.30
31.19
11.40
31.28
11.50
31.37
11.60
31.46
11.70
31.55
11.80
31.64
11.90
31.73
12.00
31.81
12.10
31.90
12.20
31.99
12.30
32.08
12.40
32.17
12.50
32.26
12.60
32.35
12.70
32.44
12.80
32.53
12.90
32.62
13.00
32.71
13.10
32.80
13.20
32.89
13.30
32.98
13.40
33.07
13.50
33.16
13.60
33.25
13.70
33.34
13.80
33.43
13.90
33.51
14.00
33.60
14.10
33.69
14.20
33.78
14.30
33.87
14.40
33.96
14.50
34.05
14.60
34.14
14.70
34.23
14.80
34.32
14.90
34.41
15.00
34.50
15.10
34.59
15.20
34.68
15.30
34.77
15.40
34.86
15.50
34.95
15.60
35.04
15.70
35.12
15.80
35.21
15.90
35.30
16.00
35.39
16.10
35.48
16.20
35.57
16.30
35.66
16.40
35.75
16.50
35.84
16.60
35.93
16.70
36.02
16.80
36.11
16.90
36.20
17.00
36.29
17.10
36.38
17.20
36.47
17.30
36.56
17.40
36.65
17.50
36.74
17.60
36.82
17.70
36.91
17.80
37.00
17.90
37.09
18.00
37.18
18.10
37.27
18.20
37.36
18.30
37.45
18.40
37.54
18.50
37.63
18.60
37.72
18.70
37.81
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
6.40
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MODEL 208 MAINTENANCE MANUAL Table 501. Calibration Table (continued)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
37.90
18.90
19.20
38.26
19.60
38.61
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
18.80
(5) (6) (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21) 4.
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
0.00LB Calibrated Value
67.10LB Calibrated Value (+0.3 or -0.3)
37.99
19.00
38.08
19.10
38.17
19.30
38.35
19.40
38.43
19.50
38.52
19.70
38.70
19.80
38.79
19.90
38.88
Add 30 gallons of fuel to the left fuel tank. Refer to Chapter 12, Fuel - Servicing. Make sure fuel is sensed in the LEFT tank. Add 30 gallons of fuel to the right fuel tank. Refer to Chapter 12, Fuel - Servicing. Make sure fuel is sensed in the RIGHT tank. Make sure that the airplane is level at 1.50 degrees, +0.25 or -0.25 degrees nose up and 0.0 degrees wings level attitude. Push the TNK SEL softkey to highlight the CURRENT TANK Þeld. Turn the inner FMS knob to select LEFT. Push the ENT button to select the tank. Make sure the CALIBRATED TOTAL value for the LEFT tank is stable and between 133 to 269 LBS . Push the TNK SEL softkey to highlight the CURRENT TANK Þeld. Turn the inner FMS knob to select RIGHT. Push the ENT button to select the tank. Make sure the CALIBRATED TOTAL value for the RIGHT tank is stable and between 133 to 269 LBS. If the values are in tolerance, the procedure is complete. If the CALIBRATED TOTAL values are not in the range, drain the fuel from the tanks and do the fuel calibration procedure again. Put the AVIONICS 1 and AVIONICS 2 switches to the OFF position. Put the BATTERY switch to the OFF position.
Fuel Calibration Data Load A.
Load the Fuel Calibration Data. (1) Disengage PFD 1, MFD, and PFD 2 circuit breakers on the avionics circuit breaker panel. (2) Remove the database cards from the bottom SD card slots on PFD 1, MFD, and PFD 2. (3) Install the SD loader card in the top SD card slot on PFD 1. (4) Start the system in conÞguration mode. (a) Engage the PFD 2 circuit breaker while the ENT button is pushed on PFD 2. (b) Release the ENT button after the words INITIALIZING SYSTEM show on PFD 2. NOTE: (c) (d)
Engage the MFD circuit breaker while the ENT button is pushed on the MFD. Release the ENT button after the words INITIALIZING SYSTEM show on the MFD. NOTE:
(e) (f)
The MFD is now in the conÞguration mode.
Engage the PFD 1 circuit breaker while the ENT button is pushed on PFD 1. Release the ENT button after the words INITIALIZING SYSTEM show on PFD 1. NOTE:
(5) (6) (7)
PFD 2 is now in the conÞguration mode.
PFD 1 is now in the conÞguration mode.
Push the NO softkey when asked "DO YOU WANT TO UPDATE THE SYSTEM FILES?". Use the FMS knobs to go to the SYSTEM group's SYSTEM UPLOAD page . Push the inner FMS knob to start the cursor.
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MODEL 208 MAINTENANCE MANUAL (8) (9) (10) (11) (12) (13) (14)
(15) (16) (17) (18) (19) (20) (21) (22)
Use the cursor to highlight "Fuel Calibration". Push the ENT button. Turn the inner FMS knob to expand the FILE menu. Highlight the "Cessna Caravan-Default Fuel Calibration" Þle. Push the ENT button. Push the LOAD softkey to start the software update. Monitor the upload status. (a) If the upload fails, push the LOAD softkey again. If the upload fails Þve times, contact Cessna Propeller Aircraft Product Support for assistance; (316) 517-5800 or Fax (316) 942-9006. (b) If the upload is successful, push to the ENT to accept the end of the upload. Push the "UPDT CFG" softkey. Select "YES" when asked "Update ConÞg Module?". Push the ENT button. When the update is complete, push the ENT button. Disengage PFD 1, MFD, and PFD 2 circuit breakers. Install the database cards in the bottom SD card slots on PFD 1, MFD, and PFD 2. Remove the SD loader card from the top SD card slot on PFD 1. Start the system in conÞguration mode. (a) Engage the PFD 2 circuit breaker while the ENT button is pushed on PFD 2. (b) Release the ENT button after the words INITIALIZING SYSTEM show on PFD 2. NOTE: (c) (d)
Engage the MFD circuit breaker while the ENT button is pushed on the MFD. Release the ENT button after the words INITIALIZING SYSTEM show on the MFD. NOTE:
(e) (f)
PFD 2 is now in the conÞguration mode.
The MFD is now in the conÞguration mode.
Engage the PFD 1 circuit breaker while the ENT button is pushed on PFD 1. Release the ENT button after the words INITIALIZING SYSTEM show on PFD 1. NOTE:
PFD 1 is now in the conÞguration mode.
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MODEL 208 MAINTENANCE MANUAL FUEL TANKS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fuel tanks in a serviceable condition.
Task 28-10-01-220 2.
Fuel Filler Assembly Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the fuel filler assembly.
B.
Special Tools (1) Medeco Key Lube or Equivalent
C.
Access (1) None
D.
Do a Detailed Inspection of the Fuel Filler Assembly. (1) Visually examine the fuel filler caps, covers, lanyard cords, and hinges for security of installation, cleanliness, corrosion, and other damage. (2) Examine the O-rings for security of installation, deterioration, cleanliness, and other damage. (3) Apply Medeco Key Lube to the inside of the fuel cap locks. (4) Insert the key and operate the lock mechanism several times and make sure that the operation is smooth. (5) Wipe off unwanted lubricant.
E.
Restore Access (1) None End of task Task 28-10-01-221 3.
Fuel Storage System Detailed Inspection A.
General (1) This task gives the information needed to do a detailed inspection of the fuel storage systems.
B.
Special Tools (1) None
C.
Access (1) Fuselage access panels and covers NOTE:
D.
The fuel access panels and covers are removed after the inspection steps for removing the fuel.
Do a Detailed Inspection of the Fuel Storage System.
WARNING: Before you do maintenance on the fuel system, you must read and understand all of the fuel system maintenance, fire precautions, and safety practices. Refer to Fuel System - Maintenance Practices and Chapter 12, Fuel – Servicing. (1)
Defuel the airplane. Refer to Chapter 12, Fuel – Servicing. (a) Remove the remaining fuel from the fuel storage areas with the fuel drain valves. Refer to Chapter 12, Fuel – Servicing.
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MODEL 208 MAINTENANCE MANUAL (2)
Remove lower wing fuel access panels 521AB, 521BB, 521DB, 521EB left, and 621AB, 621BB, 621DB and 621EB right. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation.
CAUTION: Be careful to not separate the wing skin doubler from the wing skin. (3) (4) (5) (6) E.
(a) Purge the fuel tanks. Refer to Chapter 12, Fuel – Servicing. Examine the eight (8) quantity transmitter mounting plates for condition, leaks, and security. Examine the transmitters wire harnesses and terminals at the transmitters for condition and security. Examine the tank drains for condition, leaks, and security. Examine the fuel lines, fuel shut-off-valves, and filters for condition and security.
Do a Fuel Reservoir Tank Inspection. NOTE:
(1)
(2)
This inspection includes the system and components of the inner reservoir tank cavity only. On airplanes with pods, the reservoir includes a mechanical drain that is connected to a push/pull cable on the left side of the pod.
For (a) (b) (c)
airplanes with a POD installed, remove the drain line cover. Install a plug In the drain line opening. Pull the drain valve open and examine the line connections for leaks. Examine the reservoir drain system for condition, security, correct drain valve rigging, and correct operation of the drain valve. (d) Remove the plug from the drain line opening. Examine the metal fuel lines, manifold, vent lines, and drain lines in the tank area for condition, security, and signs of leakage. NOTE:
(3) (4) (5)
The auxiliary fuel pump seal drain line is not installed on airplanes with the Airborne pump installed.
Examine the seals where fuel and vent lines go through the structure for condition and security. Examine the rubber hoses (7 each) and the hose clamps for condition, leaks, deterioration, and security. Examine the fuel pressure switch and the auxiliary pump relay for condition, security, correct wire routing, chafing of wires, and leaks at the pressure switch.
CAUTION: Before you remove fuselage access panel 253AC, make sure that all residual fuel is drained. (6)
Remove fuselage access cover 253AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (7) Examine the reservoir tank for condition, leaks, and security of installation (8) Examine the reservoir tank mounting brackets and attachment structure for condition, cracks, corrosion, and security. (9) Examine the auxiliary fuel pump, ejector boost pump, and plumbing for condition and security. (10) Examine the swing check valves for condition, security, and freedom of movement from the closed to the open position. (11) Examine the interior paint primer to make sure that it is not peeled, blistered, or separated from the surfaces of the reservoir. (a) If loose primer is found, find the cause and correct it. Refer to Fuel Tanks - Maintenance Practices. 1 If no paint primer particles are found, no further action is necessary. F.
Restore Access (1) Install fuselage access cover 253AC. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
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Install lower wing fuel access panels 521AB, 521BB, 521DB, 521EB left, and 621AB, 621BB, 621DB and 621EB right. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation.
CAUTION: Be careful to not separate the wing skin doubler from the wing skin. (3) End of task
(a) Purge the fuel tanks. Refer to Chapter 12, Fuel – Servicing. Refuel the airplane. Refer to Chapter 12, Fuel – Servicing.
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MODEL 208 MAINTENANCE MANUAL FUEL TANK SEALING - MAINTENANCE PRACTICES 1.
2.
General A.
If a leak has started in the tank or the wing has been repaired, you may need to seal the fuel tank again. These procedures provide instructions to classify leaks, repair leaks, seal the fuel tank during structural repair, and to do a integral fuel tank test.
B.
The sealant must be applied properly to make sure the fuel tank seals completely. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices for specific sealing procedures and illustrations.
Integral Fuel Bay Sealant A.
3.
Two kinds of sealants are used, one to seal the bay and the other to seal the access doors, the fuel quantity transmitters, the fuel inlet assemblies, and the fuel test receptacle. The access door sealant is more pliable and will not bond to metal as firmly as the bay sealant. The access door sealant lets the doors, the fuel quantity transmitters, etc., be removed without damage. The sealants can be identified by color. The bay sealant is white and its accelerator is a black paste. The access door sealant is red or black and its accelerator is black.
Mixing and Applying Sealant
WARNING: Do cleaning and sealing operations in a well ventilated area away from excessive heat, open flame and sparks. Do not smoke in any area where cleaning solvents and sealants are being used. WARNING: Use cleaning solvents only from approved containers. WARNING: Discard solvent-wetted cheesecloth and used solvent in specially supplied safety containers. Put reusable, solvent-saturated cleaning cloths in designated containers. WARNING: Wear rubber gloves and safety glasses or goggles during cleaning operations. When you use solvent and sealing materials, prevent contact with eyes or breaks in skin, and repeated or prolonged contact with skin. WARNING: Do not ingest solvent and sealing materials. Before eating or smoking, thoroughly wash your hands after you use solvent and sealing materials. A. 4.
Refer to specific manufacturer’s instructions and Chapter 20, Fuel Weather and High-Temperature Sealing - Maintenance Practices for procedures to mix and apply all types of sealants.
Cure Time A.
Cure time for sealants is calculated at a temperature of 77°F and 50 percent relative humidity. Table 201 contains the cure time for Type I and Type VIII sealants:
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Table 201. Cure Properties of Type I and Type VIII Sealants MINIMUM APPLICATION TIME (HOURS)
MAXIMUM TACK-FREE TIME (HOURS)
MAXIMUM CURE TIME (HOURS)
A-1/2
0.5
10
40
A-2
2.0
40
72
B-1/2
0.5
4
6
B-2
2.0
40
72
B-1/2
0.5
10
24
B-2
2.0
24
72
CLASS
TYPE I
TYPE VIII
5.
Fuel Leak Classes A.
Classes of Fuel Leaks (Refer to Figure 201). NOTE: (1) (2) (3) (4)
B.
The class of a leak is identified by the size of the monitored leak. The leak location will identify if repair is necessary before the next flight.
Stains - An area of 0.75 inch in diameter (or less) is classified as a stain. Seep - An area from 0.75 inch to 1.50 inches in diameter is classified as a seep. Heavy Seep - An area from 1.50 inches to 4.00 inches in diameter is classified as a heavy seep. Running Leak - The size will be different with location and intensity of the leak.
Leaks that must be repaired before flight are: (1) Running leaks in any area. (2) Stains, seeps or heavy seeps in a closed area. NOTE:
C. 6.
"Closed areas" means an area inside a wing leading edge or the section of a fuel tank located between the rear spar and the trailing edge.
Repair the following leaks when the airplane is grounded for other maintenance: (1) Stains, seeps or heavy seeps not in a closed area.
Sealing Fuel Leaks A.
Find Source of Leak. (1) Fuel can flow along a seam or structure of the wing for several inches. This makes it hard to find the leak source. A stained area is an indication of the leak source. (2) To find fuel leaks do a Integral Fuel Bay Test of the complete bay. (3) Another procedure to find the source of a fuel leak is as follows: (a) Remove the access doors in the area of the leak. (b) Apply a soap bubble solution to the outside of the bay in the area of the leak. (c) Use a nozzle to blow air from the inside of the bay in the area of the leak. (d) Monitor the outside of the bay for air bubbles.
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Fuel Leak Classes Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
Repair Leak. NOTE: (1) (2) (3) (4) (5) (6) (7)
7.
To see an illustration of the different sealing procedures given below, refer to Chapter 20, Fuel, Weather, Pressure and High-Temperature Sealing - Maintenance Practices.
Remove the sealant in the area of the leak. Clean the area and apply a fillet seal. Use a small paddle to push the sealant into the area of the leak. Make sure to remove all the air bubbles from the seal. If the leak occurs around a rivet, the rivet must be driven again. If the leak occurs around a bolt, loosen the bolt, torque the bolt, and seal around the nutplate. If removed, apply fay surface door sealant to the access doors, the fuel quantity transmitters, etc. If removed, install the access doors, the fuel quantity transmitters, etc. Do a fuel leak test on the fuel bay (Refer to Integral Fuel Bay Test).
Wing Sealing During Structural Repair A.
Preliminary Notes. NOTE: (1) (2) (3) (4) (5) (6) (7)
B.
To see an illustration of the different sealing procedures given below, refer to Chapter 20, Fuel, Weather, Pressure and High-Temperature Sealing - Maintenance Practices.
Seal that bay area after any repair that breaks the fuel bay seal. If it is necessary to seal the repair parts, make sure they are installed during the sealing procedures. Make sure to fay-surface-seal and fillet seal the fuel side of all the joints within the boundary of the bay which do not supply a direct fuel path out of the bay. The fuel spar flanges and the rib flanges are examples of these joints. To fay-surface-seal is to apply sealant to one mating part before the parts are assembled. When you make a fay-surface-sea, make sure to apply enough sealant so it will squeeze out completely around the joint when the parts are attached. Apply a fillet seal after the joint is fay-surface-sealed and attached. Apply a fillet seal to the edge of all riveted joints, joggles, bend reliefs, voids, rivets, or fasteners. Seal all boundaries and any other place that could become a fuel leak. It is not necessary for the fay sealant to cure before you apply the fillet seal. However, the fay sealant must not have dirt or other contaminants before you apply the fillet seal. Join the fillets laid on intersecting joints to make a continuous seal. Make sure to push the sealant into the joint to remove trapped air bubbles. Use an extrusion gun to lay a bead along the joint. Use a small paddle and remove all the trapped air to eliminate the bubbles.
Sealing Procedures.
CAUTION: Protect drain holes and fuel outlet screens when you apply sealants. NOTE:
To see an illustration of the different sealing procedures given below, refer to Chapter 20, Fuel, Weather, Pressure and High-Temperature Sealing - Maintenance Practices.
NOTE:
During structural repair, before the parts are sealed and put in position for the final installation, the parts must be drilled already, have a countersink or dimple, and be cleaned.
(1)
Remove all sealant from the area to be sealed. NOTE:
The best method to remove most of the sealant is to use a chisel-like tool made of hard fiber. The remaining sealant can be removed with aluminum wool but do not use steel wool or sandpaper.
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MODEL 208 MAINTENANCE MANUAL (2)
Leave a taper on the remaining sealant. NOTE:
(3) (4)
Use a vacuum to remove all chips, filings, dirt, etc., from the tank area. Use Methyl n-Propyl Ketone or equivalent to clean all surfaces that you will seal. Dry all cleaned surfaces with a clean cloth before the solvent evaporates. NOTE:
(5)
8.
Always put the solvent on the cloth to prevent contamination of the solvent. Do not let the solvent drip from the cloth. Dry the surfaces with clean, dry cloths until the white is removed. Never use dirty solvent.
Apply the fay-surface-sealant to one mating part and install the rivets or fasteners while the sealant is still in its work life. NOTE:
(6) (7) (8) (9) (10) (11)
When the new sealant is applied, the taper will make a scarf bond and a continuous seal.
During the sealing procedure, you must monitor the supply of mixed sealant to be sure it is in its normal work life. To make sure, use a small wooden paddle, or tongue depressor to gather a small amount of the sealant. Touch this sealant to a piece of clean sheet metal. If it bonds, you can still use the sealant. If it does not bond, the sealant is not in its permitted work life and you must discard it.
Apply a fillet seal to the repaired area on the inside of the tank. If necessary, apply a fay-surface-seal to the access door, the fuel quantity transmitter, etc. If necessary, install the door(s). Let the sealant cure. Clean the stains on the outer surface. Do a fuel tank leak test (Refer to Integral Fuel Bay Test).
Integral Fuel Bay Test A.
The fuel system has two vented, integral fuel tanks (one in each wing). The following procedures are the integral fuel bay test. (1) Remove the vent line from the vent fitting and the cap fitting. (2) Disconnect the fuel lines from the bay. (3) To one of the bay fittings, attach a water manometer capable of measuring 20 inches of water. (4) To the other bay fitting, connect a well-regulated air supply (0.5 PSI maximum, or 13.8 inches of water). NOTE: (5)
You can use nitrogen if the temperature changes during the test will have an effect on the bay.
Make sure the filler cap is installed and sealed.
CAUTION: Do not try to apply pressure to the bay without a good regulator and a positive shutoff in the supply line. Do not pressurize the fuel bay to more than 0.5 PSI or damage can occur. (6) (7) (8) (9)
Slowly apply pressure until you get 0.5 PSI. If necessary, apply a soap solution. 15 to 30 minutes are necessary to let the pressure become stable. If the bay holds the pressure for 15 minutes without a decrease in the pressure, the seal is satisfactory. (10) If there is a decrease in the pressure, seal the bay again and then do a fuel leak test.
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MODEL 208 MAINTENANCE MANUAL FUEL VENTILATION SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
The fuel ventilation system is divided into two sections. One section is routed from the fuel reservoir to a cross in left wing, with lines running outboard from cross to each fuel tank. The second section includes a fuel vent line valve in each tank and lines extending outboard to each wing tip, then aft to trailing edge of wing. Each line has a 0.040 inch diameter hole drilled in the upper surface 6.50 inches from aft end of line. The fuel vent line valve, located inside each fuel tank, is equipped with a float valve that shuts off the fuel flow to the vent line any time the fuel level in the tank rises above the level of the vent line. A relief valve in the fuel vent line valve opens at 0.7 PSI negative pressure and 1.0 PSI positive pressure to prevent fuel starvation or tank damage.
Vent Lines and Valve Removal/Installation A.
Remove the Vent Lines and Valve (Refer to Figure 201). (1) Remove the wing tip assembly. Refer to Chapter 57, Wings - Removal/Installation. (2) Remove the lower wing surface access covers adjacent to the vent lines, fuselage access cover on fuselage below reservoir and the left and right sidewall panels aft of pilot's and copilot's entrance doors. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (3) Remove the fuel from the fuel tanks. Refer to Fuel Systems - Maintenance Practices. (4) Do a purge of the fuel tanks. Refer to Fuel Systems - Maintenance Practices. (5) Disconnect tube nut, remove nut and O-ring discard ring. (6) Remove the fuel vent valve from the fuel bulkhead. (7) Remove the fuel vent line valve from the fuel tank. (8) Remove the washer from the fuel vent line valve. (9) Remove the nuts and screws from the clamps and jumper wire. (10) Disconnect the jumper wire from the clamps. (11) Loosen the clamps and disconnect the hose from the wing tip vent line and the vent line. Discard the hose. (12) Remove the screws from the clamp and disconnect the grommet from the wing rib. Remove vent line from the wing. (a) Make sure that the hole in the tube is not plugged. (13) Loosen the tube nuts and remove the union. (14) Remove the bolts and detach the retainers from the root rib.
B.
Install the Vent Lines and Valve (Refer to Figure 201). (1) Replace the union and attach the tube nuts. (2) Attach the vent line to the wing. (3) Connect the grommet to wing rib. (4) Attach the clamp to the wing rib and replace the screws.
CAUTION: The end of the wing tip vent line must project 0.50 inch, +0.03 or -0.03 inch beyond the wing trailing edge to supply even fuel flow from left and right sides of the system. (5) (6) (7)
Attach the hose to the vent line and wing tip vent line. Install and tighten the clamps. Attach the upper wires to the clamps. Install the screws and nuts. (a) Make sure the jumper wires and clamps are bonded. (b) After you install all the jumper wires in the system, check the resistance with an ohmmeter. The maximum allowable resistance must show 0.01 ohm. (8) Attach the washers to the fuel vent line valves. (9) Attach the fuel vent line valves to the fuel bulkhead. Replace the O-rings and nuts. (10) Attach and tighten the tube nut to the fuel vent line valve. (11) Seal and clean the fuel tanks. (12) Install the wing tip assemblies. Refer to Chapter 57, Wings - Removal/Installation.
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Fuel Vent System Installation Figure 201 (Sheet 1)
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Fuel Vent System Installation Figure 201 (Sheet 2)
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Fuel Vent System Installation Figure 201 (Sheet 3)
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Fuel Vent System Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (13) Install the lower wing access covers, fuselage access cover, and sidewall panels. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (14) Put fuel in the fuel tanks. Refer to Chapter 12, Fuel - Servicing. (15) Do a check for leaks. 3.
4.
5.
Union, Tubes and Cross Removal/Installation A.
Remove the Union, Tubes and Cross (Refer to Figure 201). (1) Loosen the tube nuts and disconnect tubes from the union. (2) Remove the nut and ring from the union. (3) Discard the ring and disconnect the union from the fuel bulkhead. (4) Loosen the clamps and disconnect the tubes and hoses from the cross. Discard hoses. (5) Disconnect the hoses from the tubes and cross. Remove the cross. (6) Remove the bolts. Disconnect the upper shield halves and lower shield halves from the floorboard rib. (7) Remove and discard the seals. (8) Remove the nut from the screw and disconnect the vent line clamps from the fuel line clamps.
B.
Install the Union, Tubes and Cross (Refer to Figure 201). (1) Locate the seals on the tubes. Attach the upper and lower shield halves to floorboard ribs. Replace bolts. (2) Attach the union to the fuel bulkhead. (3) Install the O-ring and nut. (4) Attach and tighten the tube nuts on the union. (5) Attach the hoses to the cross, tubes and tighten the clamps. (6) Attach and tighten the tube nut on the union. (7) Attach the hoses to the tubes and tighten the clamps. (8) Install the screws in the clamps, attach and tighten nuts on the screws. (9) Install the union and tighten the tube nuts. (10) Attach the retainers to the root ribs and install the bolts.
Seals and Reservoir Vent Lines Removal/Installation A.
Remove the Seals and Reservoir Vent Lines (Refer to Figure 201). (1) Remove the screws from the lower shield half and upper shield half. (2) Loosen the clamps and remove the vent tube from the hose. Remove and discard the hose. (3) Loosen the tube nut. (4) Disconnect the tube from the manifold.
B.
Install the Seals and Reservoir Vent Lines (Refer to Figure 201). (1) Attach the tube to the manifold and tighten the nut. (2) Install the hose on the reservoir. (3) Attach the tube to the hose and tighten the clamps. (4) Attach the seal to the tube, locate upper and lower shield halves and install the screws.
Fuel Vent Line Float Valve Test NOTE: A.
The fuel vent line float valve can become plugged causing possible fuel starvation of engine. The following procedure must be used to check the function of the fuel vent valve.
Do a valve test of the fuel vent line float. (1) Attach a rubber tube end to one wing tip vent line. (2) Do a check to make sure the fuel caps are correctly installed and the fuel selector valve is turned off. (3) Put a plug in the 0.040 inch diameter hole in the wing tip vent line located 6.50 inches from the end of the part.
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MODEL 208 MAINTENANCE MANUAL (4)
(5) (6)
Blow into the tube to slightly pressurize fuel tank (if air can be blown into tank, vent lines are open and float valve is not plugged). (a) If air cannot be blown into the fuel tank, remove the vent lines and fuel vent line valve from the system. Clean the vent lines. 1 2 Do a functional check of the fuel vent line valve. If the float valve sticks or the relief valve does not open at 0.7 negative PSI or 1.0 3 PSI, replace the fuel line vent valve. Repeat steps 5.A.(1) through 5.A.(4)(a)3 for the opposite fuel tank. Remove the tube and plug from the wing tip vent line.
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MODEL 208 MAINTENANCE MANUAL FUEL VENTILATION SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fuel ventilation system in a serviceable condition.
Task 28-10-03-710 2.
Fuel Vent Line Float Valve Operational Check NOTE:
The fuel vent line float valve can become clogged and cause possible fuel starvation of the engine. The following procedure must be used to make sure that the valve operates correctly.
NOTE:
The operational check of the left and the right fuel vent line float valve is typical.
A.
General (1) This task gives the procedures to do a functional check of the fuel vent line float valve.
B.
Special Tools (1) Tube (2) Plug
C.
Access (1) None
D.
Do a Operational Check of the Fuel Vent Line Float Valve (Refer to Figure 601). (1) Make sure that the fuel selector valve is turned off. (2) Attach a rubber tube to the end of the wing tip vent line. (3) Make sure that the fuel caps are installed correctly. (4) Put a plug in the 0.040 inch (1.01 mm) diameter hole in the wing tip vent line. (5) Blow into the tube to give a small amount of pressurization into the fuel tank. (a) If you can blow air into the fuel tank, the vent lines are open and float valve is not clogged. (b) If you can not blow air into the fuel tank, do the Fuel Vent Line Float Valve Test to examine if the vent line is plugged and/or the float valve is stuck at the closed position. Refer to Fuel Ventilation System - Maintenance Practices. (6) Remove the tube and plug from the wing tip vent line.
E.
Restore Access (1) None End of task
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Fuel Ventilation System Figure 601 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FUEL LINES, VALVES AND FILTERS - MAINTENANCE PRACTICES 1.
General A.
2.
Fuel lines, valves and filters maintenance practices consist of removal/installation and test of components.
Screen and Fuel Shutoff Valve Removal/Installation A.
Remove Screen and Fuel Shutoff Valve (Refer to Figure 201). (1) Remove lower wing access covers to gain access to fuel tanks and that segment of system located between fuel tank and fuselage. (2) Remove access cover located on bottom of fuselage below reservoir, gain access to fuselage through cargo pod. (3) Remove sidewall covers aft of pilot's and copilot's doors. (4) Defuel and purge fuel tanks. (5) Detach adapter (11) from fuel shutoff valve (4). Remove adapter (11) and O-ring (9) from fuel tank, discard O-ring (9). (6) Loosen clamp (12) and detach screen (13) from adapter (11). Discard screen (13) if blocked or damaged. (7) Detach tube (2) from shutoff valve (4). (8) Remove cotter pin (24), washer (23), and pin (22) from link (24A) and handle (24B). (9) Remove fuel shutoff valve (4) from fuel bulkhead (10). NOTE:
B.
3.
Cap all open fuel lines during removal of system components.
Install Screen and Fuel Shutoff Valve (Refer to Figure 201). (1) Replace fuel shutoff valve (4) and O-ring (9) on fuel bulkhead (10). (2) Replace screen (13) and clamp (12) on adapter (11). Install and tighten adapter on fuel shutoff valve (4). (3) Install and tighten aft fuel line (2) on fuel shutoff valve (4). (4) Attach link (24A) to handle (24B), replace pin (22), washer (23), and cotter pin (24). (5) Do an operational check of the wing area fuel selector valves, linkages, and switches. Refer to Task 28-21-00-710.
Lines and Filter Removal/Installation A.
Remove Lines and Filter (Refer to Figure 201). (1) Cut sta-straps (33) from covers (34), on early 208 Models, loosen clamps (40), remove hoses (39) from drain (41), discard hoses. On later 208 Models, remove hoses from grommets (44) and covers (34), discard hoses, or, remove screws (44D) and detach cover (44A). Remove clamps (40) and remove hoses (39) from grommets (44); discard hoses. (2) Remove screws (28) and (30), nuts (29) and (31), and clamps (26) and (27) from forward motive flow line (25) and fuel supply line (43). (3) Loosen tube nuts and detach unions (35) from motive flow lines (25) and (37) and fuel supply lines (38) and (43). (4) Detach forward fuel supply line (43) from firewall shutoff valve (48). Detach forward motive flow line (25) from elbow (62). (5) Remove screws (88) and detach upper and lower shield halves (85) and (86) from supports (91) and (92). (6) Remove seals (87) from aft motive flow and fuel supply lines (37) and (38). Discard seals (87). (7) Detach aft fuel supply line (38) from manifold (89). (8) Detach aft motive flow line (37) from motive line elbow (90). (9) Disconnect control (65) from handle (66), loosen nuts (49) and (50), and unscrew firewall shutoff valve (48) from fuel filter (52), removing nut (49) and (50) to remove valve; discard O-ring (51). (10) Detach hose (56) from elbow (55), detach elbow (55) from reducer (54). (11) Detach drain valve (64) from drain port (63). (12) Remove bolts (58), (60) and lockwashers (59), (61) from bracket (57), detach bracket (57) from fuel filter (52) and mounting brackets (46), and remove filter (52).
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 1)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 2)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 3)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 4)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 5)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 6)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 7)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 8)
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Fuel Lines, Valves, and Filters Installation Figure 201 (Sheet 9)
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MODEL 208 MAINTENANCE MANUAL B.
Install Lines and Filter (Refer to Figure 201). (1) Install and tighten aft motive flow line (37) to motive line elbow (90). (2) Install and tighten aft fuel supply line (38) to manifold (89). (3) Replace seals (87) on tubes (38) and (37). Install upper and lower seal halves (85) and (86) on reservoir support (81). Replace screws (78). (4) Install and tighten forward and aft motive flow lines (25), (37), forward and aft fuel supply lines (43) and (38) to unions (35). NOTE:
(5) (6)
(7) (8) (9) (10) (11) (12) (13) (14)
Mating surfaces of covers (34) and mating surfaces of covers to tubes (25), (37), (38), and (43) shall be cleaned and primed with Dow-Corning 1200 primer or equivalent. Let primer air dry for at least one-half hour. Seal mating surfaces with TBS-758 thermal coating or equivalent. Pot life of mixed material is four hours.
Install covers (34) and hoses (39) on motive flow lines (25) and (37), fuel supply lines (38) and (43) and drain (41). Attach sta-straps (33) to covers (34). On early 208 Models, attach hoses (39) to covers (34) and drain (41). Replace clamps (40). On later 208 Models, attach hoses (39) to covers (34). Route hoses through grommets (44) and belly skin of airplane, or, attach hoses (39) to covers (34). Route hoses (39) through grommets (44), cover (44A), clamp (44B), and belly of cargo pod (44C). Attach cover (44A) and clamp (44B) to cargo pod with screws (44D) and (44E). Attach clamp (27) to forward support clamp (32) and replace screw (30) and nut (31). Align clamps (26) and (27) and replace screw (28) and nut (29). Place firewall shutoff valve (48) to firewall (45); install nut (49). Install and tighten forward fuel supply line (43) through firewall shutoff valve (48). Attach control (65) to handle (66) on firewall shutoff valve (48). Connect forward motive flow line (25) to elbow (62), tighten tube nut. Replace nut (50) and O-ring (51) on firewall shutoff valve (48). Install fuel filter (52) on firewall shutoff valve (48). NOTE:
Turn fuel filter (52) clockwise on firewall shutoff valve boss until drain valve (64) may be installed through hole in bottom of engine cowl. Check alignment of attaching holes in bracket (57) with fuel filter (52) and firewall mounting brackets (46). Fuel filter may have to be turned in either direction on firewall shutoff valve boss in order for all three points to be in line.
(15) After locating fuel filter (52), tighten nut (50) on 0-ring (51) and filter (52). (16) Attach bracket (57) to firewall mounting brackets (46) and fuel filter (52). Install and tighten bolts (58), (60) and lockwashers (59) and (61). (17) Install and tighten drain valve (64) on drain port (63). (18) Install O-ring (53) on reducer (54), install and tighten reducer (54) on fuel filter (52). (19) Install and tighten hose (56) on elbow (55), install and tighten elbow (55) on adapter (54). NOTE:
Ensure fuel tank and reservoir seals, adapters, unions, etc. have been sealed.
(20) Replace all wing and fuselage access covers, and sidewall covers. (21) Refuel system and check for leaks. 4.
Lines and Filters Removal/Installation (208B Airplanes) A.
Remove Lines and Filters (Refer to Figure 201). (1) Remove screws (44D) and remove cover (44A). (2) Remove clamps (93) and (40) from hoses (39); discard hoses. (3) Cut sta-straps (33) from covers (34). (4) Remove clamps (96) from hoses (97) and remove drain line (94); then remove drain line (95). (5) Remove screws (28) and (30), nuts (29) and (31), and clamps (27) and (26) from forward motive flow line (25) and forward supply line (43). (6) Loosen tube nuts and detach unions (35) from motive flow lines (25) and (37) and fuel supply lines (38) and (43).
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) B.
Detach forward fuel supply line (43) from firewall shutoff valve (48). Detach forward motive flow line (25) from elbow (62). Remove screws (88) and detach upper and lower shield halves (85) and (84) from supports (91) and (92). Remove seals (87) from aft motive flow and fuel supply lines (37) and (38). Discard seals (87). Detach aft fuel supply line (38) from manifold (89). Detach aft motive flow line (37) from motive line elbow (90). Disconnect control (65) from handle (66), loosen nuts (49) and (50) and unscrew firewall shutoff valve (48) from fuel filter (52), removing nuts (49) and (50) to remove valve; discard 0-ring (51). Detach hose (56) from elbow (55) and detach elbow (55) from reducer (54). Detach drain valve (64) from drain port (63). Remove bolts (58) and (60) and lockwashers (59) and (61) from bracket (57), from fuel filter (52), and mounting brackets (46), and remove filter (52). Remove elbow (55) from reducer (54). Remove reducer (54) from filter (52) and discard O-ring (53).
Install Lines and Filters (Refer to Figure 201). (1) Install and tighten aft motive flow line (37) to motive line elbow (90). (2) Install and tighten aft fuel supply line (38) to manifold (89). (3) Replace seals (87) on lines (38) and (37). Install upper and lower seal halves (85) and (86) on reservoir support (81). Install screws (78). (4) Install and tighten forward and aft motive flow lines (25) and (37), and forward and aft fuel supply lines (43) and (38) to unions (35). NOTE:
Mating surfaces of covers (34) and mating surfaces of covers to lines (25), (37), (38), and (43) shall be cleaned and primed with Dow-Corning 1200 primer or equivalent. Let primer air-dry for at least one-half hour. Seal mating surfaces with TBS-758 thermal coating or equivalent. Pot life of mixed material is four hours.
(5) (6) (7)
Install covers (34) on motive flow lines (25) and (37), and fuel supply lines (38) and (43). Install drain line (94). Install drain line (95) and connect drain lines (95) and (94) to covers (34) using hoses (97) and clamps (96). (8) Install drain hoses (39) through grommets (44) and connect to drain lines (95) and (96) using clamps (93). Connect to drain line (41) using clamps (40). (9) Install cover (44A) using screws (44D). (10) Install clamp (26) on forward motive flow line and clamps (27) on forward fuel supply line (43) using screws (28) and (30) and nuts (29) and (31). (11) Install firewall shutoff valve (48) through firewall (45) and install nuts (49) and (50) on shutoff valve; do not tighten. (12) Install new O-ring (51) and screw filter (52) on firewall shutoff valve (48). NOTE:
(13) (14) (15) (16) (17) (18) (19) (20) (21) (22)
Turn fuel filter (52) clockwise on firewall shutoff valve boss until drain valve (64) may be installed through hole in bottom of engine cowl. Check alignment of attaching holes in bracket (57) with fuel filter (52) and firewall mounting brackets (46). Fuel filter may have to be turned in either direction on firewall shutoff valve boss in order for all three points to be in line.
After locating fuel filter (52), tighten nut (50) on O-ring (51) and filter (52). Connect forward motive flow line (25) to elbow (62) and tighten. Connect forward fuel supply line (43) to firewall shutoff valve (48) and tighten. Attach brackets (57) on firewall mounting brackets (46) using lockwashers (59) and bolts (58). Connect bracket (57) to filter (52) using lockwashers (61) and bolts (60). Install reducer (54) in filter (52) using new O-ring (53). Install elbow (55) on reducer (54). Connect hose (56) to elbow (55) and tighten. Connect hose to elbow (62). Connect control (65) to firewall shutoff valve arm (66) and adjust.
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MODEL 208 MAINTENANCE MANUAL (23) Connect drain valve (64) to drain port (63). NOTE:
Make sure fuel tank and reservoir seals, adapters, unions, etc. have been sealed.
(24) Do an operational check of the firewall shutoff valve. Refer to Task 28-21-00-710. (25) Replace all wing and fuselage access covers, and sidewall covers. (26) Refuel system and check for leaks. 5.
Firewall Mounted Fuel Filter Servicing A.
The fuel filter is found on the forward face of the firewall near the bottom of the engine compartment. You can use the 20-micron filter element again in the filter. The element can be cleaned in naptha and dried with a low-pressure, clean air source. NOTE:
The low pressure clean air source must be less than 30 PSI and oil-free.
CAUTION: Make sure that the compressed air pressure is less than 30 psi. This will help prevent damage to the filter element. CAUTION: Make sure that you do not use sharp or hard objects on the mesh surfaces. This will help prevent damage to the mesh. CAUTION: Make sure that you do not separate the individual filter elements from the perforated tube. B.
A red warning button is found on the top of the filter. If the button is out, it shows the filter element is clogged, and the filter is bypassed. When this occurs, the filter must be disassembled and the element cleaned.
CAUTION: If during any inspection, the red warning button on top of the fuel filter is out, do not fly the airplane until the source of the fuel contamination is found and stopped. Disassemble the filter and clean the screen. Make sure that you do an inspection of the interior of the reservoir tank, the interior of the wing tanks, and the engine fuel system. Refer to the Pratt and Whitney Maintenance Manual for the engine fuel system maintenance and inspection procedures. 6.
Fuel Filter Disassembly/Assembly A.
Disassemble Fuel Filter (Refer to Figure 201). (1) Remove drain valve (79) and discard O-ring (78). (2) Cut safety wire and remove drain port (77). Discard O-ring (76). (3) Detach bowl (75) from body (70). Discard O-ring (72). (4) Loosen and remove lock nuts (74), detach filter element (73) from threaded rod (80).
B.
Assemble Fuel Filter (Refer to Figure 201). NOTE: (1) (2) (3) (4)
Before installing new O-rings (72), (76), and (78), lubricate with light general purpose grease. After installation, wipe off excess grease.
Attach filter element (73) to threaded rod (80), install and tighten lock nuts (74). Replace O-ring (72) bowl (75) and O-ring (76). Install, tighten and safety wire drain port (77). Replace O-ring (78) on drain valve (79). Install and tighten drain valve (79). Push warning button (71) down until it contacts magnetic lock (flush position).
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MODEL 208 MAINTENANCE MANUAL FUEL LINES, VALVES AND FILTERS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fuel lines, valves, and filters in a serviceable condition.
Task 28-21-00-710 2.
Firewall and Wing Fuel Shutoff Valves Operational Check A.
General (1) This task gives the information needed to do an operational check of the firewall and the wing fuel shutoff valves.
B.
Special Tools (1) None
C.
Access (1) Remove wing access panels 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (2) Open (unzip) the fabric headliner (passenger) or remove the hard shelled headliner (cargo). Refer to Chapter 25, Cabin Upholstery - Maintenance Practices.
D.
Do an Operational Check of the Firewall Shutoff Valve. (1) Remove unwanted sealant from the wing access panels and the airframe structure. (2) Examine the firewall shut off valve for condition, signs of damage, leakage, security of attachment, and freedom of operation. (3) Examine the control cable for security of attachment at the valve, handle, and cable housing clamp. (4) Operate the valve control and examine for freedom of movement. (a) Make sure that the valve handle travels to the stop screw when the control is placed to the ON and OFF position. (5) Apply a release agent to the access cover and structure and install the cover with a fuel cell approved sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices. (6) Examine the firewall fuel shutoff control cable housing for condition, routing, and security. (7) Examine the operation of the lock knob for positive locking of control. (8) Examine the operation of the control for positive shutoff of fuel at the firewall. (9) With the control pulled to the OFF position, open the drain valve on the firewall fuel filter. (a) Make sure that the fuel drains until the filter is empty.
E.
Do an Operational Check of the Wing Area Fuel Selector Valves, Linkages, and Switches. (1) Examine the forward and the aft shutoff valves for condition, security, signs of damage, and leakage. (2) Examine the fuel selector cables and connecting linkage for condition, security, and signs of damage. (a) Make sure that the rubber boot is correctly seated at the end of the tele-flex cable housing. (b) Make sure that all roll pins are correctly installed and safety wired.
WARNING: Before you open the selector valves, make sure that the fuel system is closed and will not leak fuel. (3)
Operate the selectors to the ON and OFF positions. (a) Make sure that there is freedom of movement and that the valve handle contacts the stops (index screws). 1 If the valve arms do not contact the stops with the selector at the OFF or ON position, adjust the linkage. Refer to Fuel Selector Controls - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL (4) (5) (6)
Examine the fuel lines from the selector valve to the fuselage for leaks, condition, security, and correct bonding of the metal lines. Examine the rubber interconnect hoses that connect each supply line for condition, security, and excessive cold flow around the hose clamps. Examine the fuel selector off warning system switches, connectors, actuators, and electrical wiring for condition and security of installation.
F.
Restore Access (1) Close (zip) the fabric headliner (passenger) or install the hard shelled headliner (cargo). Refer to Chapter 25, Cabin Upholstery - Maintenance Practices. (2) Install wing access panels 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task Task 28-21-00-711 3.
Firewall Fuel Shutoff Valve Control Operational Check A.
General (1) This task gives the procedures to do the operational check of the firewall mounted fuel shutoff valve. The firewall mounted fuel shutoff valve is located on the cockpit side of the firewall on the lower left side.
B.
Special Tools (1) B2 sealant or equivalent
C.
Access (1) None
D.
Do a Detailed Inspection of the Firewall Fuel Shutoff Valve Control. Refer to Fuel Lines, Valves, and Filters - Maintenance Practices, Figure 201. (1) Remove the shutoff valve access cover. (a) Use a parting tool to loosen the B2 sealant. (b) Clean the old sealant from the cover and the airframe structure. (2) Examine the fuel firewall shutoff valve for condition, signs of damage, leakage, security of installation, and freedom of operation. (3) Examine the control cable for security of attachment at the firewall shutoff valve and the control handle. (a) Make sure that the cable housing clamps are installed correctly. (4) Examine the control cable housing for condition, correct routing, and security. (5) Examine the motive flow fuel line located adjacent to the shutoff valve for condition, leaks, and security. (6) Examine the main fuel line and motive flow fuel line sealing grommets for condition and security. (7) Examine the cavity drain hole for obstructions.
E.
Do an Operational Check of the Firewall Fuel Shutoff Valve Control. (1) Operate the valve control and examine for freedom of movement and that the valve handle travels to the stop screw when the control is set to the ON and OFF position. (2) Examine the operation of the lock knob for positive locking of the control. (3) Examine the operation of the control for positive fuel shutoff at the firewall. (a) With the control pulled OFF, open the drain valve on the firewall fuel filter. 1 Make sure that fuel drains until the filter is empty. (4) Install the shutoff valve access cover. (a) Apply a release agent to the panel and structure. (b) Apply an approved sealant to the panel and the structure. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (c) Install the shutoff valve access cover.
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MODEL 208 MAINTENANCE MANUAL F.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL FUEL RESERVOIR - MAINTENANCE PRACTICES 1.
General A.
The fuel reservoir is located under the floor between stations 168.70 and 194.40 on the Model 208 airplanes, and between stations 188.70 and 214.40 on Model 208B airplanes. An auxiliary fuel pump, ejector pump and a low fuel level switch are located in the reservoir. The low fuel level switch actuates a red warning light located in the annunciator panel, labeled "RESERVOlR LOW". NOTE:
2.
Airborne fuel pump (2C6-8) may be used as an alternate for Dukes fuel pump (1613-00-1) (19). Airborne pump (2C6-8) does not require a pump seal drain line. If the airplane does not have a cargo pod, remove drain line (30A) and adapter (30B). If airplane has a cargo pod remove drain lines (26A), (58), elbow (28A), and adapter. In either case after removing adapter from reservoir, plug port from which adapter was removed.
Reservoir Components Removal/Installation A.
Remove Reservoir Components (Refer to Figure 201). (1) Remove access cover in fuselage below reservoir. (2) Defuel and purge fuel system. (3) Remove bolts (28) and lockwashers (29) from plate (27). Detach cover (27) and gasket (26) from reservoir (1). (4) Detach tube (9) from adapter (8) and check valve (10) and remove tube (9) from reservoir (1). (5) Remove nut (6) and O-ring (5), detach ejector pump (4) from reservoir (1). Discard O-ring (5). (6) Remove adapter (8) and O-ring (7) from ejector pump (4). Discard O-ring (7). (7) Remove nut (30), detach check valve (10) and O-ring (11) from reservoir (1). Discard O-ring (11). (8) Loosen nut (13), remove adapter (14), nut (13), and O-ring (12) from reservoir (1). (9) Remove bolts (22) and lockwashers (21), detach auxiliary fuel pump (19) and gasket (20) from reservoir (1). Discard gasket (20) from reservoir (1). Discard gasket (20). (10) Remove nut (15), retainer (16), and O-rings (17) from float switch (18). Discard O-rings (17) and detach float switch (18) from reservoir (1). (11) Remove auxiliary pump seal drain line (30A) and adapter (30B) from reservoir (1). (12) Disconnect plug (32) from connector (31), detach pressure switch (36A) from manifold (36B.) (13) Disconnect tubes (45), (47), and (51) from manifold (36B). Remove manifold (36B) from reservoir (1). (14) Remove check valve adapters (46), reducer (50), O-rings (52) and (53) from manifold (36B). Discard O-rings (52) and (53). (15) Loosen clamps (48), detach hose (49) from fuel line (47) and adapter (54). (16) Remove screws (39) and washers (40) and detach swing check valves (38) from reservoir (1). NOTE:
If nutplate (37) is damaged, or fuel leaks are noted, it must be removed, replaced and resealed.
(17) Remove nut (23), detach drain valve (25) and O-ring (24) from plate (27). Discard O-ring (24) B.
Install Reservoir Components (Refer to Figure 201). (1) Attach drain valve (25) and O-ring (24) to plate (27). Replace and tighten nut (23). (2) Attach swing check valves (38) to reservoir (1), replace washers (40) and screws (39). (3) Attach hose (49) to tube (47) and adapter (54). Install and tighten clamps (48). (4) Replace O-rings (52) and (53), check valve adapters (46), and reducer (50) on manifold (36B). (5) Connect tubes (45), (47), and (51) to manifold (36B). (6) Connect pressure switch (36A) to manifold (36B). Attach plug (32) to connector (31). (7) Replace adapter (30B) in reservoir (1). Attach auxiliary pump seal drain line (30A) to adapter (30B). (8) Attach float switch (18) and O-ring (17) to reservoir (1). Replace O-ring (17), retainer (16), and nut (15). Tighten nut (15). (9) Attach auxiliary fuel pump (19) and gasket (20) to reservoir (1). Replace Iockwashers (21) and bolts (22).
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MODEL 208 MAINTENANCE MANUAL
Fuel Reservoir Components Installation Figure 201 (Sheet 1)
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Fuel Reservoir Components Installation Figure 201 (Sheet 2)
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Fuel Reservoir Components Installation Figure 201 (Sheet 3)
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Fuel Reservoir Components Installation Figure 201 (Sheet 4)
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Fuel Reservoir Components Installation Figure 201 (Sheet 5)
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Fuel Reservoir Components Installation Figure 201 (Sheet 6)
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MODEL 208 MAINTENANCE MANUAL (10) Attach adapter (14), nut (13), and O-ring (12) to reservoir (1) and electric fuel pump (19). Tighten nut (13). (11) Attach O-ring (7) and adapter (8) to ejector pump (4). Tighten adapter (8). (12) Attach O- ring (11), check valve (10), and nut (30) to reservoir (1). Tighten nut (30). (13) Attach tube (9) to adapter (8) and check valve (10). Tighten tube nuts. (14) Attach gasket (26) and plate (27) to reservoir (1). Replace lockwashers (29) and bolts (28). Tighten bolts (28) in sequence to 18-20 inch-pounds. (15) Refuel system and check for fuel leaks. (16) Replace fuselage access cover below reservoir.
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MODEL 208 MAINTENANCE MANUAL FUEL SELECTOR CONTROLS - MAINTENANCE PRACTICES 1.
General A.
2.
Fuel flow from left and right fuel tanks is selected by the following components: fuel selector, two knobs, two control cables, two interconnects and requried attaching parts. Fuel selector panel is labeled: FUEL TANK SELECTORS; ON LEFT OFF and ON RlGHT OFF. Knobs and selector panel are located in the overhead console between pilot and co-pilot. The interconnects in each wing adjacent to fuel shutoff valves. They are attached to the shutoff valves and cables routed through the wings, above overhead console to knobs in the overhead console. Whenever a knob is placed in a position ON or OFF, or any intermediate setting, corresponding interconnect moves both shutoff valves simultaneously to that setting.
Fuel Selector Controls Removal/Installation A.
Remove Fuel Selector Controls (Refer to Figure 201). (1) Remove access panel on bottom of wings adjacent to fuselage, and open zippers in headliner. (2) Remove safety wire (23), screws (19) and plates (18). Detach knobs (17). (3) Remove screws (16), detach controls cover (15) from bracket (20). (4) Remove cotter pin (14), washer (13), and pin (12). Detach control terminal (4) from lever (10). (5) Remove safety wire and remove roll pin (11) and detach lever (10) from shaft (9). (6) Remove screws (7), detach retainer (6), shaft (9), and washers (8) and (22) from bracket (20). (7) Remove safety wire (21) from brackets (3). Detach cable (1) from support (2). (8) Remove cotter pin (37), washer (36), and pin (35). Detach terminal (35) from interconnect link (40). (9) Remove cotter pins (34), washers (33), and pins (32). Detach interconnect (29) from handles (30).
B.
Install Fuel Selector Controls (Refer to Figure 201). NOTE:
(1) (2) (3) (4) (5)
Before installing fuel selector controls handles (30) on fuel shutoff valves (31) must be correctly indexed by moving both handles (30) forward until they firmly contact index screws (42).
Attach interconnect (29) to handles (30). Replace pins (32), washers (33), and cotter pins (34). Attach cable (1) to supports (2). Replace safety wire (21) on supports (2) and brackets (3). Install shaft (9) and washer (8) on retainer (6). Position washer (22), shaft (9), washer (8), and retainer (6) on bracket (20). Replace screws (7). Attach lever (10) to shaft (9). Align holes in shaft (9) and lever (10), replace roll pin (11) and safety wire. Attach terminal (4) to lever (10). Replace pin (12). NOTE:
(6) (7)
Attach controls cover (15) to bracket (20). Replace screws (16). Attach knobs (17) to shafts (9). Replace plates (18), screws (19), and safety wire (23). NOTE:
(8)
Place knob (17) in the ON position before proceeding to the next step.
Attach terminal (26) to aft interconnect link (40). NOTE:
(9)
Do not replace washer (13) and cotter pin (14), since cable (1) may have to be shortened or lengthened during indexing.
Ensure that handles (30) are fully forward and knob (17) is located in ON position. lt may be necessary to loosen nut (27) and turn terminal (26) in or out on fuel selector cable shaft to align holes in terminal (26) with hole in aft interconnect link (40). lf more than two turns in either direction are required, remove pin (12), loosen nut (5) and turn terminals (4) and (26) equally.
Replace pin (35), washer (36), and cotter pin (37) and tighten nut (27).
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MODEL 208 MAINTENANCE MANUAL
Fuel Selector Controls Figure 201 (Sheet 1)
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Fuel Selector Controls Figure 201 (Sheet 2)
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Fuel Selector Controls Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (10) lF NECESSARY - Replace pin (12), washer (13), and cotter pin (14), tighten nut (5). (11) Functional check system by placing handle (17) in ON and OFF positions and verifying that handles (30) contact corresponding stops (42) and (44). (12) Replace access panels on bottom of wings adjacent to fuselage. Close zippers in headliner.
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MODEL 208 MAINTENANCE MANUAL FUEL SELECTOR CONTROLS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fuel selector controls in a serviceable condition.
Task 28-23-00-640 2.
Wing Shutoff Valve Linkage Lubrication A.
General (1) This task gives the procedures to do the lubrication of the wing shutoff valve linkage.
B.
Special Tools (1) LPS 1 Lubricant, or Equivalent
C.
Access (1) Remove lower wing access panels 511AB and 611AB to get access to the wing shutoff valves. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do a Lubrication of the Wing Shutoff Valve Linkage (1) Apply lubricant to the pins that connect the interconnect link to the valve handle on all four valves. (2) Apply lubricant to the exposed part of the shutoff cable. (3) Operate the fuel selectors several times. (4) Wipe off any unwanted lubricant.
E.
Restore Access (1) Install lower wing access panels 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task
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MODEL 208 MAINTENANCE MANUAL INDICATING - DESCRIPTION AND OPERATION 1.
General A.
2.
This section covers system components utilized to indicate fuel quantity.
Description and Operation A.
Each fuel tank is equipped with four float-operated, variable-resistance fuel quantity transmitters in the following locations; (1) at the inboard end of fuel tank, (2) at the outboard end of fuel tank, (3) at the center inboard end of fuel tank, (4) at the center outboard end of fuel tank. A low fuel level switch is located adjacent to the inboard fuel quantity transmitter. Fuel quantity indicators (one for each fuel tank) are located on the upper right side of the instrument panel. The fuel quantity transmitters are connected to the electrically operated fuel quantity indicators in series. Because of dihedral angle of the wing, as fuel level drops resistance of outboard fuel quantity transmitter(s) decreases first from maximum causing a corresponding decrease in fuel gage indications.
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MODEL 208 MAINTENANCE MANUAL FUEL QUANTITY INDICATING SYSTEMS - MAINTENANCE PRACTICES 1.
General A.
Fuel quantity indicating systems maintenance practices consist of component removal/installation.
CAUTION: When performing resistance tests of the fuel quantity wiring or fuel probe transmitters, use a digital ohmmeter only. Some older analog ohmmeters may introduce high current, which will destroy fuel probe transmitters. 2.
3.
4.
Fuel Quantity Indicator Removal/Installation A.
Remove Fuel Quantity Indicator (Refer to Figure 201). (1) Remove screws (5) from instrument panel (3). Detach fuel quantity indicators (1) or (2) from instrument panel. (2) Disconnect fuel quantity indicators (1) or (2) from electrical connectors (6).
B.
Install Fuel Quantity Indicator (Refer to Figure 201). (1) Attach fuel quantity indicators (1) or (2) to electrical connectors (6). (2) Attach fuel quantity indicators (1) or (2) to instrument panel. (3) Replace screws (5) in instrument panel (3).
Low Fuel Level Switch Removal/Installation. A.
Remove Low Fuel Level Switch (Refer to Figure 201). (1) Defuel and purge fuel system. (2) Clean access cover sealant from mating surfaces of fuel bulkhead (7), connector (10), jamnut (11), and cover (14). (3) Remove screws (15) and detach cover (14) from fuel bulkhead (7). (4) Loosen jamnut (11) and connector (10) and remove low fuel level switch (9) at bracket (17).
B.
Install Low Fuel Level Switch (Refer to Figure 201 ). (1) Attach low fuel level switch (9) to bracket (17) and replace jamnut (11) and connector (10). (2) Attach cover (14) to fuel bulkhead (7) and replace screws (15). (3) Apply access cover sealant to jamnut (11) connector (10) sides, and bottom surfaces of cover (14). (4) Refuel system.
Reservoir Low Fuel Level Switch Removal/Installation A.
Remove Reservoir Low Fuel Level Switch (Refer to Fuel Reservoir - Maintenance Practices, Figure 201). (1) Remove safety wire and turn both fuel selector valves to the position. (2) Remove drain valve (25) and O-ring (24). Discard O-ring. Drain fuel in suitable containers. (3) Remove bolts (28) and lockwashers (29). Detach plate (27) and gasket (26) from fuel reservoir (1). Discard gasket (26). (4) Disconnect switch wires (30C) from connector, and remove nut (15) and washer (16). Separate nut and washer from wires (30C). (5) Remove switch (18), O-ring (17) and switch wires through port in bottom of fuel reservoir (1). Discard O-ring.
B.
Install Reservoir Low Fuel Level Switch (Refer to Fuel Reservoir - Maintenance Practices, Figure 201). (1) Replace switch wires (30C), O-ring (17) and switch (18) in fuel reservoir (1). (2) Replace new O-ring (17), washer (16) and nut (15) on switch wires (30C). Attach switch wires (30C) to connector. Replace washer (16) and nut (15) on threaded stem of switch (18). Tighten nut (15). (3) Replace new gasket (26) on fuel reservoir (1). Attach plate (27) to gasket (26) and fuel reservoir (1). (4) Install lockwashers (29) and bolts (28). Tighten bolts in a cross sequence pattern.
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MODEL 208 MAINTENANCE MANUAL
Fuel Quantity Indicating System Figure 201 (Sheet 1)
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Fuel Quantity Indicating System Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (5) (6) 5.
Replace new O-ring (24) and drain valve (25). Tighten drain valve. Turn both fuel selector valves on and safety wire. Check system for leaks.
Wing and Reservoir Low Fuel Level Switches Test A.
Test Wing and Reservoir Low Fuel Level Switches. (1) Place airplane in flight attitude (wings level, nose up 1 degree 30 minutes - refer to Chapter 8, Leveling and Weighing. (2) Defuel system. (3) Turn both fuel selector valves off. (4) Check that LEFT FUEL LOW, RIGHT FUEL LOW AND RESERVOIR low annunciator panel lights are on. (If any of the lights are not on, run continuity check of the circuit and replace bulbs if required.) (5) Partially fill each fuel bay with 20 gallons of measured fuel, continue to add fuel, if necessary until each bay contains 30 gallons of fuel. NOTE:
(6) (7) (8)
If LEFT and RIGHT annunciator lights do not shut off between the 20 and 30 gallon levels, drain fuel, purge system and replace either or both switches. After replacing defective switch(es), repeat steps (1) thru (5).
Tum both fuel selector valves ON until RESERVOIR low light turns off, then turn both fuel selector valves OFF. Slowly drain fuel from reservoir until light turns ON. Drain remaining fuel from reservoir, and measure quantity, should check from 1.95 to 2.15 gallons. NOTE:
If quantity of measured fuel exceeds 2.25 gallons, or is less than 1.75 gallons; remove and replace the reservoir fuel low level switch. Repeat step (8) to verify continuity and accuracy of replacement.
(9) Remove jacks (if utilized) and refuel airplane. (10) Check fuel system for leaks and replace any covers or panels removed during testing procedures. 6.
Fuel Quantity Indication System Calibration A.
Calibrate Fuel Quantity Indication System (Refer to Figure 202). NOTE:
Always use a screwdriver with an insulated shank when calibrating the fuel system, also, use a quality ohmmeter while conducting continuity checks of fuel system.
(1) (2)
Remove fuel system access covers from bottom of wings. Place airplane in level flight attitude (wings level, 1 degree 30 minutes nose up waterline, refer to Chapter 8, Leveling and Weighing). (3) Place fuel shutoff valves in OFF position. (4) Drain fuel tanks completely. (5) Place 2.5 gallons of fuel in each fuel tank and turn electrical power on. (6) Adjust null trimpot on each gage carefully so that needle is completely within red zone. (7) After adjusting null trimpot as noted in step (6), turn electrical power off, locate and disconnect electrical wire from center post of inboard fuel transmitter on one wing. (8) Insert a 230-ohm dummy Ioad between this wire and airplane ground, turn electrical power on. (9) Carefully adjust gain trimpot on fuel gage to center the needle on 1100 pound mark for wing being checked, turn electrical power off, remove dummy load and connect electrical wire to center post of inboard fuel level transmitter. (10) Disconnect electrical wire from center post of opposite inboard fuel level transmitter and repeat steps (8) and (9). (11) Readjust null and gain trimpots if required until both fuel gage settings are satisfactory. (12) Fill both tanks to full capacity.
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MODEL 208 MAINTENANCE MANUAL
Fuel Transmitter Resistance Check Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (13) Measure the resistance between the center post and ground on the wing being checked. The resistance should be between 224-ohms and 242-ohms. (14) If the resistance is above or below 224-ohms and 242-ohms, check connections at each transmitter and the ground at the outboard transmitter for corrosion and security. (15) If the connections between each transmitter and ground are good, a resistance check of each transmitter will be required. (16) Remove both wires from each transmitter. (17) Use Table 201 for transmitter resistance check. (18) If a faulty transmitter is found drain fuel in accordance with 28-01, A. (19) Replace transmitter and perform the calibration procedure starting with step (1). (20) Perform steps (13) thru (15) on opposite wing. (21) After calibration procedure is complete, check that all transmitter connections are secure and install access covers. Table 201. Resistance Values at Various Float Levels. PART NUMBER
EMPTY
IN. ABOVE EMPTY
FULL TANK
C668050-1103
0 ohms to 1 ohm
.93 inch / 2 to 6 ohms
41 ohms to 45 ohms
C668050-1104
0 ohms to 1 ohm
1.00 inch / 2 to 6 ohms
23 ohms to 27 ohms
C668050-1105
0 ohms to 1 ohm
1.00 inch / 3 to 7 ohms
43 ohms to 47 ohms
C668050-1106
0 ohms to 1 ohm
1.00 inch / 11 to 15 ohms
117 ohms to 123 ohms
C668050-1107
0 ohms to 1 ohm
1.00 inch / 3 to 7 ohms
43 ohms to 47 ohms
C668050-1108
0 ohms to 0.5 ohm
1.00 inch / 11 to 15 ohms
117 ohms to 123 ohms
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MODEL 208 MAINTENANCE MANUAL FUEL QUANTITY INDICATING SYSTEMS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fuel quantity indicating systems in a serviceable condition.
Task 28-41-00-710 2.
Fuel Reservoir Warning System Operational Check A.
General (1) This task gives the procedures to do a check of the low fuel warning system for the reservoir. This check is done with the engines running.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do an Operational Check of the Fuel Reservoir Warning System (Non-Garmin equipped airplanes). (1) Start the engine. Refer to Pilot’s Operating Handbook and Approved Airplane Flight Manual. (2) Set both fuel selectors to OFF. (a) Make sure that the FUEL SELECT OFF (red) annunciator comes on and a fuel selector warning horn sounds. NOTE: (3)
Make sure that the RESERVOIR FUEL LOW (red) annunciator warning light comes on with approximately one-half or less fuel remaining in the reservoir tank. NOTE:
(4) (5) E.
The horn can be shut off by pulling the START CONT circuit breaker.
With the fuel reservoir full, there is sufficient fuel for approximately 3 minutes of maximum continuous power or approximately 9 minutes at idle power.
Set both fuel selectors to ON. Shut down the engine. Refer to Pilot’s Operating Handbook and Approved Flight Manual.
Do an Operational Check of the Fuel Reservoir Warning System (Garmin equipped airplanes). (1) Start the engine. Refer to Pilot’s Operating Handbook and Approved Airplane Flight Manual. (2) Set both fuel selectors to OFF. (a) Make sure that the FUEL SELECT OFF (red) CAS message comes on and a fuel selector warning horn sounds. NOTE: (3)
Make sure that the RSVR FUEL LOW (red) CAS message comes on with approximately one-half or less fuel remaining in the reservoir tank. NOTE:
(4) (5)
The horn can be shut off by pulling the START CONT circuit breaker.
With the fuel reservoir full, there is sufficient fuel for approximately 3 minutes of maximum continuous power or approximately 9 minutes at idle power.
Set both fuel selectors to ON. Shut down the engine. Refer to Pilot’s Operating Handbook and Approved Flight Manual.
F.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL Task 28-41-00-720 3.
Fuel Quantity and Low Fuel Warning Systems Functional Check A.
General (1) This task gives the procedures to do a functional check of the fuel quantity and low fuel warning systems.
B.
Special Tools (1) Ground Electrical Power Unit (2) Digital Ohm Meter (3) 230-Ohm Dummy Load
C.
Access (1) Remove access panels 511AB and 611AB from bottom of wings. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do a Functional Check of the Fuel Quantity and Low Fuel Warning Systems (Non-Garmin Equipped).
CAUTION: When you do the resistance tests of the fuel quantity wiring or the fuel probe transmitters, use a digital ohmmeter only. Some analog ohmmeters can introduce high current, which will make the fuel probe transmitters unserviceable. NOTE: (1) (2) (3) (4) (5) (6) (7)
(8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21) (22)
Always use a screwdriver with an insulated shank when calibrating the fuel system.
Make sure that the airplane is in a level condition. Refer to Chapter 8, Leveling - Maintenance Practices. Make sure that the airplane is correctly grounded. Refer to Chapter 12, Fuel - Servicing. Connect external electrical power to the airplane. Disengage the AUX FUEL PUMP circuit breaker. Set the external power switch to BUS. Set the battery switch to ON. De-fuel the airplane. Refer to Chapter 12, Fuel - Servicing. (a) Make sure that the amber LEFT FUEL LOW and RIGHT FUEL LOW annunciator panel lights come on when approximately 25 +/-5 gallons (170 +/- 33.5 lbs.) of fuel remains in the related main fuel tank. When all fuel is drained (except unusable) from the wing tanks, position both fuel selector valves to OFF. Fully drain the reservoir tank until it is empty. Make sure that each fuel quantity gage needle is fully in the empty red zone. (a) If necessary, carefully adjust the null trimpot on each fuel quantity gage so that the needle is fully in the empty red zone. Set the external power switch to OFF. Set the battery switch to OFF. Find and disconnect the electrical wire from the center post of the inboard fuel transmitter on one wing. Install a 230-ohm dummy load between this wire and airplane ground. Set the external power switch to BUS. Set the battery switch to ON. Carefully adjust the gain trimpot on the fuel gage to put the needle to the center of the 1100 pound mark for the wing being checked. Set the external power switch to OFF. Set the battery switch to OFF. Remove dummy load and connect electrical wire to center post of inboard fuel level transmitter. Do the steps again for the opposite wing. (a) If necessary, adjust the null and the gain trimpots again until the left and the right fuel gage low end setting reads zero and the high end setting reads 1100 pounds. Set the external power switch to BUS.
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MODEL 208 MAINTENANCE MANUAL (23) Set the battery switch to ON. (24) Make sure that the LEFT FUEL LOW, RIGHT FUEL LOW, and RESERVOIR LOW annunciator panel lights are on. (a) If any of the lights are not on, do a continuity check of the circuit and replace the bulbs if necessary. (25) Fill each wing fuel bay with 20 gallons of measured fuel, then continue to add fuel, if necessary until each bay contains 30 gallons of fuel. (a) Make sure that the LEFT and the RIGHT annunciator lights go off between the 20 and 30 gallon levels. 1 If one or both lights do not go off, replace the applicable low fuel level switches. Refer to Fuel Quantity Indicating Systems - Maintenance Practices. (26) Set both fuel selector valves ON until the RESERVOIR LOW light turns off, then position both fuel selector valves OFF. (27) Slowly drain the fuel from the reservoir until the Reservoir Low light turns ON. (28) Drain the remaining fuel from the reservoir and measure the quantity. (a) Make sure that the measurement is from 1.95 to 2.15 gallons. (b) If quantity of measured fuel is more than 2.25 gallons, or is less than 1.75 gallons, remove and replace the reservoir fuel low level switch. Refer to Fuel Quantity Indicating Systems - Maintenance Practices. (29) Fill both tanks to the full capacity. Refer to Chapter 12, Fuel - Servicing. (30) Make sure that the indication on both fuel gages is FULL. (31) Set the external power switch to OFF. (32) Set the battery switch to OFF. (33) Engage the AUX FUEL PUMP circuit breaker. (34) Remove the external electrical power unit from the airplane. (35) Remove the grounding wire from the airplane. E.
Do a Functional Check of the Fuel Quantity and Low Fuel Warning Systems (Garmin equipped airplanes with CAN bus type fuel level sensors). NOTE:
All G1000 aircraft must have software version 0767.00 or later. The software version is shown on the upper right corner of the MFD on the first page shown after the MFD is powered on in normal operation.
NOTE:
If the fuel quantity indicator on the Garmin G1000 system has a red X on it during normal operation, examine the fuel quantity sensors and wiring and refer to the Garmin G1000 Line Maintenance Manual for more Garmin system troubleshooting. If the values given on the PFD are not the same as the values given in the calibration procedure, refer to the Garmin G1000 Line Maintenance Manual for troubleshooting.
(1)
Make sure that the airplane is in a level condition. Refer to Chapter 8, Leveling - Maintenance Practices. (2) Make sure that the airplane is correctly grounded. Refer to Chapter 12, Fuel - Servicing. (3) Connect external electrical power to the airplane. (4) Disengage the AUX FUEL PUMP circuit breaker. (5) Set the external power switch to BUS. (6) Set the battery switch to ON. (7) Set the AVIONICS 1 and AVIONICS 2 switches to ON. (8) Defuel the airplane. Refer to Chapter 12, Fuel - Servicing. (a) Make sure that the amber L-R FUEL LEVEL LOW CAS message comes on when approximately 25 +/-5 gallons (170 +/- 33.5 lbs.) of fuel remains in the related main fuel tank. (9) When the fuel is fully drained from the wing tanks (except unusable), set both fuel selector valves to OFF. (10) Fully drain the reservoir tank.
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MODEL 208 MAINTENANCE MANUAL (11) Make sure that the Fuel Quantity Indications (MFD - Engine and System displays) are zero (analog pointers and digital indications). (a) If there is not a zero indication, do a Fuel Quantity System Calibration (Airplanes with CAN bus type fuel level sensors). Refer to Fuel Tanks - Adjustment/Test. (12) Add or remove the fuel quantities shown in the “Fuel Qty” column of Table 601, for both the left (L) and the right (R) wing fuel tanks and make sure that you have the correct values. NOTE:
The value listed in the “fuel qty” column is the amount of fuel to add (positive values) or remove (negative values), such that it increases or decreases the existing fuel level in the tank by that amount; each row must be completed and verified separately before you go to the next row.
Table 601. Fuel Data Verification Tank (L/R)
L
L
Fuel Qty (gal)
+20
+8#
EIS Parameter L Fuel Qty Scale/Pointer (Analog) (Amber/Amber Background) QTY L LBS (Digital) (Amber Background) L Fuel Qty Scale/Pointer (Analog)
CAS Annunciation (color)
< 200 lbs L-R FUEL LOW - Amber 134* (Theoretical)
R FUEL LOW - Amber
QTY L LBS (Digital)
R
+20
R Fuel Qty Scale/Pointer (Analog) (Amber/Amber Background)
L
+10
-4**
< 200 lbs R FUEL LOW - Amber 134* (Theoretical)
R Fuel Qty Scale/Pointer (Analog)
None
L/R
L/R
##
+50
+50
Approx. 200 lbs
QTY R LBS (Digital)
200*
L Fuel Qty Scale/Pointer (Analog) (Amber/Amber Background)
< 200 lbs L FUEL LOW - Amber
QTY L LBS (Digital) (Amber Background)
L
< 200 lbs 175*
QTY R LBS (Digital) (Amber Background) R
Display Value
L-R Fuel Qty Scale/Pointer (Analog) QTY L LBS (Digital) QTY R LBS (Digital) L-R Fuel Qty Scale/Pointer (Analog) QTY L LBS (Digital) QTY R LBS (Digital) L-R Fuel Qty Scale/Pointer (Analog) QTY L LBS (Digital) QTY R LBS (Digital)
150* Approx. 200 lbs None 200* 400-600 lbs None 536* (Theoretical) 800-1000 lbs None 871* (Theoretical)
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MODEL 208 MAINTENANCE MANUAL Table 601. Fuel Data Verification (continued) Tank (L/R)
Fuel Qty (gal)
L
F
EIS Parameter
CAS Annunciation (color)
L Fuel Qty Scale/Pointer (Analog)
None
R
F
QTY R LBS (Digital)
Full 1100
QTY L LBS (Digital) R Fuel Qty Scale/Pointer (Analog)
Display Value
None
Full 1100
• •
(13) (14) (15) (16) (17) (18)
* Tolerance +/- 75 lbs. ** Remove fuel by draining into reservoir tank (approx. 4 gallons) as follows: Set the Left Fuel Selector to ON; Make sure that the “RSVR FUEL LOW” red CAS annunciation goes off; set the Left Fuel Selector to OFF. • # You can add 2 more gallons (10 gal total) if the “R FUEL LOW” amber CAS annunciation does not show. Digital display value must show 200*, if 10 gallons is added. • ## Add (14 – X) gallons, where X is the number of gallons added in the second step (row 2 of table). Set the AVIONICS 1 and AVIONICS 2 switches to OFF. Set the external power switch to OFF. Set the battery switch to OFF. Engage the AUX FUEL PUMP circuit breaker. Remove the external electrical power from the airplane. Remove the grounding wire from the airplane.
F.
Restore Access (1) Install lower wing access panels 511AB and 611AB. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task
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MODEL 208 MAINTENANCE MANUAL FUEL SELECTORS OFF WARNING SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
4.
The fuel selectors off warning system consists of two horns installed on the overhead console; three relays, one installed on the circuit breaker panel and two on the avionics panel behind the circuit breaker panel; four switches, two on each side installed adjacent to each other at the aft wing root shutoff valve of each wing; a circuit breaker on the circuit breaker panel; and an annunciator light on the annunciator panel.
The aural warning system is powered through the start circuit breaker with a circuit breaker installed in series to protect the integrity of the start system. The switches are actuated by a cam motion which gives a positive valve position indication. One switch on each side provides a ground signal to a relay when the shutoff valve is in the open position. When the switches are not providing a ground signal for the relays, they provide a ground signal during start for the FUEL SELECT OFF light and horn. The FUEL SELECT OFF light is powered by the annunciator circuit breaker. The other two switches provide a ground signal for the second horn when either shutoff valve is closed. Power is supplied to the second horn directly off the start circuit breaker via the start switch. When the start switch is engaged, a start command signal (24 VDC) is applied simultaneously to the Generator Control Unit, the second horn, and a relay. A current limiting resistor and diode are installed in series with the horn to protect the start system. The horn sounds as long as the start switch is engaged if either shutoff valve is closed. A relay is also actuated during starter engagement. The relay switch contacts provide a method of sounding the first horn and illuminating the red FUEL SELECT OFF light if either isolation valve is closed. Also the low fuel annunciator will illuminate if the tank selected is at or below 25 gallons
Warning Horns Removal/Installation A.
Remove Warning Horns (Refer to Figure 201). (1) Remove screws (16) and remove cover (15) and trim panel (14). (2) Remove screws (13) securing hinged mounting plate (12). (3) Disconnect housing plug (4) from housing cap (5). (4) Unscrew nut ring (7) and remove horn (6). (5) Open zipper in headliner. (6) Disconnect housing plug (11) from housing cap (10). (7) Unscrew nut ring (8) and remove horn (9).
B.
Install Warning Horns (Refer to Figure 201). (1) Position horn (9) through bracket and install nut ring (8). (2) Connect housing plug (11) to housing cap (10). (3) Close zipper in headliner. (4) Position horn (6) through hinged mounting plate (12) and install nut ring (7). (5) Connect housing plug (4) to housing cap (5). (6) Secure hinged mounting plate (12) using screw (13). (7) Install trim plate (14) and cover (15) using screws (16). (8) Check horns for operation.
Relay Assembly Removal/Installation A.
Remove Relay Assembly (Refer to Figure 201). (1) Remove left sidewall circuit breaker panel by removing screws securing panel to side of airplane. (2) Disconnect electrical receptacle (21). (3) Remove nuts, washers, and screws securing relay assembly (22) to panel (23).
B.
Install Relay Assembly (Refer to Figure 201). (1) Install relay assembly (22) on panel (23) using screws, washers, and nuts. (2) Connect electrical receptacle (21). (3) Install circuit breaker panel to side of airplane.
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MODEL 208 MAINTENANCE MANUAL
Fuel Selector Off Warning System Figure 201 (Sheet 1)
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Fuel Selector Off Warning System Figure 201 (Sheet 2)
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Fuel Selector Off Warning System Figure 201 (Sheet 3)
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Fuel Selector Off Warning System Figure 201 (Sheet 4)
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Fuel Selector Off Warning System Figure 201 (Sheet 5)
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5.
6.
Fuel Off Warning Switches Removal/Installation A.
Remove Fuel Off Warning Switches (Refer to Figure 201). (1) Open right cowl and disconnect airplane battery. (2) Remove lower wing access covers to gain access to that segment of system located between fuel tanks and fuselage. (3) Remove electrical leads from switches; tag for identification for reinstallation. (4) Remove nuts (30), screws (27), and remove switches (28) and actuators (26) from bracket (29).
B.
Install Fuel Off Warning Switches (Refer to Figure 201). (1) Install switches (28) and actuators (26) on bracket (29) using screws (27) and nuts (30). (2) Connect electrical leads to switches (28), then remove tags installed for identification. (3) Connect airplane battery and check switch operation. (4) Install lower Wing access covers.
Fuel Warning Switches Adjustment A.
7.
Adjust Fuel Warning Switches (Refer to Figure 201). (1) Slotted holes are provided in bracket (29) for switch adjustment. Adjust switches and when fuel selector is moved from ON to OFF position and handle (25) moves 0.50 inch the horn will sound.
Fuel Selector Off Warning System Check A.
Annunciator Test. (1) DAY-NIGHT switch in NIGHT position. (2) ENG. INST. lighting rheostat full dim. (3) Both fuel tank selectors ON. (4) Turn BATTERY switch ON and verify FUEL SELECT OFF annunciator is extinguished and there are no warning horns sounding. (5) Press ANNUN PANEL LAMP TEST switch and verify that FUEL SELECT OFF annunciator is illuminated (full bright) and two warning horns are sounding.
B.
Circuit Test (In-Flight Mode). (1) Both fuel tank selectors OFF. (2) Turn BATTERY switch ON and verify that FUEL SELECT OFF annunciator is illuminated (full bright) and one warning horn is sounding. (3) With fuel quantity less than 20 gallons per tank (LOW FUEL annunciator ON), turn left fuel tank selector ON and verify that annunciator is still illuminated and one warning horn is still sounding. (4) Turn the left fuel tank selector OFF and right fuel tank selector ON and verify that annunciator is still illuminated and one warning horn is still sounding. (5) Add fuel to left and right fuel tanks to extinguish the LEFT and RIGHT FUEL LOW annunciators. (6) With right fuel tank selector ON and left selector OFF, verify that annunciator is extinguished and neither warning horn is sounding. (7) Repeat step (6) with left fuel tank selector ON and right selector OFF.
C.
Circuit Test (Start Mode). (1) Both fuel tank selectors ON. (2) Move START switch to MOTOR and verify that FUEL SELECT OFF annunciator is extinguished and neither warning horn is sounding. (3) Turn left fuel tank selector OFF. (4) Move START switch to MOTOR and verify that FUEL SELECT OFF annunciator is illuminated and both warning horns are sounding. (5) Turn the left fuel tank selector ON and the right fuel tank selector OFF. (6) Repeat step (4). (7) Turn both fuel tank selectors OFF. (8) Repeat step (4). (9) Pull START CONT circuit breaker. (10) Turn both fuel tank selectors ON and verify that FUEL SELECT OFF annunciator is illuminated.
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MODEL 208 MAINTENANCE MANUAL (11) Reset START CONT circuit breaker. NOTE:
If the FUEL SELECT WARNING circuit breaker has popped or the START CONT circuit breaker has been pulled (possibly for ground maintenance), the FUEL SELECT OFF annunciator will be illuminated even with both fuel tank selectors ON. This is an indication that the fuel selector warning system has been deactivated.
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ICE AND RAIN PROTECTION
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MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
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30-Title 30-List of Effective Pages 30-Record of Temporary Revisions 30-Table of Contents 30-List of Tasks
30 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Mar 1/2012
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS ICE AND RAIN PROTECTION - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment, and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-00-00 30-00-00 30-00-00 30-00-00
PNEUMATIC SURFACE DEICE - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-10-00 Page 1 30-10-00 Page 1 30-10-00 Page 1
PNEUMATIC SURFACE DEICE - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-10-00 Page 101 30-10-00 Page 101
PNEUMATIC SURFACE DEICE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation for Installation of Deice Boot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation and Application of Fuel Barrier . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation and Application of Bonding Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pneumatic Deice Boots Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pneumatic Deice Flow Control Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . Pneumatic Deice Pressure Switches Removal/Installation . . . . . . . . . . . . . . . . . . . . . . Pneumatic Deice Timer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Air System Pressure Regulator Removal/Installation . . . . . . . . . . . . . . . . . . . . . Pneumatic Deice System Adjustment/Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Approved Repairs (Cold Patch). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pneumatic Deice Boot Winterization Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pneumatic Flow Control Valves Disassembly/Cleaning/Assembly Procedure . . . . .
30-10-00 Page 201 30-10-00 Page 201 30-10-00 Page 201 30-10-00 Page 201 30-10-00 Page 202 30-10-00 Page 202 30-10-00 Page 203 30-10-00 Page 208 30-10-00 Page 209 30-10-00 Page 209 30-10-00 Page 209 30-10-00 Page 213 30-10-00 Page 214 30-10-00 Page 215 30-10-00 Page 215 30-10-00 Page 217
PNEUMATIC SURFACE DEICE - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Air Pressure Regulator Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-10-00 Page 601 30-10-00 Page 601 30-10-00 Page 601
TKS ANTI-ICE SYSTEM - DESCRIPTION AND OPERATION Cargo Pod Installation. . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-00 Page 1 30-11-00 Page 1 30-11-00 Page 1 30-11-00 Page 8
TKS ANTI-ICE SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Anti-Ice System Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inboard TKS Wing Panel Pressurization Functional Check . . . . . . . . . . . . . . . . . . . . . .
30-11-00 Page 601 30-11-00 Page 601 30-11-00 Page 601 30-11-00 Page 601
G1000 AVIONICS AND TKS ANTI-ICE SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Troubleshooting Description. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Troubleshooting Preliminary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Troubleshooting Charts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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30 - CONTENTS © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE FLUID TANK COMPONENTS - MAINTENANCE PRACTICES Pod Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Anti-Ice Fluid Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Anti-Ice Fluid Tank Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Pack Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Metering Pump Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Pump Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fluid Level Sender Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Level Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . High Pressure Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Timer Box and/or Wire Bundle Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Solenoid Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sight Glass Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drain Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fluid Filler Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Contamination (Fuel) Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Contamination (Water) Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Contamination (Solids) Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-10 Page 201 30-11-10 Page 201 30-11-10 Page 201 30-11-10 Page 202 30-11-10 Page 210 30-11-10 Page 212 30-11-10 Page 214 30-11-10 Page 216 30-11-10 Page 217 30-11-10 Page 219 30-11-10 Page 221 30-11-10 Page 222 30-11-10 Page 223 30-11-10 Page 225 30-11-10 Page 227 30-11-10 Page 228 30-11-10 Page 229 30-11-10 Page 231 30-11-10 Page 234 30-11-10 Page 235
TKS ANTI-ICE FLUID TANK COMPONENTS - ADJUSTMENT/TEST Pod Installation . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Anti-Ice Fluid Tank Component Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Anti-Ice Level Sender Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-10 Page 501 30-11-10 Page 501 30-11-10 Page 501 30-11-10 Page 502 30-11-10 Page 506
TKS ANTI-ICE SYSTEM - DESCRIPTION AND OPERATION FAIRING INSTALLATION General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-11 Page 30-11-11 Page 30-11-11 Page 30-11-11 Page
TKS ANTI-ICE SYSTEM - MAINTENANCE PRACTICES FAIRING INSTALLATION . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fairing Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Accessory Bracket Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Tank Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Metering Pump Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Pump Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fluid Level Sender Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Level Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Timer Box Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Solenoid Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sight Glass Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drain Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fluid Filler Tube Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fluid Filler Port Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vent Tube Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pump Strainer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Contamination (Fuel) Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Contamination (Water) Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Fluid Contamination (Solids) Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-11 Page 201 30-11-11 Page 201 30-11-11 Page 201 30-11-11 Page 201 30-11-11 Page 204 30-11-11 Page 212 30-11-11 Page 214 30-11-11 Page 216 30-11-11 Page 218 30-11-11 Page 220 30-11-11 Page 221 30-11-11 Page 223 30-11-11 Page 225 30-11-11 Page 226 30-11-11 Page 227 30-11-11 Page 229 30-11-11 Page 231 30-11-11 Page 232 30-11-11 Page 233 30-11-11 Page 238 30-11-11 Page 239 30-11-11 Page 240 30-11-11 Page 241 30-11-11 Page 244 30-11-11 Page 245
30 - CONTENTS © Cessna Aircraft Company
1 1 1 7
Page 2 of 4 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE SYSTEM - ADJUSTMENT/TEST FAIRING INSTALLATION . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Anti-Ice System Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Level Sender Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-11 Page 501 30-11-11 Page 501 30-11-11 Page 501 30-11-11 Page 502 30-11-11 Page 506
TKS ANTI-ICE LEADING EDGE POROUS PANEL - MAINTENANCE PRACTICES . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Porous Panel Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Porous Panel Sealant Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-20 Page 201 30-11-20 Page 201 30-11-20 Page 201 30-11-20 Page 201 30-11-20 Page 211
TKS ANTI-ICE LEADING EDGE POROUS PANEL - ADJUSTMENT/TEST . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Porous Panel Purge and Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-20 Page 501 30-11-20 Page 501 30-11-20 Page 501 30-11-20 Page 502
TKS ANTI-ICE FLUID DISTRIBUTION SYSTEM - MAINTENANCE PRACTICES . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Propeller Proportioning Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Wing Proportioning Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Stabilizer Proportioning Unit and Low Pressure Switch (Tail Bracket Assembly) Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nylon Tubing Repair/Replacement. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-11-30 Page 201 30-11-30 Page 201 30-11-30 Page 201 30-11-30 Page 202 30-11-30 Page 209
PITOT AND STATIC HEATERS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-30-00 Page 1 30-30-00 Page 1
STALL WARNING HEATER - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-31-00 Page 1 30-31-00 Page 1
WINDSHIELD ANTI-ICE - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-40-00 Page 1 30-40-00 Page 1 30-40-00 Page 1
WINDSHIELD ANTI-ICE - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-40-00 Page 101 30-40-00 Page 101
WINDSHIELD ANTI-ICE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Anti-Ice Panel and Attach Bracket Removal/Installation . . . . . . . . . . . . . . Windshield Anti-Ice Electrical Receptacle Cover Removal/Installation . . . . . . . . . . . . Windshield Anti-Ice Controller Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Anti-Ice Relay Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Approved Repairs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-40-00 Page 201 30-40-00 Page 201 30-40-00 Page 201 30-40-00 Page 201 30-40-00 Page 207 30-40-00 Page 207 30-40-00 Page 207
WINDSHIELD ANTI-ICE - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Anti-Ice System Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-40-00 Page 601 30-40-00 Page 601 30-40-00 Page 601
TKS ANTI-ICE WINDSHIELD SPRAY BAR - MAINTENANCE PRACTICES . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TKS Anti-Ice Windshield Spray Bar Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . .
30-41-00 Page 201 30-41-00 Page 201 30-41-00 Page 201 30-41-00 Page 201
PROPELLER ANTI-ICE - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-60-00 Page 1 30-60-00 Page 1 30-60-00 Page 1
PROPELLER ANTI-ICE - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-60-00 Page 101 30-60-00 Page 101
30-11-30 Page 211 30-11-30 Page 213
30 - CONTENTS © Cessna Aircraft Company
Page 3 of 4 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL PROPELLER ANTI-ICE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Anti-Ice Boots (Hartzell) Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Anti-Ice Boots (McCauley) Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . Slip Ring Assembly (Hartzell) Removal/Rework/Installation . . . . . . . . . . . . . . . . . . . . . Slip Ring Alignment Check (Hartzell). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Slip Ring Run-Out Test (McCauley) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Length Inspection (Hartzell) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Length Inspection (McCauley). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Block Removal/Installation (Hartzell) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Block Removal/Installation (McCauley). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Block Assembly to Slip Ring Alignment (Hartzell) . . . . . . . . . . . . . . . . . . . . . . . . Brush Block Assembly to Slip Ring Alignment (McCauley) . . . . . . . . . . . . . . . . . . . . . . Brush Block/Slip Ring Cleaning. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Anti-Ice Timer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Anti-Ice Ammeter Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-60-00 Page 201 30-60-00 Page 201 30-60-00 Page 201 30-60-00 Page 207 30-60-00 Page 209 30-60-00 Page 209 30-60-00 Page 212 30-60-00 Page 212 30-60-00 Page 213 30-60-00 Page 213 30-60-00 Page 213 30-60-00 Page 216 30-60-00 Page 216 30-60-00 Page 216 30-60-00 Page 216 30-60-00 Page 217
TKS ANTI-ICE PROPELLER (McCauley) - MAINTENANCE PRACTICES . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller TKS Anti-Ice Feed Shoes Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . Slinger Ring and Feed Nozzle Alignment Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Feed Tube to Propeller Blade Alignment Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-61-00 Page 201 30-61-00 Page 201 30-61-00 Page 201 30-61-00 Page 205 30-61-00 Page 206
WINDSHIELD ICE INDICATOR LIGHT - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Ice Indicator Light . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-80-00 Page 201 30-80-00 Page 201 30-80-00 Page 201 30-80-00 Page 201
FLIGHT INTO KNOWN ICING CONDITIONS EQUIPMENT - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
30-90-00 Page 1 30-90-00 Page 1
30 - CONTENTS © Cessna Aircraft Company
Page 4 of 4 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 30-10-00-720
Bleed Air Pressure Regulator Functional Check
30-10-00 Page 601
30-11-00-720
TKS Anti-Ice System Functional Check
30-11-00 Page 601
30-11-00-721
Inboard TKS Wing Panel Pressurization Functional Check
30-11-00 Page 601
30-40-00-710
Windshield Anti-Ice System Operational Check
30-40-00 Page 601
30 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ICE AND RAIN PROTECTION - GENERAL 1.
Scope A.
2.
This chapter provides information on systems that detect, remove, or prevent ice formation on critical surfaces.
Tools, Equipment, and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Adhesive
Scotch Grip EC-1300L Cement
3M 3M Center St. Paul, MN 55101-0000
To bond deice boot to wing
Adhesive
EC-776 Fuel Resistant
3M
To provide barrier coat in fuel bay
Commercially Available
To clean deice boot and mating surface
Methyl-n-Propyl Ketone Cleaning Solvent
A-A-59281
Commercially Available
To clean metal surface of deice boots
Cleaning Solvent
Technical Toluol FSN A-A-59107
Commercially Available
To remove deice boots.
Isopropyl Alcohol
TT-1-735
Commercially Available
To clean outer surface of deice boots and components of pneumatic flow control valve when disassembled
Silicone Fluid
Dow Corning 200, 100cs
Dow Corning Corp. P. O. Box 997 3901 S. Saginaw Rd. Midland, MI 48686
To lubricate components of pneumatic flow valve when disconnected
ScotchBrite
7440, Heavy Duty
3M
To remove paint/finish for deice boot or TKS panel installation
ScotchBrite
7445, Very Fine
3M
To remove paint/finish for deice boot installation, or polish TKS porous panel
ScotchBrite
7448, Ultra Fine
3M
To remove paint/finish for deice boot installation, or polish TKS porous panel
Norgren 5400 S. Delaware St. Littleton, CO 81120
To winterize pneumatic deice system
Filter/Regulator/ Lubricator Assembly Paint and Lacquer Remover
TT-R-248
Commercially Available
To remove paint prior to installing deice boot
Wash Primer
WMS 30-1
Commercially Available
To remove paint prior to installing deice boot
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NAME
NUMBER
MANUFACTURER
USE
Coating
78-U-1003 Black Enamel with U-1001 Catalyst
Sterling Lacquer 3150 Brannon Ave. St. Louis, MO 63139
To dress edge of deice boot
Coating
ShineMaster Prep
BF Goodrich 1555 Corporate Woods Pky. P. O. Box 1277 Union Town, OH 44685
To prepare deice boot surface
Coating
ShineMaster
BFGoodrich
To coat deice boots
Jet Glo Enamel Polyurethane
571-510
Sherwin-Williams 630 E. 13th Andover, KS 67002
To dress edge of deice boot (alternate)
Acry Glo Enamel Polyurethane
571-010
Sherwin-Williams
To dress edge of deice boot (alternate).
Rubber Roller
2 inch wide, 2 inch diameter, Rubber
Everhard Products, Inc. 1016 9th St. SW Canton, OH 44707
To install deice boots
Roller
0.25 inch wide, 2 inch diameter, Metal Sticher
Everhard Products, Inc.
To install deice boots
Masking Tape
One Inch
Commercially Available
To mask off boot area
Commercially Available
To mark centerline of wing and boot
Carpenter’s Chalk Line Universal Repair Kit
74-451-AA
Cessna Aircraft Company Cessna Parts Distribution Department 701, CPD 2 5800 East Pawnee Road Wichita, KS 67218-5590
To repair deice boots
Pin Hole Repair Kit
74-451-AE
Cessna Aircraft Company
To repair deice boots
Multimeter
Model 260 or equivalent
Simpson Electric Co. 853 Dundee Ave. Elgin, IL 60120
To check voltage and continuity in electrical circuits
Measuring Tape
Commercially Available
Sharp Knives
Commercially Available
Cleaning cloths (Lint-free)
Commercially Available
Cleaning
Cleaning cloths (Lint-free)
Commercially Available
Cleaning
Rymple cloth
301 purified
International Paper Veratec Division 100 Elm Street Walpole, MA 02081
Cleaning
Adhesive
EA9309
Dexter Corp. Hysol Division 2850 Willow Pass Road Pittsburg, CA 94565
To bond pneumatic flow control valve pushrod
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NAME
NUMBER
MANUFACTURER
USE
Clenching Tool (3/16 Inch)
T300-112A
A S and T Aerospace Systems & Technologies Inc. 2734 Arnold Court Salina, KS 67401
To repair and replace TKS tubing
Clenching Tool (5/16 Inch)
T300-120A
A S and T Aerospace Systems & Technologies Inc.
To repair and replace TKS tubing
Clenching Tool (1/2 Inch)
T300-144A
A S and T Aerospace Systems & Technologies Inc.
To repair and replace TKS tubing
TKS System Test Cart
09301-01
A S and T Aerospace Systems & Technologies Inc.
To do the TKS porous panel purge and test procedures
Blank Olive with MN4856 Nut (3/16-Inch Tube)
P075
A S and T Aerospace Systems & Technologies Inc.
To cap the TKS tubing during tests
Nylon Ball with MN4855 Nut (5/16-Inch Tube)
03-151-07
A S and T Aerospace Systems & Technologies Inc.
To cap the TKS tubing during tests
Nylon Ball with MN6201 Nut (1/2-Inch Tube)
03-151-10
A S and T Aerospace Systems & Technologies Inc.
To cap the TKS tubing during tests
Test Harness with Breakout Box
P2697006
Cessna Aircraft Company
To do a test of the TKS system
Low-Adhesive Plastic Tape
PC 628
Shurtape Technologies
To install the TKS porous panels
Commercially Available
To install the TKS porous panels
Cessna Aircraft Company
To install TKS porous panels to the leading edges
Plastic Sheeting
Commercially Available
To catch deice fluid from the TKS panel purge procedure, or when you test of the TKS system
Plastic Guttering
Commercially Available
To catch deice fluid from the TKS panel purge procedure, or when you test of the TKS system
Aluminum Tape
Commercially Available
To attach guttering and tubes to catch deice fluid from the TKS panel purge procedure, or when you test of the TKS system
Plastic Tubes
Commercially Available
To catch deice fluid from the TKS panel purge procedure, or when you test of the TKS system
Velcro Straps Sealant
Pro-Seal 870, (U544042S) Type X, Class B
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3.
Definition A.
This chapter is divided into sections and subsections to assist maintenance personnel in locating specific systems and information. For locating specific information within the chapter, refer to Contents, located at the beginning of the chapter. (1) The section on airfoils provides information on the surface deice system (pneumatic boots) used on the airplane. (2) The section on windshield anti-ice provides information on those components used to keep ice from the windshields. (3) The section on propeller anti-ice provides information on those components used to keep ice from the propeller. (4) There are also sections for the TKS anti-ice system.
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MODEL 208 MAINTENANCE MANUAL PNEUMATIC SURFACE DEICE - DESCRIPTION AND OPERATION 1.
General A.
2.
Ice protection systems are offered as independent options or as a complete package referred to as Flight Into Known Icing Conditions. These systems are provided to prevent the formation of ice and to remove ice from various areas of the airplane. The following are available options (Refer to Figure 1). (1) Propeller anti-ice system option includes electrically heated boots bonded to the propeller blades. (2) Windshield anti-ice system option includes an electrically heated, removable windshield panel on the pilot's side of the windshield. (3) Electrical heaters option provides ice protection for pitot/static and stall warning systems. (4) Flight Into Known Icing Conditions package allows flight penetration of icing conditions as defined by the FAA. The package includes all optional ice protection systems, as well as control surface mounted electrostatic discharge wicks, standby electrical system incorporating a 75-amp alternator, pneumatic deice boots on wing leading edges, wing struts, landing gear legs (optional), cargo pod nosecap (optional), horizontal and vertical stabilizers, and an ice detector light to aid in night time ice detection on the left wing inboard leading edge. An inertial separator system is built into the engine air inlet duct to prevent particles (i.e., liquid droplets, ice crystals or snow) from entering the engine inlet plenum.
Description and Operation A.
Pneumatic deice boots, installed on the wing leading edges, wing struts, landing gear legs, cargo pod nosecap, horizontal and vertical stabilizers, are utilized to break up ice accumulation on the leading edges during flight. (1) System components include an engine compartment pressure line which leads from the engine bleed air system pressure regulator to three ejector flow control valves, three pressure switches, timer, located in left wing root area, three-position system activation switch, labeled BOOT PRESS and located on deice/anti-ice switch panel of lower left instrument panel, circuit breaker, labeled DEICE BOOT and located on left sidewall circuit breaker panel, deice pressure indicator light, mounted in annunciator panel, supply lines and pneumatically operated airfoil surface deice boots. (2) An ice detector light, flush mounted near upper left corner of windshield and directed on left inboard wing leading edge, is included in deice system to aid in night time ice detection on wing leading edges. An ice detector light switch, labeled WING LIGHT, is a spring-loaded switch which must be held in the ON (upper) position to keep the ice detector light illuminating. (a) Ice detector light components include a two- position toggle switch, labeled WING LIGHT, located on deice/anti-ice switch panel on lower left instrument panel, and a circuit breaker, labeled WING ICE DET LIGHT and located on left sidewall circuit breaker panel.
B.
Pneumatic deice system utilizes bleed air from turbine power plant as air pressure source to inflate pneumatic deice boots. A pressure regulator relief valve reduces bleed air pressure of deice system to 18 PSI.
C.
An electrical three-cycle timer is utilized to control three solenoid actuated ejector flow control valves. System is activated by a three-position momentary contact switch on the instrument panel. Deice cycle is initiated by pressing switch to BOOT PRESS (upper) position, then releasing switch. Each deice cycle has a duration of six seconds.
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Pneumatic Deice System Schematic Figure 1 (Sheet 1)
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WARNING: The absence of illumination during any one of the three sequences of a cycle indicates insufficient pressure for proper boot inflation and effective deicing ability. Additionally, any deviation from the described sequence could be an indication of a malfunction in some other portion of the system and icing conditions must be avoided. (1)
(2)
The first cycle controls the vertical fin and the horizontal stabilizer deice boots inflation. The second cycle controls the inboard wing deice boots, the cargo pod, and the landing gear fairings inflation. The third cycle controls the outboard wing and wing strut deice boots inflation. The total time for one complete deice cycle is 18 seconds. When the ejector flow control valves are in their de-energized condition, the ejector section of the valve provides vacuum necessary to maintain deice boots in a deflated condition. Each time a cycle is desired, the deice activation switch must be pressed to BOOT PRESS (upper) position and released. (a) In the event of a malfunction in the timer, which causes erratic operation of a sequence of a cycle, the switch can be held momentarily in the MANUAL (lower) position to achieve simultaneous inflation of all of the deice boots. If necessary, the system can be stopped at any point in the cycle (deflating the boots) by disengaging the circuit breaker labeled DEICE BOOT. (b) A pressure switch is installed downstream of each ejector flow control valve. These three pressure switches activate a light in the annunciator panel, allowing the pilot to verify each cycle has pressurized. Pressure switches activate at 14 to 16 PSI. (c) The pressure indicator annunciator, labeled DEICE PRESSURE, will illuminate initially within approximately three seconds after initiating a cycle and remain on for approximately three additional seconds to the end of the first sequence. Through each of the remaining two sequences of the cycle, the annunciator will remain off during pressure build up for about three seconds, then illuminate for about three seconds. If necessary, the system may be recycled six seconds after the completion of a cycle.
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MODEL 208 MAINTENANCE MANUAL PNEUMATIC SURFACE DEICE - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Deice Boots Troubleshooting Chart Figure 101 (Sheet 1)
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Deice Boots Troubleshooting Chart Figure 101 (Sheet 2)
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Deice Boots Troubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL PNEUMATIC SURFACE DEICE - MAINTENANCE PRACTICES 1.
General A.
2.
Tools, Equipment and Materials A.
3.
Pneumatic surface deice maintenance practices consist of pneumatic deice boot removal/installation, adjustment/test, approved repairs, and component removal/installation.
Refer to Ice and Rain Protection - General, for required tools, equipment and materials
Preparation for Installation of Deice Boot A.
Requirements. (1) Fill all gaps and metal mismatches with Type I aerodynamic fairing compound and sand smooth. Brush chem film on bare metal areas where chem film was removed during sanding process. (2) Adhesives, primers and coatings shall not be used beyond original expiration date, even if they have been retested and approved. Jelled or contaminated adhesives shall not be used. (3) Do not use EC776 or EC1300L adhesives for deice boot installation if adhesives have been stored for six months or more in the 80°F to 90°F range. Do not use for boot installation if the adhesives have been stored for 5 days or more at temperatures above 90°F. (4) Containers for adhesives, primers and coatings shall be kept tightly closed when the materials are not being used, unless otherwise specified. (5) Preassembly operations, such as fitting, drilling, deburring, punching, trimming, masking, etc., shall be completed prior to cleaning and bonding. Slight tensioning of deice boot may be required when reducing internal pressure to remove minor wrinkles and to obtain a smooth surface both for adhesive application and installation. An air ejector or jet pump shall be used to reduce the internal pressure of the boot to 10 Hg absolute pressure or less during cleaning and bonding of deice boot. (6) Surfaces must be clean and dry, free from dust, lint, chips, grease, oil, condensation or other moisture, as well as other contaminating substances, prior to application of adhesives, primers, coatings, and ice release promoters. (7) Deice boots and wing leading edge shall be cleaned with Methyl n-Propyl Ketone. (8) All paints, lacquers, etc., shall be removed prior to cleaning and bonding. Primed surfaces shall be cleaned with ScotchBrite pads wetted with Methyl n-Propyl Ketone, then solvent wiped. (9) Cleaning and bonding shall not be accomplished when temperature of structure, deice boots or bonding materials is below 60°F, nor when relative humidity is 90 percent or greater. (10) Faying surfaces shall be placed together while one or both surfaces exhibit an aggressive tack. Bonding must be accomplished before adhesive becomes too dry. (11) Adhesive bonds shall be free of wrinkles and entrapped air bubbles, shall not be loose at edges, nor exhibit poor adhesion. Wrinkles in boots, which prevent acceptable installation, may be removed. (12) To prevent damage to deice boots, do not use metal hand stitcher roller over areas of boot with internal tubes or wires. (13) Airplane may be flown one hour after installation/bonding deice boots, provided the boots are not operated for 48 hours following bonding. (14) Adhesives, primers and coatings shall be stirred thoroughly prior to application.
B.
Surface Treatment. (1) Waxes or wax like materials shall not be used on deice boots. Rubber protective agents/ice release promoters such as Age-Master Number 1, Icex, Acroseal, ShineMaster Prep and ShineMaster are the only acceptable surface treatments for deice boots. Refer to Chapter 12, Deicing - Maintenance Practices.
C.
Positioning. (1) Indexing marks shall be placed on metal surface outside of bonding area or a chalk line shall be snapped lengthwise down bonding area approximately on centerline of leading edge. Faying surface of boot shall be marked in a similar manner to provide for correct alignment during installation and attachment of boot. Intensify chalk line on leading edge and reference line on boot using a felt tip marking pen after first application of adhesive is thoroughly dry.
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MODEL 208 MAINTENANCE MANUAL (2) (3)
D.
Deice boot or a pattern shall be positioned on metal surface to which boot is to be bonded to provide a guide for masking and to check boot fit. Leaving an edge margin of approximately one-half inch from boot or pattern, a single strip of one-inch wide masking tape shall be applied to metal surface around periphery of boot or pattern. Masking shall be accurate so clean up time will be minimal. Boot or pattern shall then be removed.
Cleaning. NOTE: (1) (2) (3)
(4) (5)
(6) (7)
4.
All paint in masked off area shall be removed by sanding. Primed surfaces shall be cleaned with ScotchBrite pads wetted with Methyl n-Propyl Ketone, then solvent wiped. Loosened paint and remover shall be wiped off, thoroughly rinsed with clean water and dried with clean cheesecloth. All surfaces to be bonded shall be clean and dry. If no primer is present in area area, lightly abrade metal surface using ScotchBrite pads. (a) Using air ejector or jet pump, reduce pressure in boot to 10 inches of mercury (absolute) or less, smooth surface on back side of boot. (b) Using a clean cloth moistened with Methyl n-Propyl Ketone, scrub metal surface in masked off area and rough, unglazed faying surface of deice boot. Cloth must not be saturated to the point where dripping will occur. (c) Using a clean cloth, wipe Methyl n-Propyl Ketone from surfaces before evaporation to ensure oils, grease, wax, etc., will not be redeposited. Cleaning solvent must never be poured or sprayed on a structure. Final cleaning shall be accomplished immediately prior to bonding. Previously cleaned areas shall be thoroughly recleaned. As an area is being scrubbed with a moistened cloth in one hand, another clean dry cloth shall be held in opposite hand and used to dry area before solvent evaporates. Bonding procedures shall begin as soon as possible after cleaning and drying surfaces. Do not allow handling of surfaces between cleaning and bonding operations. Caution must be observed during cleaning and bonding. Solvents, adhesives, etc., are toxic and flammable. Fresh air masks and/or adequate ventilation are required for all closed areas. Structure shall be electrically grounded before beginning any cleaning or bonding operation.
Preparation and Application of Fuel Barrier A.
5.
All bare metal surfaces shall be brush coated with Iridite Chemical Film prior to adding aerodynamic smoothing compound or adhesive.
Procedure. (1) Adhesive EC776 must be thoroughly stirred prior to application as a barrier coat. A small amount of Methyl n-Propyl Ketone may be added to EC776 to achieve a more applicable consistency. One uniform coat of barrier shall be brushed over all rivet heads which penetrate the integral fuel tank and allowed to dry thoroughly until it does not have any tack. Apply a second uniform coat and allow to dry a minimum of two hours.
Preparation and Application of Bonding Material A.
Procedure. (1) Adhesive EC1300L must be thoroughly stirred prior to application. A uniform coat of adhesive shall be brushed onto masked off metal surface and onto faying surface of deice boot. When brushing adhesives on, use good quality, clean nylon brushes. Avoid hot air drafts from heaters or fans which can cause dragging and produce a very rough surface. Adhesive shall be allowed to dry thoroughly and shall not exhibit any tack. A second uniform coat of adhesive shall be brushed onto each faying surfaces and allowed to dry thoroughly and shall not exhibit tack. NOTE:
Minimum drying time is one hour at 77°F and 50 percent relative humidity. Lower temperatures and/or higher humidities require longer drying times.
NOTE:
Adhesive EC1300L may be thinned by adding 1.5 fluid ounces of Toluene to 16 ounces (1 fluid pint) of adhesive to achieve a more applicable consistency.
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MODEL 208 MAINTENANCE MANUAL (2)
6.
Dry adhesive shall be covered and kept clean until reactivated. Adhesive shall be reactivated within 48 hours by wiping lightly with clean cheesecloth, slightly moistened with toluene. Only a small area, approximately 3 inches by 18 inches or less, shall be reactivated at one time. Do not allow adhesive to become too dry before placing deice boot in contact with metal surface. Excessive rubbing or solvent usage shall be avoided to ensure adhesive will not be removed.
Pneumatic Deice Boots Removal/Installation
WARNING: Cement and solvent vapors are toxic and extremely flammable. Use only in a well ventilated area away from sparks or vapors. Excess exposure could cause injury or death. If dizziness or nausea occur, obtain fresh air immediately. Avoid contact with skin or eyes. Use solvent resistant gloves to minimize skin exposure. Use safety glasses to minimize chance of eye contact. If eye contact occurs, flush eyes with water for 15 minutes and see a physician. If skin contact occurs, wash thoroughly with soap and water. If swallowed, do not induce vomiting. See a physician immediately. WARNING: Verify aircraft is electrically grounded to prevent static sparks which could ignite solvent vapors. A.
Remove Pneumatic Deice Boot. (1) Apply toluene solvent along bond line of deice boot. Solvent will soften and undercut adhesive.
CAUTION: Do not use excessive amounts of solvent. Do not apply excessive tension to the deice boot. (2) (3) B.
Carefully apply tension to deice boot while applying solvent to bond line, then peel deice boot from airplane. Removal process should be slow enough to allow solvent to undercut the adhesive to ensure boot will not be damaged. Separate hose from deice boot.
Install Pneumatic Deice Boots (Refer to Figure 201). (1) Verify type of pneumatic deice boot to be installed. NOTE:
(2)
Clean mating surfaces of airplane and deice boot to be bonded. Refer to Preparation for Installation of Deice Boot, Cleaning. NOTE:
(3) (4) (5) (6) (7)
Fastboot pneumatic de-ice boots incorporate a bonding system with a unique pressure sensitive adhesive applied to the de-icer bond side in the manufacturing process. For fastboot installation refer to the instructions included with each boot or refer to the manufacturer’s installation instructions listed in Introduction - List of Publications.
Use removed boot bond line as a guide for cleaning paint surface to which boot adheres.
Identify position and location of deice boot on airplane. Refer to Preparation for Installation of Deice Boot, Positioning. Repeat applicable cleaning requirements. Refer to Preparation for Installation of Deice Boot, Cleaning. Apply fuel barrier. Refer to Preparation and Application of Fuel Barrier. Apply bonding material. Refer to Preparation and Application of Bonding Material. Attach hoses, connected to a vacuum source (air ejector or jet pump), to deice boot nipples with clamps.
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Deice Boot Installation Figure 201 (Sheet 1)
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Deice Boot Installation Figure 201 (Sheet 2)
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Deice Boot Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (8)
Route hose and nipple through skin. Attach vacuum to open end of hose and activate vacuum. Reduce pressure in boot to 10 inches of mercury (absolute) or less. Maintain vacuum in boot throughout entire installation process to inhibit amount of air trapped within deice boot during installation. (9) Ensure smooth outer surface of boot is clean, then roll boot up with adhesive side out, starting from end opposite air connections. (10) Position boot with reference centerline aligned with and against reference centerline on leading edge. Ensure air connections match airfoil holes. Ensure hoses are routed through correct holes to prevent improper inflation sequence. Each hose must be centered in its mating hole and not crimped. Clamp each hose to respective air supply tube to maintain proper positioning. (11) Using a clean wiping cloth dampened with toluene, reactivate adhesive on a 2 to 3 inch wide by approximately 18 inch long section on wing leading edge, outboard from air connections. Reactivate a matching section of adhesive on boot and press boot to leading edge, ensuring reference centerlines coincide and each air connection is centered on its mating hole in leading edge. (12) Using a rubber roller, roll boot down firmly against leading edge skin in reactivated area. Be careful not to trap any air under boot. Distortion of boot shall be held to a minimum. NOTE:
Reactivating adhesive is a very critical step in achieving a good bond. After cloth is thoroughly saturated with toluene, remove excess by squeezing, wringing, and/ or snapping. A properly prepared cloth should be damp, but not wet or dripping. Reactivating both surfaces will help assure 100 percent tack when pressing down and rolling boot. When in doubt, check tack with a finger before rolling boot down.
(13) Repeat steps 6.B.(10) and (11) along leading edge on inboard end of boot. Reactivate and install area around each air connection hole. Use metal stitcher roller around each air connection. (14) Complete installation of boot along leading edge. (a) Reactive adhesive on leading edge and boot 2 to 3 inches wide and an additional 2 feet to 3 feet outboard. (b) Unroll boot against leading edge, maintaining light tension on boot to prevent wrinkling. Align reference centerlines of boot and leading edge. (c) Roll boot down firmly with rubber roller. When outboard edge of boot is reached, roll it down with metal stitcher roller. (d) After entire length has been bonded at centerline of leading edge, roll reactivated area again using rubber roller. NOTE:
Installation along leading edge is best accomplished using two persons; one to hold and guide boot during installation, the other to reactivate adhesive and roll boot down.
(15) If boot should attach "off-course" (reference centerline on leading edge not coinciding with reference centerline on boot), apply Toluene with a small brush or squirt bottle to soften bond line. (a) Apply only a small amount of toluene while applying sufficient tension to peel back softened adhesive. (b) To prevent damage to the boot, avoid twisting, sharply bending, or jerking boot loose from bonded area. Allow solvent wetted area to dry thoroughly before continuing with applications. Reapply EC1300L adhesive as needed. (16) After boot is bonded along leading edge centerline, begin to reactivate adhesive on either the upper or lower surface and install remainder of boot. (a) Starting at inboard end, hold boot back to reveal bond line. (b) Using a clean cloth dampened with toluene, wipe adhesive on leading edge, 2 to 3 inches wide by approximately 18 inches long. Wipe corresponding area of boot, keeping damp cloth tight into fold of bond line. To avoid trapping air, do not allow reactivated surfaces to touch until they are rolled down. NOTE:
Maintain bond line as straight as possible, allowing bond line to be more closely monitored and aid in eliminating pockets where air can be trapped.
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MODEL 208 MAINTENANCE MANUAL (c)
Roll reactivated area firmly with rubber roller, starting at bond line and rolling with span while working toward outboard end of boot. Roll boot, leaving an open angle at end of reactivated section of boot to permit easy access for reactivating next section. (d) Constantly check for bubbles and poor adhesion as installation progresses. Rework problem areas as soon as they are discovered. (e) As outboard edge is achieved for each 2 to 3 inches of width, roll it down with metal stitcher roller. 1 If boot lifts after rolling and/or adhesive shows a "cobweb" appearance, adhesive is too wet. Wait until adhesive becomes tacky and reroll. (17) Run a toluene dampened cloth along each span edge of the boot which has previously been rolled down with rubber roller. Immediately roll edge with metal stitcher roller. (18) Using a sharp knife, trim inboard edge of boot to butt against adjacent structure. (19) Apply wash primer to area around periphery of boot. Coverage area shall include entire exposed adhesive surface and approximately 0.50 inch margin on the boot itself. Wash primer shall be allowed to dry a minimum of one hour. NOTE:
Black polyurethane enamel may be applied as an alternate to 78-U-1003 and U-1001 edge seal in applications where a greater luster and gloss is desired. Refer to Tools, Equipment and Materials approved polyurethane enamels.
(20) Using manufacturer’s instructions, mix edge seal coating components, 78-U-1003 black enamel and U-1001 catalyst, for application over washer primer. If black polyurethane enamel is utilized, apply per manufacturer’s instructions. NOTE:
Edge seal components, 78-U-1003 and U- 1001, are packaged together in a kit and are to be mixed in a ratio of two parts black enamel to one part catalyst. Edge seal is to be thoroughly stirred prior to application.
(21) Apply a uniform film of coating around periphery of the deice boot to edge seal and dress up appearance. Coverage shall include the entire wash primed area. (22) Remove the masking tape immediately following application of coating. NOTE: (23) (24) (25) (26) 7.
Edge sealing shall be accomplished after final paint to aid in protecting leading edge paint from erosion.
Attach hose to air line with clamp. Ensure bleed air tubes are sealed with RTV106 wherever they pass through a rib panel. Secure access panel and heated wing leading edge panel. Airplane may be flown one hour after installing a deice boot, provided deice boots are not operated for 48 hours following bonding.
Pneumatic Deice Flow Control Valve Removal/Installation A.
Remove Pneumatic Deice Flow Control Valve (Refer to Figure 202). NOTE: (1) (2) (3) (4) (5) (6) (7)
All three control valves are mounted on outboard side of left wing root. Each valve may be removed separately.
Remove wing root access plate 511AB, located under left wing. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Ensure airplane electrical power is OFF. Tag for identification and remove electrical wires from solenoid (39) and pressure switch (22) on flow control valve being removed. Disconnect all lines at flow control valve (41) fittings. Plug and cap all open lines and fittings. Remove flow control valve (41) fittings if flow control valve is to be replaced. Remove screws (37) securing flow control valve (41) to bracket (40). Remove flow control valve from airplane.
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MODEL 208 MAINTENANCE MANUAL B.
8.
Install Pneumatic Deice Flow Control Valve (Refer to Figure 202). (1) Ensure airplane electrical power is OFF. (2) Install fittings to flow control valve (41), if removed, and clock in proper direction. (3) Position flow control valve (41) in place. install screws (37). (4) Remove plugs and caps from lines and fittings and connect lines to flow control valve (41) fittings. (5) Connect electrical wires and remove tags. (6) Install wing root access plate 511AB under left wing. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Pneumatic Deice Pressure Switches Removal/Installation A.
Remove Pneumatic Deice Pressure Switches (Refer to Figure 202). NOTE: (1) (2) (3) (4) (5)
B.
9.
Ensure airplane electrical power is OFF. Remove wing root access plate 511AB, located under left wing. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Determine which pressure switch (22) is to be removed. Tag for identification and remove electrical wires from pressure switch. Rotate pressure switch (22) counterclockwise until fully disengaged from block (23) and remove from airplane. Plug block (23) to prevent entry of foreign material.
Install Pneumatic Deice Pressure Switches (Refer to Figure 202). (1) Ensure airplane electrical power is OFF. (2) Remove plug from block (23). (3) Position pressure switch (22) to block (23) and rotate clockwise until fully seated. (4) Identify electrical wires, remove tags and connect to pressure switch (22). (5) Install wing root access plate 511AB under left wing. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Pneumatic Deice Timer Removal/Installation A.
Remove Pneumatic Deice Timer (Refer to Figure 202). NOTE: (1) (2) (3) (4)
B.
10.
Three pressure switches are mounted in left wing root area. Each pressure switch is installed in line, downstream of flow control valves, and may be removed separately.
Timer (26) is mounted on outboard side of left wing root.
Remove wing root access plate 511AB, located under left wing. Ensure airplane electrical power is OFF. Disconnect electrical connector (27) from timer (26). Remove four screws (25) securing timer (26) to wing root and remove from airplane.
Install Pneumatic Deice Timer (Refer to Figure 202 ). (1) Align timer (26) with nutplates in wing root. (2) Install four screws (25) to secure timer (26). (3) Connect electrical connector (27) to timer (26). (4) Install wing root access plate 511AB under left wing. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
Bleed Air System Pressure Regulator Removal/Installation NOTE:
For removal/installation of pressure regulator, refer to Chapter 36, Pneumatic Distribution Maintenance Practices.
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Pneumatic Deice System Installation Figure 202 (Sheet 1)
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Pneumatic Deice System Installation Figure 202 (Sheet 2)
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Pneumatic Deice System Installation Figure 202 (Sheet 3)
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11.
Pneumatic Deice System Adjustment/Test NOTE:
Use filtered, regulated shop air to perform the following tests. Refer to Chapter 36, Pneumatic Distribution - Maintenance Practices, for procedures on connecting shop air to pneumatic system. If shop air is not available, the following warning must be complied with.
WARNING: In order to perform some of the following test procedures, the engine must be operating. Do not stand or allow anyone else to stand close to the arc of the airplanes propeller while conducting these test procedures. A.
Electrical Test. (1) Engage deice boot circuit breaker. (2) Place deice boot press switch to OFF (center) position. (3) Position battery master switch to ON. (4) Press annunciator panel test switch to check light circuit and lamps. (5) With engines running, momentarily position deice boot press switch to AUTO (up) position and verify following conditions are exhibited sequentially. (a) Boots on vertical fin and horizontal stabilizers inflate for a period of six seconds. (b) Boots on inboard wings inflate for a period of six seconds. (c) Boots on outboard wings and wing struts inflate for a period of six seconds. (d) System deactivates. (6) Repeat step 12.A.(5) and verify deice pressure light, located on annunciator panel, remains illuminated for a period of six seconds during each of the three sequential cycles and extinguishes momentarily between cycles. (7) Position and hold deice boot press switch to manual (down) and verify all boots inflate simultaneously and deice pressure light illuminates. (8) Position wing light switch to ON (up) and verify ice detector light is illuminated. (9) Shut off engine.
B.
Pneumatic Deice System Conductivity Test. NOTE:
(1) (2) (3) (4)
This test will apply a voltage charge directly to the surface of deice boot and measure resistance to airplane structure to determine if the resistance is within a specified megohm range.
Obtain a megohmmeter capable of applying 500 volts and set to 500 volts. Place a wetted cloth, approximately the size of a paper towel, over surface of deice boot. Place one probe of megohmmeter in contact with cloth. Place other probe in contact with an exposed rivet head near point of contact with first probe. NOTE:
(5) (6) (7) (8) C.
Do not apply second probe to a painted rivet head, as it needs to make good contact with structure. It may be necessary to open an inspection panel and use a screw or rivet inside wing.
Obtain three measurements along length of wing boots. Obtain two measurements along tail and strut boots. Ensure measurements are all less than 1 megohm to approximately 10 megohms. Maximum allowable measurement is 100 megohms. If a boot reading above 100 megohms is present, boot must be replaced.
Pneumatic Deice System Timer Test. (1) Check for system voltage between pins 3 and 1 of deice timer with deice switch OFF. Verify voltage is indicated. (2) Check for system voltage between pins 6 and 1 of deice timer with deice switch ON. Verify voltage is indicated.
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5) (6) (7) D.
Check for system voltage between pins 8 and 1 of deice timer with deice switch ON. Verify voltage is indicated as timer cycles. Check for system voltage between pins 9 and 1 of deice timer with deice switch ON. Verify voltage is indicated as timer cycles. Check for system voltage between pins 7 and 1 of deice timer with deice switch ON. Verify voltage is indicated as timer cycles. If system voltage is not present between any of pins tested, timer is defective and must be replaced. Refer to Pneumatic Deice Timer Removal/Installation. Place all switches in OFF position.
Air Leakage Test. NOTE: (1) (2) (3) (4)
Air leakage test can be performed in engine compartment.
Disconnect air pressure supply line from engine bleed air system pressure regulator (outlet side). Disconnect vacuum hoses from vacuum regulator, located on aft side of firewall. Remove vacuum regulator and cap line. Remove vacuum ejector and install a union and cap. Connect a source of clean air to end of supply line. NOTE:
(5)
Disconnect electrical power leads from each of three flow control valve solenoids. NOTE:
(6) (7) (8) (9) 12.
Inlet pressure must be a minimum of 18 to 20 PSIG to perform this test. Include a pressure gage in air line to observe system pressures.
For testing and troubleshooting, one flow control valve solenoid may be actuated at a time to test and isolate each system.
Apply 18 PSI pressure to system and, by means of an in-line hand operated valve, trap pressure in deice system. Observe system for leakage and verify leakage rate does not exceed a pressure drop of 3.0 PSI per minute. Ensure all deice boots inflate and no leaks are present. Remove test equipment. Lubricate all threads and connect all previously disconnected components. Remove the 28 VDC electrical source from flow valves and reconnect airplane electrical system.
Cleaning
CAUTION: Only use the instructions in this section when you clean deice/anti-ice boots. Disregard instructions which recommend petroleum base liquids (methyl n-propyl ketone, unleaded gasoline, etc.) which can harm the boot material if you allow it to soak on the material. A.
Clean boots with mild soap and water, then rinse them thoroughly with clean water. NOTE:
B.
Isopropyl alcohol or toluene can be used to remove grime which cannot be removed with soap. If isopropyl alcohol or toluene is used for cleaning, wash each area with mild soap and water, then rinse the area thoroughly with clean water and allow the boots to dry completely.
Apply a layer of Age Master No. 1 on each wing and stabilizer leading edge deicing boot in accordance with the Age Master No. 1 application instructions, every 150 flight hours or six months, regardless of the operation status or climatic conditions. Apply a coating of Age Master No. 1 to the boots in accordance with application instructions on the container and allow the boots to dry. NOTE:
Age Master No. 1 is beneficial for its ozone and weather resistance features.
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After boots have been treated with Age Master No. 1, apply a coating of ICEX II to the boots in accordance with application instructions on the ICEX II container. NOTE: (1)
ICEX II may be beneficial as an ice adhesion depressant. Both Age Master No. 1 and ICEX II are distributed by BFGoodrich Company.
Apply a layer of ICEX II every 50 flight hours or less on the leading edge deicers; apply a layer every 15 hours or less on the propeller boots.
CAUTION: ICEX II contains silicone, which lessens paint adhesion. Use care when applying ICEX II and protect the adjacent surfaces from overspray. An overspray of ICEX II will make touchup painting almost impossible. AGE MASTER No. 1 and ICEX II last approximately 50 hours on wing and stabilizer deice boots and 15 hours on propeller anti-ice boots. 13.
Approved Repairs (Cold Patch) NOTE:
14.
Surface coatings and surface refurbishing kits will not repair leaks. Use repair kit materials. When repairing deice boots utilizing patch, exercise care to prevent trapping air beneath patch. Should air blisters appear after boots have been installed for a length of time, it is permissible to cut a slit in deice boot. Slit and repair are only appropriate if blister is a result of surface ply delamination. If blister is a result of debonding or stitch line failure, this repair is not appropriate. If it is a delamination air blister, it is recommended that slit be no larger than 0.75 inch, or within 0.125 inch of a stitch line, otherwise deice boot should be replaced. An alternate method of repair is to peel deice boot back using Methyl n-Propyl Ketone solvent and reapply using normal adhesives.
A.
Repair surface damage using 74-451- AA Universal Repair Kit. Repair instructions are provided with each individual package of patches. Refer to Ice and Rain Protection - General, Tools, Equipment and Materials.
B.
Repair pinholes using 74-451-AE Pinhole Repair Kit. Repair instructions are provided with each individual package of patches. Refer to Ice and Rain Protection - General, Tools, Equipment and Materials.
Pneumatic Deice Boot Winterization Procedure A.
Winterize Deice Boots. (1) Winterizing pneumatic deice system several weeks prior to flight into freezing weather, followed by periodic maintenance during freezing season, improves reliability, while reducing chance of moisture freezing within deice boots. (2) Moisture can enter deice system through unrepaired pin size holes, cuts or abrasions when vacuum is applied to boots. Moisture can also be caused by condensation of compressor bleed air during boot activation. Accumulated moisture can pool at low points within system lines, fittings and valves and then freeze, causing blockage. Moisture in system can also result in deterioration and sticking of pneumatic deice flow control valves. (3) Proper cleaning and treating procedures protect exterior surface of boots from ultraviolet rays and ozone exposure while retarding ice adhesion during flight. Refer to BFGoodrich Service Newsletter 91-015. (4) Inspect deice boots for weathering, unglued edges and damage. Repair deice boots as required. NOTE:
(5) (6)
All holes found in deice boots should be patched, as this is the first line of defense in preventing moisture entry into deice system; however, do not attempt to extend life of deice boots in poor condition. Replace deice boots which have numerous patches, deteriorated areas or holes.
Work moisture out of strut deice boots. Work moisture from inboard and outboard wing deice boots.
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9)
Drain moisture from supply lines located inside left inboard wing access panel. Drain stabilizer deice supply line at cross fitting connection in tailcone. Perform an operational check of system after steps 15.A.(1) through (8) have been accomplished. (10) Obtain filter/regulator/lubricator specified in Tools, Equipment and Materials. An equivalent substitute may be utilized. (11) Mix a 50/50 solution of isopropyl alcohol (TT-l-735A) and Dow Corning 200, 100cs industrial grade fluid. Fill reservoir of lubricator with mixed solution. (12) Disconnect deice supply line at elbow fitting, located inside cockpit at firewall, immediately forward of pilot's position. Connect and apply shop air supply to filter/regulator/lubricator and preset regulator for a maximum of 17 to 19 PSlG. Remove shop air supply and connect filter/regulator/lubricator to disconnected line using necessary adapter fitting(s). NOTE:
Attach filter/regulator/lubricator as close to disconnected supply line as possible for optimum lubricator operation.
(13) Remove access panel at inboard leading edge of left wing to expose pneumatic deice flow control valves. Temporarily install a hose on each valve outlet port for drainage collection. (14) With regulator preset to 17 to 19 PSIG maximum pressure, connect shop air to inlet port on filter/regulator/lubricator and begin to apply 17 to 19 PSlG through lubricator containing mixed solution.
CAUTION: TT-L-735A isopropyl alcohol is flammable by itself or in mixture with Dow Corning 200, 100CS fluid. Do not allow Dow Corning 200 solution to come in contact with painted surfaces of airplane. Silicone based compounds, such as Dow Corning 200, will impair the ability to paint/refinish any surface it contacts. (15) Look for solution drainage at hoses attached to outlets of pneumatic deice flow control valves to ensure solution (mist from lubricator) is being introduced into boots. (16) Cycle deice boots for seven minutes to inject a fine mist of solution throughout deice system. Approximately 17 cycles will be required. NOTE:
Begin cycling boots as soon as shop air is connected. Wait six seconds after a cycle is complete and reactuate deice switch. When the system is not being cycled, mixture is being pumped directly through flow control valve, into discharge container. After cycling deice boots, watch for fluid weeping from pin size holes or cuts in boots. Mark locations of needed repairs. Repair holes immediately to avoid further moisture contamination.
(17) Shut off deice system and remove/disconnect lubricator from supply line. Allow supply line to drain, then reconnect supply line to filter/regulator/lubricator. (18) Using hands or rollers, work boots to aid in coating inside surfaces. Work outboard to inboard on wing and stabilizer deice boots; bottom to top on strut deice boots. (19) Operate deice boots again through five complete cycles to remove any pooling of solution. (20) Disconnect drain hoses from pneumatic deice flow control valves. Disconnect filter/regulator/ lubricator from supply line. Reconnect supply line to firewall fitting and restore deice system. (21) Install access panel on inboard leading edge of left wing. (22) Repeat procedure every 100 hours or as required throughout freezing season. (23) To increase deice boot life and improve ice removal in flight, regularly wash boots with a mild soap and warm water. If necessary, isopropyl alcohol may be used to clean boot surfaces. Apply Age Master No. 1 every 150 hours or six months, regardless of operation status or climatic conditions. Apply Icex every 50 hours, or less, (every 15 hours or less on propeller boots) during icing season.
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15.
Pneumatic Flow Control Valves Disassembly/Cleaning/Assembly Procedure A.
Disassemble/Clean Pneumatic Flow Control Valve (Refer to Figure 202). (1) Remove access panel 511AB on left inboard wing leading edge. Refer to Chapter 6, Access Plates/Panels Identification. (2) Remove flow control valve (41). Note clocking of drain port and bracket to aid in reassembly. Refer to Pneumatic Deice Flow Control Valve Removal/Installation. (3) Remove solenoid (65) from valve body assembly (57). Wrenches should be positioned on solenoid hex and on body assembly wrench flats. (4) Remove overboard reservoir assembly (63), compression washers (62) and bracket (59). Clean parts using isopropyl alcohol. (5) Remove stainless ball (61) and spring (60). Clean parts using isopropyl alcohol. (6) Remove pushrod (64) from solenoid (65), if not secured with epoxy, and clean using isopropyl alcohol. (7) Check all fittings for tightness and ability to seal against leakage. (8) Check and clean seats for compression washers (62) in overboard reservoir assembly (63).
CAUTION: Do not intermix parts between components. Parts removed from one component for cleaning must be reinstalled in the same component. (9) (10) (11) (12) (13) B.
Check seal of compression washers (62). Clean valve assembly components, except solenoid (65), using isopropyl alcohol. Check mounting brackets for cracks and damage. Check electrical wiring to solenoid assembly for evidence of wear and chafing. Using a soft brush, remove any dirt accumulated on solenoid (65), particularly in pushrod (64) area.
Assemble Pneumatic Flow Control Valves (Refer to Figure 202). (1) If pushrod (64) has not previously been epoxied to solenoid, seat pushrod (64) in solenoid (65) and epoxy using EA9309 or equivalent. Mix adhesive per manufacturer’s instructions. Remove excess adhesive from pushrod shaft and allow to cure per manufacturer’s instructions. (2) Install new compression washers (62). (3) Lubricate all valve components, except solenoid (65), using full strength Dow Corning 200, 100cs lubricant. Ensure spring (60) and stainless ball (61) are thoroughly coated. (4) Insert spring (60) and stainless ball (61) into overboard reservoir assembly (63). (5) Assemble solenoid (65), overboard reservoir assembly (63), bracket (59) and valve body assembly (57). Torque per decal (58) located on valve body. NOTE: (6)
(7)
Ensure overboard reservoir assembly (63) and bracket (59) are clocked correctly prior to torquing.
Perform and operational test of flow control valves (41). (a) Remove cover from top of solenoid (65) and depress plunger, while simultaneously checking movement of internal assembly at inlet port. Verify assembly shifts approximately 0.02 inch. (b) Verify compression washer (62) provides sufficient seal to prevent random movement of overboard reservoir assembly (63) on bracket (59). (c) Connect solenoid (65) to adjustable source of 28 VDC power. (d) Apply rated pressure of 18 PSIG to valve inlet port. Verify vacuum is present at deicer port. (e) Energize solenoid (65) and verify valve shifts such that pressurized air is directed to deicer port with none being dumped overboard. (f) Adjust power with minimum 24 VDC and maximum 32 VDC. Energize solenoid (65) at each setting and verify valve shifts satisfactorily from vacuum to pressure supply at deicer port. Reinstall valve(s) on airplane. Refer to Pneumatic Deice Flow Control Valve Removal/ Installation.
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Reinstall access panel 511AB on left inboard wing leading edge. Refer to Chapter 6, Access Plates/Panels Identification.
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MODEL 208 MAINTENANCE MANUAL PNEUMATIC SURFACE DEICE - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the pneumatic surface deice system in a serviceable condition.
Task 30-10-00-720 2.
Bleed Air Pressure Regulator Functional Check A.
General (1) This task gives the procedures to do a bleed air pressure regulator functional check.
B.
Special Tools (1) Filtered Shop Air (2) Flexible Hose (3) Pressure Gage
C.
Access (1) Open the right engine cowling door.
D.
Do the Bleed Air Pressure Regulator Functional Check. (1) For the procedures necessary to do the bleed air pressure regulator functional check, refer to Chapter 36, Pneumatic Distribution - Maintenance Practices.
E.
Restore Access (1) Close the right engine cowling door. End of task
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE SYSTEM - DESCRIPTION AND OPERATION Cargo Pod Installation 1.
General A.
2.
The TKS anti-ice system is a fluid anti-ice system to prevent ice formation on the leading edges of the wings, horizontal stabilizers, struts, vertical stabilizer, propeller, and the windshield. A monoethylene glycol/isopropyl alcohol/deionized water solution is used to anti-ice the airframe surfaces and windshield in flight. The fluid solution is a freezing point depressant that is swept rearward over the surfaces and prevents ice buildup. For a list of approved TKS anti-icing fluids, refer to Chapter 12, Replenishing - Description and Operation.
Description A.
Laser-drilled titanium panels are mounted to the leading edges of the wings, wing struts, horizontal and vertical stabilizers that give TKS ice protection for the Caravan. The propeller is protected with a fluid slinger ring and the windshield is protected with a fluid spray bar. The TKS anti-ice system is divided into two subsystems; the airframe anti-ice system and the windshield anti-ice system. Refer to Figure 1 and Figure 2. (1) Anti-ice fluid solution comes out of the airframe anti-ice system through flush-fitting laser drilled titanium leading edge panels on the wings, stabilizers, and struts. The airframe anti-ice system applies anti-icing fluid to the wing leading edge, that has three panels on each wing, two panels on each strut, and one panel on each horizontal and vertical stabilizer leading edge. The system provides full coverage of the leading edge of the wings, lift struts, horizontal and vertical stabilizer, excluding the dorsal fin. The airframe system also includes the propeller slinger application anti-icing system. (a) The outer skin of the ice protection panels are manufactured with titanium, 0.9 mm thick. Titanium provides excellent strength, durability, light weight, and corrosion resistance. (b) The panel skin is perforated by laser drilled holes, 0.0025 inches in diameter, 800 per square inch. The porous area of the titanium panels is designed to cover the stagnation point travel on the appropriate leading edge over a normal operating environment. (c) The back plates of the porous panels are manufactured with 0.7-mm thick titanium. They are formed to create reservoirs for the ice protection fluid to supply the entire porous area. A porous membrane between the outer skin and the reservoir gives even flow and distribution through the entire porous area of the panel. (d) The porous panels are bonded or attached as a cuff over a leading edge. Panels are bonded to the airframe with a two-part flexible adhesive. (e) Fluid is supplied to the panels and propeller by two positive displacement, constant volume metering pumps. The pumps give various flow rates to the panels and propeller. Single pump operation, a combined pump mode, and timed pumping provide a range of flow rates for different icing conditions. (f) The fluid passes through microfilters before it gets to the porous panels and propeller. The filter removes contaminants from the fluid and prevents panel blockage. A system of nylon tubing carries the fluid from the fluid tank to the proportioning units that divide the flow into the volumetric requirements of each panel or device supplied through the unit. The proportioning units are located in the wings, fuselage, and tail of the aircraft and feed each panel and device through nylon tubing. (g) The system has a fluid tank that gives a minimum ice protection endurance when filled. The endurance capacity is more than the endurance guidelines of AC 23.1419-2C. The tank also serves as an attachment structure for the main metering pumps, windshield pump, filters, and additional hardware. The combination of equipment creates an equipment pack assembly to ease installation. The tank assembly is mounted to the belly of the aircraft in the mid cargo pod area. Refer to Figure 5. (h) The fluid tank is equipped with a low level switch, that gives a warning annunciation at a predetermined fluid level. The annunciation level is when only 20 minutes of fluid remains in the tank with the system in the normal operation mode. (i) An external filler for the fluid tank is on the left side of the cargo pod. Refer to Chapter 12, Replenishing - Description and Operation.
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TKS Anti-Ice System Components Figure 1 (Sheet 1)
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TKS Anti-Ice System Flow Diagram Figure 2 (Sheet 1)
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TKS Anti-Ice System Instrument Panel Operation Devices With G1000 Systems Figure 3 (Sheet 1)
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TKS Anti-Ice System Instrument Panel Operation Devices Without G1000 System Figure 4 (Sheet 1)
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TKS Anti-Ice System Fluid Tank Figure 5 (Sheet 1)
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(2)
The system is operated through a series of three control switches. All modes of operation and selection for the metering pumps and the windshield pump are controlled through these devices. Refer to Figure 3 and Figure 4. (k) The operational state can be monitored with: 1 Annunciators on the instrument panel of the non G1000 TKS system. This system annunciation is done through a 3-element annunciator light array. Indications of normal operation state, cautionary state, and warning conditions will be displayed when necessary. These annunciators are independent of the other aircraft annunciators. The fluid level for this system is monitored with a 270° sweep reservoir contents gage that is coupled to a capacitive level sender. Refer to Tables 1-5 and Figure 4. 2 CAS messages and indications on the MFD with the G1000 TKS System. The windshield anti-ice system applies anti-icing fluid through a spray bar to the pilot's windshield. Refer to Chapter 30, TKS Anti-Ice Windshield Spray Bar. (a) Fluid for the windshield spray bar system comes from an on-demand gear pump that is attached to the fluid tank. The spray bar can be operated as needed to clear forward vision through the windshield.
B.
The system is configured with two main metering pumps. The pumps give both the delivery mechanism for all modes of operation of the system, and a pump backup system. The modes of operation are (1) NORMAL, (2) HIGH, (3) MAXIMUM, and (4) BACKUP. (1) HIGH mode is the design flow rate for the system and occurs when one pump is run continuously. (2) MAXIMUM mode is a flow rate that is used for a intermittent maximum icing condition, and occurs when both pumps run continuously. MAXIMUM mode is twice the flow rate of HIGH mode. (3) NORMAL mode is 66% of the HIGH or design flow rate, and happens when both pumps run for a time cycle of 17% on and 83% off. (4) The final mode is BACKUP. In the event that a pump fails, one of the pumps will be available and capable to pump the design flow rate to the system. The BACKUP system provides power to the second pump, independent of the circuit used for the other modes.
C.
The operation of the TKS anti-ice systems is controlled by three switches on the left panel. The switches are PRIMARY, MAX FLOW, and BACKUP. Refer to Figure 3 and Figure 4.
D.
The airframe and windshield spray bar anti-ice systems share the anti-icing fluid tank which is in the cargo pod. The fluid tank assembly is attached to the belly of the aircraft in the second bay area of the cargo pod. The assembly is accessible through the cargo pod doors on the left side of the pod. Refer to Figure 1, Figure 2, and Figure 5. (1) The tank anti-ice fluid level is monitored with a gage that is on the left meter panel in the cockpit on the non G1000 TKS system, and is monitored with an indication on the MFD on the G1000 system. The fluid level monitor devices show the total fluid available for operation of both the airframe and windshield spray bar anti-ice systems. The tank fluid level monitor devices are electrically operated and receive inputs from a capacitance sensing level sender probe in the fluid tank. Refer to Figure 2, Figure 3, Figure 4, and Figure 5. (2) In addition to the fluid level monitors, the tank has a low level switch. Refer to Figure 2 and Figure 5. (a) The low level switch controls operation of the ANTI-ICE annunciator CAUT (amber light) on the annunciator panel for the non G1000 TKS system. The low level switch illuminates the annunciator CAUT when there is only enough anti-ice fluid in the tank to last for approximately 20 minutes of full system continuous operation. (b) The low level switch is monitored by the G1000 system and shows the CAS messages A-ICE LOW FLUID (amber) when the anti-ice fluid is low. (3) The anti-ice fluid tank also has a sight glass that gives a visual indication of the fluid level in the tank to assist when you fill the tank. Refer to Figure 4 and Figure 5.
E.
The airframe and windshield spray bar anti-ice system have pumps, filters, and a high pressure switch that are installed on the fluid tank in the cargo pod. Refer to Figure 4. (1) The anti-ice windshield spray bar pump and the two airframe pumps are electric motor driven. (2) Filters are installed downstream of the two airframe pumps. Each filter contains a replaceable element. The filter ports are marked IN and OUT for correct plumbing connection.
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3.
A high pressure switch is installed in line with the metering pumps. Refer to Figure 1, Figure 2, and Figure 5.
F.
Proportioning units are installed in four different locations on the airplane. Refer to Figure 2. (1) A seven-place proportioning unit is found in each wing leading edge near the strut attach fitting. (2) A seven-place proportioning unit is found (3) A single-place proportioning unit is found in the feed line to the propeller, under the floor near the copilots seat. (4) A three-place proportioning unit is found on the floor of the tail cone (RBL 3.35, FS 422.75). This proportioning unit supplies the vertical stabilizer and each horizontal stabilizer.
G.
The proportioning units are metering units which supply anti-icing fluid at a predetermined flow rate for each individual porous leading edge panel. The proportioning units incorporate a manifold with calibrated capillary tubes which meter the fluid through the outlet ports. The outlets are marked 1, 2, 3, 4, 5, 6, and 7 on each wing's seven-place proportioning unit. Plumbing to the outlet ports must be connected as specified for proper operation. Refer to Figure 2.
H.
A total of three pressure switches are installed in the TKS anti-ice system plumbing. There are two low pressure switches and one high pressure switch in the system. The pressure switches transmit signals to annunciator lights on the non G1000 models and to CAS display messages on the G1000 models. (1) One high pressure switch is installed downstream of the two surface metering pumps in the cargo pod. Refer to Figure 5. (a) On the non G1000 TKS system, the high pressure switch is electrically connected to the CAUT (amber) light on the ICE FLUID annunciator. Closure of the high pressure switch will cause the ICE FLUID CAUT annunciator to come on. (b) On the G1000 system, the high pressure switch is monitored by the G1000 system and shows the CAS messages A-ICE HIGH PRESS (amber) when the anti-ice fluid pressure is high. (2) There are two low pressure switches to monitor the horizontal stabilizers leading edge panels; one pressure switch for each panel. (a) On the non G1000 TKS system, the low pressure switches are electrically connected to the WARN (red) light on the ICE FLUID annunciator. Closure of the low pressure switch will cause the ICE FLUID WARN annunciator to come on. (b) On the G1000 system, the low pressure switches are monitored by the G1000 system and shows the CAS messages A-ICE LOW PRESS (red ) when the anti-ice fluid pressure is low.
I.
There are a total of two check valves installed in the fluid anti-ice system plumbing network downstream of the metering pumps in the cargo pod. The check valves prevent reverse of fluid flow through the plumbing.
J.
There is a solenoid vale installed between the fluid tank and the windshield pump. The valve makes sure that fluid does not run back in the tank when the pump stops operation. (1) There is a strainer installed at the fluid tank for the windshield pump.
Operation A.
Operation of the TKS anti-ice system is controlled by three switches on the left meter panel. The switches are PRIMARY, MAX FLOW, and BACKUP. Refer to Figure 3 and Figure 4.
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MODEL 208 MAINTENANCE MANUAL B.
There are a total of 18 different switch position configurations possible with the three TKS ANTI-ICE system control switches. Of the 18 possible switch configurations, only six are considered normal. These switch configurations are shown in the TKS Anti-Ice System Operation Matrix in Table 1 and Table 3. NOTE:
All of the switch configurations recorded in the tables below are true when the electrical power is applied to the airplane, the ICE circuit breakers are on, and the anti-ice system starts operation.
NOTE:
The MAX FLOW switches only operate momentary when depressed.
NOTE:
Timer: • Number one comes on for 20 seconds and turns off, and repeats every 100 seconds • Number two comes on for 120 seconds and then turns off • Number three comes on for four seconds and then turns off.
NOTE:
Table 2, and Table 4 give the G1000 CAS Message Triggers and Annunciator Triggers for the G1000 and non G1000 TKS anti-ice systems, respectively. Refer to Table 2 and Table 4
NOTE:
The MAX FLOW only operates with the NORM or HIGH switch ON.
Table 1. Pumps Operation Matrix for the TKS Anti-Ice System With the G1000 Pumps Operation Matrix For the TKS Anti-Ice System With the G1000 CONTROL SWITCHES PRIMARY Off Norm
High
MAX FLOW
#2
OFF
INT
INT
OFF
ON
TRIP
OFF
TRIP
#2 Max Flow
#3 A-ICE A-ICE Wind- NORM HIGH Shield (white) (white)
OFF
ON
OFF
OFF
ON
OFF
OFF
OFF
OFF OFF
OFF
OFF
ON
INT
INT
OFF
ON
ON
OFF
ON
OFF
OFF
ON
INT
OFF
OFF ON
OFF
OFF
ON
TRIP
OFF
INT
INT
ON
ON
OFF
ON
ON
OFF
TRIP
OFF
ON
OFF
ON
OFF OFF
ON
OFF
ON
OFF
OFF OFF
OFF
OFF OFF
OFF
OFF
OFF
OFF
OFF OFF
ON
OFF OFF
ON
OFF
OFF
OFF
OFF OFF
OFF
OFF OFF
OFF
OFF
OFF
ON
INT
ON
OFF
ON
OFF
OFF
ON
OFF
ON
ON
ON
OFF
OFF OFF
OFF
OFF
ON
TRIP
ON
INT
ON
OFF
ON
ON
OFF
ON
OFF
TRIP
ON
ON
ON
OFF
OFF ON
OFF
OFF
ON
ON TRIP
TRIP
*** ***
ON ON ON ON
WindShield
#1
ON
***
ANNUNCIATORS G1000 CAS MESSAGE
#1
ON
ON
TIMERS
BACK UP
AirWindFrame Shield
ON
ON
PUMPS
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MODEL 208 MAINTENANCE MANUAL Table 1. Pumps Operation Matrix for the TKS Anti-Ice System With the G1000 (continued) Pumps Operation Matrix For the TKS Anti-Ice System With the G1000 CONTROL SWITCHES PRIMARY Off Norm
PUMPS
MAX FLOW
High
ON
#1
#2
WindShield
#1
#2 Max Flow
#3 A-ICE A-ICE Wind- NORM HIGH Shield (white) (white)
ON
INT
ON
ON
ON
OFF
ON
ON
OFF
TRIP
ON
ON
ON
ON
OFF OFF
ON
OFF
ON
ON
OFF ON
OFF
OFF OFF
OFF
OFF
OFF
ON
OFF ON
ON
OFF OFF
ON
OFF
OFF
ON
OFF ON
OFF
OFF OFF
OFF
OFF
OFF
TRIP
*** *** NOTE:
G1000 CAS MESSAGE
TRIP
TRIP
***
ANNUNCIATORS
BACK UP
AirWindFrame Shield
ON
TIMERS
INT = Intermittent
Table 2. Operation Matrix for the TKS Anti-Ice System With the G1000 TKS with G1000 CAS Message Triggers SWITCH LOW LEVEL SWITCH
LOW PRESSURE SWITCH
G1000 CAS Message HIGH PRESSURE SWITCH
A-ICE LOW PRESS (red)
A-ICE HI PRESS (AMBER)
A-LOW FLUID (AMBER)
ON
OFF
OFF
OFF
ON
OFF
OFF
OFF
ON
ON ON ON
Table 3. Operation Matrix for Field Installed TKS Anti-Ice System Without the G1000 Operation Matrix For the TKS Anti-Ice System Without the G1000 CONTROL SWITCHES PRIMARY
MAX FLOW
PUMPS
#1
#2
Wind- #1 Shield
ANTI- CAUT WARN #2 #3 Max Wind- ICE Flow Shield ON
OFF
INT
INT
OFF
ON
OFF OFF
ON
OFF
OFF
OFF
ON
OFF
OFF
OFF OFF OFF
ON
OFF
OFF
TRIP
OFF
INT
INT
OFF
ON
ON
OFF
ON
OFF
OFF
TRIP
OFF
ON
INT
OFF
OFF ON
OFF
ON
OFF
OFF
TRIP
OFF
ON
ON
ON
ON
OFF ON
ON
OFF
OFF
TRIP
OFF
ON
ON
ON
OFF OFF ON
ON
OFF
OFF
ON ON
ON ON ON
ANNUNCIATORS
BACK UP
WindOff Norm High AirFrame Shield
ON
TIMERS
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MODEL 208 MAINTENANCE MANUAL Table 3. Operation Matrix for Field Installed TKS Anti-Ice System Without the G1000 (continued) Operation Matrix For the TKS Anti-Ice System Without the G1000 CONTROL SWITCHES PRIMARY
MAX FLOW
PUMPS
TRIP
ANTI- CAUT WARN #3 #2 Max Wind- ICE Flow Shield ON
OFF OFF OFF
OFF
OFF
OFF
OFF
OFF OFF
ON
OFF OFF ON
OFF
OFF
OFF
OFF
OFF OFF
OFF
OFF OFF OFF
OFF
OFF
OFF
ON
INT
ON
OFF
ON
OFF OFF
ON
OFF
OFF
ON
ON
ON
OFF
OFF OFF OFF
ON
OFF
OFF
TRIP
ON
INT
ON
OFF
ON
ON
OFF
ON
OFF
OFF
TRIP
ON
ON
ON
OFF
OFF ON
OFF
ON
OFF
OFF
TRIP
ON
INT
ON
ON
ON
OFF ON
ON
OFF
OFF
TRIP
ON
ON
ON
ON
OFF OFF ON
ON
OFF
OFF
ON
OFF ON
OFF
OFF OFF OFF
OFF
OFF
OFF
ON
OFF ON
ON
OFF OFF ON
OFF
OFF
OFF
ON
OFF ON
OFF
OFF OFF OFF
OFF
OFF
OFF
ON ON
ON ON TRIP
TRIP
***
Wind- #1 Shield OFF
***
***
#2
OFF OFF
TRIP
ON
#1
OFF
***
ON
ANNUNCIATORS
BACK UP
WindOff Norm High AirFrame Shield ***
TIMERS
*** NOTE:
The field installed TKS without the G1000: ANTI-ICE ON annunciator is white, the CAUT annunciator is amber, and the WARN annunciator is red.
NOTE:
INT = Intermittent
Table 4. Operation Matrix for Field Installed TKS Anti-Ice System Without the G1000 TKS without G1000 Annunciator Triggers SWITCH LOW PRESSURE SWITCH
LOW LEVEL SWITCH
ANNUNCIATORS HIGH PRESSURE SWITCH
ANTI-ICE ON
CAUT
WARN
ON
ON
OFF
ON
ON
OFF
ON
OFF
ON
ON ON ON
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the TKS anti-ice system in a serviceable condition.
Task 30-11-00-720 2.
TKS Anti-Ice System Functional Check A.
General (1) This task gives the procedures to do a TKS anti-ice system functional check.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the TKS Anti-Ice System Functional Check. (1) For the procedures necessary to do a functional check of the TKS anti-ice system for airplanes with the fairing installation, do the TKS Anti-Ice System Test. Refer to TKS Anti-Ice System Adjustment/Test (Fairing Installation). (2) For the procedures necessary to do a functional check of the TKS anti-ice system for airplanes with the pod installation, do the TKS Anti-Ice Fluid Tank Component Test. Refer to TKS Anti-Ice Fluid Tank Components - Adjustment/Test (Pod Installation).
E.
Restore Access (1) None End of task Task 30-11-00-721 3.
Inboard TKS Wing Panel Pressurization Functional Check A.
General (1) This task gives the procedures to do a pressurization functional check of the Inboard TKS Wing Panels.
B.
Special Tools (1) TKS System Test Cart. Refer to Ice and Rain Protection - General.
C.
Access (1) Remove wing access panels as needed to get access to the inboard wing panel ßuid Þttings.
D.
Do the Inboard TKS Wing Panel Pressurization Functional Check. (1) Use the Porous Panel Purge and Test procedures to operate the inboard TKS wing panels at a pressure of 60 psi (413 kPa). Refer to TKS Anti-Ice Leading Edge Porous Panel - Adjustment/ Test. NOTE: (2)
When the inboard ports are pressurized, the outboard port must be capped.
With 60 psi (413kPa) applied to the inboard wing panels, examine the top and bottom trailing edges of the inboard TKS wing panels. (a) Make sure there is no separation between the trailing edges of the panel and the wing leading edge. (b) If separation between the wing panels and the wing leading edge is found, the TKS porous panel must be replaced. Refer to TKS Anti-Ice Leading Edge Porous Panel - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL E.
Restore Access (1) Remove any caps or plugs that were installed on the supply tubes. (2) Connect the panel supply tubes to the porous panels. Refer to TKS Anti-Ice Leading Edge Porous Panel - Adjustment/Test. (3) Install the access panels that were removed. End of task
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MODEL 208 MAINTENANCE MANUAL G1000 AVIONICS AND TKS ANTI-ICE SYSTEM - TROUBLESHOOTING 1.
General A.
2.
3.
This section will help you to troubleshoot the TKS system for airplanes that have the G1000 system installed. Logic troubleshooting charts in this section will guide you to the most likely cause and to some recommended actions for repair.
Troubleshooting Description A.
The TKS system has control switches located on the lower left instrument panel for the non-G1000 configuration, and in the upper left for the G1000 configuration. The non-G1000 TKS system has a fluid level gauge and three annunciators that display the necessary status, warnings and cautions. The G1000 system multifunction display units and the primary flight display units will display the necessary status, warnings, cautions, and fluid quantity information with Crew Alerting System-CAS messages, about the TKS system.
B.
A good understanding of the TKS system is important. You can get more detailed information about the TKS system for airplanes that have the G1000 avionics system installed from the G1000 Avionics System - Description and Operation.
C.
A good understanding of what the discrepancy is also very helpful. There are three different categories of TKS discrepancies. This section has an extensive set of troubleshooting procedures to use but, you can begin your troubleshooting at any procedure that is appropriate for the trouble you are working. (1) The first is the display category. This is any display that shows on the G1000 display units that was not expected, or an unexpected anti-ice quantity indication. (a) If the anti-ice quantity is incorrect, you can test the system using the appropriate procedure in this section. (b) If there is a message that is shown on one of the G1000 display units, you can troubleshoot each one individually using the correct procedure from this section. (2) The second category is the pump category. This means that there is an incorrect operation of one or more of the anti-ice pumps. (a) The check procedures for these discrepancies are found in this section of the troubleshooting. (b) If a pump is not operating correctly, it is recommended that you begin your troubleshooting at the pump section procedures first. (3) The third category is the fluid flow category. This section has the procedures to correct the distribution and flow of the fluid across the porous panels.
Troubleshooting Preliminary A.
Visual Inspections. NOTE:
(1)
It is recommended that you do a visual inspection of the anti-ice components before you do the troubleshooting procedures. The visual inspections in this section do not need to be done in order, you may look at the most likely parts and components as necessary. It is recommended that you do all of these inspections when possible.
Look at all of the nylon tube connections. Look at the condition of the nut, olive, and seal or O-ring. If you must replace a part, refer to the Model 208 Illustrated Parts Catalog.
CAUTION: The length of the nylon tube is important. Refer to the Model 208 Illustrated Parts Catalog for the correct part number. A tube with an incorrect length can make the anti-ice system not operate properly. (2) (3)
Inspect the nylon tubing for fluid leaks or evidence of obstructions. Make sure there are no cuts, holes, or kinks at any of the tubing. Do an inspection of the windshield nozzles and the spray bar. Look for any defects and deformed parts. Look for any evidence of blockage, damage, or other defects.
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MODEL 208 MAINTENANCE MANUAL (4) B.
Look at the anti-ice sight gage ball and make sure there is no evidence of blockage that could prevent fluid from entering the sight gage.
Preliminary Set Up. (1) Make sure the circuit breakers PRMRY ANTI-ICE, W/S ANTI-ICE, AND BACKUP - ANTI-ICE located on the left main circuit breaker panel are engaged. (2) Put the following switches in the switch setting or position that follows: SWITCH NAME
(3) (4) (5) (6)
4.
SETTING / POSITION
External Power (Bus) Switch
ON
Battery (DC Power) Switch
OFF
Fluid Control - Primary
OFF
Fluid Control - Max Flow
OFF
Fluid Control - Backup
OFF
Avionics Bus 1 and 2
ON
Make sure the G1000 is operating correctly and is powered up. Make sure the display units are operating. You can cycle the power to individual equipment until proper operation of all the equipment is operating correctly. Make sure the anti-ice primary switch (SI022) is in the OFF position. Make sure that the anti-ice messages are not displayed on PDF 1 and PFD 2. If an anti-ice message is shown on PFD 1 or PFD 2, refer to Table 101. From this table you can find the message and the troubleshooting chart that matches your trouble. Do a quantity check of the system as follows. (a) Put the anti-ice primary switch (SI022) in the OFF position. (b) On the G1000 system, select the MFD ENGINE page. (c) Make sure the message A-ICE GAL message is shown (white in color). (d) Next to the A-ICE GAL message, make sure you see the gallons value shown as 0.0 (green in color).
Troubleshooting Charts A.
There is an extensive set of troubleshooting charts for some of the most common problems for the anti-ice system. The table below is an index to the subject of the trouble and a link for Cesview IIi users. You can link to the chart and print them.
Table 101. G1000 Messages and Troubleshooting Index TROUBLE
REFER TO:
The G1000 MFD does not show the correct amount of anti-ice fluid remaining.
Figure 101
The A-ICE NORM message does not show or is incorrectly shown on the G1000 display.
Figure 102
The A-ICE HIGH does not show or is incorrectly shown on the G1000 display.
Figure 103
The A-ICE LOW PRESS message does not show or is incorrectly shown on the G1000 display when the primary anti-ice switch (SI022) is in the NORM position.
Figure 104
The A-ICE LOW PRESS message does not show or is incorrectly shown on the G1000 display when the backup anti-ice switch (SI024) is in the BACKUP ON position.
Figure 105
There is an A-ICE HIGH PRESS message shown on the G1000 display when the primary anti-ice switch (SI022) is in the NORM position.
Figure 106
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MODEL 208 MAINTENANCE MANUAL Table 101. G1000 Messages and Troubleshooting Index (continued) TROUBLE
REFER TO:
There is an A-ICE LOW FLUID message shown on the G1000 display when the primary anti-ice switch (SI022) is in the NORM position.
Figure 107
The windshield pump does not operate correctly.
Figure 108
Pump 1 does not operate when the primary anti-ice switch (SI022) is in the HIGH mode.
Figure 109
Pump 2 does not operate when the backup anti-ice switch (SI024) is in the BACKUP ON switch position.
Figure 110
One pump does not operate in the normal or max flow mode.
Figure 111
Pumps 1 and 2 do not operate in the normal flow mode.
Figure 112
Pumps 1 and 2 do not operate in the max flow mode.
Figure 113
The windshield pump does not go off after 4 seconds when the max flow switch (momentary switch) (SI023) is released from the windshield position.
Figure 114
Pumps 1 and 2 do not go on for 20 seconds and then go off for 100 seconds when the system is in the normal flow mode.
Figure 115
Pumps 1 and 2 do not go off after 120 seconds when the system is in the max flow mode.
Figure 116
Pumps 1 and 2 come on when the system is in the high, normal, or max mode with circuit breaker (CB309) disengaged.
Figure 117
Pumps 1 and 2 come on when the system is in the backup mode with the circuit breaker (CB410) disengaged.
Figure 118
The windshield pump comes on when the system is operated on with the circuit breaker (CB409) disengaged.
Figure 119
There is an uneven distribution of the anti-ice fluid at the porous panels.
Figure 120
There is no anti-ice fluid present at the nylon tubing.
Figure 121
There is no anti-ice fluid present at the porous panels.
Figure 122
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE FLUID TANK COMPONENTS - MAINTENANCE PRACTICES Pod Installation 1.
General A.
This section contains the removal and installation procedures for the TKS anti-ice ßuid tank and its equipment pack.
B.
The equipment pack includes the access panels, drain valve, check valves, ßuid level sender, Þlter pack, pumps, and switches.
C.
After you remove and install or replace the ßuid tank, it is necessary to do the panel purge and test procedures. Refer to TKS Leading Edge Porous Panel - Adjustment/Test.
D.
After you remove and install or replace the ßuid tank, you can calibrate the ßuid level sender, if necessary. Refer to TKS Anti-Ice Fluid Tank Components Adjustment/Test.
E.
Recommended maintenance to make sure that the TKS system operates correctly is as follows: • Operate the metering pumps each month, or when necessary, in the HIGH mode to remove the air from the ßuid system. • If the ßuid tank is removed and installed or replaced, do the porous panel purge and test procedures.
F.
Some Airplanes with the TKS anti-ice system have the G1000 avionic system installed. Table 201 shows the TKS-related circuit breakers and their reference designators.
G.
Some Airplanes with the TKS anti-ice system do not have the G1000 avionic system installed. Table 201 shows the TKS-related circuit breakers and their reference designators.
Table 201. TKS Circuit Breakers Airplanes With G1000
2.
Airplanes Without G1000
TKS Circuit Breaker
Reference Designator
TKS Circuit Breaker
Reference Designator
PRIMARY ANTI-ICE
(HC005)
PRIMARY ANTI-ICE
(CB309)
W/S
(HC015)
W/S
(CB409)
BACKUP ANTI-ICE
(HC016)
BACKUP ANTI-ICE
(CB410)
ENG INTFC
(HI013)
ANTI-ICE GAUGE
(CB310)
Tools and Equipment A.
For a list of tools and equipment, refer to Ice and Rain Protection - General.
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3.
TKS Anti-Ice Fluid Removal
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Anti-Ice Fluid (Refer to Figure 201 and Figure 202). (1) Find the drain tube outlet in the bottom of the cargo pod below the ßuid tank. (2) Put a container with a capacity of approximately 3 to 5 gallons below the drain tube outlet. (3) Open the forward-center and aft-center cargo pod doors to get access to the ßuid tank and aft bulkhead. (4) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (5) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE: (6) (7) (8)
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Push and turn (lock open) the knurled nut on the drain valve below the ßuid tank to release the ßuid. Turn and pull (lock closed) the knurled nut on the drain valve to stop the drain procedure.
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TKS Anti-Ice System Flow Diagram Figure 201 (Sheet 1)
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TKS Anti-Ice System Installation Figure 202 (Sheet 1)
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TKS Anti-Ice System Installation Figure 202 (Sheet 2)
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TKS Anti-Ice System Installation Figure 202 (Sheet 3)
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TKS Nylon Tubing Assembly Figure 203 (Sheet 1)
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TKS Nylon Tubing Assembly Figure 203 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (9)
Refer to Chapter 12, TKS Anti-Ice System - Servicing for the servicing procedures. NOTE:
You must calibrate the ßuid level sender if the MFD (G1000) or the quantity level gage (none G1000) does not read zero when the TKS ßuid tank is empty. The calibration procedures are in TKS Anti-Ice Fluid Tank Components - Adjustment/Test.
(10) Put the aft bulkhead in position in the cargo pod. (11) To install the drip pan, do the step that follows: (a) Use Type I, Class B sealant to bond the forward edge of the drip pan to the cargo pod. (12) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (13) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (14) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (15) Close the cargo pod doors.
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4.
TKS Anti-Ice Fluid Tank Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Anti-Ice Fluid Tank (Refer to Figure 201 and Figure 202). (1) Open the forward-center and aft-center cargo pod doors to get access to the ßuid tank and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the anti-ice ßuid from the ßuid tank. Refer to TKS Anti-Ice Fluid Removal in this section. (7) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (8) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE:
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
(9) Remove the aft bulkhead from the cargo pod. (10) Identify and disconnect the equipment pack electrical connectors from the airplane fuselage connector.
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MODEL 208 MAINTENANCE MANUAL (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21)
Loosen, but do not remove the hose clamp at the Þller neck on the ßuid tank assembly. Disconnect the Þller tube from the Þller neck. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Remove the bottom screws that attach the shear plates to the ßuid tank and carefully lower it. Loosen, but do not remove the hose clamps on the vent tubes at the access panel bushings. Disconnect the vent tubes at the access panel bushings. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Loosen, but do not remove the hose clamp on the drain tube at the drain valve. Disconnect the drain tube from the drain valve. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Disconnect the airplane supply line from the Þlter pack and disconnect the windshield supply line from the windshield pump. (22) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (23) Carefully remove the ßuid tank assembly from the cargo pod. NOTE:
You can remove the shear plates and the drip pan from the airplane to increase the space in the cargo pod.
(24) To remove the drip pan, do the steps that follow: (a) Use a scraper or putty knife to remove the sealant bond that attaches the forward edge of the drip pan to the cargo pod. (b) Remove the drip pan from the cargo pod. (c) Remove the sealant from the drip pan and the cargo pod ßoor. (25) Make sure that all openings and tube ends have caps installed. B.
Install the Anti-Ice Fluid Tank (Refer to Figure 201 and Figure 202). (1) Install the shear plates, if applicable. (2) Put the drip pan in position in the cargo pod, if applicable. (3) Put the ßuid tank assembly in position in the cargo pod. (4) Remove the caps from the airplane supply line and the windshield line tube ends. (5) Install new seals in the airplane supply line and the windshield couplings as shown in Figure 203. (6) Connect the airplane supply line at the Þlter pack and windshield line. (a) Safety all the tube couplings. Refer to Chapter 20, Safetying - Maintenance Practices. (7) Remove the caps from the vent tubes and access panel bushings. (8) Connect the vent tubes to the access panel bushings. (9) Tighten the hose clamps on the vent tubes at the access panel bushings. (10) Carefully lift the ßuid tank assembly, align the attach points, and install the bottom screws that attach the shear plates to the ßuid tank. (11) Remove the caps from the drain tube and drain valve. (12) Connect the drain tube to the drain valve. (13) Tighten the hose clamp on the drain tube at the drain valve. (14) Remove the caps from the Þller tube and ßuid neck. (15) Connect the Þller tube to the Þller neck. (16) Tighten the hose clamp at the Þller neck on the ßuid tank assembly. (17) Connect the electrical connectors to the airplane fuselage connector. (18) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (19) Engage the ENG INTFC circuit breaker, if applicable. (20) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (21) Supply external electrical power to the airplane. (22) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. NOTE:
You must calibrate the ßuid level sender if the MFD (G1000) or the quantity level gage (none G1000) does not read zero when the TKS ßuid tank is empty. The calibration procedures are in TKS Anti-Ice Fluid Tank Components - Adjustment/Test..
(23) Do a test of the ßuid tank components. Adjustment/Test.
Refer to TKS Anti-Ice Fluid Tank Components -
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MODEL 208 MAINTENANCE MANUAL (24) Do the panel purge and test procedures. Refer to TKS Anti-Ice Leading Edge Porous Panel Adjustment/Test. (25) Remove external electrical power from the airplane. (26) Put the aft bulkhead in position in the cargo pod. (27) To install the drip pan, do the step that follows: (a) Use Type I, Class B sealant to bond the forward edge of the drip pan to the cargo pod. (28) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (29) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (30) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (31) Close the cargo pod doors. 5.
Filter Pack Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Filter Pack (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable.
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MODEL 208 MAINTENANCE MANUAL (4) (5) (6) (7)
Disengage the ENG INTFC circuit breaker, if applicable. Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. Remove the screw and nut that attaches the bonding jumper to the bulkhead. Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE:
(8) (9) (10) (11) (12) B.
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Slowly loosen and disconnect the manifold couplings from the Þlter pack. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Remove the screws that attach the Þlter pack to the tank bracket. Remove the Þlter pack from the cargo pod.
Install the Filter Pack (Refer to Figure 201 and Figure 202). (1) Put the Þlter pack in position on the tank bracket. (2) Install the screws that attach the Þlter pack to the tank bracket. (3) Remove the caps from the tube ends. (4) Install new seals in the couplings as shown in Figure 203. (5) Connect the manifold couplings to the Þlter pack. (6) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (7) Engage the ENG INTFC circuit breaker, if applicable. (8) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (9) Supply external electrical power to the airplane. (10) Do the ßuid tank servicing as necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (11) Put the EXTERNAL POWER switch on the pilot's switch panel to the ON position. (12) Put the PRIMARY switch on the ANTI-ICE switch panel to the HIGH position. (a) Make sure that there is no ßuid leakage from the couplings. (13) Put the PRIMARY switch on the ANTI-ICE switch panel to the OFF position. (14) Put the EXTERNAL POWER switch on the pilot's switch panel to the OFF position. (15) Remove external electrical power from the airplane. (16) Put the aft bulkhead in position in the cargo pod. (17) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (18) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (19) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (20) Close the cargo pod door.
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6.
Metering Pump Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. NOTE: A.
The removal and installation of metering pump 1 and metering pump 1 are typical.
Remove the Metering Pump (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the anti-ice ßuid from the ßuid tank. Refer to TKS Anti-Ice Fluid Removal in this section. (7) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (8) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE:
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
(9) Remove the aft bulkhead from the cargo pod. (10) Identify and disconnect the electrical connector from the pump.
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MODEL 208 MAINTENANCE MANUAL (11) (12) (13) (14) B.
Slowly loosen and disconnect the tube couplings that are attached to the pump. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Remove the screws that attach the pump to the tank bracket. Remove the pump from the cargo pod.
Install the Metering Pump (Refer to Figure 201 and Figure 202). (1) Apply a light layer of TKS ßuid on the seals between the tank and the pump. (2) Put the pump in position in the pump bracket. (3) Install the screws that attach the pump to the tank bracket. (4) Remove the caps from the tube ends. (5) Install new seals in the couplings as shown in Figure 203. (6) Connect the tube couplings that are attached to the pump. (a) Safety all the tube couplings. Refer to Chapter 20, Safetying - Maintenance Practices. (7) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (8) Engage the ENG INTFC circuit breaker, if applicable. (9) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (10) Supply external electrical power to the airplane. (11) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (12) Do a test of the pump. Refer to TKS Anti-Ice Fluid Tank Components - Adjustment/Test. (13) Remove external electrical power from the airplane. (14) Put the aft bulkhead in position in the cargo pod. (15) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (16) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (17) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (18) Close the cargo pod door.
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7.
Windshield Pump Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Windshield Pump (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the anti-ice ßuid from the ßuid tank. Refer to TKS Anti-Ice Fluid Removal in this section. (7) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (8) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE: (9) (10) (11) (12)
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Identify and disconnect the electrical connector from the pump. Slowly loosen and disconnect the tube couplings that are attached to the pump. Put caps on all openings and tube ends to keep FOD out of the ßuid system.
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MODEL 208 MAINTENANCE MANUAL (13) Remove the screws and washers that attach the pump to the tank bracket. (14) Remove the pump from the cargo pod. B.
8.
Install the Windshield Pump (Refer to Figure 201 and Figure 202). (1) Put the pump in position in the pump bracket. (2) Install the screws and washers that attach the pump to the tank bracket. (3) Remove the caps from the tube ends. (4) Install new seals in the couplings as shown in Figure 203. (5) Connect the tube couplings that are attached to the pump. (a) Safety all the tube couplings. Refer to Chapter 20, Safetying - Maintenance Practices. (6) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (7) Engage the ENG INTFC circuit breaker, if applicable. (8) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (9) Supply external electrical power to the airplane. (10) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (11) Do a test of the pump. Refer to TKS Anti-Ice Fluid Tank Components - Adjustment/Test. (12) Remove external electrical power from the airplane. (13) Put the aft bulkhead in position in the cargo pod. (14) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (15) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (16) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (17) Close the cargo pod door.
Fluid Level Sender Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the Fluid Level Sender (Refer to Figure 201 and Figure 202). (1) Open the forward-center cargo pod door to get access to the sender.
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) (6) (7) (8)
Remove external electrical power from the airplane. Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. Disengage the ENG INTFC circuit breaker, if applicable. Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. Identify and disconnect the electrical wiring leads from the sender electrical posts. Remove the screws that attach the sender to the access panel. Carefully remove the sender from the access panel. (a) Do not damage the sensor. NOTE:
(9)
You can remove the screws that attach the access panel to the ßuid tank and move the access panel to help you to prevent damage to the sender.
Remove the screws that attach the access panel to the ßuid tank, if necessary. NOTE:
If you remove the access panel screws it is necessary to replace the access panel gasket.
(10) Put a cover on the opening to keep FOD out of the ßuid system. (11) Discard the gasket(s). (12) Remove the sender from the cargo pod. B.
Install the Fluid Level Sender (Refer to Figure 201 and Figure 202). (1) Remove the cover from the access panel opening. (2) Put the sender and a new gasket in position on the access panel. (a) Do not damage the sensor. (3) Install the screws that attach the sender to the access panel. NOTE:
(4) (5) (6) (7) (8) (9) (10) (11)
If you removed the access panel screws it is necessary to replace the access panel gasket.
Carefully put the access panel and new gasket in position on the ßuid tank, if applicable. Install the screws that attach the access panel to the ßuid tank, if applicable. Connect the electrical wiring leads to the sender electrical posts. Engage the ANTI-ICE GAUGE circuit breaker, if applicable. Engage the ENG INTFC circuit breaker, if applicable. Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. Supply external electrical power to the airplane. Do the ßuid tank servicing as necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. NOTE:
You must calibrate the ßuid level sender if it does not read zero when it is empty. The calibration procedures are in TKS Anti-Ice Fluid Tank Components - Adjustment/Test.
(12) Do a test of the sender. Refer to TKS Anti-Ice Fluid Tank Components - Adjustment/Test. (13) Remove external electrical power from the airplane. (14) Close the cargo pod door.
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9.
Low Level Switch Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Low Level Switch (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the anti-ice ßuid from the ßuid tank. Refer to TKS Anti-Ice Fluid Removal in this section. (7) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (8) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE: (9) (10) (11) (12)
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Identify and disconnect the electrical connector from the switch. Remove the low level switch from the ßuid tank. Put a cover over the switch opening to keep FOD out of the ßuid system.
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MODEL 208 MAINTENANCE MANUAL (13) Remove the switch from the cargo pod. B.
Install the Low Level Switch (Refer to Figure 201 and Figure 202). (1) Remove the cover from the low level switch opening. (2) Install the switch in the ßuid tank. (3) Connect the electrical connector to the switch. (4) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (5) Engage the ENG INTFC circuit breaker, if applicable. (6) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (7) Supply external electrical power to the airplane. (8) Do a test of the switch. Refer to TKS Anti-Ice Fluid Tank Components - Maintenance Practices. (9) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (10) Remove external electrical power from the airplane. (11) Put the aft bulkhead in position in the cargo pod. (12) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (13) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (14) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (15) Close the cargo pod door.
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10.
High Pressure Switch Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the High Pressure Switch (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (7) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE: (8) (9) (10) (11) (12)
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Identify and disconnect the electrical connector from the switch. Slowly loosen and disconnect the tube couplings that are connected to the switch. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Remove the screws that attach the switch to the tank bracket.
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MODEL 208 MAINTENANCE MANUAL (13) Remove the switch from the cargo pod. B.
11.
Install the High Pressure Switch (Refer to Figure 201 and Figure 202). (1) Put the switch in position in the switch bracket. (2) Install the screws that attach the switch to the tank bracket. (3) Remove the caps from the tube ends. (4) Connect the tube couplings that are connected to the switch. (a) Safety all the tube couplings. Refer to Chapter 20, Safetying - Maintenance Practices. (5) Connect the electrical connector to the switch. (6) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (7) Engage the ENG INTFC circuit breaker, if applicable. (8) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (9) Supply external electrical power to the airplane. (10) Do the ßuid tank servicing as necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (11) Do a test of the switch. Refer to TKS Anti-Ice Fluid Tank Components - Adjustment/Test. (12) Remove external electrical power from the airplane. (13) Put the aft bulkhead in position in the cargo pod. (14) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (15) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (16) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (17) Close the cargo pod door.
Timer Box and/or Wire Bundle Removal/Installation A.
Remove the Timer Box and/or Wire Bundle (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (7) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE:
(8) (9)
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. To remove the wire bundle, identify and disconnect the equipment pack electrical connectors from the equipment pack components and airplane fuselage connector. NOTE:
It is not necessary to remove the wire bundle to remove the timer box.
(10) Identify and disconnect the electrical connector from the timer box. (11) Remove the screws that attach the timer box to the tank bracket. (12) Remove the timer box from the cargo pod. B.
Install the Timer Box and/or Wire Bundle (Refer to Figure 201 and Figure 202). (1) Put the timer box in position in the timer box bracket. (2) Install the screws that attach the timer box to the tank bracket. (3) Connect the electrical connector to the timer box. (4) Identify and connect the equipment pack electrical connectors to the equipment pack components and airplane fuselage connector as applicable. (5) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (6) Engage the ENG INTFC circuit breaker, if applicable. (7) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (8) Supply external power to the airplane. (9) Do the ßuid tank servicing as necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing.
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MODEL 208 MAINTENANCE MANUAL (10) Do a test of the ßuid tank components. Refer to TKS Anti-Ice Fluid Tank Components Adjustment/Test. (11) Remove external electrical power from the airplane. (12) Put the aft bulkhead in position in the cargo pod. (13) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (14) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (15) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (16) Close the cargo pod doors. 12.
Solenoid Valve Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Solenoid Valve (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers.
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MODEL 208 MAINTENANCE MANUAL (6) (7) (8)
Remove the anti-ice ßuid from the ßuid tank. Refer to TKS Anti-Ice Fluid Removal in this section. Remove the screw and nut that attaches the bonding jumper to the bulkhead. Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE:
(9) (10) (11) (12) (13) (14) B.
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Identify and disconnect the electrical connector from the valve. Slowly loosen and disconnect the tube couplings from the valve. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Remove the screws that attach the valve to the tank bracket. Remove the valve from the cargo pod.
Install the Solenoid Valve (Refer to Figure 201 and Figure 202). (1) Put the valve in position in the valve bracket. (2) Install the screws that attach the valve to the tank bracket. (3) Remove the caps from the tube ends. (4) Install new seals in the couplings as shown in Figure 203. (5) Connect the tube couplings to the valve. (a) Safety all the tube couplings. Refer to Chapter 20, Safetying - Maintenance Practices. (6) Connect the electrical connector to the valve. (7) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (8) Engage the ENG INTFC circuit breaker, if applicable. (9) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (10) Supply external electrical power to the airplane. (11) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (12) Put the EXTERNAL POWER switch on the pilot's switch panel to the ON position. (13) Put the PRIMARY switch on the ANTI-ICE switch panel to the HIGH position. (a) Make sure that there is no ßuid leakage from the couplings. (14) Put the PRIMARY switch on the ANTI-ICE switch panel to the OFF position. (15) Put the EXTERNAL POWER switch on the pilot's switch panel to the OFF position. (16) Remove external electrical power from the airplane. (17) Put the aft bulkhead in position in the cargo pod. (18) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (19) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (20) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (21) Close the cargo pod door.
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13.
Check Valve Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Check Valve (Refer to Figure 201 and Figure 202). (1) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (7) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE: (8) (9) (10) (11) (12)
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Slowly loosen and disconnect the tube couplings that are attached to the check valve. Put caps on all openings and tube ends to keep FOD out of the ßuid system. Remove the screws that attach the check valve to the tank bracket. Remove the check valve from the cargo pod.
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MODEL 208 MAINTENANCE MANUAL B.
Install the Check Valve (Refer to Figure 201 and Figure 202). (1) Put the valve in position in the valve bracket. (2) Install the screws that attach the valve to the tank bracket. (3) Remove the caps from the tube ends. (4) Install new seals in the couplings as shown in Figure 203. (5) Connect the tube couplings to the valve. (a) Make sure that the ßuid ßow direction is correct. (b) Safety all the tube couplings. Refer to Chapter 20, Safetying - Maintenance Practices. (6) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (7) Engage the ENG INTFC circuit breaker, if applicable. (8) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (9) Supply external electrical power to the airplane. (10) Do the ßuid tank servicing as necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (11) Put the EXTERNAL POWER switch on the pilot's switch panel to the ON position. (12) Put the PRIMARY switch on the ANTI-ICE switch panel to the HIGH position. (a) Make sure that there is no ßuid leakage from the couplings. (13) Put the PRIMARY switch on the ANTI-ICE switch panel to the OFF position. (14) Put the EXTERNAL POWER switch on the pilot's switch panel to the OFF position. (15) Remove external electrical power from the airplane. (16) Put the aft bulkhead in position in the cargo pod. (17) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (18) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (19) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (20) Close the cargo pod door.
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14.
Sight Glass Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the Sight Glass (Refer to Figure 201 and Figure 202). (1) Open the forward-center cargo pod door to get access to the sight glass. (2) Remove the anti-ice ßuid from the ßuid tank. Refer to TKS Anti-Ice Fluid Removal in this section. (3) Remove external electrical power from the airplane. (4) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (5) Disengage the ENG INTFC circuit breaker, if applicable. (6) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (7) Slowly open the tube clamps that are connected to the sight glass. (8) Remove the sight glass (and ball) from the airplane. (9) Put caps on all openings and tube ends to keep FOD out of the ßuid system.
B.
Install the Sight Glass (Refer to Figure 201 and Figure 202). (1) Remove the caps from the tube ends. (2) Put the sight glass (and ball) in position in the sight glass brackets. (3) Crimp the tube clamps. (4) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (5) Engage the ENG INTFC circuit breaker, if applicable. (6) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (7) Supply external electrical power to the airplane. (8) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (9) Put the EXTERNAL POWER switch on the pilot's switch panel to the ON position. (10) Put the PRIMARY switch on the ANTI-ICE switch panel to the HIGH position. (a) Make sure that there is no ßuid leakage from the tube clamps. (11) Put the PRIMARY switch on the ANTI-ICE switch panel to the OFF position. (12) Put the EXTERNAL POWER switch on the pilot's switch panel to the OFF position. (13) Clean the ßoor and the airplane surfaces as necessary.
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MODEL 208 MAINTENANCE MANUAL (14) Close the cargo pod door. (15) Remove external electrical power from the airplane. 15.
Drain Valve Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the Drain Valve (Refer to Figure 201 and Figure 202). (1) Open the forward-center cargo pod door to get access to the drain valve below the ßuid tank. (2) Remove the anti-ice ßuid from the ßuid tank. Refer to TKS Anti-Ice Fluid Removal in this section. (3) Loosen, but do not remove the hose clamp on the drain tube. (4) Pull the drain tube off the drain valve. (5) Disconnect the drain valve from the ßuid tank. (6) Remove the drain valve from the cargo pod.
B.
Install the Drain Valve (Refer to Figure 201 and Figure 202). (1) Put the drain valve in position on the ßuid tank. (2) Connect the drain valve to the ßuid tank. (3) Connect the drain tube to the drain valve. (4) Tighten the hose clamp on the drain tube. (5) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (6) Make sure that there is no ßuid leakage from the valve. (7) Clean the ßoor and the airplane surfaces as necessary. (8) Close the cargo pod door.
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16.
Fluid Filler Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Filler Tube Duct (Refer to Figure 201 and Figure 202). (1) Open the forward-center cargo pod door to get access to the ßuid tank. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Loosen, but do not remove the hose clamps at the Þller neck and Þller internal ßange. (7) Disconnect the Þller tube duct from the Þller neck and Þller internal ßange. (8) Remove the Þller tube from the cargo pod. (9) Put caps on all openings to keep FOD out of the ßuid system.
B.
Install the Filler Tube Duct (Refer to Figure 201 and Figure 202). (1) Remove the caps from the openings. (2) Put the Þller tube duct in position on the Þller neck and Þller internal ßange. (3) Tighten the hose clamps at the Þller neck and Þller internal ßange. (4) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (5) Engage the ENG INTFC circuit breaker, if applicable.
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Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. Do the ßuid tank servicing as necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. Clean the ßoor and the airplane surfaces as necessary. Close the cargo pod door.
C.
Remove the Filler Assembly (Refer to Figure 201 and Figure 202). (1) Open the forward-center cargo pod door to get access to the ßuid tank. (2) Remove external electrical power from the airplane. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Open the Þller cap. (7) Loosen, but do not remove the hose clamp at the Þller internal ßange. (8) Disconnect the Þller tube duct from the Þller internal ßange. (9) Remove the screws, nuts, and washers that attach the Þller cap ßange, Þller internal ßange, and gaskets to the cargo pod skin. (10) Put caps on all openings to keep FOD out of the ßuid system.
D.
Install the Filler Tube Assembly (Refer to Figure 201 and Figure 202). (1) Remove the caps from the openings. (2) Install the screws, nuts, and washers that attach the Þller cap ßange, Þller internal ßange, and gaskets to the cargo pod skin. (3) Connect the Þller tube duct to the Þller internal ßange. (4) Tighten the hose clamp at the Þller internal ßange. (5) Close the Þller cap. (6) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (7) Engage the ENG INTFC circuit breaker, if applicable. (8) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (9) Do the ßuid tank servicing as necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (10) Clean the ßoor and the airplane surfaces as necessary. (11) Close the cargo pod door.
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17.
TKS Fluid Contamination (Fuel) Removal
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. CAUTION: Do not operate the windshield pump for more than 10 seconds continuously, and wait 10 seconds between pump operations before you operate the pump again. Damage to the windshield pump can occur if the pump is operated for more than the speciÞed limit. A.
Remove the TKS Fluid Contamination (Fuel) (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (3) Disengage the ENG INTFC circuit breaker, if applicable. (4) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (5) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead.
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Remove the screw and nut that attaches the bonding jumper to the aft bulkhead. Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE:
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
(8) Remove the aft bulkhead from the cargo pod. (9) Drain the TKS ßuid from the ßuid tank . Refer to TKS Anti-Ice Fluid Removal in this section. (10) Disconnect the Þlter inlet tube from the Þlter inlet manifold assembly. (a) Put the inlet tube end into a bucket with a capacity of approximately 3 to 5 gallons. (11) Disconnect the windshield pump outlet tube from the fuselage connector. (a) Put the open tube end into a bucket with a capacity of approximately 1 gallon. (12) Fill the ßuid tank with a water and mild detergent mixture. (13) Drain the water and mild detergent mixture from the ßuid tank. (14) Fill the ßuid tank with water. (15) Drain the water from the ßuid tank. (16) Remove the Fluid Level Sender. refer to Fluid Level Sender Removal/Installation in this section. (17) Clean the ßuid level sender with a water and mild detergent mixture. (18) Flush the ßuid level sender with water until no contamination shows. (19) Install the Fluid Level Sender. Refer to Fluid Level Sender Removal/Installation in this section. (20) Add 10 gallons (37.84 liters) of TKS ßuid to the ßuid tank. (21) Supply external electrical power to the airplane. (22) Put the EXTERNAL POWER switch on the pilot's switch panel to the BUS position. (23) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (24) Engage the ENG INTFC circuit breaker, if applicable. (25) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (26) Put the PRIMARY switch on the ANTI-ICE FLUID CONTROL switch panel to the HIGH position to start pump 1. (27) Put the MAX FLOW switch on the ANTI-ICE FLUID CONTROL switch panel to the AIRFRAME position, then release the switch. (a) Let the TKS ßuid ßow into the bucket at the Þlter inlet tube for one full cycle of approximately 2 minutes. (28) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch to the OFF position. (29) Put the MAX FLOW switch on the ANTI-ICE FLUID CONTROL switch panel in the WINDSHIELD position, then release the switch. NOTE:
The windshield pump will start and operate when you put the spring-loaded MAX FLOW switch to the WINDSHIELD position.
(a)
(30) (31)
(32) (33) (34) (35)
Let the TKS ßuid ßow into the bucket at the windshield pump outlet tube for one full cycle of approximately 4 seconds. Put the EXTERNAL POWER switch on the pilot's switch panel in the OFF position. Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. Disengage the ENG INTFC circuit breaker, if applicable. Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. Remove external electrical power from the airplane. If the system was operated with fuel contamination in the ßuid, replace the Þlter pack. Refer to Filter Pack Removal/Installation in this section. NOTE:
If there was fuel contamination in the ßuid, but system was not operated, it is not necessary to replace the Þlter pack.
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MODEL 208 MAINTENANCE MANUAL (36) (37) (38) (39) (40) (41) (42)
Connect the Þlter inlet tube with a new O-ring to the Þlter inlet manifold assembly. Connect the windshield pump outlet tube with a new O-ring to the fuselage connector. Supply external electrical power to the airplane. Put the EXTERNAL POWER switch on the pilot's switch panel to the BUS position. Engage the ANTI-ICE GAUGE circuit breaker, if applicable. Engage the ENG INTFC circuit breaker, if applicable. Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (43) Put the PRIMARY switch on the ANTI-ICE FLUID CONTROL switch panel to the HIGH position to start pump 1. (44) If the Þlter assembly was replaced, put the BACKUP switch on the ANTI-ICE FLUID CONTROL switch panel in the ON position to start pump 2. (a) Let the system operate for approximately 5 minutes until there are no air bubbles in the TKS ßuid ßow from the porous panels on the wing, wing strut, and horizontal and vertical stabilizer leading edges. NOTE:
If necessary, you can let the system operate for more than 5 minutes until the ßuid ßow is normal across all porous panels.
(b)
If ßuid ßow at any of the porous panels does not become normal, do the porous panel purge and test procedure for the applicable porous panel(s). Refer to TKS Leading Edge Porous Panel - Adjustment/Test. (45) If the Þlter assembly was not replaced, put the MAX FLOW switch on the ANTI-ICE FLUID CONTROL switch panel to the AIRFRAME position, then release the switch. (a) Let the TKS ßuid ßow from the porous panels on the wing, wing strut, and horizontal and vertical stabilizer leading edges for one full cycle of approximately 2 minutes. NOTE:
If necessary, you can do more MAX FLOW cycles until the ßuid ßow is normal across all porous panels.
(b)
If ßuid ßow at any of the porous panels does not become normal, do the porous panel purge and test procedure for the applicable porous panel(s). Refer to TKS Leading Edge Porous Panel - Adjustment/Test. (46) Put the PRIMARY switch on the ANTI-ICE-FLUID CONTROL switch panel to the OFF position. (47) Put the BACKUP switch on the ANTI-ICE-FLUID CONTROL switch panel to the OFF position. (48) Put the MAX FLOW switch on the ANTI-ICE FLUID CONTROL switch panel in the WINDSHIELD position, then release the switch. NOTE:
(a)
The windshield pump will start and operate when you put the spring-loaded MAX FLOW switch to the WINDSHIELD position.
Do three full cycles of approximately 4 seconds each to let the TKS ßuid ßow from the windshield spray bar. NOTE:
If necessary, you can do more cycles until the ßuid ßow is normal from the spray bar.
(49) Put the EXTERNAL POWER switch on the pilot's switch panel in the OFF position. (50) Remove external electrical power from the airplane. (51) Make sure that the ßuid tank servicing is correct. Refer to Chapter 12, TKS Anti-Ice System Servicing. (52) Make sure that the ßoor and the airplane surfaces are clean. (53) Put the aft bulkhead in position in the cargo pod. (54) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod.
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MODEL 208 MAINTENANCE MANUAL (55) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (56) Install the screws and connect the antenna coaxial cable and conduit, if applicable. (57) Close the aft-center cargo pod door. 18.
TKS Fluid Contamination (Water) Removal
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the TKS Fluid Contamination (Water) (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Put the BATTERY switch on the pilot's switch panel to the OFF position. (3) Drain the TKS ßuid from the ßuid tank . Refer to TKS Anti-Ice Fluid Removal in this section. (4) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (5) Supply external electrical power to the airplane. (6) Put the EXTERNAL POWER switch on the pilot's switch panel to the BUS position. (7) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (8) Engage the ENG INTFC circuit breaker, if applicable. (9) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (10) Put the PRIMARY switch on the ANTI-ICE FLUID CONTROL switch panel to the HIGH position to start pump 1. (11) Put the MAX FLOW switch on the ANTI-ICE FLUID CONTROL switch panel to the AIRFRAME position, then release the switch. (a) Let the TKS ßuid ßow from the porous panels on the wing, wing strut, and horizontal and vertical stabilizer leading edges for one full cycle of approximately 2 minutes. (12) If ßuid ßow at any of the porous panels is not normal, do the porous panel purge and test procedure for the applicable porous panel(s). Refer to TKS Leading Edge Porous Panel Adjustment/Test. (13) Put the PRIMARY switch on the ANTI-ICE FLUID CONTROL switch panel to the OFF position. (14) Put the EXTERNAL POWER switch on the pilot's switch panel in the OFF position. (15) Remove external electrical power from the airplane.
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MODEL 208 MAINTENANCE MANUAL (16) Make sure that the ßuid tank servicing is correct. Refer to Chapter 12, TKS Anti-Ice System Servicing. (17) Make sure that the ßoor and the airplane surfaces are clean. 19.
TKS Fluid Contamination (Solids) Removal A.
TKS Fluid Contamination (Solids) Removal (1) If the TKS ßuid contamination is a solid material, contact Cessna Customer Service, P.O. Box 7706, Wichita, Kansas 67209 USA Tele: 316-517-5800 Fax: 316-517-7271.
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE FLUID TANK COMPONENTS - ADJUSTMENT/TEST Pod Installation 1.
General A.
This section contains the test procedures that are necessary to do after TKS anti-ice fluid tank component replacement.
B.
For the purge and test procedures of the leading edge porous panels, refer to TKS Leading Edge Porous Panel - Adjustment/Test.
C.
For the test procedures for the tail bracket assembly (low pressure switches), refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices.
D.
To calibrate the fluid level sender, refer to Fluid Level Sender Calibration in this section.
E.
Recommended maintenance to keep the TKS fluid at its correct viscosity is as follows: • Operate the pumps monthly, or as necessary, in the HIGH mode until the air is removed from the fluid system. • Keep the TKS system operational at all times to keep air pockets out of the system. • If the fluid tank is removed and installed or replaced, do the porous panel purge and test procedures. NOTE:
If the fluid is too thick, the porous panels can become blocked or clogged.
F.
Some Airplanes with the TKS anti-ice system have the G1000 avionic system installed. On those Airplanes you can ignore all references to the ANTI-ICE ON, CAUT, and WARN annunciators. Table 501 shows the TKS-related circuit breakers and their reference designators.
G.
Some Airplanes with the TKS anti-ice system do not have the G1000 avionic system installed. On those Airplanes you can ignore all references to CAS messages. Table 501 shows the TKS-related circuit breakers and their reference designators.
Table 501. TKS Circuit Breakers Airplanes With G1000
2.
Airplanes Without G1000
TKS Circuit Breaker
Reference Designator
TKS Circuit Breaker
Reference Designator
PRIMARY ANTI-ICE
(HC005)
PRIMARY ANTI-ICE
(CB309)
W/S
(HC015)
W/S
(CB409)
BACKUP ANTI-ICE
(HC016)
BACKUP ANTI-ICE
(CB410)
ENG INTFC
(HI013)
ANTI-ICE GAUGE
(CB310)
Tools and Equipment A.
For a list of tools and equipment, refer to Ice and Rain Protection - General.
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3.
TKS Anti-Ice Fluid Tank Component Test
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Immediately remove (clean) or contain all the TKS fluid that is spilled. TKS fluid on the floor will cause a slip hazard. WARNING: Discard all unwanted TKS fluid and/or dirty cloths correctly. TKS fluid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS fluids. Approved fluids, in accordance with specification DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, filtered fluid in the TKS system. Contamination will cause fluid blockage and/or damage to the porous panel. NOTE:
For the tests that follow, you can disconnect the airplane supply tube from the filter manifold outlet and connect a drain tube, which will let you contain the fluid more easily, and that is how these procedures are written. Or, you can keep the airplane supply tube connected to the filter manifold outlet and use clean, dry cloths to absorb the anti-ice fluid and to clean the airplane surfaces and floor as necessary. Or, you can fabricate a fluid collector system and install it on and below the porous panels, which will contain the fluid and keep it off the floor. Recommended materials you can use are plastic sheets, tubing, aluminum tape, and rigid aluminum and/or plastic gutter material.
NOTE:
Although you can do one or more of the tests that follow, as applicable, it is necessary to do all of the tests after you have installed the fluid tank, timer box, and/or wire bundle.
NOTE:
It is easier for two persons to do these tests. One to monitor the cockpit and one to monitor the equipment pack.
A.
Prepare To Do the Fluid Tank Component Tests. (1) Remove external electrical power from the airplane. (2) Open the aft-center cargo pod door to get access to the equipment pack and aft bulkhead. (3) Disengage the ANTI-ICE GAUGE circuit breaker, if applicable. (4) Disengage the ENG INTFC circuit breaker, if applicable. (5) Disengage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (6) Remove the screw and nut that attaches the bonding jumper to the bulkhead. (7) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. NOTE:
On airplanes that have the TAS antenna installed, it is necessary to disconnect the coaxial cable and remove the screws and conduit.
Remove the aft bulkhead from the cargo pod. Find the drain tube opening in the bottom of the cargo pod below the fluid tank. (a) Put a container below the drain tube. (10) If you will do tests of the metering pumps, the high pressure switch, or the timer box, disconnect the airplane supply tube from the filter manifold outlet. (11) Put a cap or plug in the open tube end. (12) Connect a length of tubing to the filter manifold outlet. (a) Put the open tube end in the container. (8) (9)
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MODEL 208 MAINTENANCE MANUAL (13) If you will do the windshield pump test, disconnect the windshield pump outlet tube at the belly connector and place the end of the tube in the bucket. (14) Engage the ANTI-ICE GAUGE circuit breaker, if applicable. (15) Engage the ENG INTFC circuit breaker, if applicable. (16) Engage the PRIMARY ANTI-ICE, W/S, and BACKUP ANTI-ICE circuit breakers. (17) Do the test procedures as applicable. B.
Do a Test of Metering Pump 1. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure. (2) Supply external electrical power to the airplane. (3) Put the EXTERNAL POWER switch on the pilot's switch panel in the BUS position. (4) Put the PRIMARY switch on the ANTI-ICE FLUID CONTROL switch panel in the HIGH position to start pump 1. (a) Make sure that pump 1 operation starts. (b) Make sure that 1000 ml, +100 or - 100 ml discharges in the bucket in one minute. (5) Disengage the PRIMARY ANTI-ICE circuit breaker. (a) Make sure that pump 1 operation stops. (6) Make sure that there is no fluid leakage from the couplings. (7) Put the PRIMARY switch in the OFF position. (8) Put the EXTERNAL POWER switch on the pilot's switch panel in the OFF position. (9) Engage the PRIMARY ANTI-ICE circuit breaker. (10) Do the Return to Service procedures or continue the applicable test(s).
C.
Do a Test of Metering Pump 2. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure. (2) Supply external electrical power to the airplane. (3) Put the EXTERNAL POWER switch on the pilot's switch panel in the BUS position. (4) Put the BACKUP switch on the ANTI-ICE FLUID CONTROL switch panel in the ON position to start pump 2. (a) Make sure that pump 2 operation starts. (b) Make sure that 1000 ml, +100 or - 100 ml discharges in the bucket in one minute. (5) Disengage the BACKUP ANTI-ICE circuit breaker. (a) Make sure that pump 2 operation stops. (6) Make sure that there is no fluid leakage from the couplings. (7) Put the BACKUP switch in the OFF position. (8) Engage the BACKUP ANTI-ICE circuit breaker. (9) Put the EXTERNAL POWER switch on the pilot's switch panel in the OFF position. (10) Do the Return to Service procedures or continue the applicable test(s).
D.
Do a Test of the Windshield Pump. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure. (2) Supply external electrical power to the airplane. (3) Put the EXTERNAL POWER switch on the pilot's switch panel in the BUS position. (4) Put the MAX FLOW switch on the ANTI-ICE FLUID CONTROL switch panel in the WINDSHIELD position, then release the switch. NOTE:
The windshield pump will start when you put the spring-loaded MAX FLOW switch in the WINDSHIELD position and it will spray fluid on the windshield for four seconds after you release it.
Make sure that the windshield pump starts. When the pump stops make sure that a minimum of 25 ml was discharged in the collection bucket. After the windshield pump stops, disengage the W/S ANTI-ICE circuit breaker. Put the MAX FLOW switch in the WINDSHIELD position, then release the switch. (a) Make sure that the windshield pump does not operate. (a) (b)
(5) (6)
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Make sure that there is no fluid leakage from the couplings. Engage the W/S ANTI-ICE circuit breaker. Remove external electrical power from the airplane. Do the Return to Service procedures or continue the applicable test(s).
Do a Test of the Fluid Level Sender. (1) Supply external electrical power to the airplane. (2) Put the EXTERNAL POWER switch on the pilot's switch panel in the BUS position. (3) For airplanes with the G1000 do the steps that follow: (a) Put the AVIONICS 1 switch to the ON position. (b) Put the AVIONICS 2 switch to the ON position. (4) Drain the tank. (5) Make sure that there is a fluid quantity indication of E GAL on the ANTI-ICE QTY gage or that the A-ICE GAL 0.0 indication shows on the MFD display as applicable. NOTE:
You must calibrate the level sender if it does not read zero when the tank is empty.
(6) (7)
Fill the tank. Make sure that there is a fluid quantity indication of 20 GAL on the ANTI-ICE QTY gage or that the A-ICE GAL 20.8 indication shows on the MFD display as applicable. (8) Put the PRIMARY switch to the OFF position. (9) To calibrate the fluid level sender, if necessary, refer to Fluid Level Sender Calibration in this section. (10) Do the Return to Service procedures or continue the applicable test(s). F.
Do a Test of the Low Level Switch. (1) Remove the anti-ice fluid from the tank. Refer to TKS Anti-Ice Fluid Tank Components Maintenance Practices. (2) Supply external electrical power to the airplane. (3) Put the EXTERNAL POWER switch on the pilot's switch panel in the BUS position. (4) For airplanes with the G1000 do the steps that follow: (a) Put the AVIONICS 1 switch to the ON position. (b) Put the AVIONICS 2 switch to the ON position. (5) Put the PRIMARY switch on the ANTI-ICE FLUID CONTROL switch panel in the NORM position. (6) Make sure that the A-ICE LOW FLUID (amber) CAS message shows on the EICAS display or the TKS annunciator shows WARN (amber) , as applicable. (7) Put the PRIMARY switch in the OFF position (8) Add 4.0 gallons of fluid to the tank. (9) Put the PRIMARY switch on the ANTI-ICE switch panel in the NORM position. (10) Make sure that the A-ICE LOW FLUID (amber) CAS message does not show on the EICAS display or the TKS annunciator does not show WARN, as applicable. (11) Make sure that there is no fluid leakage from the couplings. (12) Put the PRIMARY switch in the OFF position (13) Do the Return to Service procedures or continue the applicable test(s).
G.
Do a Test of the High Pressure Switch. (1) Supply external electrical power to the airplane. (2) Put the EXTERNAL POWER switch on the pilot's switch panel in the BUS position. (3) For airplanes with the G1000 do the steps that follow: (a) Put the AVIONICS 1 switch to the ON position. (b) Put the AVIONICS 2 switch to the ON position. (4) Connect a pressure gage and shutoff valve to the filter manifold outlet tube. (a) Close the shutoff valve. (5) Put the PRIMARY switch on the ANTI-ICE FLUID CONTROL switch panel in the HIGH position. (6) Monitor the pressure gage. (a) Wait until the pressure gage shows a 150 psi indication. NOTE:
This makes sure that the high pressure switch is operating correctly.
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Put the PRIMARY switch in the OFF position. Put the EXTERNAL POWER switch on the pilot's switch panel in the OFF position. Slowly open the shutoff valve. Disconnect the pressure gage and shutoff valve from the tube. Do the Return to Service procedures or continue the applicable test(s).
H.
Do a Test of the Timer Box. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure. (2) Supply external electrical power to the airplane. (3) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the BUS position. (4) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel in the NORM position. (a) Make sure that each of the two pumps operate for 20 seconds, +3 or - 3 seconds. (5) Make sure that there is no fluid leakage from the fittings. (6) While the pumps are off, Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch in the AIRFRAME position. (a) Make sure that each of the two pumps operate for 120 seconds, +10 or - 20 seconds, and then do not operate for 100 seconds, +10 or -10 seconds. (7) Put the PRIMARY switch in the OFF position. (8) Do a test of the windshield pump. Refer to Do a Test of the Windshield Pump in this section. (9) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. (10) Do the Return to Service procedures or continue the applicable test(s).
I.
Do the Airplane Return to Service. NOTE:
After you have completed the applicable test(s), it is necessary to put the airplane back to its initial configuration.
(1) (2)
Make sure that there is no fluid leakage from the couplings. Make sure that all applicable connectors, fasteners, and couplings are installed correctly and safetied as necessary. Refer to Chapter 20, Safetying - Maintenance Practices. (3) Put the aft bulkhead in position in the cargo pod. (4) Turn the quarter-turn fasteners that attach the aft bulkhead to the drip pan and the cargo pod. (5) Install the screw and nut that attaches the bonding jumper to the bulkhead. (a) Make sure that there is a good electrical bond. Refer to Chapter 20, Electrical Bonding Maintenance Practices. (6) Install the screws and connect the antenna coaxial cable and conduit, if necessary. (7) Make sure that all circuit breakers are engaged. (8) Make sure that all system switches are in their initial positions. (9) Make sure that external electrical power is removed from the airplane. (10) Make sure that the cargo pod door(s) is closed. (11) Make sure that the floor and the airplane surfaces are clean.
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4.
TKS Anti-Ice Level Sender Calibration
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Immediately remove (clean) or contain all the TKS fluid that is spilled. TKS fluid on the floor will cause a slip hazard. WARNING: Discard all unwanted TKS fluid and/or dirty cloths correctly. TKS fluid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS fluids. Approved fluids, in accordance with specification DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, filtered fluid in the TKS system. Contamination will cause fluid blockage and/or damage to the porous panel. NOTE:
To make sure that the level sender is calibrated correctly, you can do this calibration procedure. The voltmeter will show 0.0 VDC when the tank is empty and 5.0 VDC, +0.1 or -0.1 VDC when the tank is full.
NOTE:
When the tank is empty and the EICAS display or the gage shows 0.0 gallons, calibration of the empty adjustment is not necessary. You can then fill the tank and calibrate the level sender full adjustment.
NOTE:
Changes in the properties of the anti-ice fluid can occur because of differences between manufacturers, or if the fluid is new (fresh), or if the fluid has gone through the tank and TKS system, or if the fluid has been in the tank too long (the TKS system has not been operated). These fluid changes can cause different results in calibration.
A.
Calibrate the Level Sender (Refer to Figure 501). (1) Make sure that the airplane is level. Refer to Chapter 8, Leveling - Maintenance Practices. (2) Open the forward-center cargo pod door. (3) Connect one lead of the voltmeter to the SEND post of the level sender. (4) Connect the other lead of the voltmeter to the NEG post of the level sender. (5) Supply external electrical power to the airplane. (6) To calibrate the level sender with an empty tank, do the steps that follow: (a) Remove the protective coating from the EMPTY adjustment screw. NOTE:
This screw is on the right side (airplane's right side) of the level sender.
(b)
(7)
Use a screwdriver to turn the EMPTY adjustment screw counterclockwise until the voltage on the voltmeter does not go lower. Turn the screw clockwise until 0.0 VDC, +0.1 or -0.1 VDC, shows on the voltmeter. To calibrate the level sender with a full tank, do the steps that follow: (a) Remove the protective coating from the FULL adjustment screw. NOTE:
This screw is on the left side (airplane's left side) of the level sender.
(b)
(8)
Use a screwdriver to turn the FULL adjustment screw until 5.0 VDC, +0.1 or -0.1 VDC, shows on the voltmeter. Remove external electrical power from the airplane.
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Level Sender Calibration Figure 501 (Sheet 1)
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Disconnect the leads of the voltmeter from the level sender. Put a protective coating on the adjustment screw(s). Close the cargo pod door. Clean the floor and the airplane surfaces as necessary.
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE SYSTEM - DESCRIPTION AND OPERATION FAIRING INSTALLATION 1.
2.
General A.
The TKS system is a fluid anti-ice system that helps prevent the formation of ice on the airplane surfaces. A monoethylene glycol/isopropyl alcohol/deionized water solution is the anti-ice fluid that is used for the TKS system. The fluid solution changes the freezing point and moves rearward on the surfaces. The surfaces the TKS system gives protection to are the wings, the horizontal stabilizers, the struts, the vertical stabilizer, the propeller, and windshield. Overspray from the propeller also protects portions of the fuselage. For a list of approved TKS anti-icing fluids, refer to Chapter 12, Replenishing - Description and Operation.
B.
Airplanes that have the optional TKS system also have the Low Airspeed Awareness (LAA) system installed. This system gives pilots a warning if the airspeed goes below 97.5 KIAS, +2 or -2 knots when operation is in icing conditions. For more data on the LAA system, refer to Low Airspeed Awareness System - Description and Operation (With TKS).
Description A.
The porous panels are laser-drilled titanium installed to the leading edges of the airplane flight surfaces. The panels give TKS ice protection for the wings, wing struts, horizontal and vertical stabilizers. A slinger ring gives ice protection to the propeller and a spray bar gives ice protection to the windshield. The TKS system is divided into two subsystems, the airframe system and the windshield system. The TKS system tank, metering pumps, and related components are installed in the fairing assembly on the bottom of the airplane fuselage. Refer to Figure 1, and Figure 2, and Figure 4. (1) The anti-ice fluid solution comes out of the airframe anti-ice system through flush-fitting titanium porous panels. A laser is used to drill the holes in the porous panels. The porous panels are installed on the leading edge of the wings, stabilizers, and struts. There are three panels on each wing, one panel on each strut, and one panel on each horizontal and vertical stabilizer leading edge. The system gives full protection of the wings leading edge, wing struts, horizontal and vertical stabilizer, but does not include the dorsal fin. The airframe system also includes the propeller slinger ring. (a) The outer skins of the ice protection panels are manufactured from 0.9 mm thick titanium. Titanium gives excellent strength, durability, light weight, and corrosion resistance. (b) The panel holes are 0.0025 inches in diameter, 800 per square inch. The porous area of the titanium panels covers the stagnation point travel on the applicable leading edge in the flight environment the airplane usually operates. (c) The back plates of the porous panels are manufactured with 0.7-mm thick titanium. The inboard wing only is 0.9-mm thick titanium. They are reservoirs for the ice protection fluid to supply all of the porous area. A porous membrane between the outer skin and the reservoir gives smooth flow and distribution through all the porous area of the panel. (d) The porous panels are bonded or attached as a cuff on a leading edge. Panels are bonded to the airframe with a two-part flexible adhesive. (e) Fluid is supplied to the panels and propeller by two positive displacement, constant volume metering pumps. The pumps give different flow rates to the panels and propeller. One pump operation, a combined pump mode, and timed operation give a range of flow rates for different icing conditions. (f) The fluid flows through microfilters before it gets to the porous panels and propeller. The filter removes contamination from the fluid and prevents panel blockage. A system of nylon tubing carries the fluid from the tank to the proportioning units that divide the flow into the volumetric requirements of each panel or device supplied through the unit. The proportioning units are located in the wings and tail of the aircraft and feed each panel and device through nylon tubing. (g) The system has a tank that gives the shortest required quantity of time for ice protection when the fluid is at the sight glass mark. The operation time quantity is more than the operation time given in AC 23.1419-2C. The tank gives a mount for the metering pumps. An isolated accessory bracket holds the windshield pump, timer box, a high pressure switch,
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TKS System Components Figure 1 (Sheet 1)
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TKS Anti-Ice System Flow Diagram Figure 2 (Sheet 1)
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TKS Anti-Ice System Instrument Panel Operation Devices Figure 3 (Sheet 1)
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TKS Fairing Components Figure 4 (Sheet 1)
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(2)
solenoid valve, check valves, and hardware for easier removal and installation. The tank assembly is installed in a fairing below the forward fuselage. Refer to Figure 1 and Figure 4 (h) The tank has a low level switch, that gives a warning annunciation at a given fluid level. The annunciation level occurs in the normal operation mode when there is only 20 minutes of fluid is in the tank with the system. (i) The external filler for the tank is on the left airplane fuselage at FS 176, WL 94.73. Refer to Chapter 12, Replenishing - Description and Operation. (j) The system operates through a series of three control switches. All modes of operation and selection for the metering pumps and the windshield pump are controlled through these devices. Refer to Figure 3 and Figure 4. (k) The serviceable condition of the TKS is monitored with CAS messages and indications that show on the G1000 displays. The windshield anti-ice system applies anti-icing fluid through a spray bar to the pilot's windshield. Refer to Chapter 30, TKS Windshield Spray Bar. (a) Fluid for the windshield spray bar system comes from an on-demand gear pump that is attached to the accessory bracket. The spray bar is operated if necessary to clear forward vision through the windshield.
B.
The system configuration has two main metering pumps. The pumps gives the supply mechanism for all modes of operation of the system, and a pump auxiliary system. The modes of operation are (1) NORMAL, (2) HIGH, (3) MAXIMUM, and (4) BACKUP. (1) NORMAL mode is 66% of the HIGH or design flow rate, and occurs when the two pumps operate for a time cycle of 17% on and 83% off. (2) HIGH mode is the design flow rate for the system and occurs when one pump is run continuously. (3) MAXIMUM mode is a flow rate that is used for a intermittent maximum icing condition, and occurs when both pumps run continuously. MAXIMUM mode is twice the flow rate of HIGH mode. (4) If there is a pump failure, The BACKUP mode gives power to the second pump. The BACKUP mode power is independent of the circuit used for the other modes.
C.
The operation of the TKS system is controlled by three switches on the left panel. The switches are PRIMARY, MAX FLOW, and BACKUP. Figure 3.
D.
The airframe and windshield spray bar anti-ice systems use the anti-icing tank which is in the fairing. The tank assembly is attached to the bottom of the aircraft. Remove the aft fairing to access the fluid tank equipment. To remove and/or install the tank, remove the forward and aft fairings. Refer to Figure 4. (1) Indications on the MFD display show the total fluid available for operation of both the airframe and windshield spray bar anti-ice systems. The tank fluid level monitor devices are electrically operated and receive inputs from a capacitance sensing level sender probe in the tank. Refer to, Figure 4. (2) In addition to the fluid level monitors, the tank has a low level switch. Refer to Figure 4. (a) The low level switch is monitored with a CAS message on the G1000 system. Refer to Table 2. (3) There is a fluid window on the left side of the fairing. At the fluid window you can see the tank sight glass that gives the fluid level indication. This can help you when you fill the tank. Refer to Figure 1 and Figure 4.
E.
The airframe and windshield spray bar anti-ice system have pumps installed on the tank and on the TKS accessory bracket in the fairing. Refer to Figure 4. (1) The anti-ice windshield spray bar pump and the two metering pumps are electric motor driven. (2) An assembly of five filters are installed downstream of the two airframe pumps. Each filter contains a replaceable element. The filter ports are marked IN and OUT for correct plumbing connection.
F.
Proportioning units are installed in four different locations on the airplane. Refer to Figure 2. (1) A seven-place proportioning unit is found in each wing leading edge outboard of the strut attach fittings. (2) A single-place proportioning unit is found in the feed line to the propeller under the floor near the copilot's seat.
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3.
A three-place proportioning unit is found on the floor of the tail cone (RBL 3.35, FS 374.75). This proportioning unit supplies the vertical stabilizer and each horizontal stabilizer.
G.
The proportioning units are metering units which supply anti-icing fluid at a predetermined flow rate for each individual porous panel. The proportioning units incorporate a manifold with calibrated capillary tubes which meter the fluid through the outlet ports. The outlets are marked 1, 2, 3, 4, 5, 6, and 7 on each wing's seven-place proportioning unit. Plumbing to the outlet ports must be connected as specified for proper operation. Refer to Figure 2.
H.
A total of three pressure switches are installed in the TKS system plumbing. There are two low pressure switches and one pressure switch in the system. The pressure switches transmit signals to show CAS messages on the G1000 displays. (1) One pressure switch is installed downstream of the two airframe metering pumps in the fairing. Refer to Figure 4. (a) The pressure switch is monitored with a CAS message on the G1000 displays. Refer to Table 2. (2) There are two low pressure switches to monitor the horizontal stabilizers leading edge panels with one pressure switch for each panel. (a) The low pressure switches are monitored with a CAS message that shows on the G1000 displays. Refer to Table 2.
I.
There are a total of two check valves installed in the fluid tube system downstream of the metering pumps in the fairing. The check valves prevent opposite fluid flow through the tube system.
J.
There is a solenoid valve installed between the tank and the windshield pump to make sure that the fluid in the tubes does not flow back in the tank when the pump is not operating.
K.
There is a strainer for the windshield pump mounted between the tank and the solenoid valve.
Operation A.
Operation of the TKS system is controlled by three FLUID CONTROL switches on the left switch panel. The switches are PRIMARY (SI022) MAX FLOW (SI023), and BACKUP (SI024). Refer to Figure 3 and Figure 4.
B.
There are a total of 18 different switch configurations possible with the three FLUID CONTROL switches. Only six of the switch configurations are usually correct. These switch configurations are shown in Table 1. NOTE:
The MAX FLOW switches only operate momentarily when pushed.
NOTE:
Timer: • Number one comes on for 20 seconds and goes off, and again each 100 seconds. • Number two comes on for 120 seconds and then goes off. • Number three comes on for four seconds and then goes off.
NOTE:
Table 1 gives the operation matrix for the FLUID CONTROL switches, pumps, and timers. Refer to Table 1.
NOTE:
Table 2 gives the CAS Message Triggers and corresponding CAS Messages. Refer to Table 2.
NOTE:
The MAX FLOW only operates with the NORM or HIGH switch ON.
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Table 1. Pumps Operation Matrix for the TKS Anti-Ice System With the G1000 Pumps Operation Matrix For the TKS Anti-Ice System With the G1000 CONTROL SWITCHES PRIMARY Off Norm
MAX FLOW
High
#1
#2 Max Flow
A-ICE A-ICE #3 Wind- NORM HIGH Shield (white) (white)
OFF
ON
OFF
OFF
ON
OFF
OFF
OFF
OFF OFF
OFF
OFF
ON
INT
INT
OFF
ON
ON
OFF
ON
OFF
OFF
ON
INT
OFF
OFF ON
OFF
OFF
ON
TRIP
OFF
INT
INT
ON
ON
OFF
ON
ON
OFF
TRIP
OFF
ON
OFF
ON
OFF OFF
ON
OFF
ON
OFF
OFF OFF
OFF
OFF OFF
OFF
OFF
OFF
OFF
OFF OFF
ON
OFF OFF
ON
OFF
OFF
OFF
OFF OFF
OFF
OFF OFF
OFF
OFF
OFF
ON
INT
ON
OFF
ON
OFF
OFF
ON
OFF
ON
ON
ON
OFF
OFF OFF
OFF
OFF
ON
TRIP
ON
INT
ON
OFF
ON
ON
OFF
ON
OFF
TRIP
ON
ON
ON
OFF
OFF ON
OFF
OFF
ON
TRIP
ON
INT
ON
ON
ON
OFF
ON
ON
OFF
TRIP
ON
ON
ON
ON
OFF OFF
ON
OFF
ON
ON
OFF ON
OFF
OFF OFF
OFF
OFF
OFF
ON
OFF ON
ON
OFF OFF
ON
OFF
OFF
ON
OFF ON
OFF
OFF OFF
OFF
OFF
OFF
#2
OFF
INT
INT
OFF
ON
TRIP
OFF
TRIP
ON ON ***
TRIP TRIP
*** ***
ON ON ON ON ON ON ***
ANNUNCIATORS G1000 CAS MESSAGE
#1
ON
ON
TIMERS
BACK UP
AirWindFrame Shield
ON
ON
PUMPS
TRIP TRIP
*** ***
WindShield
Table 2. Operation Matrix for the TKS Anti-Ice System With the G1000 TKS with G1000 CAS Message Triggers SWITCH LOW LEVEL SWITCH
LOW PRESSURE SWITCH
G1000 CAS Message HIGH PRESSURE SWITCH
ON ON ON
A-ICE LOW PRESS (red)
A-ICE HI PRESS (AMBER)
A-ICE LOW FLUID (AMBER)
ON
OFF
OFF
OFF
ON
OFF
OFF
OFF
ON
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE SYSTEM - MAINTENANCE PRACTICES FAIRING INSTALLATION 1.
2.
General A.
This section contains the removal and installation procedures for the TKS ßuid tank, ßuid tank components, Þlters assembly, and the accessory bracket.
B.
The equipment pack includes the Drain Shut Off Valve, Timer Box, Windshield Pump, Check valves, the Solenoid Valve, a Pressure Switch and a Strainer.
C.
The Þlter assembly is installed in the right side of the fairing, aft of the ßuid tank assembly.
D.
When you remove and install or replace a TKS ßuid tank, it is necessary to do the porous panel purge and test procedure. Refer to TKS Leading Edge Porous Panel - Adjustment/Test.
E.
When you remove and install, or replace a TKS ßuid tank, you can calibrate the ßuid level sender, if necessary. Refer to TKS Anti-Ice System - Adjustment/Test TKS Level Sender Calibration.
F.
For the removal and installation and test procedures for the tail bracket assembly (low pressure switches), refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices.
G.
Recommended maintenance to make sure that the TKS system operates correctly is as follows: • Operate the metering pumps each month, or when necessary, in the HIGH mode to remove the air from the ßuid system. • When you remove and install, or replace a TKS ßuid tank, do the porous panel purge and test procedures.
Tools and Equipment A.
3.
For a list of tools and equipment, (Refer to Ice and Rain Protection - General).
TKS Fairing Assembly Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths in accordance with approved procedures. A.
Remove the Aft Fairing (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Put a support below the aft fairing before you remove the screws. (5) Remove the screws that attach the aft fairing to the fore fairing.
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Remove the screws that attach the aft fairing to the airplane structure. Remove the aft fairing from the airplane.
B.
Install the Aft Fairing (1) Put the aft fairing in its position. (2) Install the screws that attach the aft fairing to the airplane structure. (3) Install the screws that attach the aft fairing to the forward faring. (4) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE.
C.
Remove the Forward Fairing (Refer to Figure 201 and Figure 204). (1) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (2) Drain the ßuid from the ßuid tank. Refer to TKS Tank Fluid Removal in this section. (3) If necessary, remove the Þlter assembly. Refer to Filter Assembly Removal/Installation in this section. (4) If necessary, remove the accessory bracket. Refer to Accessory Bracket Assembly Removal/ Installation in this section. (5) If the accessory bracket and Þlter assembly are not removed do the steps that follow: (a) Disconnect the accessory bracket wire harness from the tank wire harness. (b) Loosen the clamp that attaches the ßuid tank drain tube to the ßuid tank. (c) Remove the tube from the tank coupling. (d) Slowly loosen the nuts that attaches a ßuid tank supply tube to each of the two check valves. (e) Remove the supply tubes from the check valves. (f) Slowly loosen the nut that attaches the windshield pump discharge tube to the bulkhead coupling. (g) Remove the windshield pump discharge tube from the bulkhead coupling. (h) Slowly loosen the nut that attaches the Þlter assembly discharge tube to the bulkhead unequal tee coupling. (i) Remove the Þlter assembly discharge tube from the bulkhead unequal tee coupling. (j) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (6) Disconnect the coaxial connectors from the transponder antenna(s). (7) Remove the sump control cable from the fairing (8) Remove the sump tube from the fairing. (9) Remove the shroud drain tube from the fairing. (10) Remove the fairing and cover from around the nose gear. (11) Put a support below the forward fairing before you remove the screws. (12) Remove the screws that attach the forward fairing to the airplane structure. (13) Using a thin nonmetallic scrapper, select a location at the left forward edge and a location at the right forward edge of the cargo pod where a fuselage internal structural member exists, and perforate the seal between the fuselage and pod. (14) Using two 6 to 8 foot lengths of 0.032 stainless steel safety wire, fabricate a seal cutting tool by twisting the two pieces of wire together using safety wire pliers. (15) Feed one end of the seal cutting tool through the two existing perforations on each side of the fairing. (16) With one person inside the pod and a second on the other side of the forward fairing, wrap the cutter wire at both ends around a block of wood or similar tool to serve as handles and begin sawing through the seal. (a) Insert wooden tongue depressors, or similar tool, between fuselage and forward fairing at 1 to 2 foot increments to prevent the seal from reattaching. (17) Remove fairing from the fuselage. (18) Carefully remove unwanted sealant from the fairing and fuselage.
D.
Install the Forward Fairing (Refer to Figure 201 and Figure 204). (1) Put the forward fairing in its position. (2) Apply sealant tape to the fairing ßange. (3) Install the screws that attach the forward fairing to the airplane structure.
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(13) (14) (15) (16)
Seal fairing ßange edge to fuselage with Type I Class B sealant. Install the nose gear cover. (a) Seal cover to fuselage, nose gear spring, and fairing with Type VIII, Class B sealant. Install the nose gear fairing. Install the sump control cable in the fairing Install the sump tube in the fairing. Install the fuel shroud drain tube. (a) Seal the drain tube with Type I Class B sealant inside and outside the fairing. If necessary, install the accessory bracket. Refer to Accessory bracket Removal/Installation in this section. If necessary, install the Þlter assembly. Refer to Filter Assembly - Removal/Installation in this section. If the accessory bracket and Þlter assembly were not removed do the steps that follow: (a) Connect the accessory bracket wire harness to the tank electrical connector. 1 Attach the wire harness to appropriate locations with tie wraps. (b) Remove the caps on all the openings and tube ends of the ßuid system. (c) Put the drain tube in its position on the tank coupling. (d) Tighten the clamp that attaches the ßuid tank drain tube to the ßuid tank. (e) Put the supply tubes in their position on the check valves with new seals. (f) Slowly tighten the nuts that attaches a ßuid tank supply tube to each of the two check valves. (g) Put the windshield discharge tube in its position on the bulkhead coupling with a new seal. (h) Slowly tighten the nut that attaches the windshield pump discharge tube to the bulkhead coupling. (i) Put the Þlter assembly discharge tube in its position on the bulkhead unequal tee coupling with a new seal. (j) Slowly tighten the nut that attaches the Þlter assembly discharge tube to the bulkhead unequal tee coupling. Connect the coaxial connectors to the two Transponder antennas. Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. Do a test of the TKS system. Refer to TKS Anti-Ice System - Adjustment/Test. Install the aft fairing. Refer to Install the Aft Fairing in this section.
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4.
TKS Fluid Removal
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, if necessary, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Fluid (Refer to Figure 201 and Figure 202). (1) Remove the aft fairing to get access to the ßuid tank assembly. Refer to Remove the Aft Fairing in this section. (2) Put a container with a capacity of approximately 3 to 5 gallons below the drain tube outlet. NOTE: (3) (4) (5) (6) (7)
If necessary, a longer drain tube can be temporarily connected to the drain outlet to prevent ßuid spill. The longer drain tube causes the ßuid to drain more quickly.
Remove the safety wire from the drain valve Push the lever to the open position on the drain valve to release the ßuid. Pull the valve closed to stop the drain procedure. Safety the drain valve with wire. Refer to Chapter 12, TKS Anti-Ice System - Servicing for the servicing procedures. NOTE:
You must calibrate the ßuid level sender if the primary ßight display (G1000) does not read zero when the TKS ßuid tank is empty. Refer to TKS Anti-Ice System Adjustment/Test, TKS Level Sender Calibration.
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TKS System Installation Figure 201 (Sheet 1)
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TKS System Installation Figure 201 (Sheet 2)
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TKS System Installation Figure 201 (Sheet 3)
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TKS System Installation Figure 201 (Sheet 4)
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TKS System Installation Figure 201 (Sheet 5)
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TKS Nylon Tubing Assembly Figure 202 (Sheet 1)
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TKS Nylon Tubing Assembly Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (8) 5.
Install the aft fairing. Refer to Install the Aft Fairing in this section.
Filter Assembly Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Filter Assembly (Refer to Figure 201 and Figure 202). (1) Remove external electrical power form the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (5) Loosen the nut on Þlter input tube at the tee on the accessory bracket. (6) Drain ßuid in bucket. (7) Slowly loosen the nut that attaches input manifold elbow to the Þlter assembly. (8) Remove the elbow from the Þlter assembly. NOTE: (9)
Use the removed elbow if new Þlter assembly is installed.
Slowly loosen the nut that attaches the output manifold elbow to the Þlter assembly.
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MODEL 208 MAINTENANCE MANUAL (10) Remove elbow from the Þlter assembly. NOTE: (11) (12) (13) (14) B.
Use the removed elbow if new Þlter assembly is installed.
Put caps on all openings and tube ends to keep FOD out of the ßuid system. Remove the screws that attach the Þlter assembly to the inner wall of the fairing. Remove the Þlter assembly from the fairing. If you are to install a new Þlter, remove the Þlter from the bracket.
Install the Filter Assembly (Refer to Figure 201 and Figure 202). (1) If you are to install a new Þlter, install the Þlter on the bracket. (2) Put the Þlter assembly in its position on the inner wall of the fairing. (3) Install the screws that attach the Þlter assembly to the fairing. (4) Remove the caps from the tube ends. (5) Install new seals on the tubes and manifold elbows as shown in Figure 202. (6) Install the input manifold elbow if removed. (7) Put the input tube in its position on the input manifold elbow. (8) Slowly tighten the nut that attaches the tube to the input manifold elbow. (9) Install the output manifold elbow if removed. (10) Put the discharge tube in its position on the output manifold elbow. (11) Slowly tighten the nut that attaches the discharge tube to the output manifold elbow. (12) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (13) Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (14) Supply external electrical power to the airplane. (15) Put the EXTERNAL POWER switch (S17) on the circuit breaker switch panel to the BUS position. (16) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel to the HIGH position. (a) Make sure that there is no ßuid leakage from the couplings. (17) Purge air from the TKS system. Refer to TKS Anti-Ice Leading Edge Porous Panel - Adjustment/ Test. (18) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch on the left switch panel to the OFF position. (19) Put the EXTERNAL POWER switch on the circuit breaker switch panel to the OFF position. (20) Remove external electrical power from the airplane. (21) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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6.
TKS Accessory Bracket Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Accessory Bracket (Refer to Figure 201 and Figure 202). (1) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (2) Remove external electrical power from the airplane. (3) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (4) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (5) Make sure that the ßuid is drained from the ßuid tank . Refer to TKS Fluid Removal in this section. (6) Disconnect the electrical connector from the tank electrical harness. (7) Slowly loosen the nut that attaches the outlet tube to the bulkhead unequal tee. (8) Remove the tube from the T-Þtting. (9) Loosen the clamp that attaches the ßuid tank drain tube to the strainer assembly. (10) Remove the tube from the strainer. (11) Slowly loosen the nuts that attach the ßuid tank supply tubes to each of the two check valves. (12) Remove the tube from each of the two check valves. (13) Slowly loosen the nut that attaches the discharge tube to the windshield pump coupling. (14) Remove the discharge tube from the windshield pump.
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MODEL 208 MAINTENANCE MANUAL (15) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (16) Remove the screws and washers that attach the accessory bracket to the bottom of the fairing. (17) Remove the accessory bracket from the fairing. B.
Install the Accessory Bracket (Refer to Figure 201 and Figure 202). (1) Put the accessory bracket in its position on the bottom of the fairing. (2) Install the screws and washers that attach the accessory bracket to the fairing. (3) Remove the caps from the tube ends. (4) Install new seals in the couplings as shown in Figure 202. (5) Put the tube in its position on the bulkhead unequal tee. (6) Slowly tighten the nut that attaches the tube to the unequal tee. (7) Put the drain tube in its position on the strainer. (8) Slowly tighten the clamp that attaches the drain tube to the strainer. (9) Put the tubes in their position on each of the two check valves. (10) Slowly tighten the nut that attach the ßuid tank supply tubes to each of the two check valves. (11) Put the discharge tube in its position the windshield pump. (12) Slowly tighten the nut that attaches the discharge tube to the windshield pump coupling. (13) Connect the electrical connector to the tank electrical harness. (14) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (15) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (16) Do a test of the TKS system. Refer to TKS Anti-Ice System - Adjustment/Test. (17) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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7.
TKS Fluid Tank Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Fluid Tank (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Put the BATTERY switch (SC005) on the circuit breaker switch panel, in the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the forward and aft TKS fairings. Refer to TKS Fairing Assembly Removal/Installation in this section. (5) Make sure that the ßuid is drained from the ßuid tank . Refer to TKS Fluid Removal in this section. (6) Identify and disconnect the ßuid tank electrical connectors from the airplane fuselage connector. (7) Remove ßoor covering. Refer to FLOOR COVERING/CONTROL COLUMN COVER Maintenance Practices. (8) Remove ßoor panels, (232BC), (231DL), and (232BR). Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (9) Loosen, but do not remove the Þller tube clamp at the Þller neck on the ßuid tank. (a) Remove the Þller tube from the Þller neck.
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MODEL 208 MAINTENANCE MANUAL (10) Loosen, but do not remove the vent hose clamps at the necks on the ßuid tank. (a) Remove the vent tubes from the vent necks. (11) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (12) Slowly loosen the nut that attaches the ßuid tank supply tube to each of the two pump outlet tees. (a) Remove the tubes from each of the two pump outlet tees. (13) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (14) Put a support below the ßuid tank before you remove the screws. (15) Remove the bottom screws from the fore, aft, left and right shear plates that attach the ßuid tank to the fuselage structure. (16) Loosen the strap T-bolt on the forward strap assembly of the ßuid tank. (17) Loosen the strap T-bolt on the aft strap assembly of the ßuid tank. (18) Disconnect the forward and aft strap assemblies. (19) Carefully lower the ßuid tank assembly. (20) Make sure that all openings and tube ends have caps installed. B.
Install the Fluid Tank (Refer to Figure 201 and Figure 202). (1) Remove the caps from the components that follow: • Pump outs • Drain tube • Filler neck • Vent neck. (2) Remove the caps from the vent tubes. (3) Carefully lift the ßuid tank assembly. (a) Align the shear plates to the attach points. (4) Install the bottom screws that attach the ßuid tanks fore, aft, left, and right shear plates to the airplane structure. (5) Put each of the two straps together with its related T-bolt. (6) Torque the T-bolt in each of the two tank straps to 20 inch-pounds (2.25 N-m). (7) Put the Þller tube in its position on the Þller neck of the ßuid tank. (8) Wrap tape around the tube end. (9) Tighten the tube clamp that attaches the Þller tube to the Þller neck. (a) Make sure that the clamp is positioned on the Þller tube tape before you tighten the clamp. (10) Put the vent tubes in position on the vent necks of the ßuid tank. (11) Tighten the tube clamps that attach the vent tubes to the vent necks. (12) Put the drain tube in its position on the strainer. (13) Tighten the clamp that attaches the ßuid tank drain tube to the strainer assembly. (14) Put the supply tubes on each of the two pump outlet tees with new seals in their positions. Refer to Figure 202. (15) Tighten the nuts that attach the supply tubes to each of the two pump outlet tees on the ßuid tank. (16) Connect all the electrical connectors to the airplane fuselage connector. (17) Install the forward fairing. Refer to Install the Forward Fairing in this section. (18) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. NOTE:
You must calibrate the ßuid level sender if the primary ßight display (G1000) does not read zero when the TKS ßuid tank is empty. Refer to TKS Anti-Ice SystemAdjustment/Test, TKS Level Sender Calibration.
(19) Do a test of the ßuid tank components. Refer to TKS Anti-Ice System - Adjustment/Test. (20) Install the aft fairing. Refer to Install the Aft Fairing in this section. (21) Install the access panels (232BC), (231DL), and (232BR). Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (22) Install the ßoor covering.
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8.
Metering Pump Assembly Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, if necessary, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. NOTE: A.
The removal and installation of metering pump 1 and metering pump 2 are typical.
Remove the Metering Pump (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (5) Make sure that the ßuid is drained from the ßuid tank . Refer to TKS Fluid Removal in this section. (6) If necessary, remove the accessory bracket to get access to the pump. Refer to TKS Accessory Bracket Removal/Installation in this section. (7) Identify and disconnect the electrical connectors from the pump. (8) Slowly loosen the nut that attaches the output tube to the pump outlet tee. (9) Remove the tube from the pump. (10) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (11) Remove the screws that attach the pump to the ßuid tank bracket.
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MODEL 208 MAINTENANCE MANUAL (12) Remove the pump from the bracket. B.
Install the Metering Pump (Refer to Figure 201 and Figure 202). (1) If installing the same pump that was removed, install new seals on the pump. (2) Apply a light layer of TKS ßuid on the seals between the ßuid tank and the pump hose adapter. (3) Put the pump in its position in the pump bracket. (4) Install the screws that attach the pump to the ßuid tank bracket. (a) Make sure that the ground terminal is installed under one of the screws. (5) Install the electrical connectors to the pump. (6) Remove the caps from the tube ends. (7) Install new seals in the couplings as shown in Figure 202. (8) Put the output tube in its position on the pump. (9) Tighten the nut that attaches the output tube to the pump. (10) If removed, install the accessory bracket. Refer to TKS Accessory Bracket Removal/Installation in this section. (11) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (12) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (13) Do a test of the pump. Refer to TKS Anti-Ice System - TKS Anti-Ice System, Do a Test of Metering Pump 1 and Do a Test of Metering Pump 2 (14) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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9.
Windshield Pump Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Windshield Pump (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Make sure that the ßuid is drained from the ßuid tank . Refer to TKS Fluid Removal in this section. (5) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (6) Identify and disconnect the electrical connector from the pump. (7) Slowly loosen the nut that attaches the supply tube to the pump. (8) Remove the supply tube from the pump. (9) Slowly loosen the nut that attaches the discharge tube to the pump coupling. (10) Remove the discharge tube from the pump. (11) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (12) Remove the screws and washers that attach the pump to the accessory bracket. (13) Remove the pump from the accessory bracket.
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MODEL 208 MAINTENANCE MANUAL B.
10.
Install the Windshield Pump (Refer to Figure 201 and Figure 202). (1) Put the pump in its position on the accessory bracket. (2) Install the screws, washers, and spacers that attach the pump to the accessory bracket. (3) Connect the electrical connector to the pump. (4) Remove the caps from the tube ends. (5) Install new seals in the couplings as shown in Figure 202. (6) Put the supply hose in its position on the pump. (7) Tighten the pump coupling nut to attach the supply tube to the pump. (8) Put the discharge hose in its position on the pump. (9) Tighten the pump coupling nut to attach the discharge tube to the pump. (10) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (11) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (12) Do a test of the windshield pump. Refer to TKS Anti-Ice System - Adjustment/Test, Do a Test of the Windshield Pump. (13) Install the aft fairing. Refer to Install the Aft Fairing in this section.
Fluid Level Sender Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the Fluid Level Sender (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the cockpit ßoor covering. Refer to FLOOR COVERING/CONTROL COLUMN COVER - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL (5)
Remove the cockpit ßoor access panel 232BC .
WARNING: Do not remove hoses under pressure. This procedure will result in release of refrigerant into the atmosphere. Removing hoses under pressure may also result in personal injury if hose ends are not restrained. (6) (7) (8) (9)
If necessary, disconnect and move the air conditioning lines to get access to the level sender. Refer to Chapter 21, R134A Air Conditioning - Maintenance Practices, Air Conditioning Plumping Removal/Installation. Identify and disconnect the electrical wiring leads from the sender electrical posts. Remove the screws that attach the sender to the ßuid tank access panel. Carefully remove the sender from the access panel. NOTE:
Do not damage the sender sensor.
(10) Install a temporary cover on the opening to keep FOD out of the ßuid system. (11) Discard the gasket. B.
Install the Fluid Level Sender (Refer to Figure 201 and Figure 202). (1) Remove the temporary cover from the access panel opening. (2) Put the sender and a new gasket in their position on the access panel.
(3) (4) (5) (6) (7) (8)
(9)
NOTE:
Be careful not to damage the sensor.
NOTE:
Be careful to not push the three wires down through the hole.
Install the screws that attach the sender to the access panel. Connect the electrical wiring leads to the sender negative, positive, and send posts. If necessary, connect the air conditioning lines. Refer to Chapter 21, R134A Air Conditioning Maintenance Practices, Air Conditioning Plumping Removal/Installation. Install the cockpit ßoor access panel,232BC. Install the cockpit ßoor covering. Refer to Chapter 25, FLOOR COVERING/CONTROL COLUMN COVER - Maintenance Practices. Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. NOTE:
You must calibrate the ßuid level sender if it does not read zero when it is empty. Refer to TKS Anti-Ice System - Adjustment/Test.
(10) Do a test of the sender. Refer to TKS Anti-Ice System - Adjustment/Test.
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11.
Low Level Switch Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Low Level Switch (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the aft TKS fairing. Refer to TKS Fairing Assembly - Removal/Installation in this section. (5) If necessary, remove the accessory bracket. Refer to Accessory Bracket Assembly Removal/ Installation in this section. (6) Identify and disconnect the electrical connector from the switch. (7) Remove the low level switch from the ßuid tank. Refer to Figure 201. NOTE: (8)
A deep socket modiÞed to accommodate the switch wiring is necessary.
Put a cover on the switch opening to keep FOD out of the ßuid system.
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MODEL 208 MAINTENANCE MANUAL B.
Install the Low Level Switch (Refer to Figure 201 and Figure 202). (1) Remove the cover from the low level switch opening. (2) Mark the switch where the wires exit to identify the top of the switch. (3) Put Type I Class B sealer on the switch threads. (4) Install the switch in the ßuid tank. Refer to Figure 201. NOTE:
(5) (6) (7)
(8)
Make sure that you install the switch correctly so the mark you made on the switch is on top. If you install the ßoat correctly, it moves vertically.
Connect the electrical connector to the switch. If removed, install the accessory bracket. Refer to Accessory Bracket Assembly Removal/Installation in this section. Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. NOTE:
You must calibrate the ßuid level sender if the primary ßight display (G1000) does not read zero when the TKS ßuid tank is empty. Refer to TKS Anti-Ice System Adjustment/Test, TKS Level Sender Calibration.
(9)
Do a test of the level switch. Refer to TKS Anti-Ice System - Adjustment/Test, Do a Test of the Low Level Switch. (10) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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12.
Pressure Switch Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Pressure Switch (Refer to Figure 201 and Figure 202). (1) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (2) Remove external electrical power from the airplane. (3) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (4) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (5) Identify and disconnect the electrical connector from the switch. (6) Disconnect the pallet output tube. (7) Put the pallet output tube in bucket to drain tubing. (8) Slowly loosen and disconnect the coupling nut that attaches the input tube to the switch. (9) Remove the tube from the switch. (10) Slowly loosen and disconnect the coupling nut that attaches the output tube to the switch. (11) Remove the tube from the switch. (12) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (13) Remove the screws that attach the switch bracket to the equipment pack. (14) Remove the switch and bracket.
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MODEL 208 MAINTENANCE MANUAL (15) Remove the bracket from the switch. B.
13.
Install the Pressure Switch (Refer to Figure 201 and Figure 202). (1) Put the switch in its position in the switch bracket. (2) Install the screws and spacers that attach the switch to the bracket. (3) Put the switch and bracket in its position on the equipment pack. (4) Install the screws that attach the switch and bracket to the equipment pack. (5) Remove the caps from the tube ends. (6) Install new seals for the input tube. (7) Put the input tube in its position on the couplings. (8) Tighten the coupling nut that attaches the input tube to the switch. (9) Install new seals for the output tube. (10) Put the output tube in its position on the couplings. (11) Tighten the coupling nut that attaches the output tube to the switch. (12) Connect the electrical connector to the switch. (13) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (14) Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (15) Do a test of the switch. Refer to TKS Anti-Ice System - Adjustment/Test, Do a Test of the Pressure Switch. (16) Install the aft fairing. Refer to Install the Aft Fairing in this section.
Timer Box Removal/Installation A.
Remove the Timer Box (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (5) Identify and disconnect the electrical connector from the timer box. (6) Remove the screws that attach the timer box to the accessory bracket. (7) Remove the timer box.
B.
Install the Timer Box and/or Wire Bundle (Refer to Figure 201 and Figure 202). (1) Put the timer box in its position on the accessory bracket. (2) Install the screws that attach the timer box to the accessory bracket. (a) Make sure that you install the grounding terminal under one of the screws. (3) Connect the electrical connector to the timer box. (4) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (5) Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (6) Do a test of the ßuid tank components. Refer to TKS Anti-Ice System - Adjustment/Test. (7) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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14.
Solenoid Valve Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Solenoid Valve (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (5) Remove the ßuid from the ßuid tank. Refer to TKS Fluid Removal in this section. (6) Identify and disconnect the electrical connector from the valve. (7) Slowly loosen and disconnect the solenoid valve coupling nut that attaches the input tube to the valve. (8) Remove the tube from the valve. (9) Slowly loosen and disconnect the valve coupling nut that attaches the output tube to the valve. (10) Remove the tube from the valve. (11) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (12) Remove the screws that attach the solenoid valve to the accessory bracket. (13) Remove the solenoid valve.
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MODEL 208 MAINTENANCE MANUAL B.
Install the Solenoid Valve (Refer to Figure 201 and Figure 202). (1) Put the solenoid valve in its position in the valve bracket. (2) Install the screws that attach the valve to the ßuid tank bracket. (3) Remove the caps from the tube ends. (4) Install new seals in the couplings as shown in Figure 202. (5) Put the input tube in its position on the coupling. (6) Tighten the coupling nut that attaches the input tube to the valve. (7) Put the output tube in its position on the coupling. (8) Tighten the coupling nut that attaches the output tube to the valve. (9) Connect the electrical connector to the valve. (10) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (11) Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (12) Put external electrical power on the airplane. (13) Put the EXTERNAL POWER switch (S17) on the circuit beaker switch panel to the BUS position. (14) Put the MAX FLOW switch to the WINDSHIELD position. (a) Make sure that there is no ßuid leakage from the couplings. (b) Make sure that ßuid comes out of the windshield spray bar. (15) Put the EXTERNAL POWER switch on the circuit breaker switch panel to the OFF position. (16) Remove external electrical power from the airplane. (17) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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15.
Check Valve Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Check Valve (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (5) Remove ßuid from the TKS tubing. (a) Disconnect the pallet output tube. (b) Put the pallet output tube in bucket to drain tubing. (6) Slowly loosen the coupling nut that attaches the output tube to the check valve. (7) Remove the output tube from the check valve. (8) Slowly loosen the coupling nut that attaches the supply tube to the check valve. (9) Remove the supply tube from the check valve. (10) Put caps on all openings and tube ends to keep FOD out of the ßuid system. (11) Remove the screw that attaches the check valve clamp to the accessory bracket. (12) Remove the check valve from the equipment pack.
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MODEL 208 MAINTENANCE MANUAL B.
Install the Check Valve (Refer to Figure 201 and Figure 202). (1) Put the valve and its clamp in their position on the accessory bracket. (a) Make sure that the ßuid ßow direction is correct. (2) Install the screw and spacer that attaches the valve clamp to the accessory bracket. (3) Remove the caps from the tube ends. (4) Install new seals on the tube ends as shown in Figure 202. (5) Put the input tube in its position on the coupling. (6) Tighten the coupling nut that attaches the input tube to the valve. (7) Put the output tube in its position on the couplings. (8) Tighten the coupling nut that attaches the output tube to the valve. (9) Safety all the tube couplings. Refer to Chapter 20, Safetying - Maintenance Practices. (10) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (11) If necessary, do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (12) Put external electrical power on the airplane. (13) Put the EXTERNAL POWER switch (S17) on the circuit breaker switch panel to the BUS position. (14) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel to the HIGH position. (15) Put the BACKUP switch in the ON position. (a) Make sure that there is no ßuid leakage from the check valve couplings. (b) Make sure that ßuid ßows from the outlet tube. (16) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel to the OFF position. (17) Put the BACKUP switch in the OFF position. (18) Connect the output tube. (19) Put the EXTERNAL POWER switch on the circuit breaker switch panel to the OFF position. (20) Remove external electrical power from the airplane. (21) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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16.
Sight Glass Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the Sight Glass (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the forward and aft fairings. Refer to TKS Fairing Assembly Removal/Installation in this section. (5) Make sure that the ßuid is drained from the ßuid tank . Refer to TKS Fluid Removal in this section. (6) Slowly open the tube clamps that are connected to the sight glass. (7) Remove the sight glass (and ball) from the ßuid tank. (8) Put caps on all openings and tube ends to keep FOD out of the ßuid system.
B.
Install the Sight Glass (Refer to Figure 201 and Figure 202). (1) Remove the caps from the tube ends. (2) Put the sight glass (and ball) in its position in the sight glass brackets. (3) Crimp the tube clamps. (4) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (5) Install the forward fairing. Refer to Install the Forward Fairing in this section. (6) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (a) Make sure that the sight glass tubes do not leak. (b) Make sure that the ball moves freely in the tube. (7) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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17.
Drain Valve Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the Drain Valve (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (5) Remove the safety wire from the valve. (6) Remove the ßuid from the ßuid tank. Refer to TKS Fluid Removal in this section. (7) Loosen the drain tube nut. (8) Remove drain tube from the drain valve. (9) Remove the nut and washer that attach the drain valve to the accessory bracket. (10) Remove the valve from the accessory bracket.
B.
Install the Shutoff Valve (Refer to Figure 201 and Figure 202). (1) Put the drain valve in its correct position on the accessory bracket. (2) Install the valve with the nut and washer. (3) Put the drain tube in its position on the valve. (4) Tighten the nut that attaches the drain tube to the valve. (5) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (6) Safety wire the valve in the closed position. (7) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (a) Make sure that there is no ßuid leakage from the valve. (8) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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18.
Fluid Filler Tube Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. Refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, if necessary, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Filler Tube (Refer to Figure 201, and Figure 203). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the ßoor coverings necessary to get access to the Þller tube. Refer to Floor Covering/ Column Cover - Maintenance Practices. (5) Remove ßoor panels (251AL), (251BL), (251CL), (232BC), and (251HL), to get access to the Þller tube. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (6) At the left side of the cabin interior at FS 176.00, WL 94.73 remove airplane side wall. (7) Loosen, but do not remove the hose clamp at the ßuid tank Þller neck. (8) Remove the Þller tube from the Þller neck. (9) At the left side of the cabin interior at FS 176.00, WL 94.73 loosen the Þller tube clamp on the Þller port sleeve. (10) Remove the Þller tube from the Þller port sleeve.
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TKS Filler and Vent Tube Installation Figure 203 (Sheet 1)
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Fuel Sump Assembly Installation Figure 204 (Sheet 1)
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Fuel Sump Assembly Installation Figure 204 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (11) Cut the tie wraps at the tube mounts and remove the tie wraps from the airplane. (a) Make sure that you remove the cut tie wraps from the airplane. (12) Carefully remove the Þller tube from the airplane ßoor structure. (13) Put caps on all openings to keep FOD out of the ßuid system. B.
Install the Filler Tube (Refer to Figure 201, and Figure 203). (1) Remove the caps from the openings. (2) Prepare the end of the hose to Þt on the Þller port sleeve: (a) Unwrap the string from the hose that will go over the Þller port sleeve. (b) Use pliers to pull the wire so that on the part of the hose that will go over the Þller port sleeve, the wire is straight. Make sure that you do not make a tight bend in the wire. 1 2 Make sure that you do not damage the tube wall. (c) Cut off the part of the string and wire that go past the end of the tube. (3) Prepare the end of the hose to Þt on the Þller neck: (a) Unwrap the string from the hose that will go over the Þller neck. (b) Use pliers to pull the wire so that on the part of the hose that will go over the Þller neck, the wire is straight. 1 Make sure that you do not make a tight bend in the wire. 2 Make sure that you do not damage the tube wall. (c) Cut off the part of the string amnd wire that go past the end of the tube. (4) Install the Þller tube through the correct openings in the ßoor structure. NOTE:
(5) (6) (7) (8) (9) (10) (11) (12) (13) (14)
(15) (16) (17) (18)
The correct routing of the Þller tube has openings with grommets or mounts installed to prevent damage to the tube.
Put the Þller tube in its position on the Þller port sleeve. Wrap silicone tape around tube where clamp is installed. Put the tube clamp in its correct position on the tube where the tape is wrapped. Tighten the tube clamp at the Þller port sleeve. (a) Make sure that the clamp is positioned on the Þller tube tape before you tighten the clamp. (b) Make sure that the clamp tightening screw is not positioned over the wire. Put the Þller tube in its position on the ßuid tank Þller neck. Wrap silicone tape around tube where clamp is installed. Put the tube clamp in its correct position on the tube where the tape is wrapped. Tighten the hose clamp on the ßuid tank Þller neck. (a) Make sure that the clamp is positioned on the Þller tube tape before you tighten the clamp. (b) Make sure that the clamp tightening screw is not positioned over the wire. Install tie wraps at the tube mounts. Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing. Install the ßoor panels (251AL), (251BL), (251CL), (252BC), and (251HL). Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. Install the ßoor coverings. Refer to Floor Covering/Column Cover - Maintenance Practices. Install the aft fairing. Refer to Install the Aft Fairing in this section.
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19.
Fluid Filler Port Assembly Removal/Installation
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, if necessary, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. A.
Remove the Filler Port Assembly (Refer to Figure 201, and Figure 202). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Open the Þller cap. (5) At the left side of the cabin interior at FS 176.00, WL 94.73 loosen the Þller tube clamp on the Þller port sleeve. (6) Remove the Þller tube from the Þller port sleeve. (7) Remove the screws, nuts, and washers that attach the Þller port plate, Þller port sleeve, and gaskets to the airplane skin. (8) Put caps on all openings to keep FOD out of the ßuid system.
B.
Install the Filler Port Assembly (Refer to Figure 201, and Figure 202). (1) Remove the caps from the openings. (2) Put the Þller port sleeve gasket in its position at the cabin interior.
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Hold the Þller port sleeve, Þller port plate gasket and Þller port plate in their position. NOTE:
Make sure that the Þller port plate is positioned with the slots at the top and bottom so the cap will Þt correctly.
(4)
Install the screws, nuts, and washers that attach the Þller port sleeve, Þller port plate gasket and Þller port plate to the airplane skin. (5) Put the Þller tube in its position on the Þller port sleeve. (6) Wrap silicone tape around tube where clamp is installed. (7) Put the tube clamp in its correct position on the tube where the tape is wrapped. (8) Tighten the tube clamp at the Þller port sleeve. (9) Close the Þller cap. (10) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (11) Do the ßuid tank servicing if necessary. Refer to Chapter 12, TKS Anti-Ice System - Servicing.
20.
Vent Tube Removal/Installation A.
Remove the Vent Tube (Refer to Figure 201, and Figure 203). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Remove the ßoor covering. Refer to Chapter 25, Floor Covering/Column Cover - Maintenance Practices. (5) Remove the ßoor panels (251CL), (251BL), (251HL), (231DL), (232BC), (232DR), AND (232BR). Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (6) At the left side of the cabin interior at FS 176.00, WL 94.73 remove airplane side wall. (7) Cut the tie wraps at the tube mounts. (8) Loosen the clamps at the ßuid tank, vent weldment, and at the vent tube tee. (9) Remove the vent tube from the airplane.
B.
Install the Vent Tube (Refer to Figure 201, and Figure 203). (1) Remove the caps from the openings. (2) Install the vent tube through the correct openings in the ßoor structure. NOTE:
The correct routing of the Þller tube has openings with grommets installed to prevent damage to the tube.
(3) (4) (5) (6) (7) (8) (9) (10)
Put the vent tube in its position on the vent weldment. Tighten the tube clamp at the vent weldment. Put the vent tube in its correct position on the ßuid tank vent port. Tighten the hose clamp at the ßuid tank vent port. Put the vent tubes in their correct position on the vent tube tee. Tighten the tube clamps at the vent tube tee. Install tie wraps at the tube mounts. Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (11) Install the ßoor panels (251CL), (251BL), (251HL), (231DL), (232BC), (232DR), AND (232BR). Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (12) Install the ßoor coverings. Refer to Floor Covering/Column Cover - Maintenance Practices. (13) Install the aft fairing. Refer to Install the Aft Fairing in this section.
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21.
Pump Strainer Removal/Installation A.
Remove the Pump Strainer (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (4) Drain the ßuid tank. (5) Remove the clamp from the weld assembly. (6) Cut and remove the safety wire. (7) Unscrew strainer from weld assembly. (8) Check and make sure that there is not any foreign objects in the weld assembly.
B.
Install the Pump Strainer (Refer to Figure 201). (1) With a new O-ring and seal, screw the strainer on the weld assembly. NOTE: (2) (3) (4) (5) (6)
Make sure that the seal is next to the hex head on the strainer.
Put the clamp in its correct position on the weld assembly. Tighten the clamp on the weld assembly. Safety wire the clamp. Service the tank Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE.
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22.
TKS Fluid Contamination (Fuel) Removal
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, if necessary, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 202. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. CAUTION: Do not operate the windshield pump for more than 10 seconds continuously, and wait 10 seconds between pump operations before you operate the pump again. Damage to the windshield pump can occur if the pump is operated for more than the speciÞed limit. A.
Remove the TKS Fluid Contamination (Fuel) (Refer to Figure 201). (1) Remove the aft fairing. Refer to Remove the Aft Fairing in this section. (2) Remove external electrical power from the airplane. (3) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (4) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (5) Drain the TKS ßuid from the ßuid tank . Refer to TKS Fluid Removal in this section. (6) Disconnect the Þlter inlet tube from the Þlter inlet manifold assembly. (a) Put the inlet tube end into a bucket with a capacity of approximately 3 to 5 gallons.
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19)
(20) (21)
(22) (23)
Disconnect the windshield pump outlet tube from the fuselage connector. (a) Put the open tube end into a bucket with a capacity of approximately 1 gallon. Fill the ßuid tank with a water and mild detergent mixture. Drain the water and mild detergent mixture from the ßuid tank. Fill the ßuid tank with water. Drain the water from the ßuid tank. Remove the Fluid Level Sender. refer to Fluid Level Sender Removal/Installation in this section. Clean the ßuid level sender with a water and mild detergent mixture. Flush the ßuid level sender with water until no contamination shows. Install the Fluid Level Sender. Refer to Fluid Level Sender Removal/Installation in this section. Add 10 gallons (37.84 liters) of TKS ßuid to the ßuid tank. Supply external electrical power to the airplane. Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel to the HIGH position to start pump 1. Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch (SI023) on the left switch panel to the AIRFRAME position, then release the switch. (a) Let the TKS ßuid ßow into the bucket at the Þlter inlet tube for one full cycle of approximately 2 minutes. Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch to the OFF position. Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch (SI023) on the left switch panel to the WINDSHIELD position, then release the switch. NOTE:
The windshield pump will start and operate when you put the spring-loaded MAX FLOW switch to the WINDSHIELD position.
(a)
(24) (25)
(26) (27)
Let the TKS ßuid ßow into the bucket at the windshield pump outlet tube for one full cycle of approximately 4 seconds. Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. Remove external electrical power from the airplane. If the system was operated with fuel contamination in the ßuid, replace the Þlter assembly. Refer to Filter Assembly Removal/Installation in this section. NOTE:
If there was fuel contamination in the ßuid, but system was not operated, it is not necessary to replace the Þlter assembly.
(28) (29) (30) (31)
Connect the Þlter inlet tube with a new O-ring to the Þlter inlet manifold assembly. Connect the windshield pump outlet tube with a new O-ring to the fuselage connector. Supply external electrical power to the airplane. Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. (32) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (33) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel to the HIGH position to start pump 1.
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MODEL 208 MAINTENANCE MANUAL (34) If the Þlter assembly was replaced, put the ANTI-ICE-FLUID CONTROL, BACKUP switch (SI024) on the left switch panel to the ON position to start pump 2. (a) Let the system operate for approximately 5 minutes until there are no air bubbles in the TKS ßuid ßow from the porous panels on the wing, wing strut, and horizontal and vertical stabilizer leading edges. NOTE:
If necessary, you can let the system operate for more than 5 minutes until the ßuid ßow is normal across all porous panels.
(b)
If ßuid ßow at any of the porous panels does not become normal, do the porous panel purge and test procedure for the applicable porous panel(s). Refer to TKS Leading Edge Porous Panel - Adjustment/Test. (35) If the Þlter assembly was not replaced, put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch (SI023) on the left switch panel to the AIRFRAME position, then release the switch. (a) Let the TKS ßuid ßow from the porous panels on the wing, wing strut, and horizontal and vertical stabilizer leading edges for one full cycle of approximately 2 minutes. NOTE:
If necessary, you can do more MAX FLOW cycles until the ßuid ßow is normal across all porous panels.
(b)
If ßuid ßow at any of the porous panels does not become normal, do the porous panel purge and test procedure for the applicable porous panel(s). Refer to TKS Leading Edge Porous Panel - Adjustment/Test. (36) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch to the OFF position. (37) Put the ANTI-ICE-FLUID CONTROL, BACKUP switch to the OFF position. (38) Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch to the WINDSHIELD position, then release the switch. NOTE:
(a)
The windshield pump will start and operate when you put the spring-loaded MAX FLOW switch to the WINDSHIELD position.
Do three full cycles of approximately 4 seconds each to let the TKS ßuid ßow from the windshield spray bar. NOTE:
If necessary, you can do more cycles until the ßuid ßow is normal from the spray bar.
(39) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. (40) Remove external electrical power from the airplane. (41) Make sure that the ßuid tank servicing is correct. Refer to Chapter 12, TKS Anti-Ice System Servicing. (42) Make sure that the ßoor and the airplane surfaces are clean. (43) Install the aft fairing. Refer to Remove the Aft Fairing in this section.
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23.
TKS Fluid Contamination (Water) Removal
WARNING: For health and environmental data, refer to the applicable Material Safety Data Sheet (MSDS). WARNING: If TKS ßuid is spilled, immediately remove (clean) or contain all the TKS ßuid. TKS ßuid on the ßoor causes a dangerous condition. WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS ßuid off the ßoor which helps prevent injury to personnel. WARNING: TKS ßuid is a hazardous material. You must discard all unwanted TKS ßuid and/or dirty cloths. refer to approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are usually 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 pounds for each gallon. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. A.
Remove the TKS Fluid Contamination (Water) (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Drain the TKS ßuid from the ßuid tank . Refer to TKS Fluid Removal in this section. (4) Do the ßuid tank servicing. Refer to Chapter 12, TKS Anti-Ice System - Servicing. (5) Supply external electrical power to the airplane. (6) Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. (7) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (8) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel to the HIGH position to start pump 1. (9) Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch (SI023) on the left switch panel to the AIRFRAME position, then release the switch. (a) Let the TKS ßuid ßow from the porous panels on the wing, wing strut, and horizontal and vertical stabilizer leading edges for one full cycle of approximately 2 minutes. (10) If ßuid ßow at any of the porous panels is not normal, do the porous panel purge and test procedure for the applicable porous panel(s). Refer to TKS Leading Edge Porous Panel Adjustment/Test. (11) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch to the OFF position. (12) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. (13) Remove external electrical power from the airplane. (14) Make sure that the ßuid tank servicing is correct. Refer to Chapter 12, TKS Anti-Ice System Servicing. (15) Make sure that the ßoor and the airplane surfaces are clean.
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24.
TKS Fluid Contamination (Solids) Removal A.
TKS Fluid Contamination (Solids) Removal (1) If the TKS ßuid contamination is a solid material, contact Cessna Customer Service, P.O. Box 7706, Wichita, Kansas 67209 USA Tele: 316-517-5800 Fax: 316-517-7271.
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE SYSTEM - ADJUSTMENT/TEST FAIRING INSTALLATION 1.
General A.
This section includes the test procedures that are necessary to do after a TKS system component replacement.
B.
For the purge and test procedures of the leading edge porous panels, refer to TKS Anti-Ice Leading Edge Porous Panel - Adjustment/Test.
C.
For the removal and installation and test procedures for the tail bracket assembly (low pressure switches), refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices.
D.
To calibrate the fluid level sender, refer to Fluid Level Sender Calibration in this section.
E.
Recommended maintenance to keep the TKS fluid at its correct viscosity is as follows: • Operate the pumps monthly, or as necessary, in the HIGH mode until the air is removed from the fluid system. • Keep the TKS system operational at all times to keep air pockets out of the system. • If the fluid tank is removed and installed or replaced, do the porous panel purge and test procedures. NOTE:
2.
If the fluid is too thick, the porous panels can become blocked or clogged.
Tools and Equipment A.
For a list of tools and equipment, refer to Ice and Rain Protection - General.
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3.
TKS Anti-Ice System Test
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you operate the TKS system, put plastic sheets or absorbent cloths below the porous panels. This keeps the TKS fluid off the floor which helps prevent injury to personnel. WARNING: Slowly loosen the coupling that is connected to the component of the TKS system before you remove components. It is possible that the system continues to have pressure. WARNING: Immediately remove (clean) or contain all the TKS fluid that is spilled. TKS fluid on the floor will cause a slip hazard. WARNING: Discard all unwanted TKS fluid and/or dirty cloths correctly. TKS fluid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS fluids. Approved fluids, in accordance with specification DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, filtered fluid in the TKS system. Contamination will cause fluid blockage and/or damage to the porous panel. NOTE:
For the tests that follow, you can disconnect the discharge tube from the filter manifold outlet and connect a drain tube, which will let you contain the fluid more easily, and that is how these procedures are written. Or, you can keep the discharge connected to the filter manifold outlet and use clean, dry cloths to absorb the anti-ice fluid and to clean the airplane surfaces and floor as necessary. Or, you can fabricate a fluid collector system and install it on and below the porous panels, which will contain the fluid and keep it off the floor. Recommended materials you can use are plastic sheets, tubing, aluminum tape, and rigid aluminum and/or plastic gutter material.
NOTE:
Although you can do one or more of the tests that follow, if applicable, it is necessary to do all of the tests after you have installed the fluid tank, timer box, and/or wire bundle.
NOTE:
It is easier for two persons to do these tests. One to monitor the cockpit and one to monitor the equipment pack.
A.
Prepare To Do the TKS Anti-Ice System Test, (Refer to TKS Anti-Ice System - Maintenance Practices, Figure 201). (1) Remove external electrical power from the airplane. (2) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (3) Remove the aft fairing. Refer to TKS Fluid Tank - Maintenance Practices, Remove the Aft Fairing. (4) Find the drain tube connected to the shut off valve, extending aft of the accessory bracket. (a) Put a bucket with a capacity of approximately 3 to 5 gallons below the drain tube outlet. (5) If you will do tests of the metering pumps, the high pressure switch, or timer box, disconnect the discharge tube from the filter manifold outlet.
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Put a cap or plug in the open tube end. Connect a length of tubing to the filter manifold outlet. (a) Put the open tube end in the bucket. (8) If you are to do the windshield pump test, disconnect the windshield pump outlet tube from the fuselage connector. (a) Put the open tube end in the bucket. (9) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (10) Do the test procedures if applicable. B.
Do a Test of Metering Pump 1. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure. (2) For the correct CAS message that shows for the applicable TKS system switch position refer to, TKS System - Description and Operation, Table 1. (3) Supply external electrical power to the airplane. (4) Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. (5) Put the AVIONICS 1 switch to the ON position. (6) Put the AVIONICS 2 switch to the ON position. (7) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel in the HIGH position to start pump 1. (a) Make sure that pump 1 operation starts. (b) Make sure that 1000 ml, +100 or - 100 ml discharges in the bucket in one minute. (8) Disengage the PRIMARY ANTI-ICE circuit breaker on the left circuit breaker panel. (a) Make sure that pump 1 operation stops. (9) Make sure that there is no fluid leakage from the couplings. (10) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch in the OFF position. (11) Engage the PRIMARY ANTI-ICE circuit breaker on the left circuit breaker panel. (12) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. (13) Do the Return to Service procedures or continue the applicable test(s).
C.
Do a Test of Metering Pump 2. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure. (2) For the correct CAS message that shows for the applicable TKS system switch position refer to, TKS System - Description and Operation, Table 1. (3) Supply external electrical power to the airplane. (4) Put the EXTERNAL POWER switch (SC006) on the pilot's switch panel in the BUS position. (5) Put the AVIONICS 1 switch to the ON position. (6) Put the AVIONICS 2 switch to the ON position. (7) Put the ANTI-ICE-FLUID CONTROL, BACKUP switch (SI024) on the left switch panel in the ON position to start pump 2. (a) Make sure that pump 2 operation starts. (b) Make sure that 1000 ml, +100 or - 100 ml discharges in the bucket in one minute. (8) Disengage the BACKUP ANTI-ICE circuit breaker on the left circuit breaker panel. (a) Make sure that pump 2 operation stops. (9) Make sure that there is no fluid leakage from the couplings. (10) Put the BACKUP switch in the OFF position. (11) Engage the BACKUP ANTI-ICE circuit breaker. (12) Put the EXTERNAL POWER switch on the pilot's switch panel in the OFF position. (13) Do the Return to Service procedures or continue the applicable test(s).
D.
Do a Test of the Windshield Pump. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure.
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For the correct CAS message that shows for the applicable TKS system switch position refer to, TKS System - Description and Operation, Table 1. Supply external electrical power to the airplane. Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. Put the AVIONICS 1 switch to the ON position. Put the AVIONICS 2 switch to the ON position. Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch (SI023) on the left switch panel in the WINDSHIELD position, then release the switch. NOTE:
The windshield pump will start when you put the spring-loaded MAX FLOW switch in the WINDSHIELD position and it will spray fluid on the windshield for four seconds after you release it.
Make sure that the windshield pump starts. When the pump stops make sure that a minimum of 25 ml was discharged in the collection bucket. Disengage the W/S ANTI-ICE circuit breaker on the left circuit breaker panel. Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch on the left switch panel in the WINDSHIELD position, then release the switch. (a) Make sure that the windshield pump does not operate. Make sure that there is no fluid leakage from the couplings. Engage the W/S ANTI-ICE circuit breaker on the left circuit breaker panel. Remove external electrical power from the airplane. Do the Return to Service procedures or continue the applicable test(s). (a) (b)
(8) (9) (10) (11) (12) (13) E.
Do a Test of the Fluid Level Sender. (1) Supply external electrical power to the airplane. (2) Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. (3) Put the AVIONICS 1 switch to the ON position. (4) Put the AVIONICS 2 switch to the ON position. (5) Drain the tank. Refer to TKS Fluid Tank - Maintenance Practices, TKS Tank Fluid Removal. (a) Make sure that the fluid quantity indication on the MFD is, A-ICE GAL 0.0. NOTE: (6) (7) (8)
F.
You must calibrate the level sender if it does not read zero when the fluid tank is empty.
Fill the tank. (a) Make sure that the fluid quantity indication on the MFD is, A-ICE GAL 19.0. To calibrate the fluid level sender, if necessary, refer to Fluid Level Sender Calibration in this section. Do the Return to Service in this section, or continue the applicable test(s).
Do a Test of the Low Level Switch. (1) Supply external electrical power to the airplane. (2) Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. (3) Put the AVIONICS 1 switch to the ON position. (4) Put the AVIONICS 2 switch to the ON position. (5) Drain the tank. Refer to TKS Fluid Tank - Maintenance Practices, TKS Tank Fluid Removal. (6) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch on the left switch panel in the NORM position. (7) Make sure that the A-ICE NORM (white) CAS message shows on the EICAS display. (8) Make sure that the A-ICE LOW FLUID (amber) CAS message shows on the EICAS display. (9) Put the PRIMARY switch in the OFF position (10) Add 4.0 gallons to the tank. (11) Put the PRIMARY switch to the NORM position.
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MODEL 208 MAINTENANCE MANUAL (12) Make sure that the A-ICE LOW FLUID (amber) CAS message does not show on the EICAS display. (13) Put the PRIMARY switch to the OFF position. (14) Service the TKS system. Refer to Chapter 12, TKS Anti-Ice System - Servicing for the servicing procedures. NOTE:
You must calibrate the fluid level sender if the primary flight display (G1000) does not read zero when the TKS fluid tank is empty. Refer to TKS Anti-Ice System Adjustment/Test, TKS Level Sender Calibration.
(15) Do the, Return to Service in this section, or continue the applicable test(s). G.
Do a Test of the Pressure Switch (High). (1) Make sure that the aft fairing is removed. Refer to TKS Anti-Ice System - Maintenance Practices (Fairing Installation), Remove the Aft Fairing . (2) Supply external electrical power to the airplane. (3) Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. (4) Put the AVIONICS 1 switch to the ON position. (5) Put the AVIONICS 2 switch to the ON position. (6) Connect a pressure gage and shutoff valve to the filter outlet tube. (a) Close the shutoff valve. (7) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel in the HIGH position for intervals of 10 to 15 seconds. (8) Monitor the pressure gage for a 150 psi indication after a short time period. (a) Make sure that the A-ICE HIGH PRESS (amber) CAS message shows on the EICAS display. NOTE: (9) (10) (11) (12) (13)
H.
This makes sure that the high pressure switch is operating correctly.
Put the PRIMARY switch in the OFF position. Put the EXTERNAL POWER switch on the circuit beaker switch panel in the OFF position. Slowly open the shutoff valve to release pressure in the system. Disconnect the pressure gage and shutoff valve from the tube. Do the, Return to Service in this section, or continue the applicable test(s).
Do a Test of the Timer Box. (1) Make sure that there is enough fluid in the tank to keep the pump from running dry during the test procedure. (2) For the correct CAS message that shows for the applicable TKS system switch position refer to, TKS System - Description and Operation, Table 1. (3) Supply external electrical power to the airplane. (4) Put the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel in the BUS position. (5) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel in the NORM position. (a) Make sure that each of the two pumps operate for 20 seconds, +3 or - 3 seconds and then do not run for 100 seconds, + 10 or - 10 seconds. (6) Make sure that there is no fluid leakage from the couplings. (7) While the pumps are off, Put the ANTI-ICE-FLUID CONTROL, MAX FLOW switch in the AIRFRAME position. (a) Make sure that each of the two pumps operate for 120 seconds, +10 or - 20 seconds. (8) Put the PRIMARY switch in the OFF position. (9) Do a test of the windshield pump. Refer to Do a Test of the Windshield Pump in this section. (10) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. (11) Do the Return to Service procedures or continue the applicable test(s).
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Do the Return to Service. NOTE: (1) (2) (3) (4) (5) (6) (7) (8) (9)
4.
After you have completed the applicable test(s), it is necessary to put the airplane back to its initial configuration.
Make sure that there is no fluid leakage from the couplings. Make sure that all applicable connectors, fasteners, and couplings are installed correctly. Make sure that the drain valve is closed and safety with wire. Refer to Chapter 20, Safetying Maintenance Practices. Install the screws and connect the antenna coaxial cable and covers, if necessary. Make sure that all the circuit breakers are engaged. Make sure that all the system switches are in their initial positions. Make sure that the external electrical power is removed from the airplane. Make sure that the aft fairing is installed. Refer to TKS Anti-Ice System - Maintenance Practices (Fairing Installation), Install the Aft Fairing . Make sure that the floor and the airplane surfaces are clean.
TKS Level Sender Calibration
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Immediately remove (clean) or contain all the TKS fluid that is spilled. TKS fluid on the floor will cause a slip hazard. WARNING: Discard all unwanted TKS fluid and/or dirty cloths correctly. TKS fluid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS fluids. Approved fluids, in accordance with specification DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal CAUTION: Use only clean, filtered fluid in the TKS system. Contamination will cause fluid blockage and/or damage to the porous panel. NOTE:
To make sure that the level sender is calibrated correctly, you can do this calibration procedure. The voltmeter will show 0.0 VDC when the fluid tank is empty and 5.0 VDC, +0.1 or -0.1 VDC when the fluid tank is full.
NOTE:
When the fluid tank is empty and the EICAS display shows 0.0 gallons, calibration of the empty adjustment is not necessary. You can then fill the fluid tank and calibrate the level sender full adjustment.
NOTE:
Changes in the properties of the anti-ice fluid can occur because of differences between manufacturers, or if the fluid is new (fresh), or if the fluid has gone through the fluid tank and TKS system, or if the fluid has been in the fluid tank too long (the TKS system has not been operated). These fluid changes can cause different results in calibration.
A.
Calibrate the Level Sender (Refer to Figure 501). (1) Make sure that the airplane is level. Refer to Chapter 8, Leveling - Maintenance Practices. (2) Remove the cockpit floor covering to access floor panel 232BC. (3) Remove the cockpit floor access panel 232BC. Refer to FLOOR COVERING/CONTROL COLUMN COVER - Maintenance Practices.
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Level Sender Calibration Figure 501 (Sheet 1)
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WARNING: Do not remove hoses under pressure. This procedure will result in release of refrigerant into the atmosphere. Removing hoses under pressure may also result in personal injury if hose ends are not restrained. (4) (5) (6) (7) (8) (9)
If necessary, disconnect and move the air conditioning lines to get access to the level sender. Refer to Chapter 21, R134A Air Conditioning - Maintenance Practices, Air Conditioning Plumping Removal/Installation. Move aside the rubber nipples that cover the level sender posts. Connect one lead of the voltmeter to the SEND post of the level sender. Connect the other lead of the voltmeter to the NEG post of the level sender. Supply external electrical power to the airplane. To calibrate the level sender with an empty fluid tank, do the steps that follow: (a) Remove the protective layer from the EMPTY adjustment screw. NOTE:
This screw is on the right side (airplane's right side) of the level sender.
Use a screwdriver to turn the EMPTY adjustment screw counter clockwise until the voltage that shows on the voltmeter does not go lower. (c) Turn the screw clockwise until 0.0 VDC, +0.1 or -0.1 VDC, shows on the voltmeter. (10) To calibrate the level sender with a full fluid tank, do the steps that follow: (a) Remove the protective layer from the FULL adjustment screw. (b)
NOTE:
This screw is on the right side (airplane's right side) of the level sender.
(b)
(11) (12) (13) (14) (15) (16) (17) (18)
Use a screwdriver to turn the FULL adjustment screw until 5.0 VDC, +0.1 or -0.1 VDC, shows on the voltmeter. Remove external electrical power from the airplane. Disconnect the leads of the voltmeter from the level sender. Cover the level sender posts with the rubber nipples. Put a protective layer on the adjustment screw(s). Install the cockpit floor access panel 232BC. Refer to FLOOR COVERING/CONTROL COLUMN COVER - Maintenance Practices. If necessary, connect the air conditioning lines. Refer to Chapter 21, R134A Air Conditioning Maintenance Practices, Air Conditioning Plumping Removal/Installation. Install the cockpit floor covering to access floor panel 232BC. Clean the floor and the airplane surfaces as necessary.
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE LEADING EDGE POROUS PANEL - MAINTENANCE PRACTICES 1.
General A.
This section contains the removal and installation procedures for the TKS anti-ice porous panels, which include the wing, wing strut, and horizontal and vertical stabilizer leading edges. The sealing procedures for the porous panels are also included in this section. The procedures apply to the cargo pod and the fairing TKS system installation.
B.
After a porous panel is replaced, it is necessary to do the purge and test procedures. Those procedures are in TKS Anti-Ice Leading Edge Porous Panel - Adjustment/Test.
C.
Recommended maintenance to keep the TKS ßuid at its correct viscosity is as follows: • Operate the pumps monthly, or as necessary, in the HIGH mode until the air is removed from the ßuid system. • Keep the TKS system operational at all times to keep air pockets out of the system. • If the ßuid tank is removed and installed or replaced, do the porous panel purge and test procedures. NOTE:
2.
If the ßuid is too thick, the porous panels can become blocked or clogged.
Tools and Equipment A.
For a list of tools and equipment, refer to Ice and Rain Protection - General.
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. 3.
TKS Porous Panel Removal/Installation A.
Remove the Porous Panel (Refer to Figure 201 and Figure 202).
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WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that high pressure is still in the system. CAUTION: Do not use MEK, acetone, paint thinner, or similar chlorinated solvents on the porous panels. To prevent damage, only use water and detergent, and/or isopropyl alcohol, AV gas, industrial methylated spirit, and approved anti-ice ßuid on the porous panel surfaces. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal . CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 203. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. CAUTION: After you remove and before you install a porous panel, apply low-adhesive tape on the panel to give it protection. CAUTION: Before you remove or install a porous panel, apply low-adhesive tape on the skin adjacent to the panel to give the skin protection. CAUTION: Be careful when you remove and install the porous panels. The panels are easily damaged. Use nonmetallic tools, if possible, to prevent tears, gouges, scratches, and other damage. NOTE:
The panel purge and test procedures are only necessary after you install a replacement porous panel.
NOTE:
The removal and installation of the porous panels are typical.
(1) (2)
(3)
Remove external electrical power from the airplane. Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. To remove the wing porous panels, remove wing access panels 501AB and 501DB left inboard, 503CB left center, 503GB and 503JB left outboard, or 601AB and 601DB right inboard, 603CB right center, or 603GB and 603JB right outboard as applicable. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation.
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TKS Anti-Ice System Flow Diagram Figure 201 (Sheet 1)
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TKS Porous Panel Installation Figure 202 (Sheet 1)
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TKS Porous Panel Installation Figure 202 (Sheet 2)
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TKS Porous Panel Installation Figure 202 (Sheet 3)
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TKS Nylon Tubing Assembly Figure 203 (Sheet 1)
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TKS Nylon Tubing Assembly Figure 203 (Sheet 2)
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CAUTION: Before you remove or install a porous panel, apply low-adhesive tape on the skin adjacent to the panel to give the skin protection. (4) (5)
To remove the stabilizer porous panels, remove tailcone access panel 373BL, 374BR, and 341C . Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. Apply tape on the leading edge next to the sealant along the panel edges. NOTE:
(6) (7)
Carefully remove the sealant along the panel edges. Remove the rivets and/or screws, as applicable, that attach the panel to the leading edge. NOTE:
(8)
The wing and wing strut panels have only rivets installed. The vertical stabilizer panels has rivets and screws installed.
Disconnect the tubing from the coupling if it is accessible. NOTE:
(9)
The tape will give a guide to help apply new sealant when the new panel is installed.
On the vertical stabilizer panel and strut panel the panel must be removed to get access to the tubing.
Install caps on all tube ends to keep FOD out of the TKS system.
CAUTION: While pulling the panel away from the leading edge, be careful not to damage the leading edge by prying against it. (10) Carefully use a ßexible-blade knife (putty knife) and your hands to pull the panel away from the leading edge. (11) Carefully remove the panel from the leading edge. (12) Remove all remaining sealant from the leading edge. B.
Install the Porous Panel (Refer to Figure 201 and Figure 202). (1) Put the porous panel in it correct position on the leading edge. (2) Apply masking tape or equivalent) on the airplane skin around the perimeter of the porous panels. NOTE: (3) (4) (5)
This will allow you to Þllet seal around the edges of the porous panels.
Drill the rivet and screw holes to match the existing hole locations on the leading edge. Remove the caps from the tube ends. Install new seals in the tubing ends as shown in Figure 203.
CAUTION: Before you install or remove a porous panel, apply low-adhesive tape on the panel to give it protection. (6)
Do the porous panel sealant procedures. Refer to TKS Porous Panel Sealant Procedure in this section. (7) Hold the panel in position near the leading edge and connect the tubing. (a) Tighten the coupling with your Þngers. (b) Continue to tighten the coupling with a wrench approximately 180 degrees more. (8) Align the panel and install temporary fasteners to hold the panel in position. (9) Push the panel against the leading edge with enough pressure to cause the sealant to squeeze out along the edges. (10) Keep applying pressure until the tape strips or ratcheting straps are applied to hold the panel in position. (a) Make sure that the panel edges are against the leading edge skin.
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CAUTION: To spread the load where ratcheting straps are used, use metal angles to protect the trailing edge. Place blankets or foam between the metal angles and the trailing edge. Do not use too much force when you tighten the straps. (11) Use tape strips or racheting straps, as necessary, to hold the panel tightly against the leading edge. NOTE:
Since more sealant is used on the inboard wing panels it is necessary to use racheting straps for these panels around the wing to maintain pressure on the panel while the sealant cures.
(12) Install the rivets and/or screws, as applicable, that attach the panel to the leading edge. NOTE:
The wing and wing strut panels only use rivets. The stabilizer panels have rivets and screws installed.
(a) (b) (c)
Examine the rivet heads and/or screw heads for correct installation. Use Type X, Class B sealant to apply a shank seal to the rivets. On the vertical stabilizer panel only, torque the screws to 12 to 15 inch-pounds (1.35 to 1.69 N-m). (13) Use Type I, Class B sealant to Þllet seal around the edges of the porous panel. Refer to Chapter 20, Fuel, Weather, and High Temperature Sealing. (a) Apply the sealant in the 0.12 inch area between the tape and the porous panel. Refer to Figure 202. NOTE:
Make sure that the sealant touches the edge of the porous panel at all positions.
(14) After the sealant is cured, remove all tape and racheting straps. (15) Electrically bond all porous panels. Refer to Chapter 20, Electrical Bonding- Maintenance Practices. (16) Do the panel purge and test procedures. Refer to TKS Anti-Ice Leading Edge Porous Panel Adjustment/Test. (17) Do a leak check follows: (a) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (b) Supply external electrical power to the airplane. (c) Put the EXTERNAL POWER switch (S17) on the circuit breaker switch panel to the BUS position. (d) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch (SI022) on the left switch panel to the HIGH position. Make sure that there is no ßuid leakage from the couplings. 1 (e) Put the ANTI-ICE-FLUID CONTROL, PRIMARY switch on the left switch panel to the OFF position. (f) Put the EXTERNAL POWER switch on the circuit breaker switch panel to the OFF position. (18) Clean the ßoor and the airplane surfaces as necessary. (19) Install wing access panels 501AB and 501DB left inboard, 503CB left center, 503GB and 503JB left outboard, or 601AB and 601DB right inboard, 603CB right center, or 603GB and 603JB right outboard as applicable. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (20) Install tailcone access panels 373BL, 374BR, and 341C, if applicable. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (21) Remove external electrical power from the airplane.
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4.
TKS Porous Panel Sealant Procedure
CAUTION: Do not use MEK, acetone, lacquer thinner, or similar chlorinated solvents on the porous panels. To prevent damage, only use water and detergent, and/or isopropyl alcohol, AV gas, industrial methylated spirit, AV turbine ßuid, ethyl alcohol, and approved ice protection ßuid on the porous panel surfaces. A.
Apply the Sealant (Refer to Figure 202). (1) Lightly abrade the airplane leading edge with a ScotchBrite pad. (2) Use isopropyl alcohol to scrub the area of the leading edge with a sponge or short-bristle brush where the panel will be installed. (a) Use a clean lint-free cloth to dry the area before the cleaning solution evaporates. (3) Use a clean lint-free cloth that is wet with isopropyl alcohol to clean the leading edge. (a) Use a clean lint-free cloth to dry the area before the cleaning solution evaporates. (4) Use a clean lint-free cloth that is wet with isopropyl alcohol to clean the aft surface of the panel. (a) Use a clean lint-free cloth to remove debis and contaminants from the surface of the panel. (5) Use a clean lint-free cloth that is wet with isopropyl alcohol to clean the surface again. NOTE:
Wipe in approximately 12 to 15 inch sections.
(a)
(6)
Use a clean lint-free cloth to dry the area immediately before the isopropyl alcohol evaporates. Wipe the surface of the panel with a clean lint-free cloth to make sure that there is no debis and contaminants. NOTE:
(7) (8)
If the cloth shows signs of debris or contaminants clean the area again.
On all panels except the inboard wing panel, Use Type X, Class B sealant to seal the void on the upper and lower backshell joggle, down the center (0.50-inch (12.7 mm) fay seal), and around the feed inlet and air bleed valve. (Refer to Figure 202). On the inboard wing panels, apply Type X, Class B sealant to all of the back surface of the panel, and to the feed input and bleed valve. NOTE:
On the inboard wing panels approximately 65 ounces (1.92 l) of sealant is needed to make sure there are no air pockets between the panel and the wing leading edge.
(9)
Use Type X, Class B sealant to shank seal around each rivet and/or screw hole in the leading edge skin. (10) Install the porous panel on the leading edge. Refer to TKS Porous Panel Removal/Installation in this section.
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE LEADING EDGE POROUS PANEL - ADJUSTMENT/TEST 1.
General A.
This section contains the procedures to remove (purge) the air from the porous panels on the wing, wing strut, and horizontal and vertical stabilizer leading edges. The procedures apply to the cargo pod and the fairing TKS system installation.
B.
The panel purge and test procedures are only necessary after you install a replacement porous panel, or if you remove and install or replace the ßuid tank.
C.
The function of the panel purge and test procedures is to remove most of the air and make the membrane in the panel completely wet and to make sure that there is no leakage from the panel and its connections to the tubing.
D.
Recommended maintenance to keep the TKS ßuid at its correct viscosity is as follows: • Operate the pumps monthly, or as necessary, in the HIGH mode until the air is removed from the ßuid system. • Keep the TKS system operational at all times to keep air pockets out of the system. • If the ßuid tank is removed and installed or replaced, do the porous panel purge and test procedures. NOTE:
2.
If the ßuid is too thick, the porous panels can become blocked or clogged.
Tools and Equipment A.
For a list of tools and equipment, refer to Ice and Rain Protection - General.
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3.
Porous Panel Purge and Test
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that high pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 501. This will help to prevent ßuid leakage from the coupling. Refer to TKS Anti-Ice Fluid Distribution System - Maintenance Practices for Nylon Tubing Repair/Replacement. NOTE:
It is necessary that you have access to clean dry cloths, 30 gallons of approved TKS ßuid, a TKS system test cart with connection hardware, 75 psi Þltered shop air (to use with a test cart), and a container with a capacity of three to Þve gallons.
NOTE:
You can fabricate a ßuid collector system, which will contain the ßuid and keep it off the ßoor. Recommended materials you can use are plastic sheets, tubing, aluminum tape, and rigid aluminum and/or plastic gutter material.
A.
Do the Panel Purge and Test (Refer to Figure 501 and Figure 502). (1) Remove external electrical power from the airplane. (2) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (3) On airplanes that have the G1000 system installed, disengage the ENG INTFC circuit breaker on the avionics circuit breaker panel. (4) On airplanes that do not have the G1000 system installed, disengage the ANTI-ICE GAGE circuit breaker on the left circuit breaker panel. (5) To get access to the tubing for the vertical panel at the proportioning unit in the tail bracket assembly, remove tailcone access panel 320A . Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation.
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TKS Nylon Tubing Assembly Figure 501 (Sheet 1)
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TKS Nylon Tubing Assembly Figure 501 (Sheet 2)
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Leading Edge Porous Panel Purge Figure 502 (Sheet 1)
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Leading Edge Porous Panel Purge Figure 502 (Sheet 2)
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Leading Edge Porous Panel Purge Figure 502 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (6) (7)
To get to the horizontal panel Þttings, remove the access panels 373BL and 374BR Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation as applicable. To get access to the wing panel Þttings, remove wing access panels 501BB, 501EB, 503AB, 503DB, and 503HB left, or 601BB, 601EB, 603AB, 603DB, and 603HB right as applicable. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. NOTE:
(8) (9)
On airplanes that have a radome installed on the right wing, you can only remove wing access panel 603BB.
Attach the ßuid collector system below the area of the panel that will be purged. Disconnect the ßuid inlet tube from the panel Þtting. NOTE:
The inboard and outboard wing panels have two supply tubes. Install a tee Þtting and connect the test cart to both Þttings.
NOTE:
The strut panel has two membranes (upper and lower). Each membrane has a ßuid supply tube. You can install a tee Þtting and connect the test cart supply tube to each membrane.
(10) Connect the test cart ßuid supply tube to the ßuid inlet Þttings. NOTE:
On the vertical stab panel and strut panel the test cart must be connected to the tubing, since you cannot get access to the Þttings. Label the strut tubing to make sure that it is connected to the correct Þttings when the test is complete.
(11) Follow the operation and safety instructions that are supplied with the test cart. (a) Use the Panels Installed on Aircraft TKS Panel Test section of the TKS Ice Protection Panel Flow Check Procedure Using TKS System Test Cart publication. (12) Slowly start the ßuid ßow through the panel at 10 psi. (a) Correct any leaks as needed. (13) Set the ßuid pressure to 20 psi. (14) When ßuid starts to come through panel pores along the entire length of the panel set the pressure as appropriate to each panel. Refer to Table 1 (15) Increase the pressure at the test cart outlet as follows: Table 501. Panel Purge Pressure Table PANEL PURGE PRESSURE TABLE Panel
Pressure (Maximum)
Inboard wing (Note 1)
60 psi (Note 2)
Middle wing
60 psi
Outboard wing (Note 1)
65 psi
Strut
75 psi
Horizontal Stabilizer
65 psi
NOTE 1: If the panels are purged individually, the opposite port must be capped. NOTE 2: During the annual purge, take the pressure to 60 psi. (16) Make sure that the ßuid ßows from the porous panels as follows: (a) The ßuid ßows from the porous panels without any dry spots. (b) The ßuid ßows evenly and not in streams.
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MODEL 208 MAINTENANCE MANUAL (c)
There are no areas with clusters of bubbles. NOTE:
(17) (18) (19) (20) (21) (22) (23) (24) (25) (26) (27)
(28) (29) (30)
The bubbles will be very small and continue to ßow even after you wipe ßuid across the area.
Stop the ßuid supply. Wait until the ßuid pressure is released. Slowly disconnect the test cart ßuid supply tube from the ßuid inlet Þtting. Remove the cap or plug from the tube end, if applicable. Install a new seal(s) in the coupling(s) as shown in Figure 501. Connect the panel supply tube(s) to the proportioning unit port. Remove the ßuid collector system from the airplane. Clean the ßoor and the airplane surfaces as necessary. Install wing access panels 503AB and 503BB left, or 603AB and 603BB right as applicable. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. Install tailcone access panel 320A. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. On airplanes that have the G1000 system installed, engage the ENG INTFC circuit breaker on the avionics circuit breaker panel. On airplanes that do not have the G1000 system installed, engage the ANTI-ICE GAGE circuit breaker on the left circuit breaker panel. Remove external electrical power from the airplane, if applicable.
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE FLUID DISTRIBUTION SYSTEM - MAINTENANCE PRACTICES 1.
General A.
This section contains the removal and installation procedures for the TKS anti-ice ßuid distribution system, which include the nylon tubing and connections and the ßuid proportioning units installed in the wings and fuselage. The tail bracket assembly includes the proportioning unit and low pressure switches for the horizontal and vertical stabilizers. The procedures in this section apply the cargo pod and the fairing installations of the TKS system.
B.
Recommended maintenance to keep the TKS ßuid at its correct viscosity is as follows: • Operate the pumps monthly, or as necessary, in the HIGH mode until the air is removed from the ßuid system. • Keep the TKS system operational at all times to keep air pockets out of the system. • If the ßuid tank is removed and installed or replaced, do the porous panel purge and test procedures. NOTE:
2.
If the ßuid is too thick, the porous panels can become blocked or clogged.
Tools and Equipment A.
For a list of tools and equipment, refer to Ice and Rain Protection - General.
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3.
TKS Propeller Proportioning Unit Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 204. This will help to prevent ßuid leakage from the coupling. NOTE:
The propeller proportioning unit is installed on the right side at FS179.36.
A.
Remove the Propeller Proportioning Unit (Refer to Figure 201, Figure 202, and Figure 203). (1) Remove external electrical power from the airplane. (2) Disengage the PRIMARY ANTI-ICE and BACKUP ANTI-ICE circuit breakers on the left circuit breaker panel. (3) To get access to the propeller proportioning unit, remove ßoorboard access panels 232DR and 252GR, as applicable. Refer to Chapter 6, Access Plates and Panels IdentiÞcation. (4) Slowly loosen and disconnect the inlet and outlet tubes from the proportioning unit. (5) Put caps on all tube ends to keep FOD out of the ßuid system. (6) Remove the proportioning unit from the airplane.
B.
Install the Propeller Proportioning Unit (Refer to Figure 201, Figure 202, and Figure 203). (1) Put the proportioning unit in position in the airplane. (a) Make sure that the arrow on the proportioning unit points in the direction of ßuid ßow. (2) Remove the caps from the tube ends. (3) Install new seals in the couplings as shown in Figure 204. (4) Connect and tighten the inlet and outlet tubes to the proportioning unit.
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TKS Anti-Ice System Flow Diagram Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
TKS Fluid Distribution Installation Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
TKS Fluid Distribution Installation Figure 202 (Sheet 2)
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TKS Proportioning Unit Installation Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
TKS Proportioning Unit Installation Figure 203 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7)
Engage the PRIMARY ANTI-ICE and BACKUP ANTI-ICE circuit breakers on the left circuit breaker panel. Supply external electrical power to the airplane. Put the EXTERNAL POWER switch on the circuit breaker switch panel in the ON position. NOTE:
(8) (9) (10) (11) (12) (13)
For airplanes that have G1000, the EXTERNAL POWER switch reference designator is (SC006) and for airplanes that do not have G1000, the reference designator is (S17).
Put the ANTI-ICE-FLUID FLOW, PRIMARY switch (SI022) on the left switch panel in the HIGH position. (a) Make sure that there is no ßuid leakage from the couplings. Put the ANTI-ICE-FLUID FLOW, PRIMARY switch on the left switch panel in the OFF position. Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. Clean the ßoor and the airplane surfaces as necessary. Install the ßoorboard access panels 232DR and 252GR, as applicable. Refer to Chapter 6, Access Plates and Panels IdentiÞcation. Remove external electrical power from the airplane.
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4.
TKS Wing Proportioning Unit Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 204. This will help to prevent ßuid leakage from the coupling. NOTE:
The wing proportioning units are installed at WS170.60
NOTE:
The removal and installation of the wing proportioning units are typical.
NOTE:
The wing proportioning units also supply ßuid to the wing struts.
A.
Remove the Wing Proportioning Unit (Refer to Figure 201, Figure 202, and Figure 203). (1) Remove external electrical power from the airplane. (2) Disengage the PRIMARY ANTI-ICE circuit breaker on the left circuit breaker panel. (3) Remove wing access panels 503AB and 503BB left, or 603AB and 603BB right. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. NOTE: (4) (5)
On airplanes that have a radome installed on the right wing, you can only remove wing access panel 603BB.
Remove the screws, washers, and spacers that attach the proportioning unit to the airplane structure. Identify, slowly loosen, and disconnect the inlet and outlet tubes from the proportioning unit.
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MODEL 208 MAINTENANCE MANUAL (6) (7)
(8) B.
Put caps on all tube ends to keep FOD out of the ßuid system. To Þnd the applicable proportioning unit port(s) and its related porous panel, refer to the table that follows: TKS POROUS PANEL
PROPORTIONING UNIT PORT
Left Wing Inboard Panel
Port 1 and Port 5
Left Wing Center Panel
Port 2
Left Wing Outboard Panel
Port 3 and Port 4
Left Wing Strut Panel
Port 6 and Port 7
Right Wing Inboard Panel
Port 1 and Port 5
Right Wing Center Panel
Port 2
Right Wing Outboard Panel
Port 3 and Port 4
Right Wing Strut Panel
Port 6 and Port 7
Remove the proportioning unit from the airplane.
Install the Wing Proportioning Unit (Refer to Figure 201, Figure 202, and Figure 203). (1) Put the proportioning unit in position in the airplane. (2) Remove the caps from the tube ends. (3) Install new seals in the couplings as shown in Figure 204. (4) Connect and tighten the inlet and outlet couplings to the proportioning unit. (5) Install the screws, washers, and spacers that attach the proportioning unit to the airplane structure. (6) Engage the PRIMARY ANTI-ICE circuit breaker on the left circuit breaker panel. (7) Supply external electrical power to the airplane. (8) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the ON position. NOTE:
(9) (10) (11) (12) (13) (14)
For airplanes that have G1000, the EXTERNAL POWER switch reference designator is (SC006) and for airplanes that do not have G1000, the reference designator is (S17).
Put the ANTI-ICE-FLUID-FLOW, PRIMARY switch on the left switch panel in the HIGH position. (a) Make sure that there is no ßuid leakage from the couplings. Put the ANTI-ICE-FLUID-FLOW, PRIMARY switch on the left switch panel in the OFF position. Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. Clean the ßoor and the airplane surfaces as necessary. Install wing access panels 503AB and 503BB left, or 603AB and 603BB right. Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. Remove external electrical power from the airplane.
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5.
TKS Stabilizer Proportioning Unit and Low Pressure Switch (Tail Bracket Assembly) Removal/ Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS ßuid that is spilled. TKS ßuid on the ßoor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS ßuid off the ßoor. This will help to prevent injury to personnel. WARNING: Discard all unwanted TKS ßuid and/or dirty cloths correctly. TKS ßuid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS ßuids. Approved ßuids, in accordance with speciÞcation DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal. CAUTION: Use only clean, Þltered ßuid in the TKS system. Contamination will cause ßuid blockage and/or damage to the porous panel. CAUTION: Do not use the seals again after you loosen or disconnect a tube coupling. Replace the 3/16-inch and 5/16-inch sealing ring and/or 1/2-inch O-ring, as applicable, when you assemble a tube coupling. Examine the seal for damage and make sure that it is in the correct position in the coupling as shown in Figure 204. This will help to prevent ßuid leakage from the coupling. NOTE: A.
The tail bracket assembly is installed at FS415.20 (208B), and FS367.20 (208).
Remove the Tail Bracket Assembly (Refer to Figure 201, Figure 202, and Figure 203). (1) Remove external electrical power from the airplane. (2) Disengage the PRIMARY ANTI-ICE and BACKUP ANTI-ICE circuit breakers on the left circuit breaker panel. (3) Remove the tailcone access panel 320A . Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (4) Disconnect the airplane electrical connector on the bracket assembly. (5) Remove the screws, washers, and nuts that attach the electrical connector to the bracket assembly. (6) Identify and slowly loosen and disconnect the inlet and outlet tubes from the tail bracket assembly and quickly put the tube ends in the container. (7) Remove the screws and washers that attach the tail bracket assembly to the airplane structure. (8) Remove the tail bracket assembly from the airplane.
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MODEL 208 MAINTENANCE MANUAL B.
Install the Tail Bracket Assembly (Refer to Figure 201, Figure 202, and Figure 203). (1) Put the tail bracket assembly in its correct position on the airplane. (2) Install the screws that attach the tail bracket assembly to the airplane. (3) Install new seals on the tubing ends. Refer to Figure 204. (4) Connect the inlet and outlet tubes to tail bracket and tighten. (5) Put the electrical connector in its correct position on the tail bracket assembly. (6) Install the screws, washers, and nuts that attach the electrical connector to the bracket assembly. (7) Connect the low pressure switches electrical connector to the electrical connector on the tail bracket assembly. (8) Engage the PRIMARY ANTI-ICE and BACKUP ANTI-ICE circuit breakers on the left circuit breaker panel. (9) Supply external electrical power to the airplane. (10) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the BUS position. NOTE:
For airplanes that G1000, the EXTERNAL POWER switch reference designator is (SC006) and for airplanes that do not have G1000, the reference designator is (S17).
(11) Put the ANTI-ICE FLUID FLOW BACKUP (SI022) switch on the left switch panel in the ON position. (12) On airplanes with the G1000 system installed, monitor the EICAS display. (a) Make sure that the applicable A-ICE LOW PRESS red CAS message comes on. (13) On airplanes that do not have the G1000 system installed, monitor the anti-ice annunciators. (a) Make sure that the red anti-ice WARN annunciator comes on. (14) Let the system operate for a minimum of one minute to purge air from the porous panels and start the ßow of anti-ice ßuid from the panels. (15) Make sure that there is not any ßuid leakage from the couplings. (16) Put the ANTI-ICE FLUID FLOW BACKUP (SI022) switch on the left switch panel in the OFF position. (17) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. (18) Remove external electrical power to the airplane. (19) Clean the ßoor and airplane surface as necessary. (20) Install the tailcone access panel 320A . Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. C.
Remove the Low Pressure Switch(es) (Refer to Figure 201. (1) Remove external electrical power from the airplane. (2) Disengage the PRIMARY ANTI-ICE and BACKUP ANTI-ICE circuit breakers on the left circuit breaker panel. (3) Remove the tailcone access panel 320A . Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation. (4) Disconnect the electrical connector from the pressure switch. (5) Slowly disconnect the inlet and outlet tubes from the pressure switch assembly and quickly put the inlet tube end in the container. (6) Remove the screws and spacers that attach the pressure switch to the bulkhead. (7) Remove the pressure switch from the airplane.
D.
Install the Low Pressure Switch(es) (Refer to Figure 201. (1) Put the switch in its correct position on the bulkhead. (2) Install the screws and spacers that attach the switch to the bulkhead. (3) Install new seals on the tubing ends. Refer to Figure 204. (4) Connect and tighten the inlet and outlet tubes to the pressure switch assembly. (5) Connect the electrical connector to the switch. (6) Engage the PRIMARY ANTI-ICE and BACKUP ANTI-ICE circuit breakers on the left circuit breaker panel. (7) Supply external electrical power to the airplane. (8) Put the EXTERNAL POWER switch on the circuit breaker switch panel in the BUS position. NOTE:
For airplanes that G1000, the EXTERNAL POWER switch reference designator is (SC006) and for airplanes that do not have G1000, the reference designator is (S17).
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MODEL 208 MAINTENANCE MANUAL (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) 6.
Put the ANTI-ICE FLUID FLOW BACKUP (SI022) switch on the left switch panel in the ON position. On airplanes with the G1000 system installed, monitor the EICAS display. (a) Make sure that the applicable A-ICE LOW PRESS red CAS message comes on. On airplanes that do not have the G1000 system installed, monitor the anti-ice annunciators. (a) Make sure that the red anti-ice WARN annunciator comes on. Let the system operate for a minimum of one minute to purge air from the porous panels and start the ßow of anti-ice ßuid from the panels. Make sure that there is not any ßuid leakage from the couplings. Put the ANTI-ICE FLUID FLOW BACKUP (SI022) switch on the left switch panel in the OFF position. Put the EXTERNAL POWER switch on the circuit breaker switch panel in the OFF position. Remove external electrical power to the airplane. Clean the ßoor and airplane surface as necessary. Install the tailcone access panel 320A . Refer to Chapter 6, Access Plates and Panels IdentiÞcation - Description and Operation.
Nylon Tubing Repair/Replacement A.
All plumbing used in the ßuid anti-ice system is ßexible nylon tubing connected with special compression-type couplings. Three different sizes of tubing are used.
B.
The couplings used to connect sections of nylon tubing are metallic compression-type couplings, which include a machined coupling end, an olive (ferrule), a nut, and an elastomeric sealing ring (Refer to Figure 204). The couplings used to connect the tubing to all the anti-ice porous panels and the cuff, fairing, and tail bracket low-pressure switches are stainless steel. All other couplings used to connect the ßuid anti-ice system components are made of aluminum. To help to prevent electrolytic corrosion, aluminum couplings must be assembled only to aluminum couplings and stainless steel couplings must be assembled only to stainless steel or titanium couplings. The compression-type couplings used in the ßuid anti-ice system are not interchangeable with standard-type couplings (AN or MS). To make sure that the correct couplings are used, refer to the Model 208 Illustrated Parts Catalog to Þnd the correct coupling.
C.
When you do maintenance of the ßuid distribution system, examine all tubing for kinks, cuts, abrasion, crushing, or other indications of damage. The nylon tubing can discolor to a light straw color with age. Discoloration of the tubing is normally not a cause for rejection. Any damaged or deteriorated tubing found must be replaced with tubing and couplings of the correct size and type (aluminum or stainless steel). Additionally, when the tube couplings are disconnected from components or other couplings, it is recommended to install a new seal in the coupling.
D.
Minimum bend radii for ßuid tubing at ambient temperature is shown in Table 202. In some areas of the airplane, tighter bend radii is necessary. In these areas, you can use a heat gun to bend the tubing as follows:
Table 201. Minimum Bend Radii for Fluid Tubing (At Ambient Temperature) TUBING OUTSIDE DIAMETER
MINIMUM BEND RADII
1/2 Inch
3.0 Inches
5/16 Inch
2.0 Inches
3/16 Inch
1.5 Inches
(1)
Put the tubing in position to Þnd the necessary bend radius. NOTE:
You can use a piece of soft wire as a pattern.
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CAUTION: Be careful not to overheat, burn, or collapse the tubing, which can prevent the correct ßow of ßuid and/or ßuid leakage. (2)
Hold the tubing and carefully apply heat to the area where the bend will be. NOTE:
(3) (4)
The tube will become ßexible at approximately 300°F.
After the radius is correct, hold the tube in that position and let it become cool (ambient temperature). When the tube is at ambient temperature, carefully examine it to make sure that the it has not been heated too much (burned) and that there is no restriction of ßuid ßow through the tubing.
CAUTION: Do not use the coupling to clench the olive to the ßuid tubing. Use only speciÞed clenching tools to do the clenching operation. Also, do not torque the couplings too much during the repair or replacement procedure. If the couplings leak, install new seals as necessary. E.
When tubing repair or replacement is necessary, the olive must always be clenched (swaged) to the tubing as a separate operation. Use approved clenching tools before you assemble the coupling. For clenching tools, refer to Ice and Rain Protection - General.
F.
When installing a new Þtting, refer to Table 202 for Tubing Dimension and Table 203 for Torque Requirements.
Table 202.
Tubing Dimension Aluminum Alloy / Stainless Steel Fittings on Nylon Tubes
Tube Outside Diameter (OD) in Inches
Olive Distance (Dimensions A) (+0.05 inch or - 0.05 inch)
Clenching Torque for Aluminum Fittings (+10% or -10%) (inch pounds)
Clenching Torque for Steel Fittings (+10% or -10%) (inch pounds)
3/16
0.22
70
90
5/16
0.22
120
200
1/2
0.38
250
N/A
Table 203. Torque Requirements Aluminum Alloy / Stainless Steel Fittings on Nylon Tubes Tube Outside Diameter (OD) in Inches
Tightening Torque (Reference) (+10% or -10%) (inch pounds)
3/16
28
5/16
48
1/2
63
G.
The recommended procedure for replacement of damaged or deteriorated ßuid tubing is to replace the full length of tubing from coupling to coupling. However, in areas of the airplane where this type of replacement is not possible, it is permitted to do a tube repair as an alternative to complete tubing replacement. (1) To repair a damaged section of tubing, cut out the damaged area and replace it with a new section of tubing. Connect the new tubing to the other tube ends with straight couplings.
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MODEL 208 MAINTENANCE MANUAL (2)
All new connections used in the repair must be tested for leaks. Operate the system before you install the access panels. NOTE:
If bulk tubing is installed, it is necessary to attach an identiÞcation tag to each end of the tube with the tube part number labeled on it for tube identiÞcation.
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TKS Nylon Tubing Assembly Figure 204 (Sheet 1)
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TKS Nylon Tubing Assembly Figure 204 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL PITOT AND STATIC HEATERS - DESCRIPTION AND OPERATION 1.
General A.
For information on pitot and static heaters, refer to Chapter 34, Pitot/Static System - Description and Operation.
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MODEL 208 MAINTENANCE MANUAL STALL WARNING HEATER - DESCRIPTION AND OPERATION 1.
General A.
For information on the stall warning heater, refer to Chapter 27, Stall Warning System - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL WINDSHIELD ANTI-ICE - DESCRIPTION AND OPERATION 1.
General A.
2.
An optional heated anti-ice panel is provided to prevent the formation of ice on the windshield. The system consists of a removable electrically heated glass panel mounted to the base of the windshield in front of the pilot.
Description and Operation A.
The sides of the electrically heated glass panel are supported by rubber covered frames that hold the panel off the windshield. The lower mounting bracket is hinged and hinges are spring-loaded so panel may be easily removed for storage. The system is controlled by a three-position toggle switch located on the left switch control panel. When the switch is placed in the AUTO position, electrical current flows to the anti-ice panel and is allowed to heat to a maximum of 136°F. The controller then opens the relay in the circuit which cuts off current to the controller and allows the anti-ice panel to cool down to a minimum of 129°F. The controller will then close the relay which starts the heating cycle over again. Therefore, when the switch is in the AUTO position, the anti-ice panel cycles on and off to assist in preventing ice formation on the windshield panel. In the event of a malfunction in the system control circuitry, the switch can be held in the momentary MANUAL position to achieve windshield anti-icing. A light in the annunciator panel, illuminates to indicate the system is operating. The system is protected by two pull-off type circuit breakers, a control circuit breaker, labeled W/S ANTI-ICE CONT and a heater circuit breaker, labeled W/S ANTI-ICE. Circuit breakers are installed in the left sidewall switch and circuit breaker panel.
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MODEL 208 MAINTENANCE MANUAL WINDSHIELD ANTI-ICE - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure 101.
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Windshield Anti-Ice Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL WINDSHIELD ANTI-ICE - MAINTENANCE PRACTICES 1.
General A.
2.
Windshield anti-ice maintenance practices include component removal/installation.
Windshield Anti-Ice Panel and Attach Bracket Removal/Installation NOTE:
A.
Airplanes 20800381 and On and 208B1088 and On, and those incorporating SK208-146, have a larger anti-ice panel installed. The larger anti-ice panel has an additional attach rod, and a different electrical connector.
Remove Windshield Anti-Ice Panel and Attach Bracket (Refer to Figure 201). (1) Make sure electrical power is off. (2) Disconnect the connector from the receptacle that is found at the center lower windshield, then install the cover over the receptacle. (3) Disconnect the shock chord hook or the attach rod(s). (4) Move the spring-loaded lever on the release assembly to release the hinge pins, then remove the anti-ice panel. (5) Stow the anti-ice panel in the storage bag for protection. (6) Attach the anti-ice panel to the sidewall of the airplane, aft of the cargo door. (a) Stretch the storage bag attach chord over the anti-ice panel, then hook the attach cord into the bracket. (7) Remove the screws, washers, nuts and spacer, that attach the attach bracket to the windshield retainer. NOTE: (8)
B.
Do not let the washers and nuts fall behind the instrument panel when you remove the windshield anti-ice attach bracket.
Remove attach bracket from the anti-ice panel, then remove the screws, washers and nuts.
Install the Windshield Anti-Ice Panel and Attach Bracket (Refer to Figure 201). (1) To install the attach bracket, align the attach bracket with the holes in windshield retainer, then install the screws, washers and nuts. (2) To install the anti-ice panel, do the steps that follow: (a) Put the anti-ice panel in the center of the attach bracket. (b) Insert the inboard hinge pin into the hole in the attach bracket. (c) Retract the hinge pin on the other end of release assembly with the spring-loaded lever. (d) Release the spring-loaded lever to let the hinge pin engage into the hole in the attach bracket. (e) Do a check that the anti-ice panel is installed correctly. (3) Attach the shock chord hook or the attach rod(s) to the airplane. (4) Make that electrical power is off.
CAUTION: Accidental operation of the heated anti-ice panel for an extended period of time, without the engine in operation, will cause damage to the anti-ice panel and crazing of the windshield. (5) (6) 3.
Remove the cover from receptacle, then connect connector to the receptacle. Do the Windshield Anti-Ice System Operational Check. Refer to Windshield Anti-Ice - Inspection/ Check.
Windshield Anti-Ice Electrical Receptacle Cover Removal/Installation A.
Remove the Windshield Anti-Ice Electrical Receptacle Cover (Refer to Figure 201). (1) Make sure that the electrical power is off. (2) Disconnect the connector from the receptacle in the cover. (3) Remove the screws, washers and nuts that attach the cover to the windshield retainers. (4) Remove the nut that attaches the receptacle to the cover.
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Windshield Anti-Ice Panel Installation Figure 201 (Sheet 1)
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Windshield Anti-Ice Panel Installation Figure 201 (Sheet 2)
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Windshield Anti-Ice Panel Installation Figure 201 (Sheet 3)
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Windshield Anti-Ice Panel Installation Figure 201 (Sheet 4)
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Windshield Anti-Ice Panel Installation Figure 201 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7) B.
Put a tag on the wires from the receptacle to identify them. Remove the wires from the receptacle. Remove the cover.
Install the Windshield Anti-Ice Electrical Receptacle Cover (Refer to Figure 201). (1) Put the wires through the hole in the electrical receptacle cover, then install the wires in the receptacle. (2) Remove the tags from the wires. (3) Put the receptacle in the cover, then attach with the nut. (a) Use FULACRYL FS3606T to make a seal around the perimeter of the cover, electrical receptacle and the hole in the windshield retainer where the anti-ice wires route. NOTE: (4) (5)
4.
5.
6.
Seal is applied to keep moisture out of the cabin area.
Install the screws, washers and nuts that attach the cover to the windshield retainers. Connect the connector to the receptacle.
Windshield Anti-Ice Controller Removal/Installation A.
Remove the Windshield Anti-Ice Controller (Refer to Figure 201). (1) Make sure electrical power is off. (2) Remove the circuit breaker panel to get access to the anti-ice controller that is found in the electrical equipment box behind the circuit breaker panel. (3) Put a tag on the wires from the controller. (4) Remove the controller wires. (5) Remove the screws that attach the controller to the electrical equipment box, then remove the controller from the airplane.
B.
Install the Windshield Anti-Ice Controller (Refer to Figure 201). (1) Make sure that electrical power is off. (2) Align the controller with the nutplates in the electrical equipment box, then install with the screws. (3) Remove the identification tags from wires, then install the wires to the controller. (4) Install the circuit breaker panel.
Windshield Anti-Ice Relay Removal/Installation A.
Remove the Windshield Anti-Ice Relay (Refer to Figure 201). (1) Make sure electrical power is off. (2) Remove the circuit breaker panel to get access to the anti-ice relay that is found in the electrical equipment box. (3) Remove the plug(s) from the anti-ice relay mounting base. (4) Remove the screw(s) that attach the relay to the airplane. (5) Remove the relay from the airplane.
B.
Install the Windshield Anti-Ice Relay (Refer to Figure 201). (1) Make sure electrical power is off. (2) Attach the relay with the screw(s). (3) Connect the plug(s) into the relay. (4) Install the circuit breaker panel.
Approved Repairs A.
Windshield Anti-Ice Panel Repair (Refer to Figure 202). NOTE:
(1)
There are two different methods to attach the upper side of the anti-ice panel. Earlier serial number airplanes had a rubber bungee system and this was later changed to a steel rod system.
Make sure electrical power is off.
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Windshield Anti-Ice Panel Repair Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2) (3)
Disconnect the electrical connector from the receptacle that is found at the lower, center part of the windshield. Remove the screws, washers and nuts that attach the lower bracket and release assembly, remove the lower bracket and release assembly. NOTE:
If you replace only one side of the frame, use extreme precaution when you remove the hinge block of the release assembly from the remaining side. This is necessary so you do not damage the copper strip that is found between the hinge block and the frame channel.
CAUTION: Be careful so you do not damage the glass. (4)
Remove the side frame assembly. NOTE:
You can use a thin putty knife to help in the removal of the side frame assembly.
Use methyl n-propyl ketone to remove all the copper tape or sealer from the glass, if necessary. Clean the glass with alcohol, then a clean cloth. Pull off the backing and put the new self-sticking copper tape (3M, 1181) in position the on outboard face of the glass. (8) Make sure there is a 0.06-inch space between the top and side edges of the glass. (9) Make sure there is a 1.0-inch overlap of copper tape from the bottom edge of the glass. (10) Tin the outer 0.75 inches of the overlap of the copper strip(s). Refer to Model 208 Series Wiring Diagram Manual, Chapter 20, Soldering - Maintenance Practices. (11) Put the three 2601445-13 rubber spacers on the inside length of the glass retainer channel. (12) Make sure they are equally distanced between the ends of the channel.
(5) (6) (7)
NOTE:
There will be one spacer in the center and one spacer near each end. The spacers are to prevent contact with the glass and the rivet tails.
(13) Use Type 1, Class 2A adhesive to make a bond between the new 2601445-2 and/or-3 frame assembly to the glass. Refer to Chapter 20, Fuel, Weather and High- Temperature Sealing Maintenance Practices. (14) Align the top edges of the glass and the frame. (15) Put the hinge blocks in position, then attach the components for the windshield anti-ice panel. (16) Drill two holes in each new tinned copper to align the hinge block. (17) Make an electrical bond between the following components. Refer to Chapter 20, Electrical Bonding - Maintenance Practices. (a) The lower bracket to the 2601445-2 and -3 frame assemblies. (b) The 2601445-2 and -3 frame assemblies to the copper strips.
CAUTION: Make sure the electrical lead is attached in the clamp. (c)
The ground terminal to the frame assembly. NOTE:
Make sure that none of the screws extend into a position where the windshield can be scratched.
(18) Remove the rod end and nut from the 2601445-10 rod assembly, then install it through the spring guide assembly. (19) Install the nut and the rod end on the rod assembly. (20) Connect the connector to the receptacle that is found at the center lower windshield.
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MODEL 208 MAINTENANCE MANUAL (21) If the panel has not been modified per SK208-113, or on Airplanes 20800001 thru 20800232, or on Airplanes 208B0001 thru 208B0353, you must also do the step that follows: (a) Remove the existing screw, nut and washer in the upper windshield located at approximately LBL 18.25. 1 Replace the existing screw with a screw combination that forms the mount for the rod assembly. 2 Put the washer and nut on both the outside and inside of the windshield. 3 Make sure you leave 0.22 inches of screw shank between the outer nut and the head of the screw.
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MODEL 208 MAINTENANCE MANUAL WINDSHIELD ANTI-ICE - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the windshield anti-ice system in a serviceable condition.
Task 30-40-00-710 2.
Windshield Anti-Ice System Operational Check A.
General (1) This task gives the procedures to do an operational check of the windshield anti-ice system (non TKS airplanes).
B.
Tools and Equipment (1) None
C.
Access (1) None
D.
Do the Windshield Anti-Ice System Operational Check. (1) For Airplanes 20800001 thru 20800381 and 208B0001 thru 208B1087, and incorporating SK208-113 (a) Put the battery switch to the ON position. (b) Put the windshield anti-ice switch to the AUTO position for one minute. (c) Make sure that the anti-ice panel temperature increases. NOTE:
(2)
The anti-ice panel will feel warm to the touch.
(d) Put the battery switch to the OFF position. (e) Put the windshield anti-ice switches to the OFF position. For Airplanes 20800382 and On and 208B1088 and On and incorporating SK208-146 (a) Put the battery switch to the ON position. (b) Put the ammeter selector switch to the BATT position. (c) Make sure that the W/S ANTI-ICE PRI, W/S ANTI-ICE SEC and W/S ANTI-ICE CONT circuit breakers are engaged. NOTE:
(d) (e) (f) (g) (h) (i) (j)
Each time you move a switch as follows, there will be a change in the ammeter indication and illumination of the WINDSHIELD ANTI-ICE annunciator. If you do not see a change when you move the switch(es), then record the difference and continue the test.
Put the W/S PRIMARY switch to the AUTO position, then record the time. Put the W/S SECONDARY switch to the AUTO position. Make sure that the WINDSHIELD ANTI-ICE annunciator goes off in less than 120 seconds from the time that you put the W/S PRIMARY switch to the AUTO position. If you do not see a change in the ammeter indication when you move the switch(es), then record the difference and continue the test. Momentarily put the W/S PRIMARY switch to the MANUAL position. Momentarily put the W/S SECONDARY switch to the MANUAL position. Put the BATTERY switch to the OFF position.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE WINDSHIELD SPRAY BAR - MAINTENANCE PRACTICES 1.
General A.
This section contains the removal and installation procedures for the TKS anti-ice windshield spray bar.
B.
Recommended maintenance to keep the TKS fluid at its correct viscosity is as follows: • Operate the pumps monthly, or as necessary, in the HIGH mode until the air is removed from the fluid system. • Keep the TKS system operational at all times to keep air pockets out of the system. • If the fluid tank is removed and installed or replaced, do the porous panel purge and test procedures. NOTE:
2.
Tools and Equipment A.
3.
If the fluid is too thick, the porous panels can become blocked or clogged.
For a list of tools and equipment, refer to Ice and Rain Protection - General.
TKS Anti-Ice Windshield Spray Bar Removal/Installation
WARNING: For health and environmental data, review the applicable Material Safety Data Sheet (MSDS). WARNING: Before you disconnect components of the TKS anti-ice system, slowly loosen the coupling that is connected to the component to be removed because it is possible that pressure is still in the system. WARNING: Immediately remove (clean) or contain all the TKS fluid that is spilled. TKS fluid on the floor will cause a slip hazard. WARNING: Before you operate the TKS system during this procedure put plastic sheets or absorbent cloths under the porous panels to keep the TKS fluid off the floor. This will help to prevent injury to personnel. WARNING: Correctly discard all unwanted TKS fluid and dirty cloths. TKS fluid is a hazardous waste and must be discarded in accordance with approved procedures. CAUTION: Use only approved TKS fluids. Approved fluids, in accordance with specification DTD 406B, are normally 80% to 85% mono-ethylene glycol, 5% isopropyl alcohol, and 10% to 20% de-ionized water. Fluid density is approximately 9.2 lbs/gal. CAUTION: Use only clean, filtered fluid in the TKS system. Contamination will cause fluid blockage and/or damage to the porous panel. A.
Remove the Windshield Spray Bar (Refer to Figure 201 and Figure 202). (1) Remove external electrical power from the airplane. (2) Disengage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY • W/S ANTI-ICE • BACKUP ANTI-ICE. (3) On airplanes that have the G1000 system installed, disengage the ENG INTFC circuit breaker on the left circuit breaker panel.
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TKS Anti-Ice System Flow Diagram Figure 201 (Sheet 1)
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TKS Anti-Ice System Windshield Spray Bar Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (4)
On airplanes that do not have the G1000 system installed, disengage the ANTI-ICE GAUGE circuit breaker on the left circuit breaker panel. (5) Open the engine cowl doors. (6) Disconnect the hose fitting from the spray bar assembly. (7) Loosen the spray bar backnut that attaches the spray bar fitting to the bracket. (8) Install caps on all tube ends to keep FOD out of the TKS system. (9) Remove the screws that attach the deflector, P-clips, and spray bar to the cowl deck. (10) Remove the deflector, P-clips, and spray bar from the airplane. B.
Install the Windshield Spray Bar (Refer to Figure 201 and Figure 202). (1) Put the deflector, P-clips, and spray bar in position on the cowl deck. (a) Make sure that the P-clips are not on the spray bar holes. (2) Install the screws that attach the deflector, P-clips, and spray bar to the cowl deck. (3) Remove the caps from the tube ends. (4) Connect the hose fitting to the spray bar assembly. (5) Tighten the spray bar backnut that attaches the spray bar fitting to the bracket. (6) Use Type I, Class B sealant to apply a fillet seal along the forward edge of the deflector. (7) Engage the circuit breakers on the left circuit breaker panel that follow: • PRIMARY ANTI-ICE • W/S ANTI-ICE • BACKUP ANTI-ICE. (8) On airplanes that have the G1000 system installed, engage the ENG INTFC circuit breaker on the left circuit breaker panel. (9) On airplanes that do not have the G1000 system installed, engage the ANTI-ICE GAUGE circuit breaker on the left circuit breaker panel. (10) Supply external electrical power to the airplane. (11) Put the EXTERNAL POWER switch on the pilot's switch panel in the BUS position. (12) Put the MAX FLOW switch on the ANTI-ICE FLUID CONTROL switch panel in the WINDSHIELD position, then release the switch. NOTE:
The windshield pump will start when you put the spring-loaded MAX FLOW switch in the WINDSHIELD position and it will continue for four seconds after you release it.
(a) Make sure that fluid flows from the spray bar. (b) Make sure that there is no fluid leakage from the fitting. (13) Remove external electrical power from the airplane. (14) Clean the floor and the airplane surfaces as necessary. (15) Close the engine cowl doors.
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MODEL 208 MAINTENANCE MANUAL PROPELLER ANTI-ICE - DESCRIPTION AND OPERATION 1.
General A.
2.
Model 208 propeller is protected from ice by an electrical anti-ice system.
Description and Operation A.
The system is of an electrothermal type, consisting of electrically heated de-ice boots bonded to each propeller blade, a slip ring assembly for power distribution to the propeller de-ice boots, a brush block assembly to transfer electrical power to the rotating slip ring and a timer, installed on a diagonal brace on the aft, left side of the firewall, to cycle electrical power to the de-ice boots in proper sequence. A three position toggle switch labeled PROP, located on the de- ice/anti-ice switch panel, on lower left instrument panel, controls the engine propeller de-ice systems. Two circuit breakers; a heater circuit breaker, labeled PROP ANTI-ICE protects the propeller anti-ice heating circuit, and a control circuit breaker, labeled PROP ANTI-ICE CONTROL protects the propeller anti-ice timer circuit. Both circuit breakers are located on the left sidewall circuit breaker panel. A propeller anti-ice ammeter, located on the left instrument panel, indicates amperage for the propeller anti-ice system. The anti-ice system applies heat to the surfaces of the propeller blades where ice would normally adhere. This heat, plus centrifugal force and the blast from the airstream, removes accumulated ice. Because excessive anti-ice heat may damage the propeller blades which are constructed of composite materials, an oil pressure switch in the electrical circuit is utilized to prevent the propeller anti-ice from being turned on without the engine operating. When the PROP anti- ice switch is placed in the AUTO (upper) position, the timer controls electrical power through the brush block and slip ring to the three propeller anti-ice boots in intervals of 9O seconds on and 90 seconds off. The anti-ice system is off when the switch is placed in the center position. In the event of a malfunction in the propeller anti-ice timer circuit, the switch can be held in the momentary MANUAL (lower) position to bypass the timer circuit and achieve propeller anti-icing. Operation of the propeller anti-ice system can be checked by monitoring the prop anti-ice ammeter. The anti-ice system is operating correctly when ammeter is indicating within the green arc. The anti-ice system incorporates lightning protection which consist of a metal oxide varistor mounted on the left upper forward side of the firewall, and three overspark brackets installed beneath the brush block terminal strips.
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MODEL 208 MAINTENANCE MANUAL PROPELLER ANTI-ICE - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system understanding. Refer to Figure101.
B.
The propeller anti-ice timer has a 90 second ON cycle and a 90 second OFF cycle. but does not have a automatic reset function. When the propeller anti-ice switch is placed in the OFF position during a OFF cycle the timer will stop at that point in its 90 second cycle. When the switch is placed in the AUTO position, the system will remain OFF until the remainder of the 90 second cycle is completed. Also if the anti-ice switch is turned OFF during a ON cycle the timer will stop at this point in its 90 second cycle. When the anti-ice switch is placed in the AUTO position, the system will remain ON until the remainder of the ON cycle is completed. The completion of these cycle times should not be misinterpreted as caused by a faulty timer.
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Propeller Anti-Ice Toubleshooting Chart Figure 101 (Sheet 1)
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Propeller Anti-Ice Toubleshooting Chart Figure 101 (Sheet 2)
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Propeller Anti-Ice Toubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL PROPELLER ANTI-ICE - MAINTENANCE PRACTICES 1.
General A.
2.
Propeller anti-ice maintenance practices consist of propeller anti-ice boot removal/installation and electrical components removal/installation.
Propeller Anti-Ice Boots (Hartzell) Removal/Installation A.
Remove Propeller Anti-Ice Boots (Refer to Figure 201).
WARNING: Cement and solvent vapors are toxic and extremely flammable. Use only in a well ventilated area away from sparks and vapors. Excess exposure could cause injury or death. If dizziness or nausea occur, obtain fresh air immediately. Avoid contact with skin or eyes. Use solvent- resistant gloves to minimize skin exposure. Use safety glasses to minimize chance of eye contact. If eye contact occurs, flush eyes with water for 15 minutes and see a physician. If skin contact occurs wash thoroughly with soap and water. If swallowed, do not induce vomiting. See a physician immediately. (1) (2) (3) (4)
B.
Ensure that airplane electrical power is off. Remove all large tie straps (5) securing electrical leads to propeller and hub. Remove small tie strap (2) securing the two pin housing (6) and disconnect two pin housing (6) on the anti-ice boot leads. To remove or loosen installed anti-ice boots, use Toluol to soften the cement line. Apply a minimum amount of this solvent to the cement line as tension is applied to peel back the boot. The removal should be slow enough to allow the solvent to undercut the cement so that parts will not be damaged.
Install Propeller Anti-Ice Boots (Refer to Figure 201). (1) Clean propeller surface to be bonded methyl n-propyl ketone. For final cleaning, wipe solvent film off quickly with a clear, dry cloth before it has time to dry. (2) Draw a line on centerline of leading edge of propeller blade. (3) Position boot on propeller so bottom of boot is 1.0 inch, +0.031 or -0.031 inch from propeller hub. (4) Position boot centerline (boot centerline is indicated by embossed marks 3/8 inch long at each end of the boot on the breeze surface) over propeller leading edge centerline. These marks may be transferred to the boot side using a silver pencil. (5) Slide inboard end of boot centerline 0.25 inch toward forward side of prop and mark this dimension. (6) Slide outboard end of boot centerline 0.25 inch toward forward side of prop and mark this dimension. (7) Draw a line between two marks established in steps 5 and 6. This is the line to be used to center boot on the propeller leading edge centerline. (8) Mask off an area 1/2 from each side and outer end of boot and remove boot. (9) Mix EC-1300L cement (Minnesota Mining & Mfg. Co.) thoroughly. Surfaces shall be above 60°F prior to applying cement. During periods of high humidity, care shall be taken to prevent moisture condensation due to cooling effect of evaporating solvent. This can be done by warming the area with a heat gun or heat lamp. Apply one even brush coat of EC-1300L cement to cleaned composite surface. Allow to air dry for a minimum of one hour, then apply a second even brush coat of EC-1300L cement. (10) Moisten a clean cloth with methyl n-propyl ketone and clean bond surface of boot, changing cloths frequently to avoid contamination of cleaned area. (11) Apply one even coat of EC-1300L cement to bond surface of boot. It is not necessary to cement more than half of boot strap.
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Propeller Anti-Ice Boots Installation Figure 201 (Sheet 1)
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Propeller Anti-Ice Boots Installation Figure 201 (Sheet 2)
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Propeller Anti-Ice Boots Installation Figure 201 (Sheet 3)
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Propeller Anti-Ice Boots Installation Figure 201 (Sheet 4)
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Propeller Anti-Ice Boots Installation Figure 201 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL (12) Using a silver colored pencil, mark a centerline along leading edge of propeller blade and a corresponding centerline on cemented bond surface of boot. (13) Reactivate surface of cement using a clean, link-free cloth, heavily moistened with Toluol. Avoid excessive rubbing of cement, which would remove cement. (14) Position boot centerline on line established in step 7, starting at hub end at the position marked. Tack boot centerline to line established in step 7. If boot is allowed to get off center, pull up with a quick motion and replace properly. Roll firmly along centerline with a rubber roller. (15) Gradually tilting roller, work boot carefully over either side of blade contour to avoid trapping air in pockets. (16) Roll outwardly from centerline to edges. If excess material at edges tends to form wrinkles, work them out smoothly and carefully with fingers. (17) Apply one even coat of EC-539 (Minnesota Mining & Mfg. Co.), mixed per manufacturers instructions, around edges of installed boot. (18) Remove masking tape from propeller and clean surface of propeller by wiping with a clean cloth dampened with Toluol. (19) Test anti-ice boots for continuity by using an ohmeter. Connect leads from ohmmeter to anti- ice boot wire connector (6) terminal. There should be a reading between 2.5 and 3.5 ohms. (20) Install electrical connectors, wire harness leads, and tie-straps in the following order. (a) Connect anti-ice boot wire connector (6) to slip ring wire connector (3). (b) Install small tie- strap (2) between leads and around assembled connector. (Do not tighten). (c) Install large tie-straps (5) through small tie-strap (2) securing connector and around prop blade clamp. (Do not tighten). (d) Route slip ring wire leads under upper tie-strap (5). (e) Install second small tie- strap (4) securing slip ring wire leads to upper large tie-strap (5). (f) Position connector assembly on hub to eliminate slack in anti-ice boot leads. (g) Install tie-strap (1) around prop blade over anti-ice boot leads. (h) Tighten all tie-straps. 3.
Propeller Anti-Ice Boots (McCauley) Removal/Installation A.
Remove Propeller Anti-ice Boots (Refer to Figure 201).
CAUTION: Cushion the jaws of any pulling tool (vise grips, pliers, etc.) to prevent damage to the boot, unless the boot is to be scrapped. (1) (2)
If boot is to be scrapped, strip from blade without solvent. Cut sta-straps (42) and disconnect electrical lead (46).
CAUTION: When removing boots from a complete propeller assembly, care must be taken to prevent solvent from leaking into the propeller hub and causing damage to the seals. The blade being worked on should be pointed down so all excess solvent will run to the outboard tip of the blade. As an extra precautionary measure, the hub and blade area should be masked. Do not use any sharp objects which might scratch the blade when removing the boot. (3) (4) (5) (6)
Using methyl n-propyl ketone or toluol to soften adhesion line between the anti-ice boot and propeller blade, start at one corner, loosen enough of the boot to grasp with vise grips, pliers, or similar tool. Apply a steady pull to remove boot; pull the boot from the blade slowly and carefully while continuing to use a liberal amount of methyl n-propyl ketone or toluol to soften the adhesion line. Remove all residual cement from blade. Use solvents with caution as mentioned above. Visually inspect propeller blade for damage or deterioration. Check for corrosion, cracks, dents or nicks. If defects are found propeller must be repaired by an authorized propeller repair station.
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Install Propeller Anti-ice Boots (Refer to Figure 201). (1) Outline area to be masked using a red pencil according to dimensions shown in Figure 204. A template or the anti-ice boot may be used. Electrical leads must be aligned with terminal bracket. (2) Once the anti-ice boot is positioned, mark an area 1/2 inch ouside the boot perimeter. Using masking tape, mask around the outline. (3) Clean entire masked area throughly with methyl n-propyl ketone or acetone. For final cleaning, quickly wipe off solvent with a clean, dry, lint-free cloth to avoid leaving a film. (4) Apply a second layer of masking tape to cover an additional 1/8 inch inside of previously masked area.
CAUTION: Cleanliness of metal and rubber parts cannot be overemphasized. Only very clean surfaces will ensure maximum adhesion. (5)
Moisten a clean cloth with methyl n-propyl ketone or acetone. Clean unglazed (back) surface of the anti-ice boot. Change the cloth frequently to avoid contamination of the clean area. NOTE:
(6) (7) (8)
(9)
(10) (11)
(12) (13) (14) (15)
To prevent curling of the anti-ice boot edges, apply masking tape to the edges on the smooth side before applying cement to the fabric impressioned side. Remove tape from the anti-ice boot before starting installation.
Installation should be made at room temperature (60°-75°F). Apply one even brush coat of cement to the clean, masked surface of the propeller blade and the fabric impressioned side of the anti-ice boot. Allow cement to air dry for a minimum of one hour at 40°F or above, when the relative humidity is less than 75%. If humidity is 75% to 90%, additional drying time will be required to cure cement. Do not apply cement if relative humidity is higher than 90%. After cement is dry (not tacky), apply a second even, brush coat to the anti-ice boot. Then immediately apply an even brush coat of cement to the clean masked off area of the propeller. Timing is important because the cement on both surfaces must reach the tacky stage at the same time. When cement is tacky on both surfaces, locate anti-ice boot electrical leads with terminal bracket installed. Tack the anti-ice boot center line to the leading edge of the blade, starting at the inboard end working toward the tip. If cement dries, use methyl n-propyl ketone or toluol as necessary. If boot is allowed to get off center, pull up with a quick motion and re-apply. If cement is removed from either surface, completely remove the boot and re-apply cement, as in step (6) and (7). Use methyl n-propyl ketone or toluol as necessary to re-install boot. When correctly positioned, press firmly with rubber or wooden hand roller along full length of the leading edge to form a tight bond. Gradually tilt roller over either side of leading edge contour to avoid trapping air. Roll from leading edge of blade toward the tip. Work all excess boot material out to perimeter before moving to the next section. If excess material at boot edges tends to pucker, use fingers to carefully work puckers smooth. Remove masking tape applied in step (4). Check the electrical resistance between the boot leads, reading should be between 3.60 maximum, 3.26 minimum. Check for intermittent open circuits by tensioning the anti-ice boot leads while measuring resistance. Also press lightly on entire boot heating element surface and in area adjacent to leads. Resistance must not fluctuate. Mix two parts of Sunbrite 78-U-1003 brushable black enamel with one part enamel catalyst U1001-C.
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CAUTION: It is imperative that the masking steps as described in steps (3) and (4) be followed. This will ensure the sealer will be applied to both the adhesive and 1/8 inch of bare metal. If the adhesion line and sealer line start at the same point water will be allowed to seep underneath the adhesion line, resulting in an ineffective seal. (16) Apply one, even, brush coat of sealer to the area around the boot covering the 1/8 inch of bare metal and adhesive along with a masked off area of 1/8 inch of the anti-ice boot. Remove masking tape as sealer is brushed on, otherwise, seal will pull up along with the tape. Allow sealer to dry. (17) Allow 12 hours minimum cure time and 24 hours before actually operating the anti-ice system. 4.
Slip Ring Assembly (Hartzell) Removal/Rework/Installation A.
Remove Slip Ring Assembly (Refer to Figure 201). (1) Ensure that airplane electrical power is off. (2) Remove propeller. (3) Tag and remove electrical wires from the three terminal strips (16) on spinner bulkhead. (4) Remove and retain buttonhead screws (10), beveled washers (9), washers (8), and nuts (7) securing slip ring (18) to spinner bulkhead. (5) Remove slip ring (18) carefully working electrical leads through holes in spinner bulkhead.
B.
Rework Slip Ring Assembly (Refer to Figure 202). NOTE:
The slip rings can be reworked down to a minimum height of 0.187 inch. If this dimension is exceeded then slip ring assembly will have to be replaced.
(1)
Check slip rings (2) for surface damage (gouges, pits, etc.). If damage exists, proceed with the following steps. (2) Tag and remove wires from slip ring assembly (2). (3) Check that flatness of mounting surface is within 0.005 inch overall. (4) Locate assembly concentrically in a lathe so that runout does not exceed 0.002 inch in 360degree rotation. (5) Using a light cut for a smooth finish (29-35 micro- inches), cut no deeper than required to remove surface damage. (6) Ensure that contact surfaces of slip ring (2) is parallel within 0.005 inches and flat within 0.005 inches overall; deviation is not to exceed 0.002 inches in any four inch interval of slip ring travel. (7) Undercut epoxy insulation between and around slip rings 0.020 to 0.030 inches as necessary. Outer edge of slip ring holder must be undercut to same dimension as insulation. (8) Deburr slip ring edges. (9) Check Insulation resistance between slip rings and to metal holder using a mega-ohmeter. Applying 500 VDC, resistance must be a minimum of 0.5 megohms after one minute. (10) Reconnect wires that were removed in step (2). C.
Install Slip Ring Assembly (Refer to Figure 201). (1) Insert slip ring electrical leads through holes in spinner bulkhead and align screw holes. (2) Install buttonhead screws (10), beveled washers (9), washers (8), and nuts (7). (3) Install propeller. NOTE:
5.
Before proceeding with slip ring installation, the following slip ring alignment must be accomplished.
Slip Ring Alignment Check (Hartzell) A.
Check Slip Ring Alignment (Refer to Figure 202). (1) Securely attach a dial indicator gage to brush block mounting bracket and place pointer on slip ring surface.
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Brush Face Alignment Figure 202 (Sheet 1)
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Brush Face Alignment Figure 202 (Sheet 2)
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Rotate propeller slowly by hand checking slip ring (2) surfaces to ensure they deviate no more from a true plane than 0.008 TIR (true indicator reading) and 0.002 in any four inch interval of slip ring travel. Vary torque on screws between 40 and 100 inch-pounds to obtain flatness required. NOTE:
(3)
If slip ring (2) runout is within limits specified, no corrective action is required. If runout is not within limits specified, slip ring will have to be replaced. NOTE:
6.
Care must be taken to exert a uniform push or pull on propeller to avoid a considerable error in readings.
Check alignment of brushes (1) in brush block assembly (3) with slip ring assembly (2).
Slip Ring Run-Out Test (McCauley) A.
Test Slip Ring Run-Out (Refer to Figure 202 ). NOTE:
Removal and installation of slip ring assemblies on McCauley propellers for Pratt and Whitney engine installations are limited to authorized FAA approved propeller repair stations or A & P mechanics that have completed the C703 series propeller assembly/disassembly training offered by McCauley and incorporate McCauley Technical Report 722 as revised.
CAUTION: Excessive slip ring runout will result in severe arcing between slip ring and brushes, and cause rapid brush wear. If allowed to continue, this condition will result in rapid deterioration of slip ring and brush contact surfaces, and lead to eventual failure of propeller de-icing system. (1) (2) (3) (4) 7.
To check this condition, ensure the slip ring face run-out does not exceed 0.008 inches (total indicator reading). Securely attach a dial indicator to the brush block mounting bracket and place pointer on slip ring surface. Rotate propeller through 360 degrees of rotation observing the dial indicator for a total indicator reading not to exceed 0.008 inches. If the reading exceeds 0.008 inches, slip ring assembly must be removed and remachined.
Brush Length Inspection (Hartzell) A.
Inspect Brush Length (Refer to Figure 203 ). NOTE: (1) (2) (3) (4) (5) (6)
Inspect brushes and clean slip ring in accordance with Inspection Time Limits set forth in Chapter 5.
Ensure that airplane electrical power is off. Remove right nose cap half. Remove washers (1) and nuts (2) securing brush block assembly bracket (3) to engine. Position brushes (4) so that ends of brushes extend 0.0625 inch from brush block assembly module (5). Place marks on a straightened paper clip (6) 0.36 inch from end, and 1.39 inches from end. Position paper clip (6) through slot in brush block bracket (3) and into hole in brush block assembly module (5). NOTE:
(7)
The brushes (4) may or may not be equipped with rods.
Observe appropriate mark on paper clip (6), 0.36 inch mark (with rods) or 1.39 inch mark (without rods). If appropriate mark disappears into brush block assembly module (5) brush block assembly module (5) must be replaced.
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8.
Brush Length Inspection (McCauley) A.
Inspect Brush Length (Refer to Figure 203). NOTE: (1) (2) (3) (4) (5) (6) (7) (8) (9)
9.
Ensure that airplane electrical power is off. Remove right nose cap half. Remove nuts (2) and washers (3) securing bracket (1) to engine. Insert a small diameter feeler gage (5) into the slots provided on the sides of the block holder assembly. With the feeler gage inserted, push down on the brush until it bottoms out on the gage (5). Measure the distance between the holder and top surface of the brush. If brush measures 0.094 inches or less, the brush and brush block assembly (4) should be replaced. Reinstall brush block assembly (4) on engine using washers (3) and nuts (2). Torque 28 to 32 inch- pounds. Reinstall right nose cap half.
Brush Block Removal/Installation (Hartzell) A.
Remove Brush Block (Refer to Figure 201). (1) Ensure that airplane electrical power is off. (2) Tag to identify and disconnect electrical wires from brush block (29). (3) Remove screws (30), washers (26), and nuts (25) securing brush block (29) to mount (27). (4) Remove shim (28) between brush block (29) and mount (27). Remove brush block (29) from airplane.
B.
Install Brush Block (Refer to Figure 201). NOTE: (1) (2)
10.
Inspect brushes and clean slip ring in accordance with Inspection Time Limits set forth in Chapter 5.
Before proceeding with brush block installation, ensure that brush block alignment has been accomplished.
Ensure that airplane electrical power is off. Insert shim (28) between brush block (29) and mount (27) then install screws (30), washers (26), and nuts (25).
Brush Block Removal/Installation (McCauley) A.
Remove Brush Block (Refer to Figure 201). (1) Ensure that airplane electrical power is off. (2) Tag to identify and disconnect electrical leads (54). (3) Remove screws (51) washers (52) spacers (60) and shims (59) then remove brush block (58).
B.
Install Brush Block (Refer to Figure 201). NOTE: (1) (2)
Before proceeding with brush block installation, brush block alignment must be accomplished.
Ensure that airplane electrical power is off. Insert screws (51) with washers (52) through bracket (55) then install spacers (60) and shims (59) and screw into brush block (58) but do not tighten.
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Brush Length Inspection Figure 203 (Sheet 1)
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Brush Length Inspection Figure 203 (Sheet 2)
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11.
Brush Block Assembly to Slip Ring Alignment (Hartzell)
CAUTION: Ensure that slip ring alignment has been accomplished before attempting to align brushes on slip ring. A.
Align Brush Block Assembly to Slip Ring Attachment (Refer to Figure 202). NOTE:
(1) (2) (3) (4) (5)
12.
The clearance between brush block (3) and slip ring (2) must be 0.064 inch, +0.015 or -0.015 inch. The brushes are to be lined up with slip ring so that entire face of each brush (1) is in contact with slip ring (2) throughout the full 360 degrees of slip ring rotation. The brushes must contact slip ring at an angle of 2 degree from perpendicular to slip ring surface, measured toward the direction of rotation of slip ring. Brush projection can be adjusted by loosening hardware attaching the brush block and holding the brushes in desired location while retightening hardware. Slotted holes are provided. To center brushes on slip ring, a shim made of a series of laminates is provided and may be peeled for proper alignment. Layers of metal 0.003 inch are used to make up shims which are approximately 0.20 thick overall. Shims may also be fabricated locally.
Brush Block Assembly to Slip Ring Alignment (McCauley) A.
13.
Keep brushes retracted in brush block until slip ring and propeller assemblies have been installed. In order to get smooth, efficient and quiet transfer of electric power from brushes to slip ring, brush alignment must be checked and adjusted, to meet the following requirements.
Align Brush Block Assembly to Slip Ring Attachment (Refer to Figure 202 ). (1) Check that brushes are aligned with slip ring. This may be accomplished by adding or removing shims (59). (2) Brushes must be lined with slip ring so the entire face of each brush is in contact with the slip ring throughout the full 360 degrees of slip ring rotation. (3) This may be accomplished by loosening screws (51) and adjusting in slotted holes in bracket (55). (4) At the same time the distance between the face of the brush block assembly and the slip ring must be 0.064 inch, +0.015 or -0.015 inch. (5) Torque screws (51) 25 to 30 inch-pounds and safety wire. (6) Remove tags and connect electrical leads (54) to terminal block (53).
Brush Block/Slip Ring Cleaning
CAUTION: Accumulations of engine oil and carbon dust particles on the slip ring, and brush block can have detremental effects, and cause accellerated wear. A. 14.
Cleaning Brush Block and Slip Ring. (1) Refer to INTRO - List of Publications, Chapter 30 - Deice for cleaning procedures.
Propeller Anti-Ice Timer Removal/Installation A.
Remove Propeller Anti-Ice Timer (Refer to Figure 201). (1) Ensure that airplane electrical power is off. (2) Disconnect electrical plug (22) from timer (23). (3) Remove two screws (24) securing timer to diagonal brace on aft side of firewall. (4) Remove timer from airplane.
B.
Install Propeller Anti-Ice Timer (Refer to Figure 201). (1) Ensure that airplane electrical power is off. (2) Install timer (23) to nutplates using screws (24).
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MODEL 208 MAINTENANCE MANUAL (3) 15.
Connect electrical plug (22) to timer (23).
Propeller Anti-Ice Ammeter Removal/Installation A.
Remove Propeller Anti-Ice Ammeter (Refer to Figure 201). (1) Ensure that airplane electrical power is off. (2) Unscrew bezel (34) and remove along with 0-ring (33). (3) Remove body (31) forward out of instrument panel. (4) To remove propeller anti- ice ammeter from airplane, unsolder wires from solder lugs on ammeter.
B.
Install Propeller Anti-Ice Ammeter (Refer to Figure 201). (1) Solder wires to solder lugs of propeller anti-ice ammeter (31). (2) Insert ammeter body (31) through hole in instrument panel from forward side. (3) Install 0-ring (33) and screw bezel (34) onto threads of body (31).
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Anti-Ice Boot Location Figure 204 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL TKS ANTI-ICE PROPELLER (McCauley) - MAINTENANCE PRACTICES 1.
General A.
2.
The TKS anti-ice propeller maintenance practices consist of propeller anti-ice feed shoe removal and installation procedures, and the spinner bulkhead anti-ice components removal, installation and adjustments procedures.
Propeller TKS Anti-Ice Feed Shoes Removal/Installation A.
Remove the Propeller Anti-ice Feed Shoes (Refer to Figure 201).
WARNING: Cement and solvent vapors are toxic and extremely ßammable. Use these chemicals only in a well ventilated area away from sparks and vapors. Excess exposure could cause injury or death. If dizziness or nausea occur, get to fresh air immediately. Avoid contact with skin or eyes. Use solvent-resistant gloves to minimize skin exposure. Use safety glasses to protect your eyes from chemicals. If you get chemicals in your eyes, ßush your eyes with water for 15 minutes and see a physician immediately. If you get chemicals on your skin, wash thoroughly with soap and water. If you swallow chemicals, do not induce vomiting. See a physician immediately. CAUTION: The jaws of any tool (vise grips, pliers, etc.) that you use to pull on the feed shoe must be cushioned to prevent damage to the feed shoe, unless the feed shoe is to be scrapped. (1) (2) (3)
Make sure all electrical power switches are in OFF position. Remove the spinner screws and Þber washers from the spinner, and remove the spinner from the spinner bulkhead. If the feed shoe is to be discarded, remove it from the propeller blade without solvent. NOTE:
The feed shoe can tear or come off in pieces.
CAUTION: When you remove feed shoes from a propeller assembly, be careful not to let solvent leak into the propeller hub and cause damage to the seals. The blade that is being worked on must be pointed down so all excess solvent will run to the outboard tip of the blade. As an extra precautionary measure, the hub and blade area must be masked. Do not use any sharp objects which might scratch the blade when you remove the feed shoe. (4) (5) (6) (7)
Use methyl n-propyl ketone or toluene to soften the adhesion line between the anti-ice feed shoe and the propeller blade. Start at one corner and loosen enough of the feed shoe to grasp it with vise grips, pliers, or similar tools. Apply a steady pull to remove the feed shoe; pull the feed shoe from the blade slowly and carefully while you continue to use methyl n-propyl ketone or toluene to soften the adhesion line. Remove all residual cement from the blade. Use solvents with caution as mentioned above. Visually do an inspection of the propeller blade for damage or deterioration. Look for corrosion, cracks, dents or nicks. If defects are found, the propeller must be repaired by an authorized propeller repair station.
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MODEL 208 MAINTENANCE MANUAL B.
Install the TKS Anti-ice Propeller Feed Shoes (Refer to Figure 201). (1) Prepare the Propeller Blades for the Feed Shoe Installation. (a) Trim each propeller blade feed shoe if necessary. (b) Mark the inboard center line on the leading edge of the propeller blades. (c) Use a template or hand Þt the anti-ice feed shoe to the blades so that the feed shoe center line is on the leading edge center line of the propeller blade. Refer to Figure 201. (d) Mark an area 1/2 inch outside the feed shoe perimeter on the propeller blade with a red pencil. Refer to Figure 201. (e) Use the red pencil line as the perimeter of the area on each propeller blade to be masked. (f) Install masking tape around the outline.
CAUTION: It is necessary that the masking steps described be followed so the sealer will be applied to both the cement and 1/8 inch of bare metal. If the cement line and sealer line start at the same point, water will seep under the cement line and cause an unserviceable seal. Refer to Figure 201. (g) (h) (i)
Remove all paint inside of the masked off area on propeller blades painted with lacquer. On propeller blades painted with polyurethane, lightly sand inside the masked off area with 400 grit sandpaper. Clean all of the masked area on the propeller blades thoroughly with methyl n-propyl ketone or acetone. Quickly clean the solvent from the propeller blades with a clean, dry, lint-free cloth so that you do not leave a Þlm. Apply a second layer of masking tape on the propeller blades to cover an additional 1/8 inch of bare metal area inside of the previously masked area. Refer to Figure 201.
CAUTION: The metal and rubber parts must be clean. Only very clean surfaces will cause maximum bond of the cement. (2)
Apply Cement to the Feed Shoes and the Propeller Blades. (a) Lightly sand the bond surface of the feed shoe with sandpaper to cause maximum bond. (b) Moisten a clean cloth with methyl n-propyl ketone or acetone. Clean the bond surface of the anti-ice feed shoe. Change the cloth frequently to avoid contamination of the clean area. NOTE: (c) (d) (e) (f) (g)
(h)
You can use masking tape to prevent any curl of the anti-ice feed shoe edges when you apply cement to the back side of the feed shoe.
Apply masking tape to the breeze side of the feed shoe edges. Let approximately 1/4 inch of the tape overhang the edge of the feed shoe. Lay the feed shoe on a piece of cardboard with the bond side up. Tape the feed shoe onto the cardboard with the 1/4 inch overlap of masking tape. Make sure to thoroughly mix the cement. Apply one even brush coat of 1300L or EC1403 cement to the clean, masked surface of the propeller blade and to the fabric impression side of the anti-ice feed shoe. Apply cement at the room temperature of 60°- 75°F. Allow the cement to air dry for a minimum of one hour at 40°F or above, when the relative humidity is less than 75%. If the humidity is 75% to 90%, additional drying time will be necessary to cure the cement. Do not apply cement if the relative humidity is higher than 90%. After the cement is dry (not tacky), apply a second even brush coat to the anti-ice feed shoe. Then immediately apply an even brush coat of cement to the clean masked off area of the propeller. Timing is important because the cement on both surfaces must reach the tacky stage at the same time.
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Install the TKS Propeller Anti-ice Feed Shoes Figure 201 (Sheet 1)
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Install the TKS Propeller Anti-ice Feed Shoes Figure 201 (Sheet 2)
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(4)
Install Feed Shoes to the Propeller Blades. (a) Remove the anti-ice feed shoe from the cardboard, and the masking tape from the anti-ice feed shoe before you start the installation. (b) When the cement is tacky on both the propeller blade and feed shoe surfaces, put the anti-ice feed shoe center line to the center line on leading edge of the blade. Start with the inboard end of the blade and work toward the tip. 1 If the cement dries, use methyl n-propyl ketone or toluene as necessary until the cement is tacky. 2 If the feed shoe is off center, pull it up with a quick motion and install it again. 3 Use methyl n-propyl ketone or toluene as necessary when you have to install the feed shoe again. (c) When the feed shoe is correctly in place, use a rubber or wooden hand roller and press Þrmly on the full length of the leading edge to form a tight bond. (d) Gradually push the roller over each side of the leading edge contour to avoid trapping air. Roll from the leading edge of the propeller blade toward the tip. Work all excess feed shoe material out to the perimeter before you move to the next section. If there is excess material at the feed shoe edges that tends to pucker, use your Þngers and carefully work puckers smooth. (e) Remove the masking tape from the propeller blades. Apply Sealer to the Feed Shoes and Propeller Blades (Refer to Figure 201). (a) Mix two parts of Sunbrite 78-U-1003 brushable black enamel with one part of enamel catalyst U-1001-C.
CAUTION: It is necessary that the masking steps described be followed so the sealer will be applied to both the cement and 1/8 inch of bare metal. If the cement line and sealer line start at the same point, water will seep under the cement line and cause an unserviceable seal. (Refer to Figure 201.) (b)
3.
Apply one, even, brush coat of sealer to the area around the feed shoe and make sure you cover the 1/8 inch of bare metal and adhesive along with the masked off area of 1/8 inch of the anti-ice feed shoe. Remove the masking tape as the sealer is brushed on, otherwise, the sealer will pull up along with the tape. Let the sealer dry.
Slinger Ring and Feed Nozzle Alignment Check A.
Examine the TKS Feed Nozzle to Slinger Ring Alignment (Refer to Figure 202). (1) Measure the distance between the slinger ring feed nozzle and the slinger ring channel. NOTE: (2)
(3)
The feed nozzle must extend into the slinger ring channel 0.1 to 0.15 inches and have an edge distance of 0.1 to 0.15 inches from the slinger ring. (Refer to Figure 202.)
Adjust the slinger ring feed nozzle to direct the ßuid stream to land in the slinger ring channel as necessary. (a) Bend and rotate the feed nozzle at the Þtting to align the nozzle with the slinger ring channel. (b) Extend the feed nozzle into the slinger ring channel 0.1 to 0.15 inches with and edge distance of 0.1 to 0.15 inches from the slinger ring. (Refer to Figure 202.) Rotate the propeller slowly by hand and make sure the distance between the slinger ring and the feeder tube is in alignment tolerance. Adjust the feed nozzle as necessary to get good alignment.
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4.
Feed Tube to Propeller Blade Alignment Check A.
Adjust the propeller feed tubes to direct the anti-ice ßuid stream on to the propeller blades. (Refer to Figure 203.) NOTE: (1) (2)
The feed tubes are inside of the propeller spinner.
Make sure that each feed tube is over the second groove of the adjacent feed shoe with the propeller in full Þne pitch. Cycle the propeller through a range of movement to verify that each tube has a 3/16 inch clearance from the propeller boot.
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Slinger Ring and Feed Nozzle Alignment Check Figure 202 (Sheet 1)
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Feed Tubes to Propeller Blade Alignment Check Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL WINDSHIELD ICE INDICATOR LIGHT - MAINTENANCE PRACTICES 1.
2.
General A.
This section gives information necessary to remove and install the windshield ice indicator light on the cockpit glareshield cover.
B.
A single light is on the glareshield cover at FS 112.56 and LBL 2.22. The light is controlled by the Day-Night switch on the instrument panel. The area of the windshield that the ice indicator light effects, looks different when ice is on the windshield.
Tools and Equipment A.
3.
For a List of Tools and Equipment, refer to Ice and Rain Protection - General.
Windshield Ice Indicator Light A.
Remove the Windshield Ice Indicator Light (Refer to Figure 201). (1) Remove the electrical power from the airplane. (2) Remove the screws that attach the mount assembly to the glareshield cover. (3) Remove the screws that attach the glareshield cover, as necessary, to pull the glareshield cover up to get access to the lamp assembly wires. (4) Cut the tie wraps that hold the lamp assembly wires to the wire bundle. (5) Disconnect the lamp assembly wires at the wire housing. Refer to the Model 208 Wiring Diagram Manual. (6) Remove the windshield ice indicator light from the airplane.
B.
Install the Windshield Ice Indicator Light (Refer to Figure 201). (1) Put the mount assembly in position to the glareshield cover. (2) Put the windshield lamp assembly wires through the mount assembly and the glareshield cover. (3) Install 0.5 inches of heat shrinkable tubing on the wires where they come out of the base of the mount assembly. (4) Connect the lamp assembly wires at the wire housing. Refer to the Model 208 Wiring Diagram Manual. (5) Use tie wraps to attach the lamp assembly wires to the wire bundle as necessary. (6) Install the screws that attach the glareshield cover to the airplane. (7) Install the screws that attach the mount assembly to the glareshield cover. (8) Make sure that the mount assembly tube is perpendicular to the windshield.
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MODEL 208 MAINTENANCE MANUAL
Windshield Ice Indicator Light Installation Figure 201 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FLIGHT INTO KNOWN ICING CONDITIONS EQUIPMENT - DESCRIPTION AND OPERATION 1.
General A.
The flight into known icing conditions equipment packages permit flight penetration of icing conditions as defined by the FAA. The packages include all the previously mentioned ice protection systems and either a pneumatic/electrically heated deicing system, or a TKS anti-ice system. (1) The pneumatic anti-ice system includes deicing boots on the leading edges of the wings, wing struts, and horizontal and vertical stabilizers. It also includes an electrically heated removable windshield deicing panel, electrically heated propeller deicing boots and anti-ice ammeter, electrically heated pitot-static and stall warning systems with optional right heated pitot-static system, and an icing low airspeed awareness system (110 KIAS). Also included are a standby electrical system that uses a 75-amp alternator, control surface mounted static discharge wicks, and an ice detection light for detection of ice buildup on the leading edge of the left wing at night. (2) The TKS anti-ice system is a freezing point depressant fluid anti-ice system to prevent ice formation on the leading edges of the wings, horizontal stabilizers, struts, vertical stabilizer, propeller, and the windshield. A monoethylene glycol/isopropyl alcohol/deionized water solution is used to anti-ice/de-ice the airframe surfaces and windshield in flight. The fluid solution is a freezing point depressant that is swept rearward over the surfaces and prevents ice buildup. This system also has a right and a left electrically heated pitot-static and a low airspeed awareness system (97.5 KIAS). The model 208B airplane with the TKS installed, uses vortex generators (VG's) installed on the wings to improve air flow on control surfaces. The model 208B and model 208 with the TKS installed, also have VG's installed on the flaps. The equipment package has included a standby electrical system that uses a 75-amp alternator, a control surface mounted static discharge wicks, and an ice detection light for detection of ice buildup on the leading edge of the left wing at night. Refer to Chapter 12, Replenishing Description and Operation.
B.
For information on the following items, refer to the specified locations. (1) Wing, Wing Strut, Horizontal and Vertical Stabilizers Deice System, refer to Pneumatic Surface Deice - Description and Operation or to TKS Anti-Ice - Description and Operation. (2) Windshield Anti-ice System, refer to Windshield Anti-Ice - Description and Operation or to TKS Anti-Ice - Description and Operation. (3) Propeller Anti-ice System, refer to Propeller Anti-Ice - Description and Operation or to TKS AntiIce-Description and Operation. (4) Pitot and Static Heaters, refer to Chapter 34, Pitot/Static System - Description and Operation. (5) Icing Low Airspeed Awareness System (110 KIAS), refer to Chapter 34, Pitot/Static System Description and Operation. (6) Ice Detector Light, refer to Chapter 33, Ice Detector Light - Maintenance Practices. (7) Electrostatic Discharger Installation, refer to Chapter 23, Static Discharging- Maintenance Practices. (8) Standby Electrical System, Refer to Chapter 24, Standby Electrical System - Description and Operation.
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INDICATING/ RECORDING SYSTEMS
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MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
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31-10-00
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Mar 1/1999
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31-Title 31-List of Effective Pages 31-Record of Temporary Revisions 31-Table of Contents
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS INDICATING/RECORDING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31-00-00 Page 1 31-00-00 Page 1 31-00-00 Page 1 31-00-00 Page 1
INSTRUMENT AND CONTROL PANELS - MAINTENANCE PRACTICES. . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Control Panels Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lower Left Switch Panel Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31-10-00 Page 201 31-10-00 Page 201 31-10-00 Page 201 31-10-00 Page 201
HOURMETER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hourmeter Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hourmeter Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31-20-00 Page 201 31-20-00 Page 201 31-20-00 Page 201 31-20-00 Page 201 31-20-00 Page 201
CLOCK - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Standard Clock Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Optional (Astro Tech) Clock Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31-21-00 Page 201 31-21-00 Page 201 31-21-00 Page 201 31-21-00 Page 201
L-3 COMMUNICATIONS F1000 FLIGHT DATA RECORDER SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Data Recorder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Impact Switch (5G) Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Accelerometer Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Potentiometer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FDR Buffer/Amp Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31-30-00 Page 201 31-30-00 Page 201 31-30-00 Page 201 31-30-00 Page 201 31-30-00 Page 201 31-30-00 Page 204 31-30-00 Page 204
L-3 COMMUNICATIONS F1000 FLIGHT DATA RECORDER SYSTEM - ADJUSTMENT/ TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Data Recorder Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Underwater Locator Device (ULD) Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31-30-00 Page 501 31-30-00 Page 501 31-30-00 Page 501 31-30-00 Page 505
L-3 COMMUNICATIONS FA2100 COCKPIT VOICE/DATA RECORDER SYSTEM DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31-31-00 Page 1 31-31-00 Page 1 31-31-00 Page 1
L-3 COMMUNICATIONS FA2100 COCKPIT VOICE/DATA RECORDER SYSTEM MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cockpit Voice/Data Recorder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Impact Switch (5G) Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Accelerometer Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Potentiometer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CVR Adapter Assembly (Sum Amplifiers) Removal/Installation . . . . . . . . . . . . . . . . . . Area Microphone Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MASTER WARNING AND ANNUNCIATOR PANEL - DESCRIPTION AND OPERATION General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL MASTER WARNING AND ANNUNCIATOR PANEL - MAINTENANCE PRACTICES. . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Annunciator Panel Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Annunciator Panel Lamp(s) Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Annunciator Indicator Light Lens Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . Annunciator Indicator Light Housing Assembly Removal/Installation . . . . . . . . . . . . . Day/Night Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lamp Test Switch Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fire Detector Test Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL INDICATING/RECORDING - GENERAL 1.
Scope A.
2.
This chapter provides information on the annunciator panel and instrument panel switches.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Multimeter
Model 260 or equivalent
Simpson Electric Co. 853 Dundee Ave. Chicago, IL 60120
To check voltage and continuity in the electrical circuits.
3.
Definition A.
This chapter is divided into sections and subsections to assist maintenance personnel in locating specific components and information. For locating specific information within the chapter, refer to the Table of Contents at the beginning of the chapter.
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MODEL 208 MAINTENANCE MANUAL INSTRUMENT AND CONTROL PANELS - MAINTENANCE PRACTICES 1.
General A.
2.
The instrument panels discussed in this section includes the left and right flight control panels and lower left switch panel. The left and right flight control panels and the light, de-ice/anti-ice switch panel are located on the main instrument panel. The left and right flight control panel(s) are attached to panel slides and are designed to slide aft and out of the instrument panel, then tilt for maintenance to be performed without completely removing the panels.
Flight Control Panels Removal/Installation A.
Remove Flight Control Panels (Refer to Figure 201). (1) Ensure all electrical power is off.
CAUTION: Do not pull panel to the point where there is unnecessary strain on wiring and hoses connected to the instruments. (2) (3) (4) (5) (6) B.
3.
Remove screws (7) securing flight control panels and pull panel aft. Tag and disconnect electrical connectors, hoses and associated wiring from components as required to allow panels to be pulled aft far enough for maintenance work. If panels are to be completely removed from airplane, tag and disconnect all electrical connectors, hoses and associated wiring from instruments, postlights, and switches. On left flight control panel (1) only, remove screws attaching control column collar (4) to left flight control panel. Remove cotter pins (9) and pins (6) attaching panels to panel slides (10) and remove from airplane.
Install Flight Control Panels (Refer to Figure 201). (1) If panel(s) have been completely removed from airplane, position flight control panel(s) and attach to panel slides (10) using pins (6) and cotter pins (9). (2) Connect hoses, electrical connectors, and associated wiring to instruments, postlights, and switches. Remove all tags. (3) Push panel(s) forward and secure with screws (7). (4) On left flight control panel, position control column collar (4) and secure with screws.
Lower Left Switch Panel Removal/Installation A.
Remove Lower Left Switch Panel (Refer to Figure 201). (1) Ensure all electrical power is off. (2) Remove screws attaching knee bumper to stationary instrument panel. (3) Remove alternate static source valve. (4) Remove screws attaching switch panel to stationary instrument panel. (5) Disconnect electrical wires from postlights and remove postlights. (6) Pull switch panel aft as necessary to perform required maintenance.
B.
Install Lower Left Switch Panel (Refer to Figure 201). (1) Position switch panel, install postlights and connect electrical wires to postlights. (2) Install screws securing switch panel to stationary instrument panel. (3) Install alternate static source valve. (4) Install knee bumper.
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Instrument Panel Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL HOURMETER - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
4.
An oil pressure type hourmeter is optional equipment on Airplanes 20800001 thru 20800055, and 208B0001 thru 208B0087. Airplanes 20800056 and On, and 20800001 thru 20800055 incorporating SK208-9, and 208B0087 and On, have an airflow type hourmeter installed.
The oil pressure type hourmeter is activated by an electrical signal from engine oil pressure switch. Airflow type hourmeter operates when air flows across a switch tab located on wing, which closes an electrical circuit activating hourmeter.
Hourmeter Removal/Installation A.
Remove Hourmeter (Refer to Figure 201). (1) Disconnect airplane battery. (2) Disconnect electrical leads (2) and (3) from hourmeter (1). (3) Remove screws and remove hourmeter.
B.
Install Hourmeter (Refer to Figure 201). (1) Connect electrical leads (2) and (3) to hourmeter (1). (2) Position hourmeter in instrument panel and install screws. (3) Connect airplane battery.
Hourmeter Switch Removal/Installation A.
Remove Hourmeter Switch (Refer to Figure 201). (1) Disconnect airplane battery. (2) Remove screws securing cover plate (11). (3) Remove screws (7) and washers (6) from wire (2) and ground line (8). (4) Remove nuts (4), washers (5), and screws (9) and remove switch from cover plate.
B.
Install Hourmeter Switch (Refer to Figure 201). (1) Remove and discard spring from hourmeter switch (10). NOTE: (2) (3) (4) (5)
Hourmeter switch (10) will not operate without removing the spring prior to installation.
Install switch (10) on cover plate (11) using screws (9), washers (5), and nuts (4). Connect wire (2) and ground wire (8) to switch (10) using washers (6) and screws (7). Install cover plate (11) on wing using screws. Connect airplane battery.
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Hourmeter Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL CLOCK - MAINTENANCE PRACTICES 1.
General A.
2.
The 12 hour clock is located on the upper left corner of the instrument panel. It is electrically operated and consists of an hour hand, minute hand, and second hand. The Astro Tech LC-2 Quartz Chronometer is offered as optional equipment. It is a precision, solid-state time keeping device which will display to the pilot the time-of-day, the calendar date, and the elapsed time interval between a series of selected events, such as in flight check points or legs of a cross country flight, etc. These three modes of operation function independently and can be alternately selected for viewing on the four digit liquid crystal display (LCD) on the front face of the instrument, three pushbutton type switches directly below the display control all time keeping functions. The digital display features an internal light (back light) to ensure good visibility under low cabin lighting conditions or at night. The intensity of the backlight is controlled by the light rheostat. In addition, the display incorporates a test function which allows checking that all elements of the display are operating. To activate the test function, press the LH and RN buttons at the same time. Refer to Pilot's Operating Handbook for operation.
Standard Clock Removal/Installation A.
Remove Standard Clock (Refer to Figure 201). (1) Ensure all electrical power is off. (2) Tag and disconnect electrical wires from back of clock. (3) Remove screws (1) and remove clock.
B.
Install Standard Clock (Refer to Figure 201). (1) Position clock and install screws (1).
CAUTION: Power wire must be connected to "+" terminal of clock or internal damage to unit will result. (2) 3.
Connect electrical wires to clock. Remove tags.
Optional (Astro Tech) Clock Removal/Installation A.
Remove Optional Clock (Refer to Figure 201). (1) Ensure all electrical power is off. (2) Disconnect clock wires by disconnecting three pin housing. (3) Remove screws (1) and remove clock.
B.
Install Optional Clock (Refer to Figure 201). (1) Position clock and install screws (1). (2) Connect electrical three pin housing. (3) Set clock in accordance with Pilot's Operating Handbook.
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Clock Installation Figure 201 (Sheet 1)
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Clock Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL L-3 COMMUNICATIONS F1000 FLIGHT DATA RECORDER SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
The L-3 Communications F1000 Flight Data Recorder (FDR) System consists of a solid state data recorder, impact switch (5G), accelerometer, FDR buffer/amp and a potentiometer. Maintenance practices consists of removal and installation of the FDR and it's components.
Flight Data Recorder Removal/Installation A.
Remove Flight Data Recorder (Refer to Figure 201). (1) Remove electrical power from the airplane. (2) Disengage the FDR circuit breaker on the left circuit breaker panel. NOTE: (3) (4) (5) (6) (7)
B.
3.
The FDR must be removed from the aft side of the avionics shelf.
In the tailcone, on the avionics shelf cut and remove the safety wire between the knurl nuts located on the mounting tray. Loosen the knurl nuts and unlatching locking mechanism located on the mounting tray. Disconnect the pitot/static quick-disconnects from the FDR. Carefully move the flight data recorder out of the mounting tray. Remove the FDR from the airplane.
Install Flight Data Recorder (Refer to Figure 201). (1) From the aft side of the avionics shelf, put the flight data recorder on the mounting tray and slide in until the electrical connectors are firmly engaged. (2) Connect the pitot/static lines to the FDR. (3) Tighten and safety wire the knurl nuts together on the mounting tray. Refer to Chapter 20, Safetying - Maintenance Practices. (4) Engage the FDR circuit breaker on the left circuit breaker panel. (5) Restore electrical power to the airplane.
Impact Switch (5G) Removal/Installation A.
Remove Impact Switch (5G) (Refer to Figure 201). (1) Remove electrical power from the airplane.
CAUTION: The CVR and FDR are linked to the same impact switch circuit. If one circuit breaker is disengaged, the other circuit breaker must be disengaged. (2) (3) (4) (5) B.
4.
Disengage the FDR circuit breaker and if installed, the CVR circuit breaker on the left circuit breaker panel. In the tailcone, on the avionics shelf, disconnect the electrical connector from the impact switch (5G). Remove the screws from the impact switch (5G). Remove the impact switch (5G) from the airplane.
Install Impact Switch (5G) (Refer to Figure 201). (1) Put the impact switch (5G) in position at the aft avionics shelf and install with screws. (2) Connect the electrical connector to the impact switch (5G). (3) Engage the FDR circuit breaker on the left circuit breaker panel. (4) Restore electrical power to the airplane.
Accelerometer Removal/Installation A.
Remove Accelerometer (Refer to Figure 201). (1) Remove electrical power from the airplane. (2) Disengage FDR circuit breaker, located on the left circuit breaker panel. (3) At FS 263.96 and RBL 5.00, remove the fasteners from the hard shell headliner panel.
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F1000 Flight Data Recorder Installation Figure 201 (Sheet 1)
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F1000 Flight Data Recorder Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (4) (5) (6) (7) B.
5.
6.
Remove the hard shell headliner from the airplane. Disconnect the electrical connector from the accelerometer. Remove the screws from the accelerometer bracket assembly. Remove accelerometer from the airplane.
Install Accelerometer (Refer to Figure 201). (1) Put the accelerometer in position and secure to bracket assembly with screws. (2) Connect the electrical connector to the accelerometer. (3) Put the hard shell headliner panel into position and install using fasteners. (4) Engage FDR circuit breaker on left circuit breaker panel. (5) Restore electrical power to the airplane.
Potentiometer Removal/Installation A.
Remove Potentiometer (Refer to Figure 201). (1) Remove electrical power from the airplane. (2) Disengage FDR circuit breaker on left circuit breaker panel. (3) At FS 263.96 and RBL 5.00, remove the fasteners from the hard shell headliner panel. (4) Remove the hard shell headliner panel from the airplane. (5) Disconnect the electrical connector to the potentiometer. (6) Loosen the set screws that attach the lever assembly to the potentiometer shaft. (7) Loosen screws that hold the potentiometer to the bracket assembly. (8) Turn the cleats so that the potentiometer can be removed from the bottom of the bracket assembly (9) Remove the potentiometer from the airplane.
B.
Install Potentiometer (Refer to Figure 201). (1) Make sure the flaps are retracted. (2) Rotate the potentiometer shaft counter-clockwise as viewed from above. (3) Put the potentiometer in the bracket assembly. (4) Tighten cleats only enough to hold in place. (5) Put the lever assembly on the potentiometer shaft and tighten the set screws. (6) Tighten the set screws that attach the potentiometer to the lever assembly. (7) With the cleats loose, turn the potentiometer to obtain a resistance of 4.00 kΩ, + or - .10 kΩ between terminals 2 and 3. (8) Tighten the cleats. (9) Cycle the flaps to make sure the system operates without binding or interference. (10) Put the hard shell headliner panel into position and install using fasteners. (11) Engage the FDR circuit breaker, located on the left circuit breaker panel. (12) Restore electrical power to the airplane.
FDR Buffer/Amp Removal/Installation A.
Remove FDR Buffer/Amp (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5)
The FDR Buffer/Amp is located on the forward side of the pedestal.
Remove electrical power from the airplane. Disengage the FDR and CVR circuit breaker on the left circuit breaker panel. Remove electrical connectors from the CVR adapters. Remove the electrical connector from the FDR buffer/amp. Remove the screws from both CVR adapters located at the FDR buffer/amp bracket.
CAUTION: The internal components of the FDR buffer/amp are mounted to the lid of the FDR buffer/amp mounting box. Use caution when removing not to damage the mounted electronic components. (6) (7)
Remove the screws from the lid on the FDR buffer/amp. Remove the FDR buffer/amp from the airplane.
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Install FDR Buffer/Amp (Refer to Figure 201). (1) Put the FDR buffer/amp lid on the mounting box and secure using screws. (2) Connect the electrical connector to the FDR buffer/amp. (3) Attach the CVR adapters to the FDR buffer/amp mounting bracket using screws. (4) Connect the electrical connector to each CVR adapter with screws.
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MODEL 208 MAINTENANCE MANUAL L-3 COMMUNICATIONS F1000 FLIGHT DATA RECORDER SYSTEM - ADJUSTMENT/TEST 1.
General A.
2.
This section provides information to perform an operational check of the L-3 Communications Flight Data Recorder.
Flight Data Recorder Operational Check A.
Required Equipment. (1) IBM compatible personal laptop computer with Read Out Support Equipment, (ROSE) software version 3.0 or higher installed. (2) Quatech PCMCIA card installed with Quatech PCMCIA card-to-15 pin adapter cable. (3) 15 socket-25 pin FDR extraction cable. (4) Gyro tilt table with mount for KI 256. (5) Extender Cable for KI 256. (6) Inclinometer with current calibration. (7) Pitot Static Test Set, (LAVERSAB Model 6520 or equivalent). (8) External Vacuum Source for KI 256 Gyro. (9) Pinglite Model PL-3 Tester. (10) Ground Power Unit (GPU).
B.
Perform an Operational Test. (1) Apply electrical power to the airplane. (2) Engage the FDR circuit breaker, located on the left circuit breaker panel. (a) Make sure the FDR FAIL annunciator goes off. (3) Disengage the FDR circuit breaker, located on the left circuit breaker panel. (a) Make sure the FDR FAIL annunciator comes on. (4) Turn the BUS 2 PWR, AVN BUS 1 and AVN BUS 2 switches to OFF. (a) Make sure the FDR receives power from BUS 1 PWR bus. (5) Put the impact switch (5G) to the OPEN position. (a) Make sure the impact switch lamp comes on. (b) Make sure there is no power to the FDR. NOTE:
(6) (7) (8)
Turn on the IBM compatible laptop computer. (a) Make sure the laptop computer has the PCMCIA card installed and is connected to a 15 pin adapter cable. Connect the adapter cable to the 15 socket-25 pin FDR extraction cable. Attach the 25 pin FDR extraction cable to the connector on the face of the FDR. NOTE:
(9)
Manual reset of the impact switch's reset switch will cause the switched output to be turned on and the impact switch lamp to go off. Upon impact, the impact will cause the output to be turned off and the impact switch lamp to come on.
After the 25 pin FDR extraction cable is connected, system will cease recording and FDR FAIL annunciator will come on.
Run the ROSE software. Refer to ROSE operators manual. NOTE:
During testing, the correct aircraft configuration file corresponding to the aircraft being monitored must be loaded into ROSE.
(10) Set the aircraft to the following configuration on the laptop. (a) Report/Analyze/Display Data. (b) Maintenance Quick Check - 208 Caravan. (11) Choose OK. (12) Select Monitor Data Display.
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MODEL 208 MAINTENANCE MANUAL (13) Choose OK. NOTE:
After selecting OK, a prompt asks if you want to create a Monitor Data Disk File. Choose NO and the ROSE Report Display Screen will appear.
(14) Set the pilot's Baro Set to 29.92. (15) Adjust Pitot/Static Test set pressure for each test point specified in Table 501. Table 501. Pressure Altitude REFERENCE ALTITUDE FEET
FDR READ OUT ALTITUDE FEET
PILOT'S ALTITUDE FEET
FDR SYSTEM ERROR FEET
ALLOWABLE RECORDED ERROR FEET
-1,000
+100 or -100
0
+100 or -100
2,000
+100 or -100
5,000
+100 or -100
10,000
+150 or -150
15,000
+150 or -150
20,000
+300 or -300
25,000
+300 or -300
26,000
+300 or -300
25,000
+300 or -300
20,000
+300 or -300
15,000
+150 or -150
10,000
+150 or -150
5,000
+100 or -100
2,000
+100 or -100
0
+100 or -100
-1,000
+100 or -100 (16) Adjust the Pitot/Static Test set airspeed for each test point specified in Table 502.
Table 502. Airspeed REFERENCE AIRSPEED KNOTS
FDR READ OUT AIRSPEED KNOTS
PILOT'S AIRSPEED KNOTS
FDR SYSTEM ERROR KNOTS
ALLOWABLE RECORDED ERROR KNOTS
70
+10 or -10
80
+10 or -10
90
+10 or -10
100
+10 or -10
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MODEL 208 MAINTENANCE MANUAL Table 502. Airspeed (continued) REFERENCE AIRSPEED KNOTS
FDR READ OUT AIRSPEED KNOTS
PILOT'S AIRSPEED KNOTS
FDR SYSTEM ERROR KNOTS
ALLOWABLE RECORDED ERROR KNOTS
110
+10 or -10
120
+10 or -10
130
+10 or -10
140
+10 or -10
150
+10 or -10
160
+10 or -10
172
+10 or -10
180
+10 or -10 (17) Manually slew the heading card on the pilot's HSI per each test point in Table 503.
Table 503. Magnetic Heading REFERENCE HEADING DEGREES
FDR READ OUT HEADING DEGREES
PILOT'S HEADING DEGREES
FDR SYSTEM ERROR DEGREES
ALLOWABLE RECORDED ERROR DEGREES
360
+2 or -2
045
+2 or -2
135
+2 or -2
180
+2 or -2
225
+2 or -2
270
+2 or -2
315
+2 or -2 (18) Move the accelerometer momentarily so that UP/DWN arrow is pointed down. (a) Monitor the PC laptop for vertical acceleration data. (b) Make sure the ROSE report display reading is within +0.6 or -0.6 G's. (19) Put the UP/DWN arrow in a horizontal attitude. (a) Monitor the PC laptop for vertical acceleration data. (b) Make sure the ROSE report display reading is within +0.6 or -0.6 G's. (20) Make sure the UP/DWN arrow is pointing up. (a) Monitor the PC laptop for vertical acceleration data. (b) Make sure the ROSE report display reading is within +0.6 or -0.6 G's. (21) Mount KI 256 Gyro on tilt table. (22) Make sure the pitch angles correspond with Table 504.
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Table 504. Pitch Angle ANGLE
DEGREES
REFERENCE PITCH UP/DOWN
FDR READ OUT
DEGREES
DEGREES
FDR SYSTEM ERROR DEGREES
ALLOWABLE RECORDED ERROR DEGREES
-30
+2 or -2
-20
+2 or -2
-10
+2 or -2
0
+2 or -2
+10
+2 or -2
+20
+2 or -2
+30
+2 or -2
NOTE 1: Pitch down is represented by a negative number on the ROSE report display. Pitch up is represented by a positive number. (23) Make sure the roll angles correspond with Table 505. Table 505. Roll Angle ANGLE
DEGREES
REFERENCE ROLL LEFT/RIGHT
FDR READ OUT
DEGREES
DEGREES
FDR SYSTEM ERROR DEGREES
ALLOWABLE RECORDED ERROR DEGREES
30L
+2 or -2
20L
+2 or -2
10L
+2 or -2
0
+2 or -2
10R
+2 or -2
20R
+2 or -2
30R
+2 or -2
NOTE 1: Roll left is represented by a negative number on the ROSE report display. Roll right is represented by a positive number. (24) Put the flaps into the positions of 0 degrees, 10 degrees, 20 degrees and 30 degrees and make sure each position displays on the ROSE report display. (25) Depress the pilot, copilot and hand mic transmit buttons. Make sure there is a change of status displayed on the ROSE report display. (26) Engage and disengage the autopilot and make sure a change of status is displayed on the ROSE report display. (27) Push the FIRE DETECT test switch on the instrument panel and make sure a change of status is displayed on the ROSE report display. (28) Run engine per each test point specified in Table 506.
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Table 506. Engine Torque REFERENCE TORQUE FOOT-POUNDS
FDR READ OUT TORQUE FOOT-POUNDS
PILOT'S TORQUE FOOT-POUNDS
FDR SYSTEM ERROR FOOTPOUNDS
ALLOWABLE RECORDED ERROR FOOT-POUNDS
600
30
800
40
1000
50
1200
60
1400
70
1600
80
1800
90
1865
93.25 (29) Run engine per each test point specified in Table 207.
Table 507. Propeller RPM REFERENCE PROPELLER RPM
FDR READ OUT PROPELLER RPM
PILOT'S PROPELLER RPM
FDR SYSTEM ERROR RPM
ALLOWABLE RECORDED ERROR RPM
1400
70
1500
75
1600
80
1700
85
1800
90
1900
95
3.
Underwater Locator Device (ULD) Test A.
Do an Underwater Locator Device (ULD) Test. (1) Remove the Velcro collar from the Pinglite Model PL-3 Tester. (2) Apply the spring end of the Pinglite Model PL-3 Tester firmly to the ULD. NOTE: (3)
The ULD is physically attached to the Locator Beacon.
Make sure the center spring is making contact with the center pin of the water switch. (a) With the center pin shorted to the center pin of the water switch, depress and hold the remaining three springs of the Pinglite PL-3 Tester against the ULD body. (b) Listen for audible pinging and make sure the LED flashes with each output pulse indicating its operation. (c) If audible pinging is not heard, or the LED does not flash, remove the FDR and return to the Radio Lab for retest and/or rejection of the Underwater Acoustic Beacon.
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MODEL 208 MAINTENANCE MANUAL L-3 COMMUNICATIONS FA2100 COCKPIT VOICE/DATA RECORDER SYSTEM - DESCRIPTION AND OPERATION 1.
2.
Description A.
The L3 Communications FA2100 Cockpit Voice/Data Recorder (CV/DR) System has a Cockpit Voice Recorder (CVR) that records a minimum of 120 minutes of cockpit audio and a Flight Data Recorder (FDR) that can record airframe and engine parameters for a minimum of 25 hours. (1) The CVR can simultaneously record the four cockpit audio input streams and convert them to a digital format. (2) The four audio streams channels are as follows: (a) The Channel 1 input is the from the speaker audio (b) The Channel 2 input is the from the copilot audio (c) The Channel 3 input is the from the pilot audio (d) The Channel 4 input is the from the area microphone audio. (3) The FDR records data from the DCU from an ARINC 717 bus at 256 words per second. (a) The FDR has storage for a minimum of 25 hours of flight data. (b) It takes about five minutes or less to download the flight data. (c) The FDR unit is in an international orange case. (d) The FDR has a Crash Survivable Memory Unit (CSMU). 1 This unit has the solid state flash memory for the recording medium. 2 The data is separate from the voice data the CVR records.
B.
The L3 Communications FA2100 CV/DR has the system components that follow: (1) The Data Collector Unit (DCU) (GSD 41) is in the tailcone between FS 356.99 and FS 384.22 and WL 93.52. (2) The recorder unit is in the tailcone section on the avionics shelf between FS 370.80 and FS 386.35 and WL 127.03. (3) The underwater locator device (ULD) is on the front panel of the recorder unit. (a) The ULD also has a battery installed with the device. (4) The impact switch is installed below the avionics shelf below the CV/DR in the tailcone. (5) The three axis accelerometer is installed at FS 210.84, WL 132.16, and LBL 29.23 (6) The area microphone is above the audio panel on the instrument panel at FS 118.00 and WL 123.06. (7) The flap position sensor potentiometer (RC300) is installed at FS 210.84 and WL 136.92. (8) The CVR adapters (sum amplifiers) are at the forward side of the pedestal at FS 114.40 and WL 95.00. (9) There is a Ground Support Equipment (GSE) connector under the front panel of the CV/DR to facilitate maintenance functions.
Operation A.
The CV/DR receives electrical power from the CV/DR circuit breaker on the left circuit breaker panel. (1) When the CV/DR receives electrical power a built-in-test (BIT) is initiated. (a) The CV/DR continuously monitors its serviceability after the initial BIT is complete. (b) If there is a CVR failure a discrete is sent to the Integrated Avionics Unit I (IAU) and then to the PFD that shows a CVR FAIL (white) Crew Alert Message (CAS) on the display. (c) The CV/DR CVR receives voice data from the CVR adapter unit (summing amplifier). (d) The CV/DR FDR receives data from the DCU on an ARINC 717 bus. (e) If there is a FDR failure a discrete is sent to the DCU that then sends the data on a ARINC 429 bus to the PFD. The PFD shows a FDR FAIL (white) CAS message on the display. (2) If there is a FDR failure the CVR continues to record cockpit voice data correctly. NOTE:
The FDR and CVR data is recorded separately in the CV/DR. Failure of one of these two CV/DR functions does not cause an interrupt of the other function and correct operation continues.
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MODEL 208 MAINTENANCE MANUAL B.
The DCU processes Models 208/208B system parameters for the CV/DR FDR. (1) The DCU receives electrical power from the DCU circuit breaker on the avionics circuit breaker panel. (2) The DCU acquires the data on a ARINC 429 bus from the IAU 1 and sends the data on a ARINC 717 bus to the CV/DR. (3) The GEA 71 Engine/Airframe Unit monitors data from the for flap position sensor, the three axis accelerometer and from airplane airframe and engine and sends the signals to IAU 1 and IAU 2 on an ARINC 485 bus.
C.
Underwater Locator Device (ULD) (1) The ULD is also referred to as an underwater acoustic beacon. (2) The ULD has a battery that has an expected life of six years. (3) Check with the ULD supplier for the data that follows: • Servicing • Recertification • Correct procedures to handle the lithium batteries • Correct procedures to dispose of the lithium batteries.
D.
Impact Switch (1) The impact switch disconnects electrical power from the CV/DR if there is a crash with a G-force that is more than 5 Gs. The recorded CV/DR data can automatically erase if the recorder was to continue to operate.
E.
3-Axis Accelerometer (1) The 3-axis accelerometer is a hermetically sealed instrument which simultaneously measures acceleration along three axes, vertical, lateral and longitudinal. (2) The accelerometer receives electrical power from the CVDR circuit breaker on the left circuit breaker panel. (3) The accelerometer sends data signals to the GEA 71 for recording by the CV/DR FDR.
F.
The Flap Position Sensor Potentiometer (1) The flap position sensor gives flap position data to the GEA 71 for recording by the CV/DR FDR.
G.
The Area Microphone (1) The area microphone can pick up and record cockpit voice signals. (2) The area microphone is installed in the top area of the instrument panel.
H.
CVR Adapters (1) The CVR adapters are sum amplifiers for the pilot (CVR channel 3) and copilot (CVR channel 2). (2) Channel 3 and Channel 4 voice signals are sent to the CV/DR.
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MODEL 208 MAINTENANCE MANUAL L-3 COMMUNICATIONS FA2100 COCKPIT VOICE/DATA RECORDER SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
The L-3 Communications FA21000 Cockpit Voice/Data Recorder (CV/DR) System consists of a solid state Cockpit Voice Recorder (CVR), Flight Data Recorder (FDR), impact switch (5G), 3-axis accelerometer, CVR adapter, and a flap position potentiometer. Maintenance practices consists of removal and installation of the CV/DR and it's components.
Cockpit Voice/Data Recorder Removal/Installation A.
Remove the Cockpit Voice/Data Recorder (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Disengage the CVDR circuit breaker on the left circuit breaker panel. (3) Disengage the DCU circuit breaker on the avionics circuit breaker panel. (4) In the tailcone, between FS 370.80 and FS 386.35 on the avionics shelf cut and remove the safety wire between the CV/DR knurl nuts. (5) Loosen the knurl nuts and unlatching locking mechanism located on the mounting tray. (6) Carefully move the CV/DR out of the mounting tray. NOTE: (7)
B.
3.
4.
The CV/DR must be removed from the aft side of the avionics shelf.
Remove the CV/DR from the airplane.
Install the Cockpit Voice/Data Recorder (Refer to Figure 201). (1) From the aft side of the avionics shelf, put the flight data recorder on the mounting tray and slide in until the electrical connectors are firmly engaged. (2) Tighten and safety wire the knurl nuts together on the mounting tray. Refer to Chapter 20, Safetying - Maintenance Practices. (3) Engage the CVDR circuit breaker on the left circuit breaker panel.
Impact Switch (5G) Removal/Installation A.
Remove the Impact Switch (5G) (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Disengage the CVDR circuit breaker on the left circuit breaker panel. (3) Disengage the DCU circuit breaker on the avionics circuit breaker panel. (4) At the impact switch (5G) in the tailcone between FS 370.80 and FS 386.35, below the CV/DR avionics shelf, disconnect the electrical connector (PT305). (5) Remove the screws from the impact switch (5G). (6) Remove the impact switch (5G) from the airplane.
B.
Install the Impact Switch (5G) (Refer to Figure 201). (1) Put the impact switch (5G) in its correct position below the aft avionics shelf and install with screws. (2) Attach the impact switch shelf with the screws. (3) Connect the electrical connector to the impact switch (5G). (4) Engage the CVDR circuit breaker on the left circuit breaker panel. (5) Engage the DCU circuit breaker on the avionics circuit breaker panel.
Accelerometer Removal/Installation A.
Remove the Accelerometer (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Disengage the CVDR circuit breaker on the left circuit breaker panel. (3) Disengage the DCU circuit breaker on the avionics circuit breaker panel. (4) At FS 210.84, WL 132.16, and LBL 29.23, remove the fasteners from the hard shell headliner panel. (5) Remove the hard shell headliner from the airplane.
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MODEL 208 MAINTENANCE MANUAL
FA21000 Cockpit Voice Data Recorder Components Installation Figure 201 (Sheet 1)
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FA21000 Cockpit Voice Data Recorder Components Installation Figure 201 (Sheet 2)
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FA21000 Cockpit Voice Data Recorder Components Installation Figure 201 (Sheet 3)
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FA21000 Cockpit Voice Data Recorder Components Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (6) (7) (8) B.
5.
Disconnect the electrical connector (PC302) from the accelerometer. Remove the screws from the accelerometer bracket assembly. Remove the accelerometer from the airplane.
Install the Accelerometer (Refer to Figure 201). (1) Put the accelerometer in its correct position on the bracket assembly and secure to bracket assembly with screws. (2) Attach the accelerometer on the bracket assembly with screws. (3) Connect the electrical connector to the accelerometer. (4) Put the hard shell headliner panel into position and install using fasteners. (5) Engage the CVDR circuit breaker on the left circuit breaker panel. (6) Engage the DCU circuit breaker on the avionics circuit breaker panel.
Potentiometer Removal/Installation A.
Remove the Potentiometer (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Disengage the CVDR circuit breaker on the left circuit breaker panel. (3) Disengage the DCU circuit breaker on the avionics circuit breaker panel. (4) At FS 210.84, WL 136.92, AND LBL 14.23, remove the fasteners from the hard shell headliner panel. (5) Remove the hard shell headliner panel from the airplane. (6) Disconnect the electrical connector (PC303) from the potentiometer. (7) Loosen the set screws that attach the lever assembly to the potentiometer shaft. (8) Loosen screws that hold the potentiometer to the bracket assembly. (9) Turn the cleats so that the potentiometer can be removed from the bottom of the bracket assembly (10) Remove the potentiometer from the airplane.
B.
Install Potentiometer (Refer to Figure 201). (1) Make sure that the flaps are rigged. Refer to Chapter 27, Flap Rigging Guide- Adjustment/Test. (2) Attach an inclinometer to one of the two flaps at WS 68.00 approximately 34 inches from the inboard edge of the flap. (3) Put the potentiometer in the bracket assembly. (4) Tighten cleats only enough to hold in place. (5) Put the lever assembly on the potentiometer shaft and tighten the set screws. (6) Tighten the set screws that attach the potentiometer to the lever assembly. (7) Apply external electrical power to the airplane. (8) Put the AVIONICS 1 and AVIONICS 2 switches on the left switch panel to the ion position. (9) At Primary Flight Display 1 make sure that system software version 767.05 or later is installed. (10) With the flaps in the up position, set the inclinometer to 0 degrees. (11) Set the flap lever to 20 degrees. (12) Push and hold the ENT button on the PDF 1 display and at the same time push in the PDF1 circuit breaker on the avionics circuit breaker panel. NOTE: (13) (14) (15) (16)
You can release the ENT button when the INITIALIZING SYSTEM message shows on the display.
Use the outer FMS knob to scroll to the GEA page. Use the inner FMS knob to scroll to the GEA STATUS page. Push the ANLG softkey. Locate the channel 3A FLAPS POSITION field. NOTE:
This shows the flap position value sensed by the flap position potentiometer.
(17) Loosen the flap position potentiometer cleats. (a) Turn the potentiometer until the value in the FLAPS POSITION field is the same as the inclinometer (0 degrees + .3 or -.3 degrees). (b) Carefully tighten the potentiometer cleats and make sure the potentiometer does not turn.
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MODEL 208 MAINTENANCE MANUAL (18) Cycle the flaps and make sure that they operate with not interference. (19) Make sure that the PDF1 FLAP POSITION field value shows the same as the inclinometer at the positions that follow: (a) UP (b) 10 degees (c) 20 degrees (d) UP. (20) Remove the inclinometer from the flap. (21) Put the AVIONICS 1 and AVIONICS 2 switches on the left switch to the OFF position. (22) Remove external electrical power from the airplane. 6.
CVR Adapter Assembly (Sum Amplifiers) Removal/Installation A.
Remove the CVR Adapters (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6)
B.
7.
8.
CVR Adapters are located on the forward side of the pedestal. The removal of the pilot and copilot adapters are identical.
Remove external electrical power from the airplane. Disengage the CVDR circuit breaker on the avionics circuit breaker panel. Disengage the DCU circuit breaker on the left circuit breaker panel. Remove electrical connector (PI528 pilot, PI529 copilot) from the applicable CVR adapter. Remove the screws and washers from the CVR adapter bracket. Remove the CVR adapter from the airplane.
Install the CVR Adapters (Refer to Figure 201). (1) Put the CVR adapter in its correct position on the CVR adapter bracket. (2) Attach the CVR adapter to the bracket with screws and washers. (3) Connect the electrical connector to the CVR adapter. (4) Engage the CVDR circuit breaker on the avionics circuit breaker panel. (5) Engage the DCU circuit breaker on the left circuit breaker panel.
Area Microphone Removal/Installation A.
Remove the Area Microphone (Refer to Figure 201). (1) Remove external electrical power from the airplane. (2) Disengage the CVDR circuit breaker on the avionics circuit breaker panel. (3) Disengage the DCU circuit breaker on the left circuit breaker panel. (4) Remove instrument panel components if necessary to get access to the area microphone. (a) If necessary, remove the primary flight display or multi-fuction display. Refer to Chapter 34, Garmin Display Unit (GDU) - Maintenance Practices. (5) Disconnect the electrical connector (JI527) from the area microphone. (6) Remove the screws that attach the area microphone to the bracket. (7) Remove the area microphone from the airplane.
B.
Install the Area Microphone (Refer to Figure 201). (1) Put the area microphone in its correct position on the bracket. (2) Install the screws that attach the area microphone to the bracket. (3) Connect the electrical connector to the area microphone. (4) If removed, install instrument panel components. (a) If removed, install the primary flight display or multi-function display. Refer to Chapter 34, Garmin Display Unit (GDU) - Maintenance Practices. (5) Engage the CVDR circuit breaker on the avionics circuit breaker panel. (6) Engage the DCU circuit breaker on the left circuit breaker panel.
Maintenance Practices A.
For maintenance practices related to the CV/DR, refer to the manufacturer.
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MODEL 208 MAINTENANCE MANUAL MASTER WARNING AND ANNUNCIATOR PANEL - DESCRIPTION AND OPERATION 1.
General A.
2.
The annunciator panel is located in the upper portion of the instrument panel and provides emergency and normal operational information to the flight crew.
Description and Operation A.
The annunciator panel is has three different colored lenses. (1) The red lenses include the ENGINE FIRE, OIL PRESS LOW, GENERATOR OFF, EMERGENCY PWR LEVER, VOLTAGE LOW, VACUUM LOW, RESERVOIR FUEL LOW, DOOR WARNING, BATTERY OVERHEAT (optional) and FUEL SELECT OFF. (2) The amber lenses include the AUX FUEL PUMP ON, FUEL PRESS LOW, STARTER ENERGIZED, LEFT FUEL LOW, RIGHT FUEL FLOW LOW, STBY ELEC PWR ON (optional), INVERTER lNOP, BATTERY HOT (optional), CHIP DETECTOR, STBY ELEC PWR lNOP (optional), A/P B.C. (3) The green lenses include the IGNITION ON, WINDSHIELD ANTI-ICE (optional), and DE-ICE PRESSURE (optional).
B.
Protection for annunciator panel is given by two circuit breakers attached on the left sidewall circuit breaker panel ANNUN PANEL. When the Standby Alternator is installed, one annunciator circuit breaker is removed and this annunciator supply comes from the Keep Alive #2 circuit breaker in the Electrical Power Box through the Standby Power Switch. The annunciator panel will stay on until the Master Switch and Standby Power Switches are turned off. This is a reminder to turn the Standby Power Switch to OFF, this removes the annunciator and Alternator Control Unit drain from the battery.
C.
There is an annunciator panel day/night selector switch, a press-to-test annunciator lamp switch, and a press-to-test fire detect switch installed adjacent to the left end of the annunciator panel. When in the NIGHT position, the day/night switch gives variable intensity down to a preset minimum dim level for the green lamps and some of the amber lamps (non-dimmable amber lamps are: AUX FUEL PUMP ON, FUEL PRESS LOW, and BATTERY HOT). This variable intensity is controlled by the ENG INST light rheostat.
D.
The annunciator lamp test switch is used to test the annunciator lights. The fire detect switch will illuminate the Engine Fire annunciator light and also cause the fire warning horn to sound if the system is operational.
E.
The Altair ADAS+ Engine Trend Monitoring System uses a divided switch-light installed at the top of the left instrument panel to indicate when the system senses an unwanted condition. The switchlight has a white half and an amber half. The Altair ADASd Engine Trend Monitoring System uses a CAS message to display system conditions. For more information on the Altair ADAS+ Engine Trend Monitoring System, refer to Chapter 77 Altair ADAS+ Engine Trend Monitoring System - Description and Operation. For more information on the Altair ADASd Engine Trend Monitoring System, refer to Chapter 77 Altair ADASd Engine Trend Monitoring System - Description and Operation
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MODEL 208 MAINTENANCE MANUAL MASTER WARNING AND ANNUNCIATOR PANEL - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Master warning and annunciator panel maintenance practices consist of annunciator panel components removal/installation.
Annunciator Panel Removal/Installation A.
Remove Annunciator Panel (Refer to Figure 201). (1) Remove seven screws (1) securing cowl deck cover (2) to cowl deck assembly and remove cover. (2) Disconnect connector (4) from annunciator panel (3). (3) Remove six screws (5) securing annunciator panel to instrument panel, and then remove by pulling the annunciator panel back and up through the cowl deck access opening.
B.
Install Annunciator Panel (Refer to Figure 201). (1) Install annunciator panel down and through the cowl deck access opening and secure to instrument panel using six screws (5). (2) Connect connector (4) to annunciator panel (3) and perform operational check. (3) Install cowl deck cover (2) to cowl deck assembly and secure using seven screws (1).
Annunciator Panel Lamp(s) Removal/Installation A.
Remove Annunciator Panel Lamp(s) (Refer to Figure 201). (1) Push in on face of module (18), release the module and allow it to pop out. (2) Pull module out to limit of hinged retainer and allow it to rotate down 90 degrees. (3) Remove bulb (17) from retainer.
B.
Install Annunciator Panel Lamp(s) (Refer to Figure 201). (1) Install bulb (17) into retainer. (2) Rotate module (18) up and press in until it catches and release it. NOTE:
4.
Each module contains two bulbs, and, if necessary, with one defective bulb, the module will remain sufficiently illuminated.
Annunciator Indicator Light Lens Removal/Installation A.
Remove Annunciator Indicator Light Lens (Refer to Figure 201). (1) Push in on face of module (18), then release module and allow it to pop out. (2) Pull module out to limit of hinged retainer (23) and move hinged retainer (23) from side side while continuing to pull. (3) Module (18) should pull free from housing assembly (26). (4) Pry lens retainer (21) up over reflector assembly flanges on top and bottom of module assembly (18) and pull lens retainer (21) off of module assembly (18). (5) Remove nomenclature filter (22B), colored filter (22A) and cover lens (22) from back side of lens retainer (21). Note removal sequence of filters.
B.
Install Annunciator Indicator Light Lens (Refer to Figure 201). (1) Install cover lens (22), colored filter (22A) and nomenclature filter (22B) into the lens.' retainer (21). Make sure stenciled lens read correctly with right side up. (2) Position lens retainer (21) to module assembly (20). (3) Press lens retainer (21) on module assembly (20). (4) Position hinged retainer (23) over retaining spring (19) in housing assembly (26).
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MODEL 208 MAINTENANCE MANUAL
Annunciator Panel Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Annunciator Panel Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
CAUTION: To avoid damaging retaining spring (19) be sure hinged retainer (23) is properly positioned in housing assembly (26) and not against retaining spring (19). If retaining spring (19) is inadvertently damaged, replace housing assembly (26). (5)
Push module (18) into housing assembly (26) until it snaps into a locked position. NOTE:
5.
If module becomes disengaged and will not lock when pressed, it may be necessary to pull module out slightly and press in again.
Annunciator Indicator Light Housing Assembly Removal/Installation A.
Remove Annunciator Indicator Light Housing Assembly (Refer to Figure 201 ). (1) Locate housing assembly to be replaced and remove all three modules in the stack.
CAUTION: PC boards are soldered in place. Be sure PC boards are not moved or damaged when removing or installing housing assemblies. (2) (3)
Remove two screws (27) and nuts (24) securing PC boards, spacers (25), and housing assemblies (26). Pull housing assembly (26) out of stack. NOTE:
B.
In order to remove housing assembly (26) it may be necessary to remove more than one housing assembly in the stack. Insulating spacers may become dislodged when removing housing assemblies. It may be necessary to loosen screws securing modules in adjacent stacks to ease removal of housing assemblies.
Install Annunciator Indicator Light Housing Assembly (Refer to Figure 201).
CAUTION: Be sure not to damage PC board when installing housing assemblies and spacers. (1) (2)
Position housing assembly (26) to annunciator panel and slowly push housing assembly (26) into stack. Install spacers (25). NOTE:
(3) (4) (5) (6)
It may be necessary to loosen screws securing modules in adjacent stacks to allow casing to be separated to ease installation of housing assemblies and spacers.
Align housing assemblies and spacers with PC boards and casing using awls. Install screws (27) and nuts (24). Tighten any adjacent screws that may have been loosened to ease removal and installation. Install three modules in stack as follows: Position hinged retainer (23) over retaining spring (19) to housing assembly (26).
CAUTION: To avoid damaging retaining spring (19) be sure hinged retainer (23) is positioned to housing assembly (26) and not against retaining spring (19). (7) 6.
Push module (18) into PC board until it snaps into a locked position.
Day/Night Switch Removal/Installation A.
Remove Day/Night Switch (Refer to Figure 201). (1) Remove seven screws (1) securing cowl deck cover (2) to cowl deck assembly and remove cover.
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) B.
7.
8.
Remove six screws (5) securing annunciator panel to instrument panel, and then remove by pulling the annunciator panel back and up through the cowl deck access opening. Remove decorator nut (6) from instrument panel and retain washers (7) and (8). Bring switch (13) back and up through cowl deck access opening. Remove electrical leads (12) and tag for reinstallation.
Install Day/Night Switch (Refer to Figure 201). (1) Install electrical leads (12) and remove identification tags. (2) Install switch (13) in instrument panel and install washers (7) and (8) and decorator nut (6) to secure switch to instrument panel. (3) Install annunciator panel down and through the cowl deck access opening, secure to instrument panel using six screws (5), and perform operational check. (4) Install cowl deck cover (2) to cowl deck assembly and secure using seven screws (1).
Lamp Test Switch Removal/Installation A.
Remove Lamp Test Switch (Refer to Figure 201). (1) Remove seven screws (1) securing cowl deck cover (2) to cowl deck assembly and remove cover. (2) Remove six screws (5) securing annunciator panel, and then remove by pulling annunciator panel back up and through the cowl deck access opening. (3) Remove decorator nut (6) from instrument panel and retain washers (7) and (8) for reinstallation. (4) Bring switch (14) back and up through cowl deck access opening. (5) Unsolder electrical leads from switch and tag for reinstallation.
B.
Install Lamp Test Switch (Refer to Figure 201 ). (1) Solder electrical leads to switch (14) and remove identification tags. (2) Install switch (14) in instrument panel and install washers (7) and (8) and decorator nut (6) to secure switch to instrument panel. (3) Install annunciator panel (3) in instrument panel, secure using six screws (5), and perform operational check. (4) Install cowl deck cover (2) to cowl deck assembly and secure using seven screws (1).
Fire Detector Test Switch Removal/Installation A.
Remove Fire Detector Test Switch (Refer to Figure 201). (1) Remove seven screws (1) securing cowl deck cover (2) to cowl deck assembly and remove cover. (2) Remove six screws (5) securing annunciator panel, and then remove by pulling annunciator panel back up and through the cowl deck access opening. (3) Remove decorative nut (16) and nut (6) from instrument panel and retain washer (8). (4) Bring switch (15) back up and through cowl deck access opening. (5) Unsolder electrical leads from switch (15) and tag for reinstallation.
B.
Install Fire Detector Test Switch (Refer to Figure 201 ). (1) Solder electrical leads to switch (15) and remove identification tags. (2) Install switch (15) in instrument panel and install washer (8), nut (6), and decorative nut (16) to secure switch to instrument panel. (3) Install annunciator panel (3) in instrument panel and secure using screws (5). (4) Install cowl deck cover (2) to cowl deck assembly and secure using seven screws (1).
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32 CHAPTER
LANDING GEAR
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
32-00-00
Pages 1-2
Jan 3/2005
32-10-00
Pages 101-104
Aug 1/1995
32-10-00
Pages 201-220
Mar 1/2012
32-10-00
Pages 501-505
May 5/2003
32-10-00
Pages 601-618
Mar 1/2012
32-10-00
Pages 701-703
Sep 3/1996
32-20-00
Pages 101-104
Aug 1/1995
32-20-00
Pages 201-214
Apr 1/2010
32-20-00
Page 601
Jun 1/2011
32-20-00
Pages 701-703
Sep 3/1996
32-20-01
Pages 201-210
Apr 1/2010
32-20-02
Pages 201-210
Jan 2/2006
32-20-02
Page 601
Jun 1/2011
32-20-03
Pages 201-202
Jan 2/2006
32-40-00
Page 1
Aug 1/1995
32-40-00
Pages 101-104
Aug 1/1995
32-40-00
Pages 201-214
Mar 1/2012
32-40-00
Pages 601-604
Jun 1/2011
32-40-00
Page 701
Sep 4/2001
32-40-00
Page 801
Aug 1/1995
32-41-00
Pages 201-203
May 5/2003
32-42-00
Pages 201-203
Aug 1/1995
32-50-00
Pages 101-103
Aug 1/1995
32-50-00
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Aug 1/1995
32-Title 32-List of Effective Pages 32-Record of Temporary Revisions 32-Table of Contents 32-List of Tasks
32 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS LANDING GEAR - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-00-00 32-00-00 32-00-00 32-00-00
Page 1 Page 1 Page 1 Page 2
MAIN LANDING GEAR - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-10-00 Page 101 32-10-00 Page 101
MAIN LANDING GEAR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Gear Assembly Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Gear Spring Removal/Installation (Airplanes Without Cargo Pod) . . . . . . . . . . . Main Gear Spring Removal/Installation (Airplanes With Cargo Pod). . . . . . . . . . . . . . Main Gear Fairing Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Gear Axle Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Spring Puller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-10-00 Page 201 32-10-00 Page 201 32-10-00 Page 201 32-10-00 Page 201 32-10-00 Page 201 32-10-00 Page 208 32-10-00 Page 211 32-10-00 Page 215 32-10-00 Page 217 32-10-00 Page 217
MAIN LANDING GEAR - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wheel Toe-In Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wheel Camber Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-10-00 Page 501 32-10-00 Page 501 32-10-00 Page 501 32-10-00 Page 501
MAIN LANDING GEAR - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Center-Spring and Main Gear-Spring Interface Area Special Detailed (Corrosion Inspection and Repair) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Axle Special Detailed Inspection (SID) . . . . . . . . . . . . . . . . . . . . . .
32-10-00 Page 601 32-10-00 Page 601 32-10-00 Page 601 32-10-00 Page 609 32-10-00 Page 615
MAIN LANDING GEAR - CLEANING/PAINTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ReÞnishing High Stressed Steel Shot Peened Surfaces. . . . . . . . . . . . . . . . . . . . . . . . . ReÞnishing Steel, Aluminum and Magnesium Components . . . . . . . . . . . . . . . . . . . . .
32-10-00 Page 701 32-10-00 Page 701 32-10-00 Page 701 32-10-00 Page 701 32-10-00 Page 703
NOSE LANDING GEAR - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-20-00 Page 101 32-20-00 Page 101
NOSE LANDING GEAR - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drag Link Spring Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drag Link Spring Support Liner Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Drag Link Spring Inspection/Repair . . . . . . . . . . . . . . . . . . . . . . . . Nose Wheel Grease Seal Bore And Cup Backing Bore Surface Rework . . . . . . . . . . Nose Gear Assembly Replacement Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-20-00 Page 201 32-20-00 Page 201 32-20-00 Page 201 32-20-00 Page 206 32-20-00 Page 207 32-20-00 Page 211 32-20-00 Page 213
NOSE LANDING GEAR - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Wheel Grease Seal Bore and Cup Backing Bore Surfaces. . . . . . . . . . . . . . . . .
32-20-00 Page 601 32-20-00 Page 601 32-20-00 Page 601 32-20-00 Page 601
NOSE LANDING GEAR - CLEANING/PAINTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ReÞnishing High Stressed Steel Shot Peaned Surfaces. . . . . . . . . . . . . . . . . . . . . . . . . ReÞnishing Steel, and Aluminum Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-20-00 Page 701 32-20-00 Page 701 32-20-00 Page 701 32-20-00 Page 701 32-20-00 Page 701
32 - CONTENTS © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL NOSE GEAR SHOCK STRUT - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear Shock Strut Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear Shock Strut Disassembly/Assembly (Airplanes 20800134 and On, 208B0099 and On, and Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Incorporating SK208-51) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear Shock Strut Disassembly/Assembly (Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Except Airplanes Incorporating SK208-51) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-20-01 Page 201 32-20-01 Page 201 32-20-01 Page 201
SHIMMY DAMPENER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shimmy Dampener Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shimmy Dampener Disassembly/Assembly (Airplanes 20800394 and On and 208B1140 and On) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shimmy Dampener Disassembly/Assembly (Airplanes 20800001 thru 20800393 and 208B0001 thru 208B1139). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-20-02 Page 201 32-20-02 Page 201 32-20-02 Page 201
SHIMMY DAMPENER - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shimmy Damper Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-20-02 Page 601 32-20-02 Page 601 32-20-02 Page 601
NOSE GEAR FAIRING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear Fairing Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-20-03 Page 201 32-20-03 Page 201 32-20-03 Page 201
WHEELS AND BRAKES - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-40-00 Page 1 32-40-00 Page 1
WHEELS AND BRAKES - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32-40-00 Page 101 32-40-00 Page 101
WHEELS AND BRAKES - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tire Mounting Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing Tires and Tubes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brake System Replenishing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brake System Bleeding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Wheel and Tire Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Wheel, Tire and Tube Disassembly/Reassembly . . . . . . . . . . . . . . . . . . . . . . . . . . Brake Backplate and Pressure Plate Removal/Installation . . . . . . . . . . . . . . . . . . . . . . Brake Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brake Assembly Disassembly/Reassembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . New Brake Burn-In . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Wheel Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Wheel Tire and Tube Disassembly/Reassembly . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection and Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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WHEELS AND BRAKES - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brakes Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brake System Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Wheels and Tires Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Wheel and Tire Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . Brakes Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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WHEELS AND BRAKES - CLEANING/PAINTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wheel Preparation and Painting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL WHEELS AND BRAKES - APPROVED REPAIRS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Approved Repair Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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BRAKE MASTER CYLINDER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Master Cylinder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Master Cylinder Disassembly/Reassembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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PARKING BRAKE VALVE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Parking Brake Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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NOSE GEAR STEERING - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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NOSE GEAR STEERING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear System Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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LIST OF TASKS 32-10-00-220
Main Landing Gear Detailed Inspection
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32-10-00-221
Center-Spring and Main Gear-Spring Interface Area Special Detailed (Corrosion Inspection and Repair)
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32-10-00-240
Main Landing Gear Axle Special Detailed Inspection (SID)
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32-20-00-220
Nose Landing Gear Detailed Inspection
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32-20-02-720
Shimmy Damper Functional Check
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32-40-00-220
Brakes Detailed Inspection
32-40-00 Page 601
32-40-00-610
Brake System Servicing
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32-40-00-222
Main Landing Gear Wheels and Tires Detailed Inspection
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32-40-00-224
Nose Landing Gear Wheel and Tire Detailed Inspection
32-40-00 Page 603
32-40-00-710
Brakes Operational Check
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MODEL 208 MAINTENANCE MANUAL LANDING GEAR - GENERAL 1.
Scope A.
2.
This chapter gives maintenance information for the landing gear and associated components. Also included in this chapter are procedures to clean and paint the nose and main gear components.
Tools, Equipment and Materials NOTE:
Equivalent substitutes can be used for the items that follow:
NAME
NUMBER
MANUFACTURER
USE
Washers
153-00100
Cessna Aircraft Company Cessna Parts Distribution 5800 East Pawnee P. O. Box 1521 Wichita, KS 67218 USA
To provide the specified torque range for the main tire and wheel nuts.
Corrosion Protection Primer Base Activator Reducer
Type II 10P30-5 EC-275 TR-115
Akzo Nobel Aerospace Coatings One East Water St. Waukegan, IL 60085
To lubricate the main landing gear mating surfaces of the spring, center spring, attach trunnion and pin.
Lubricant
Permatex Anti-Seize Lubricant
Commercially Available
To protect the nose gear wheel and bearings from corrosion.
Protectant
Royco 103
Royal Lubricants Co., Inc. River Road East Hanover, NJ 07936
To protect the internal surfaces of the main and nose gear wheels.
Grease
AMS/OIL GHD grease
Get in touch with Local Distributor
Alternate grease for lubricating the wheel bearings and zerk fittings.
Glass Bead
MIL-G-9954 (Size 10 or Size 13)
Cataphote Incorporated 1001 Underwood Drive Jackson, Miss 39208
Paint removal on the main gear and center springs.
Sealant Semkit, (2.5 oz.)
Type 1, Class B-1/2
Cessna Aircraft Company
To fillet seal the main gear spring at the center spring
Depth micrometer (pin type) with a tolerance of ±0.001 inch
Commercially available
To measure the damage to the landing gear springs.
4X magnifying glass
Commercially available
To inspect the landing gear springs.
Trunnion Puller
2609084-1
Get in touch with Cessna Propeller Aircraft Product Support at (316) 517-5800 or Fax (316) 942-9006.
To remove the attach trunnion from the center spring.
Main Landing Gear Spring Puller
T2680002-17
Get in touch with Cessna Propeller Aircraft Product Support at (316) 517-5800 or Fax (316) 942-9006.
To remove the main landing gear spring from the center spring.
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NAME
NUMBER
MANUFACTURER
USE
Hydraulic Jack
RLS201
Simplex Division of Templeton, Kenly & Co., Inc. 2525 Gardner Road Maywood, IL 60153-3719
To operate the T2680002-17 main landing gear spring puller (two required).
Hydraulic Hand Pump
P42
Simplex Division of Templeton, Kenly & Co., Inc.
To operate the hydraulic jacks used with the T2680002-17 main landing gear spring puller.
Penetrating Oil Basin
T2680002-18
Get in touch with Cessna Propeller Aircraft Product Support at (316) 517-5800 or Fax (316) 942-9006.
To use in between the center spring and main landing gear spring.
Penetrating Oil
Kroil
Kano Laboratories, Inc. P.O. Box 110098 Nashville, TN 37222-0098
To use in the basin between the center spring and main landing gear spring.
3.
Definition A.
This chapter is divided into sections to help maintenance personnel find information. Use the Table of Contents to help locate a particular subject. A brief definition of the sections in this chapter is as follows: (1) The section on the main landing gear gives instructions to troubleshoot, adjust, paint or clean and do maintenance practices for the main landing gear. (2) The section on the nose landing gear gives instructions to troubleshoot, inspect or check, paint or clean and do maintenance practices for the nose landing gear. (3) The section on the wheels and brakes gives description and operation information, and instructions to troubleshoot, adjust and test, and do maintenance practices for the main gear brake system. (4) The section on the nose gear steering gives instructions to troubleshoot, adjust and test, paint or clean, or do maintenance practices, and gives approved repair procedures for the nose gear steering system and related components.
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MODEL 208 MAINTENANCE MANUAL MAIN LANDING GEAR - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been prepared to aid the maintenance technician in system understanding. Refer to Figure 101.
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Main Landing Gear Troubleshooting Chart Figure 101 (Sheet 1)
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Main Landing Gear Troubleshooting Chart Figure 101 (Sheet 2)
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Main Landing Gear Troubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL MAIN LANDING GEAR - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives description, removal, and installation instructions for the main landing gear. For the maintenance procedures for the main gear wheel and brakes, refer to Wheels and Brakes Maintenance Practices. For the maintenance procedures for the nose landing gear, refer to Nose Landing Gear - Maintenance Practices.
Description A.
The Þxed, tricycle landing gear assembly has a tubular main gear and a steerable nose gear. Each main gear has a hydraulically operated disk-type brake (with a fairing), a two-piece gear spring fairing and a gear-to-fuselage fairing. (1) The tubular main gear has a center spring tube and two outer spring tubes. The center spring tube is attached to each outboard spring tube through a trunnion assembly. The trunnion assembly is attached to the fuselage at two points on each side of the lower fuselage structure. Because of the vertical loads when you land or taxi the airplane, the center and outboard spring tubes turn on the longitudinal axes around the four attached points. Each trunnion assembly uses a bearing and race, and a bearing cap attached by two cap bolts. This lets the main gear be easily removed for gear replacement, or for ßoat installation. NOTE:
3.
Tools, Equipment and Materials A.
4.
The three-piece steel tube assembly has a slightly larger diameter for the Model 208B than the Model 208. The design is identical for the Model 208 and 208B assemblies.
For a list of necessary tools, equipment and materials, refer to Landing Gear - General.
Main Gear Assembly Removal/Installation A.
Remove the Main Gear Assembly (Refer to Figure 201). (1) Remove the main gear fairings. Refer to Main Gear Fairing Removal/Installation. (2) Remove the cargo pod, if installed (Airplanes 20800001 thru 20800395 and 208B00001 thru 208B01170). Refer to Chapter 25, Cargo Pod - Maintenance Practices. (3) Remove only the main gear outer cover, if a cargo pod is installed (Airplanes 20800396 and On and 208B01171 and On). (a) Remove the screws that attach the main gear outer cover. (b) Remove the main gear outer cover. (4) Use jacks to lift the airplane. Refer to Chapter 7, Jacking - Maintenance Practices. (5) Put a support below the main gear assembly at the attach trunnion.
CAUTION: Make sure that the gear support can hold a weight of approximately 400 pounds (180 kg). This will help prevent damage to the equipment. (6) (7) (8) (9) (10)
Use an ink marker and put marks on the left and right trunnions to refer to during installation. Drain the brake system. Disconnect the brake line at the bulkhead Þtting in the center spring tunnel. Put a cap on the open end of the brake line. Carefully loosen the bearing cap bolts. (a) Remove the bearing cap bolts, washers, and shims that attach the bearing cap to the fuselage attach Þtting. Keep the shims for installation. 1 (b) While the gear assembly is on the supports, lower it as an assembly to the ground. (11) Remove the bearings from the bearing races.
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Main Landing Gear Installation Figure 201 (Sheet 1)
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Main Landing Gear Installation Figure 201 (Sheet 2)
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Main Landing Gear Installation Figure 201 (Sheet 3)
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Main Landing Gear Installation Figure 201 (Sheet 4)
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Main Landing Gear Installation Figure 201 (Sheet 5)
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CAUTION: Do not try to remove the bearing races unless it is absolutely necessary. This will help prevent damage to the attach trunnion. The bearing races are press-Þtted onto the attach trunnion. CAUTION: If the landing gear is to be removed and put in storage for long periods, such as during extended ßoat operations, it is recommended that the bearings be removed and kept separate. Before the bearings are put in storage, a protective grease layer must be applied to all the bare surfaces of the bearings and races. B.
Install the Main Gear Assembly (Refer to Figure 201). (1) Put a support below the main landing gear assembly below the airplane. (2) Install the bearings on each bearing race. (3) Lift the landing gear assembly so the bearings align with the cutouts in the fuselage attach Þttings. NOTE:
(4)
Put (a) (b) (c)
If you are installing the same bearing cap that you removed, wait to install the shims until after you examine the gap between the bearing cap and the fuselage attach Þtting.
the bearing caps in their position at the fuselage attach Þtting. Install the bearing cap bolts and washers hand tight on each bearing cap. Torque the bearing cap bolts from 770 to 950 inch-pounds (87 to 107 N-m). Use a feeler gage to make sure that the gap is less than 0.001 inch (0.025 mm) between the bearing cap and the fuselage attach Þtting. NOTE:
(d)
Shims of different thickness are available to use if the gap is more than 0.001 inch (0.025 mm). Refer to the Illustrated Parts Catalog.
If the gap between the bearing cap(s) and the fuselage attach Þtting is more than 0.001 inch (0.025 mm) do the steps that follow: Remove the bearing cap bolts, washers, and bearing cap(s) from the fuselage attach 1 Þtting. Put a shim on each of the bearing cap bolts between the bearing cap(s) and the 2 fuselage attach Þtting.
3 4 5 6 7
NOTE:
Use the thinnest shim necessary to make the gap between the bearing caps and the fuselage attach Þtting less than 0.001 inch (0.025 mm) when the bearing cap bolts are torqued.
NOTE:
Make sure that the shims are the same thickness on both the inboard and outboard sides of each bearing cap.
NOTE:
Do not use more than one layer of shims.
Put the bearing cap(s) in their position at the fuselage attach Þtting. Install the bearing cap bolts and washers hand tight on each bearing cap. Torque the bearing cap bolts from 770 to 950 inch-pounds (87 to 107 N-m). Use a feeler gage to make sure that the gap is less than 0.001 inch (0.025 mm) between the bearing cap and the fuselage attach Þtting. If the gap is more than 0.001 inch (0.025 mm), do the steps again to install thicker shims until the gap is less than 0.001 inch (0.025 mm). NOTE:
(5) (6) (7)
Do not use more than one layer of shims.
Connect the brake line at the bulkhead Þtting in the center spring tunnel. Bleed the brake system. Refer to Wheels and Brakes - Maintenance Practices. Remove the airplane from the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
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Install the main landing gear fairings. Refer to Main Gear Fairing Removal/Installation. Install the cargo pod (Airplanes 20800001 thru 20800395 and 208B0001 thru 208B1170), if it was removed. (10) Airplanes with a cargo pod installed, if only the main outer cover was removed (Airplanes 20800396 and On and 208B1171 and On), do the steps that follow: (a) Install the main gear outer cover with the screws. (b) Make sure that the drain hole is clear of blockage after the main gear outer cover is installed. 5.
Main Gear Spring Removal/Installation (Airplanes Without Cargo Pod)
CAUTION: Removal/Installation of the Main Gear Spring must be done at 72°F, +20 or - 20°F (22° C, +11 or -11°C). Damage to the spring can occur from incorrect tolerance readings. A.
Remove the Main Gear Spring (Refer to Figure 201 and Figure 202). (1) Chock the wheel of the main landing gear on the opposite side of the airplane of the main gear spring to be removed. (2) Remove the main gear fairings of the main gear spring to be removed. Refer to Main Gear Fairing Removal/Installation. (3) Open the brake bleeder valve at the brake caliper and drain the brake system. (4) Disconnect the brake line at the brake caliper. (5) Put a cap on the open end of the brake line. (6) Remove the clamps that attach the brake line to the main gear spring. (7) Put an aircraft jack below the fuselage attach Þtting at the jack point on the side the main gear spring is to be removed. (8) Remove the bearing cap bolts and bearing caps on the side the main gear spring is to be removed. (9) Remove the Þllet seal at the point where the main gear spring goes into the center spring. Refer to Figure 201 (10) Remove the Þllet seal at the point where the center gear spring goes in to and out of the trunnion. Refer to Figure 201 (11) Use the jack to carefully lift the airplane until there is sufÞcient clearance between the attach trunnion and fuselage attach Þtting to remove the bolt and pin. (12) Remove the Þllet seal from the bolt head, washers and nut that hold the pin in the trunnion. (13) Remove the nut, washers and bolt that hold the pin in the attach trunnion. NOTE:
Keep the washers for installation.
(14) Put a support below the attach trunnion. (15) Use a drift punch to remove the pin from the attach trunnion. NOTE:
The pin serial number placard is installed on the trunnion. Once the pin is removed, it must be identiÞed so it will be installed on the same airplane and in the same location. If the pin is replaced with a new pin, the placard must also be replaced. A new placard must have the pin part number and the serial number stamped on it, and then it must be installed where the previous placard was installed. The pin serial number is the airplane's serial number, followed by an L for the left side or an R for the right side, then the sequence number of the pin. For example, the Þrst replacement pin for an airplane serial number 208B970 on the left side would be SNB970L-2. Refer to CAB03-7 for instructions on the placard installation.
(16) Twist and pull the main gear spring to remove it from the center spring. (a) If the main gear spring cannot be removed, install the main landing gear spring puller. Refer to Main Landing Gear Spring Puller. NOTE:
Cessna Propeller Aircraft Product Support is the source to get instructions to fabricate the main landing gear spring puller.
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CAUTION: Make sure to be careful when the main gear spring is removed. This will help prevent damage to the landing gear during the removal process. CAUTION: Make sure that the force applied by each jack is not more than 12 tons. This will help prevent damage to the main gear. CAUTION: Do not apply heat to the main gear to remove the main gear springs. If you do, you can damage the main gear. 1
Increase the hydraulic jack (load) pressure until the main gear spring is removed or until a maximum force of 12 tons is applied by both jacks.
WARNING: Make sure the jacks extend equally when you apply the force. This will help prevent injury to personnel and damage to the equipment. NOTE:
The maximum pressure for the recommended jacks is 5000 PSI.
a
If necessary, use spacers to extend the reach of the jacks while the main gear spring is removed. b If the main gear spring still does not move when the maximum pressure is applied, get in touch with Cessna Propeller Aircraft Product Support for assistance; 316-517-5800 or Fax 316-942-9006. 2 Make sure that the jacks, spacers and the main gear spring do not fall during the main gear spring removal. (17) If necessary, remove the attach trunnion. Refer to Remove the Main Gear Assembly. (a) If you cannot remove the attach trunnion, you will need to use a special tool. NOTE: B.
Cessna Propeller Aircraft Product Support, (316) 517-5800 or Fax (316) 9429006, is the source to get a special tool or the instructions to fabricate the tool.
Install the Main Gear Spring (Refer to Figure 201 and Figure 202). (1) Examine for gouging, chaÞng or corrosion on the faying surfaces of the main gear spring and the center spring. Refer to Main Landing Gear - Inspection/Check. (a) If gouging, chaÞng or corrosion is found, prepare the damaged area for measurement. Refer to Center Spring and Main Gear Spring Interface Area Special Detailed Inspection and Repair. (2) Clean the unpainted surfaces of the main gear spring and the center spring with isopropyl alcohol. (3) Make sure that the main gear spring interior and exterior unpainted surfaces, except for the faying surface with the center spring, have a layer of polyurethane corrosion protection primer. (a) Fill and drain the interior of the main gear spring with Type II corrosion protection primer (zinc chromate). (b) If necessary, apply a layer of Type II corrosion protection primer as a spray to the unpainted surfaces of the exterior of the main gear spring except for the faying surface with the center spring. Refer to Landing Gear - General. (c) Apply a fay seal of Type I, Class B to the spring plug. Refer to Figure 201 Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices (d) Install the spring plug in to the main spring. NOTE: (e)
Do not obstruct or seal over the pin holes in the main spring.
Install a Þllet seal of Type I, Class B around the top of the plug and the spring edge. Refer to Figure 201 and Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices
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(5)
(6)
Make sure that the center spring interior unpainted surface, except for the faying surface with the main gear spring, has a layer of Type II corrosion protection primer (zinc chromate). Refer to Landing Gear - General. (a) Apply a layer of Type II corrosion protection primer as a spray to the unpainted surfaces of the interior of the center spring except for the faying surface with the main gear spring. Immediately before the installation of the main gear spring, axle Þtting and pin, do the steps that follow. (a) Clean the faying surfaces with isopropyl alcohol. (b) Use a brush and apply a Type II corrosion-protection primer to the faying surfaces that are the shaded areas shown in Figure 201 and Landing Gear - General. If necessary, install the attach trunnion on the center spring. NOTE:
(7) (8)
The Type II corrosion-protection primer on the faying surface must be wet during the installation. Refer to Landing Gear - General.
Make sure that the attach trunnion is held in position, and the airplane is lifted sufÞciently at the jack point of the fuselage attach Þtting to give clearance to install the pin in the attach trunnion. Twist and push the main gear spring to install it in the center spring. NOTE:
The Type II corrosion-protection primer on the faying surface must be wet during the installation. Refer to Landing Gear - General.
(9) Align the attach holes for the pin. (10) Use a brush and apply a Type II corrosion-protection primer to the pin, refer to Figure 201 and Landing Gear - General. (11) Use a nonmetallic hammer to tap the pin through the aligned holes in the attach trunnion, center spring and the main gear spring. NOTE:
The Type II corrosion-protection primer on the faying surface must be wet during the installation. Refer to Landing Gear - General.
(12) Shank seal the bolt that holds the attach pin in position with Type I, Class B sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices (13) Install the bolt, washers and nut that hold the pin in the attach trunnion. (a) Torque the nut from 30 to 40 inch-pounds (3.4 to 4.5 N-m). (b) Apply a Þllet seal around the bolt head, washers and nut. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices (14) Lower the airplane at the jack point of the fuselage attach Þtting until the bearings touch the recesses in the fuselage attach Þtting. (15) Use the bearing cap bolts to install the bearing caps. (a) Torque the bearing cap bolts from 770 to 950 inch-pounds (87 to 107 N-m). (16) Use methyl n-propyl ketone (or equivalent) to clean the outside of the main gear spring. (17) Use Type 1, Class B-1/2 sealant and apply a Þllet seal around the center spring at the attach trunnion. Refer to Figure 201 and Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices. (18) Use Type 1, Class B-1/2 sealant and apply a Þllet seal around the main gear spring at the center spring. Refer to Figure 201 and Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices. (19) If removed, install the axle Þtting on the main gear spring while the primer is still wet. NOTE:
The Type II corrosion-protection primer on the faying surface must be wet during the installation.
(20) Use Type 1, Class B-1/2 sealant and apply a Þllet seal around the main gear spring at the axle Þtting. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (21) Remove the aircraft jack from the fuselage attach Þtting.
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MODEL 208 MAINTENANCE MANUAL (22) Install the clamps that attach the brake line to the main gear spring. NOTE:
The upper clamp must be installed no closer than 0.75 inch (19.05 mm) to the trunnion. Refer to Figure 201.
(23) Connect the brake line Þtting at the brake caliper. (24) Bleed the brake system. Refer to Wheels and Brakes - Maintenance Practices. (25) Install the main gear fairings. Refer to Main Gear Fairing Removal/Installation. 6.
Main Gear Spring Removal/Installation (Airplanes With Cargo Pod)
CAUTION: Removal/Installation of the Main Gear Spring must be done at 72°F, +20 or -20°F (22°C, +11 or -11°C). Damage to the spring can occur from incorrect tolerance readings. A.
Remove the Main Gear Spring (Refer to Figure 201 and Figure 202). (1) Chock the wheel of the main landing gear on the opposite side of the airplane of the main gear spring to be removed. (2) Remove the main gear fairings of the main gear spring to be removed. Refer to Main Gear Fairing Removal/Installation. (3) Open the brake bleeder valve at the brake caliper and drain the brake system. (4) Disconnect the brake line at the brake caliper. (5) Put a cap on the open end of the brake line. (6) Disconnect the brake line at the fuselage. (7) Put a cap on the open end of the brake line. (8) Remove the clamps that attach the brake line to the main gear spring. (9) Remove the brake line from the airplane. (10) Prepare the aft bearing cap for jacking. (a) Remove the bolts from the aft bearing cap. (b) Turn the bearing cap away from the attach trunnion. (c) Install the outboard bolt of the bearing cap. (11) Put an aircraft jack below the fuselage attach Þtting at the jack point on the side the main gear spring is to be removed. (12) Remove the forward bearing cap bolts and bearing cap on the side that the main gear spring is to be removed from. (13) Remove the main gear strut cover on the side from which the main gear spring is to be removed. NOTE:
The main gear strut cover is in the cargo pod and outboard of the center spring cover.
(14) Put a board and a jack between the center spring and the ßoor of the cargo pod to hold the complete gear assembly (trunnion and gear).
CAUTION: Make sure that you put a board on the ßoor of the cargo pod below the jack. This will help prevent damage to the cargo pod. CAUTION: When you put the jack between the center spring and the cargo pod, be careful not to cause damage to the center spring. (15) Use the ßoor jack to carefully lift the airplane until the tire is approximately 2 inches above the ground. (16) Put aircraft jacks below the wings (in addition to the gear support jacks). (17) Use the jack in the cargo pod to lower the complete gear assembly (trunnion and gear) until there is sufÞcient clearance between the attach trunnion and fuselage attach Þtting to remove the bolt and pin. (a) If necessary, remove the outboard bolts from the center spring cover to get the necessary clearance.
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MODEL 208 MAINTENANCE MANUAL (18) Remove the Þllet seal at the point where the main gear spring goes into the center spring. Refer to Figure 201 (19) Remove the Þllet seal at the point where the center gear spring goes in to and out of the trunnion. Refer to Figure 201 (20) Make sure that the attach trunnion is out of the bearing caps. (21) Remove the Þllet seal from the bolt head, washer and the nut that hold the pin in the trunnion. (22) Remove the nut, washer and bolt that hold the pin in the attach trunnion. NOTE:
Keep the washers for installation.
(23) Use a slide hammer to remove the pin from the aft side of the attach trunnion. NOTE:
The pin serial number placard is installed on the trunnion. Once the pin is removed, it must be identiÞed so it will be installed on the same airplane and in the same location. If the pin is replaced with a new pin, the placard must also be replaced. A new placard must have the pin part number and the serial number stamped on it, and then it must be installed where the previous placard was installed. The pin serial number is the airplane's serial number, followed by an L for the left side or an R for the right side, then the sequence number of the pin. For example, the Þrst replacement pin for an airplane serial number 208B970 on the left side would be SNB970L-2. Refer to CAB03-7 for instructions on the placard installation.
(24) Twist and pull the main landing gear spring to remove it from the center spring. (a) If the main landing gear spring cannot be removed, use a main landing gear spring puller to remove the main gear spring. Refer to Main Landing Gear Spring Puller. NOTE:
Cessna Propeller Aircraft Product Support, 316-517-5800 or Fax 316-942-9006, is the source to get the instructions to fabricate the main landing gear spring puller.
CAUTION: Make sure to be careful when the main gear spring is removed. This will help prevent damage to the landing gear during the removal process. CAUTION: Make sure that the force applied by each jack is not more than 12 tons. This will help prevent damage to the equipment. CAUTION: Do not apply heat to the main gear to remove the main gear springs. If you do, you can damage the main gear. 1
Increase the hydraulic jack (load) pressure until the main gear spring is removed or until a maximum force of 12 tons is applied by both jacks.
WARNING: Make sure the jacks extend equally when you apply the force. This will help prevent injury to personnel and damage to the equipment. NOTE:
The maximum pressure for the recommended jacks is 5000 PSI.
a
2
If necessary, use spacers to extend the reach of the jacks while the main gear spring is removed. If the main gear spring still does not move when the maximum pressure b is applied, get in touch with Cessna Propeller Aircraft Product Support for assistance; (316) 517-5800 or Fax (316) 942-9006. Make sure that the jacks, spacers and the main gear spring do not fall during the main gear spring removal.
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MODEL 208 MAINTENANCE MANUAL (25) If necessary, remove the attach trunnion. Refer to the Main Gear Assembly. (a) If you cannot remove the attach trunnion, you will need to use a special tool (trunnion puller). NOTE: B.
Cessna Propeller Aircraft Product Support, (316) 517-5800 or Fax (316) 9429006, is the source to get a special tool or the instructions to fabricate the tool.
Install the Main Gear Spring (Refer to Figure 201 and Figure 202). (1) Examine for gouging, chaÞng or corrosion on the faying surfaces of the main gear spring and the center spring. Refer to the Main Landing Gear - Inspection/Check. (a) If gouging, chaÞng or corrosion is found, prepare the damaged area for measurement. Refer to Center Spring and Main Gear Spring Interface Area Special Detailed Inspection and Repair. (2) Clean the unpainted surfaces of the main gear spring and the center spring with isopropyl alcohol. (3) Make sure that the main gear spring interior and exterior unpainted surfaces, except for the faying surface with the center spring, have a layer of polyurethane corrosion protection primer. (a) Fill and drain the interior of the main gear spring with Type II corrosion protection primer (zinc chromate). (b) If necessary, apply a layer of Type II corrosion protection primer as a spray to the unpainted surfaces of the exterior of the main gear spring except for the faying surface with the center spring. Refer to Landing Gear - General. (c) Apply a fay seal of Type I, Class B to the spring plug. Refer to Figure 201 Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices (d) Install the spring plug in to the main spring. NOTE:
Do not obstruct or seal over the pin holes in the main spring.
(e)
(4)
(5)
(6)
Install a Þllet seal of Type I, Class B around the top of the plug and the spring edge. Refer to Figure 201 and Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices Make sure that the center spring interior unpainted surface, except for the faying surface with the main gear spring, has a layer of Type II corrosion protection primer (zinc chromate). (a) Apply a layer of Type II corrosion-protection primer as a spray to the unpainted surfaces of the interior of the center spring except for the faying surface with the main gear spring. Immediately before the installation of the main gear spring, axle Þtting, attach trunnion and pin do the steps that follow. (a) Clean the faying surfaces with isopropyl alcohol. (b) Use a brush and apply a Type II corrosion protection primer to the faying surfaces, shaded areas shown in Figure 201. Refer to Landing Gear - General. If necessary, install the attach trunnion on the center spring. NOTE:
(7) (8) (9)
The Type II corrosion-protection primer on the faying surface must be wet during the installation. Refer to Landing Gear - General.
Make sure that the jack in the cargo pod and the complete assembly, including the attach trunnion and gear, is held in position. Make sure that the airplane is lifted sufÞciently below the wings and at the jack point of the fuselage attach Þtting to give clearance to install the pin in the attach trunnion. Twist and push the main gear spring to install it in the center spring. NOTE:
The Type II corrosion-protection primer on the faying surface must be wet during the installation. Refer to Landing Gear - General.
(10) Align the attach holes for the pin. (11) Use a brush and apply a Type II corrosion-protection primer to the pin, refer to Figure 201 and Refer to Landing Gear - General.
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MODEL 208 MAINTENANCE MANUAL (12) Use a nonmetallic hammer to tap the pin wet through the aligned holes in the attach trunnion and the main gear spring. NOTE:
The Type II corrosion-protection primer on the faying surface must be wet during the installation. Refer to Landing Gear - General.
(13) Shank seal the bolt that holds the attach pin in position with Type I, Class B sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices (14) Install the bolt, washers and nut that hold the pin in the attach trunnion. (a) Torque the nut from 30 to 40 inch-pounds (3.4 to 4.5 N-m). (b) Apply a Þllet seal around the bolt head, washers and nut. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices (15) Use the jack in the cargo pod to lift the complete gear assembly (trunnion and gear) until the bearings touch the recesses in the fuselage attach Þtting. (16) Use the bearing cap bolts to install the forward bearing cap. (a) Torque the bearing cap bolts from 770 to 950 inch-pounds (87 to 107 N-m). (17) Remove the jack in the cargo pod from the airplane. (a) If necessary, install the outboard bolts in the center spring cover. (18) Install the main gear strut cover on the side the main gear spring is to be removed. NOTE:
The main gear strut cover is in the cargo pod and outboard of the center spring cover.
(19) Use methyl n-propyl ketone (or equivalent) to clean the outside of the main gear spring. (20) Use Type 1, Class B-1/2 sealant and apply a Þllet seal around the center spring at the attach trunnion. Refer to Figure 201 and Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices. (21) Use Type 1, Class B-1/2 sealant and apply a Þllet seal around the main gear spring at the center spring. Refer to Figure 201 and Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices. (22) If removed, install the axle Þtting wet on the main gear spring. NOTE:
The Type II corrosion protection primer on the faying surface must be wet during the installation.
(23) Use Type 1, Class B-1/2 sealant and apply a Þllet seal around the main gear spring at the axle Þtting. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (24) Remove the jacks from below the wings. (25) Use the ßoor jack at the fuselage attach Þtting to carefully lower the aircraft. (26) Remove the aircraft jack from the fuselage attach Þtting. (27) Use the bearing cap bolts to install the aft bearing cap. (a) Remove the outboard bolt of the aft bearing cap. (b) Put the bearing cap in position. (c) Install the bearing cap bolts of the aft bearing cap. (d) Torque the bearing cap bolts from 770 to 950 inch-pounds (87 to 107 N-m). (28) Connect the brake line at the fuselage. (29) Install the clamps that attach the brake line to the main gear spring. NOTE:
The upper clamp must be installed no closer than 0.75 inch (19.05mm) to the trunnion. Refer to Figure 201.
(30) Connect the brake line Þtting at the brake caliper. (31) Bleed the brake system. Refer to Wheels and Brakes - Maintenance Practices. (32) Install the main gear fairings. Refer to Main Gear Fairing Removal/Installation.
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7.
Main Gear Fairing Removal/Installation A.
Remove the Main Gear Fairing (Refer to Figure 202). (1) Remove the screws that attach the axle Þtting fairing to the lower end of the main spring lower fairing. (2) Loosen the nuts and bolts that attach the axle Þtting to the main-gear-spring. Refer to Figure 201. (3) Remove the axle Þtting fairing from the airplane. (4) Remove the screws on the inboard split line of the main-gear-spring lower fairing. (5) Remove the screws that attach the main-gear-spring upper fairing to the main-gear-spring lower fairing. (6) Remove the main-gear-spring lower fairing from the airplane. (7) Remove the gear-to-fuselage fairing screws and the screws on the lower split line of the gearto-fuselage fairing. (8) Remove the main-gear-spring upper fairing from the airplane. (9) Remove the gear-to-fuselage fairing from the airplane. (10) Remove the bolts that attach the stiffeners to the belly of the airplane (208B only). (11) Remove the stiffeners from the airplane (208B only). (12) Remove the bolts and washers that attach the center spring cover and cover stiffeners, if applicable, to the belly of the airplane. (13) Remove the center spring cover and cover stiffeners, if applicable from the airplane.
B.
Install the Main Gear Fairing (Refer to Figure 202). (1) Put the center spring cover and cover stiffeners, if applicable, in position on the belly of the airplane. (2) Install the bolts and washers that attach the center spring cover and cover stiffeners, if applicable, to the belly of the airplane. (3) Put the stiffeners in position on the belly of the airplane (208B only). (4) Install the bolts that attach the stiffeners to the belly of the airplane (208B only). (5) Put the gear-to-fuselage fairing in position. (6) Install the gear-to-fuselage fairing screws and the screws on lower split line of the gear-tofuselage fairing to attach the gear-to-fuselage fairing. (7) Align the main gear spring upper fairing with the gear-to-fuselage fairing. (8) Install only the upper screw that attaches the gear-to-fuselage fairing to the main-gear-spring upper fairing. (9) Align the main gear spring lower faring with the main-gear-spring upper fairing. Refer to Figure 202. NOTE:
The head proÞle of the screws is lower on new installations of the main-gear-spring upper fairing to the main-gear-spring lower fairing for Airplanes 2080533 and On or 208B02295 and On. If a new main-gear-spring upper fairing is to be installed, it should include the screws with the lower head proÞle.
NOTE:
For the main-gear-spring upper fairing, the hard Line-X coating can be cracked or missing. The coating only serves a cosmetic function.
(10) Install the upper screws that attach the main-gear-spring upper fairing to the main-gear-spring lower faring. (11) Install the inboard screws that attach the main-gear-spring lower faring. (12) Put the axle Þtting fairing in position with its slots behind the heads of the nuts and bolts that attach the axle Þtting to the main gear spring. Refer to Figure 201. (13) Install the screws that attach the axle Þtting fairing to the main gear spring lower faring. (14) Put the axle Þtting fairing in position to remove the force on the main-gear-spring lower fairing. (15) Make sure that the axle Þtting fairing slots are still in position behind the heads of the nuts and bolts that attach the axle Þtting to the main gear spring. Refer to Figure 201. (a) Torque the nuts on the bolts that attach the axle Þtting to the main gear spring. Refer to Torque Data - Maintenance Practices
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Main Landing Gear Fairing Installation Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
8.
Main Gear Axle Removal/Installation A.
Remove the Main Gear Axle (Refer to Figure 201). (1) Remove the main wheel. Refer to Wheels And Brakes - Maintenance Practices. (2) Remove the nuts and bolts that attach the torque plate and axle to the axle Þtting.
CAUTION: Make sure to put a support below the brake caliper and torque plate. This will help prevent damage to the brake line. (3) B.
Look at and record the initial position of the shim to make the camber check easier.
Install the Main Gear Axle. (1) Before you install the main gear axle, do a visual inspection of the inner surface for corrosion. NOTE:
(2) (3)
(4) (5) 9.
The main gear axle is a time limited part. You should verify the current number of landings, inspection status and replacement time before you install the main gear axle. Refer to Chapter 4, Replacement Time Limits and Typical Inspection Time Limits for the applicable inspection criteria.
Put the shim, axle and torque plate in position against the axle Þtting. (a) Make sure that the position of the shim is the same as the initial position when it was removed. Install the bolts, spacers, washers and nuts that attach the shim, axle and torque plate to the axle Þtting. (a) Torque the nuts on the bolts that attach the axle and torque plate to the axle Þtting. Refer to Torque Data - Maintenance Practices Install the main wheel. Refer to Wheels And Brakes - Maintenance Practices. Do a wheel camber check. Refer to Main Landing Gear - Adjustment/Test, Wheel Camber Check.
Main Landing Gear Spring Puller A.
Install the Main Landing Gear Spring Puller (Refer to Figure 201, Figure 203 and Table 201). (1) Remove the main wheel. Refer to Wheels And Brakes - Maintenance Practices. (2) Remove the main gear axle. Refer to Main Landing Gear - Maintenance Practices. (3) Remove the bolts, strap, plate (if applicable), washers and nuts that attach the axle Þtting to the main gear spring. (4) Remove the axle Þtting from the main gear spring. (5) Do all the steps of the main gear spring removal through the step to remove the pin from the attach trunnion. Refer to Main Gear Spring Removal/Installation. (6) Install the penetrating oil basin around the center spring and main gear spring joint. NOTE:
(a) (b) (c) (d) (e)
The penetrating oil basin procedure is optional. It can improve the main gear spring removal success rate. Cessna Propeller Aircraft Product Support, 316-517-5800 or Fax 316-942-9006, is the source to get the oil basin or the instructions to fabricate it.
Remove the sealant at the interface of the springs. Apply Type V Class E sealant to the mating surfaces of the oil basin halves. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. Put the oil basin in position around the joint. Install the bolts, washers and nuts that attach the oil basin to the springs. Apply Type V Class E sealant to the joints between the sides of the oil basin and the springs. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. NOTE:
(f)
The sealant will hold the penetrating oil in the oil basin.
Make sure that the drain plug is installed in the lower side of the oil basin.
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Main Landing Gear Spring Puller Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (g) (h) (i)
If necessary, remove the Þll plug and Þll the oil basin with Kroil penetrating oil. Install the Þll plug in the oil basin. Let the penetrating oil soak into the joint between the center spring and the main gear spring. NOTE:
If the penetrating oil soaks into the joint from 24 to 48 hours, it can improve the success rate for the main landing gear spring removal.
(j)
(7)
When it is time to remove the main gear spring, do the steps that follow: Remove the drain plug from the oil basin. 1 2 Drain the Kroil penetrating oil from the oil basin. 3 Install the drain plug in the oil basin. (k) Remove the bolts, washers and nuts that attach the oil basin to the springs. (l) Remove the oil basin from the springs. (m) Clean the area around the joint. Install the main landing gear spring puller on the main gear spring. NOTE:
Cessna Propeller Aircraft Product Support, 316-517-5800 or Fax 316-942-9006, is the source to get the instructions to fabricate the main landing gear spring puller.
CAUTION: Make sure to be careful when the main landing gear spring puller is installed. This will help prevent damage to the main gear spring during the installation process. (a) (b) (c) (d) (e) (f) (g) (h)
(8)
Put the support plate in position on the gear spring. Put the push plate in position on the gear spring. Align the plates of the spring puller on the ßoor. Loosely install the top spacer bolts, spacers and nuts (three places) in the spring puller. Put the spring puller in position on the gear spring. Loosely install the bottom spacer bolts, spacers and nuts (two places) in the spring puller. Loosely install the axle bolts and nuts (two places) in the spring puller. Install the puller in the spring puller. 1 Put the guard in position in the puller. 2 Put the puller with the guard in position between the plates of the spring puller. 3 Install the puller bolt and nut in the spring puller. a Torque the puller bolt from 40 to 60 foot pounds (54.2 to 81.3 N-m). (i) Tighten the spacer and axle bolts. Torque the bolts from 15 to 25 ft lb (20.3 to 33.9 N-m) 1 Put two similar hydraulic jacks in position between the push plate and the support plate. NOTE:
The correct position for the jacks between the plates is aligned with the marks on the plates.
(9) Connect both hydraulic jacks to the same pump. (10) If each jack can apply more than 12 tons of force, connect both jacks to the same pressure gage. NOTE: B.
This is the end of the main landing gear spring puller installation procedures. The puller can help during the main gear spring removal procedures.
Remove the Main Landing Gear Spring Puller (Refer to Figure 203). (1) If the gear spring is separated from the center spring and on the ßoor, do the steps that follow: (a) Remove the push plate and support plate from the gear spring. (b) Loosen the spacer and axle bolts. (c) Hold the puller and guard in position and remove the puller bolt and nut. (d) Remove the puller and guard from the spring puller. (e) Remove the axle bolts and nuts (two places) from the spring puller. (f) Remove the gear spring from the spring puller.
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MODEL 208 MAINTENANCE MANUAL (g)
(2)
Hold the plates in position and remove the spacer bolts, spacers, and nuts (Þve places) from the spring puller. If the gear spring is not separated from the center spring, do the steps that follow: (a) Remove the hydraulic jacks, if necessary. (b) Loosen the spacer bolts and axle bolts. (c) Hold the puller and guard in position and remove the puller bolt and nut. (d) Remove the puller and guard from the spring puller. (e) Remove the axle bolts and nuts (two places) from the spring puller. (f) Remove the spring puller from the gear spring. (g) Put the spring puller on the ßoor. (h) Remove the push plate and support plate from on the gear spring. (i) Hold the plates in position and remove the spacer bolts, spacers and nuts (Þve places) from the spring puller.
Table 201. Main Landing Gear Spring Puller Parts Name
Number
Comments
Spacer Bolt
NAS464-8-86
5 Required
Spacer
2680002-6
5 Required
Spacer Nut
MS21245-L8
5 Required
Puller Bolt
NAS464-17-82
1 Required
Puller Nut
MS21245-L16
1 Required
Axle Bolt
NAS464-7-86
2 Required
Axle Nut
MS21245-L7
2 Required
Penetrating Oil Basin
T2680002-15
1 Required
Basin Plug
MS21913D8
2 Required
Basin Bolt
AN4-10A
4 Required
Basin Washer
NAS1149F0432P
8 Required
Basin Nut
AN315-4R
4 Required
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MODEL 208 MAINTENANCE MANUAL MAIN LANDING GEAR - ADJUSTMENT/TEST 1.
2.
General A.
Correct main wheel alignment is important for maintaining tire wear within acceptable limits, and should be checked whenever excessive or abnormal wear is noted on tires.
B.
To correct alignment problems, special shims are utilized which change wheel camber. These shims are used in conjunction with Figure 501 and Figure 502 to produce correct camber under various airplane weights. These values are the sum of the values shown in the chart. Positive values will produce positive camber or toe-in. Negative values will produce negative camber or toe-out. Measurements are taken on wheel flange.
Wheel Toe-In Check A.
3.
Check Procedures (Refer to Figure 501). (1) Ensure airplane is sitting on a level surface. (2) Ensure tires are properly inflated. Refer to Chapter 12, Tires - Servicing. (3) Place main gear wheels on aluminum plates approximately 18.0 inches square. The plates should be resting on greased aluminum plates of the same dimension. (4) Establish airplane centerline on floor surface by dropping plumb bob line from center of forward jack point (located on forward nose gear drag link spring support) and from center of tail tiedown bracket (located on lower side of aft tail cone). Chalk a line on floor between two plumb bob points. (5) Using intersecting arc method, establish second line perpendicular to airplane centerline just forward of main gear tires and chalk line. (6) Using squares, wood blocks and long straightedge as shown in Figure 501, set up straightedge parallel to second chalk line just below level of axle nut. (7) Carefully roll airplane forward until tires just touch straightedge. (8) Place two marks on wheel flanges just below wheel nut level eleven inches apart. (9) Place carpenter’s square against straightedge, just outboard of wheel flange marks, and determine dimensions X and Y. Toe-in (for one wheel) is the difference between the two dimensions (i.e. Y-X). Compare this dimension with the chart on Figure 501, Sheet 3. (10) If toe-in dimension is not within specified tolerance, determine from charts on Figure 501 which shim (or shim combination and shim orientation) will establish specified tolerance.
Wheel Camber Check A.
Wheel camber is measured by reading a protractor level held vertically against the outboard flanges of the wheel. Refer to Figure 501, Sheet 3 for camber check procedures and Figure 502 for a weight versus camber chart.
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Main Wheel Alignment Figure 501 (Sheet 1)
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Main Wheel Alignment Figure 501 (Sheet 2)
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Main Wheel Alignment Figure 501 (Sheet 3)
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Weight Versus Camber Chart Figure 502 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL MAIN LANDING GEAR - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the main landing gear in a serviceable condition.
Task 32-10-00-220 2.
Main Landing Gear Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the main landing gear.
B.
Special Tools (1) Airplane Jacks (2) Tail Stand
C.
Access (1) None NOTE:
D.
The main landing gear fairings are removed during the inspection.
Do a Main Landing Gear Detailed Inspection. Refer to Main Landing Gear - Maintenance Practices, Figure 201. (1) Examine the left and the right main landing gear fairings for cracks, wear, loose rivets, distortion and broken or missing attachment hardware. NOTE:
For the main-gear-spring upper fairing, the hard Line-X coating can be cracked or missing. The coating only serves a cosmetic function.
(a)
(2) (3) (4) (5) (6) (7) (8) (9) E.
Make sure that no more than two plies are exposed on the main-gear-spring upper fairing. If two or more plies are exposed on the main-gear-spring upper fairing, it is recommended that you replace the fairing at or before the next scheduled maintenance interval. Remove the left and the right main landing gear fairings. Refer to Main Landing Gear Maintenance Practices, Main Gear Fairing Removal/Installation. Use jacks to lift the airplane. Refer to Chapter 7, Jacking - Maintenance Practices. Examine the main gear springs and attach trunnions for loose trunnion cap bolts, cracks, and corrosion. Make sure that the main gear springs and attach trunnions are correctly attached to the center spring and fuselage attach Þtting. Examine the brake lines for leaks. Make sure the brake lines are correctly attached to the main gear spring. Examine the axle Þttings for cracks, corrosion, pits, and any other obvious damage. Make sure that the axle Þttings are correctly attached to the main gear spring.
Do a Main Landing Gear Spring Inspection and Repair (Refer to Figure 601). (1) Examine the sealant at the main landing gear at the following locations for cuts, deterioration, dimensions and separation from the surfaces. Refer to Main Landing Gear Installation, Figure 201. (a) The main gear spring and the center spring interface area. (b) The center spring and trunnion interface area (inside and outside). (2) If no damage to the sealant is visible, continue with the inspection. (3) If there is damage to the sealant, do the Center Spring and Main Gear Spring Interface Area Special Detailed Inspection and Repair in this section. (4) Examine the left and the right main gear springs and the visible areas of the center spring for gouging, chaÞng, and corrosion. (5) Examine the visible areas of the center spring and main gear spring interface area for corrosion.
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Main Gear Spring Figure 601 (Sheet 1)
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Main Gear Spring Figure 601 (Sheet 2)
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Main Gear Spring Figure 601 (Sheet 3)
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Main Gear Spring Figure 601 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (6)
(7) (8) F.
If no gouging, chaÞng, or corrosion is found on the left and the right main gear springs and the visible areas of the center spring, and no corrosion is found at the interface area, do the steps that follow. (a) Lower the airplane and remove the jacks. Refer to Chapter 7, Jacking - Maintenance Practices. (b) Do the restoration at the end of this task. If corrosion is found at the visible interface area, do the Center Spring and Main Gear Spring Interface Area Special Detailed Inspection and Repair in this section. If gouging, chaÞng, or corrosion is found on the left and the right main gear springs and the visible areas of the center spring, continue with this inspection.
Prepare the Damaged Area for a Measurement. (1) Use abrasive cloths and brushes to clean the damaged area. NOTE:
(2)
For normal cleaning procedures the abrasive cloths are 180 grit or Þner. If it is necessary to remove heavy layers of scale or oxides, a steel brush or 150 grit abrasive cloth can be used.
Remove the paint in the damaged area.
CAUTION: Make sure that you do not use a chemical stripper on the main gear spring. (a) G.
Remove only enough paint to get the correct measurement of the damaged area.
Measure the damaged area. (1) Find if there is local or circumferential damage. NOTE:
(2)
Local damage is deÞned as damage that extends less than 0.50 inch wide in the circumferential direction. The length of the damage in the longitudinal direction is not limited. If more than one local damage area exists around a gear spring crosssection, the repair for circumferential damage is to be used. Circumferential damage is deÞned as damage that extends beyond a width of 0.50 inch in the circumferential direction. The length of the damage in the longitudinal direction is not limited.
Use a pin-type depth micrometer to measure the difference between the undamaged spring surface and the deepest portion of the damage. Record the measurement of the difference. (Refer to Figure 601 and Table 601.)
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Table 601. Maximum Repair Depth Gear Assembled MAXIMUM REPAIR DEPTH
REPAIR LOCATION Center Spring
Outer Surface
0.050 inch
Main Gear Spring
Outer Surface - Zone 1 Local Damage
0.050 inch
Outer Surface - Zone 1 Circumferential Damage
0.025 inch
Outer Surface - Zone 2 Local Damage
0.040 inch
Outer Surface - Zone 2 Circumferential Damage
0.020 inch
NOTE:
(a) (b)
The depth micrometer must have a tolerance of +0.001 or -0.001 inch to make this measurement.
When the measurement is less than or equal to 0.005 inch, no repair is necessary. When the measurement is more than 0.005 inch, but is less than or equal to the maximum repair depth, repair the damage. The permitted spring repair depths are related to the width of the damage in the 1 circumferential direction. If there is local damage in Zone 2, then the maximum repair dimensions are a 0.040 inches deep and 0.50 inches wide in the circumferential direction. If there is local damage in Zone 1, then the maximum repair dimensions are b 0.050 inches deep and 0.50 inches wide in the circumferential direction. c If there is circumferential damage in Zone 2, then the maximum repair depth is 0.020 inches. d If there is circumferential damage in Zone 1, then the maximum repair depth is 0.025 inches. Repair the damaged area. 2
CAUTION: Make sure to remove only the necessary amount of material from the damaged area. Do not increase the depth of the damaged area when you remove the material. This will help prevent the replacement of springs that can be repaired. CAUTION: Use small hand-held type tools to do the repair procedure. Do not blend in one area for a long time. This will prevent damage from too much heat in one area of the spring material. a b c
Use a blending procedure to repair the damage and to get a smooth length-todepth ratio between the damage and the adjacent area. Make sure that only enough material is removed to get a lengthwise blending transition ratio of 20 to 1 in the longitudinal direction. Make sure that only enough material is removed to get a blending transition of width-to-depth ratio of 5 to 1 in the circumferential direction.
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MODEL 208 MAINTENANCE MANUAL d
Use a pin type depth micrometer to measure the depth of the repaired area. Refer to Table 601 for the permitted depths. NOTE:
(c)
A micrometer with a tolerance of +0.001 or -0.001 inch is necessary to make this measurement.
When the measurement is more than the maximum repair depth, replace the spring. Refer to Main Landing Gear - Maintenance Practices.
H.
Do a Magnetic Particle Inspection of the Repaired Area of the Main Gear Spring(s) and/or Center Spring for Cracks. Refer to the Model 208, Nondestructive Testing Manual, Part 8, Main Gear Spring. (1) When the magnetic particle inspection of the repaired area of the main gear spring(s) is complete, do the steps that follow. (a) If cracks are found, replace the main gear spring(s). Refer to Main Landing Gear Maintenance Practices. (b) If no cracks are found, and only the top region of Zones 1 and/or 2 were repaired, apply touch up paint to the main gear spring as necessary. Refer to the Main Landing Gear Cleaning/Painting. (c) If no cracks are found, and the bottom region of Zones 1 and/or 2 were repaired, do a shot peening procedure. Use size 330, cast steel shot and do a shot peening procedure to a Almen intensity 1 of 0.12 - 0.16A on the repaired area. Refer to Model 208 Structural Repair Manual Chapter 51, Shot Peening of Ferrous and Nonferrous Metals. (2) When the magnetic particle inspection of the repaired area of the center spring is complete, do the steps that follow. (a) If cracks are found, replace the center spring. Refer to Main Landing Gear - Maintenance Practices. (b) If no cracks are found, and the center spring was repaired, do a shot peening procedure. Use size 330, cast steel shot and do a shot peening procedure to a Almen intensity 1 of 0.12 - 0.16A on the repaired area. Refer to Model 208 Structural Repair Manual Chapter 51, Shot Peening of Ferrous and Nonferrous Metals. (3) Use Type 1, Class B-1/2 sealant and apply a Þllet seal around the main gear spring at the center spring. Refer to Main Landing Gear Installation, Figure 201. (4) Lower the airplane and remove the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
I.
Do a Landing Gear Free Play Check. (1) Jack the airplane, refer to Chapter 7. (2) Remove the wheel from the axle. (3) Install a magnetic base dial indicator on the uppermost point of the center spring outboard of the trunnion Þtting. Refer to Figure 604. (4) Position dial indicator so that tip is 2.0 inches (50.8 mm) from outboard end of the center spring. (5) Zero the dial indicator. While observing the dial indicator, lift the end of the axle by hand until dial indicator pointer does not increase any more. (6) Make sure that the total dial indicator reading does not change by more than 0.0140 inch (0.3556 mm). (a) If reading is acceptable, restore airplane. (7) If indicator reading exceeds 0.0140 inch (0.3556 mm), disassemble main gear spring from center spring. (8) Using a micrometer, make sure that the difference between the outer diameter of the main gear spring and the inner diameter of the center spring does not exceed 0.0133 inch (0.3378 mm). (9) If tolerance is exceeded, replace defective parts. (10) Restore the airplane.
J.
Restore Access (1) Install the left and the right main landing gear fairings. Refer to Main Landing Gear - Maintenance Practices, Main Gear Fairing Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL Task 32-10-00-221 3.
Center-Spring and Main Gear-Spring Interface Area Special Detailed (Corrosion Inspection and Repair) A.
General (1) This task gives the procedures to do a detailed inspection and repair of the center spring and main gear spring interface area.
B.
Special Tools (1) Airplane Jacks (2) Tail Stand
C.
Access (1) None NOTE:
D.
Do a Center Spring and Main Gear Spring Interface Area Detailed Inspection and Repair (Refer to Figure 602 and Figure 603). (1) Remove and disassemble the main landing gear assembly. Refer to Main Landing Gear Maintenance Practices.
(2)
E.
The main landing gear fairings are removed during the inspection.
NOTE:
The main landing gear fairings are removed during the landing gear removal procedure.
NOTE:
The airplane is lifted on jacks during the landing gear removal procedure.
Examine the center spring and main gear spring interface area for gouging, chaÞng, or corrosion. (a) If no gouging, chaÞng, or corrosion is found, install the main landing gear assembly. Refer to Main Landing Gear - Maintenance Practices. (b) If gouging, chaÞng, or corrosion is found, prepare the damaged area for a measurement.
Prepare the Damaged Area of the Interface Area for a Measurement. (1) Use abrasive cloths and brushes to clean the damaged area. NOTE:
(2)
For normal cleaning procedures the abrasive cloths are 180 grit or Þner. If it is necessary to remove heavy layers of scale or oxides, a steel brush or 150 grit abrasive cloth can be used.
Remove the paint in the damaged area.
CAUTION: Make sure that you do not use a chemical stripper on the main gear spring. (a) F.
Remove only enough paint to get the correct measurement of the damaged area.
Measure the Damaged Area of the Interface Area. Refer to Figure 602 and Table 602.
Table 602. Maximum Repair Depth Interface Area REPAIR LOCATION Center Spring
MAXIMUM REPAIR DEPTH Inner Surface - Interface Area
DIAMETER
0.050 inches
Model 208
2.703 inches (max)
Model 208B
2.794 inches (max)
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MODEL 208 MAINTENANCE MANUAL Table 602. Maximum Repair Depth Interface Area (continued) REPAIR LOCATION Main Spring
(1)
Outer Surface - Interface Area
0.050 inches
Model 208
2.698 inches (min)
Model 208B
2.789 inches (min)
The recommended transducer for this inspection should have a diameter of less than 0.5" (12.5 mm) and an operating frequency of 5-10 MHz.
Calibrate the thickness gauge in accordance with the manufacturers instructions. NOTE:
(a)
Use a 4130 or 4340 stepwedge. The stepwedge must have thickness steps that represent the wall thickness of the tubes. If a stepwedge is not available, use a landing gear tube for the calibration.
Do the steps that follow to calibrate with a landing gear tube: Measure the wall thickness at each end of the tube with a tube micrometer. 1 Record the values for reference. 2 3 Calibrate the gauge using the thickness values found with the tube micrometer. NOTE:
(3)
DIAMETER
Use an ultrasonic thickness gauge to measure and record the wall thickness of the main gear spring and center spring. NOTE:
(2)
MAXIMUM REPAIR DEPTH
The transducer must be placed in the same location on the tube that the measurements were made.
Measure the main gear spring and center spring wall thickness as follows. Refer to Figure 602. NOTE:
Each of the three locations on the spring will have three measurements taken at 120 degree intervals around the spring circumference. There will be a total of nine measurements on each spring end.
(a) (b) (c)
(4)
Place the transducer on the main spring interface area. Record the wall thickness. Reposition the transducer 120 degrees around the circumference from the Þrst location and take a measurement. (d) Record the wall thickness (e) Reposition the transducer another 120 degrees and take a measurement.. (f) Record the wall thickness (g) Repeat the measurement procedure with the center spring. (h) If more than 20 percent of the measurements on the center spring in the interface area are more than the maximum wall thickness, replace the center spring. Refer to Main Landing Gear - Maintenance Practices. (i) If more than 20 percent of the measurements on the main gear spring in the interface area are less than the minimum diameter, replace the main gear spring. Refer to Main Landing Gear - Maintenance Practices. Use a pin type depth micrometer to measure the difference between the undamaged spring surface and the deepest portion of the damage. Record the measurement of the difference. NOTE:
(a)
A micrometer with a tolerance of +0.001 or -0.001 inch (+0.025 or -0.025 mm) is necessary to make this measurement.
If the depth of damage is greater than the permitted maximum repair depth, replace the center spring or main gear spring. Refer to Main Landing Gear - Maintenance Practices.
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Main Gear Interface Area Figure 602 (Sheet 1)
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Main Gear Interface Area Figure 602 (Sheet 2)
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Main Gear Interface Area Figure 602 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (b)
If the measurement is less than or equal to the maximum repair depth, repair the damage. NOTE:
1
The permitted interface area spring repair depth is 0.050 inches (1.270 mm) or less.
Repair the damaged area.
CAUTION: Make sure to remove only the necessary amount of material from the damaged area. Do not increase the depth of the damaged area when you remove the material. This will help prevent the replacement of springs that can be repaired. CAUTION: Use small hand-held type tools to do the repair procedure. Make sure not to stay in one spot for a long time. This will help to prevent too much heat in one area of the spring material. a b c d
Use a blending procedure to repair the damage and to get a smooth length-todepth ratio between the damage and the adjacent area. Make sure that only enough material is removed to get a lengthwise blending transition ratio of 20 to 1 in the longitudinal direction. Make sure that only enough material is removed to get a blending transition of width-to-depth ratio of 5 to 1 in the circumferential direction. Use a pin type depth micrometer to measure the depth of the repaired area. Refer to Table 602 for the permitted depths. NOTE:
(c) G.
A micrometer with a tolerance of +0.001 or -0.001 inch (+0.025 or -0.025 mm) is necessary to make this measurement.
When the measurement is more than the maximum repair depth, replace the spring. Refer to Main Landing Gear - Maintenance Practices.
Do a Magnetic Particle Inspection of the Repaired Area of the Interface Area of the Main Gear Spring(s) and/or Center Spring for Cracks. Refer to the Model 208, Nondestructive Testing Manual, Part 8, Main Gear Spring. (1) When the magnetic particle inspection of the repaired area of the interface area of the main gear spring(s) is complete, do the steps that follow. (a) If cracks are found, replace the main gear spring(s). Refer to Main Landing Gear Maintenance Practices. (b) If no cracks are found, and the main gear spring interface area was repaired, do a shot peening procedure. Use size 330, cast steel shot and do a shot peening procedure to a Almen intensity 1 of 0.12 - 0.16A on the repaired area. Refer to Model 208 Structural Repair Manual Chapter 51, Shot Peening of Ferrous and Nonferrous Metals. (2) When the magnetic particle inspection of the repaired area of the interface area of the center spring is complete, do the steps that follow. (a) If cracks are found, replace the center spring. Refer to Main Landing Gear - Maintenance Practices. (b) If no cracks are found, and the center spring interface area was repaired, do a shot peening procedure. Use size 330, cast steel shot and do a shot peening procedure to a Almen intensity 1 of 0.12 - 0.16A on the repaired area. Refer to Model 208 Structural Repair Manual Chapter 51, Shot Peening of Ferrous and Nonferrous Metals.
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MODEL 208 MAINTENANCE MANUAL (3)
Install the main landing gear assembly. Refer to Main Landing Gear - Maintenance Practices. NOTE:
(4) H.
The airplane is lowered, and the jacks and tail stand are removed during the landing gear installation procedure.
Record the necessary airplane and inspection information on the Model 208 Main Landing Gear Separation Data Form and send it to the address or fax it to the number on the form.
Restore Access (1) None NOTE:
The main landing gear fairings are installed during the landing gear installation procedure.
End of task Task 32-10-00-240 4.
Main Landing Gear Axle Special Detailed Inspection (SID) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the main landing gear axle in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do a Special Detailed Inspection of the Main Landing Gear Axle. NOTE: (1)
The main landing gear axle is replaced at 10,000 landings.
Examine the main landing gear axle for fatigue and cracks in the main landing gear axle radius. Refer to the Model 208 Nondestructive Testing Manual, Part 8, Magnetic Particle, Main Landing Gear Axle - Description And Operation.
E.
Restore Access (1) None End of task
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Model 208 Main Landing Gear Separation Data Form Figure 603 (Sheet 1)
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Model 208 Main Landing Gear Separation Data Form Figure 603 (Sheet 2)
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Free Play Check Figure 604 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL MAIN LANDING GEAR - CLEANING/PAINTING 1.
2.
General A.
This procedure provides cleaning and painting instructions for main landing gear components. These components are constructed of aluminum alloy, steel alloy and magnesium alloy. The following anticorrosion treatment processes are applied at time of manufacture: (1) Aluminum Alloy Parts: (a) Chemically degreased. (b) Chemically film-conversion-coated. (c) Epoxy primed. (d) Top coated with Polyurethane paint. (2) Steel Alloy Parts: (a) Chemically degreased. (b) Epoxy primed. (c) Top coated with Polyurethane paint or white polyester powder coat. (3) Magnesium Alloy Parts: (a) Chemically degreased. (b) Chemically film-conversion-coated. (c) Epoxy primed. (d) Top coated with Polyurethane paint.
B.
For an illustration of main landing gear construction materials, refer to Figure 701.
Tools, Equipment and Materials A.
Refer to Landing Gear - General for a list of required tools, equipment and materials. NOTE:
3.
Follow the directions of the manufacturer or supplier for storing, mixing, and applying spray wash primers, brush chem-film primers, epoxy primers, and topcoats.
Refinishing High Stressed Steel Shot Peened Surfaces
WARNING: The main gear legs and the main gear center tube outer surfaces are shot peened during final manufacture and prior to the application of protective coatings. The shot peened surface is thin and must not be disturbed or damaged. Do not use chemical strippers of any kind to remove paint from shot peened surfaces. Chemical strippers have acids that may cause hydrogen embrittlement. Also, do not sand or sand blast these surfaces. A.
Refinishing minor nicks and scratches of high stressed shot peened surfaces. (1) Using a soft, hand steel wire brush or Scotch Brite pad (no power tools), gently remove loose paint, scale, and rust. Use good judgement to prevent damaging the shot peened surface. For large areas, the preferred method is media stripping. This includes glass bead, plastic bead, or wheat starch. (2) If the primer is damaged in an area larger than the size of a dime, the area should be hand solvent cleaned with an approved solvent and reprimed with epoxy primer. (3) Spot paint the damaged areas with Polyurethane paint.
B.
Complete refinishing of high stressed steel shot peened surfaces. (1) To return these components to their factory finish, remove all paint by media stripping using glass bead, plastic bead, or wheat starch. (2) Remove all media and loose particles with compressed air, clean with an approved solvent and prime with epoxy primer. (3) Apply Polyurethane topcoat.
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Main Landing Gear Construction Materials Figure 701 (Sheet 1)
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4.
Refinishing Steel, Aluminum and Magnesium Components A.
Refinishing of minor nicks and scratches. (1) Feather sand edges of finish around the effected area with 320 grit sandpaper, followed by 400 grit. Avoid sanding through the primer if possible. (2) For steel parts, if primer is damaged in an area larger than the size of a dime, the area should be hand solvent cleaned with an approved solvent then reprimed with epoxy primer. (3) For aluminum and magnesium parts. If primer is damaged in an area larger than the size of a dime, the area should be hand solvent cleaned with an approved solvent then apply a spray wash primer or brush chem-film primer. (a) After spray wash primer or brush chem-film primer has dried for at least 30 minutes, apply epoxy primer. (4) Spot paint damaged areas with Polyurethane paint.
B.
Complete refinishing. (1) Degrease and remove sealants and heavy soil with approved solvents. (2) Strip original finish or part following recommendations of stripper manufacturer.
CAUTION: Do not allow stripper to come in contact with the main landing gear legs, center tube or axle spindles. Wear protective clothing and avoid contact with skin. Use of chemicals requires good ventilation and good fire safety practices. (3) (4) (5) (6) (7)
Use steel wire brush, Scotch Brite or fine aluminum oxide paper to remove any remaining loose paint, scale and rust. Hand solvent clean with an approved solvent. Apply epoxy primer to steel parts. Apply spray wash primer or brush chem-film primer to aluminum or magnesium parts. Apply epoxy primer to aluminum or magnesium parts. Apply Polyurethane topcoat.
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MODEL 208 MAINTENANCE MANUAL NOSE LANDING GEAR - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Nose Landing Gear Troubleshooting Chart Figure 101 (Sheet 1)
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Nose Landing Gear Troubleshooting Chart Figure 101 (Sheet 2)
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Nose Landing Gear Troubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL NOSE LANDING GEAR - MAINTENANCE PRACTICES 1.
2.
General A.
This section gives procedures for the removal/installation of the drag link spring and drag link spring support liner. Also included in this section are procedures to inspect and repair the drag link spring, surface rework instructions for the grease seal bore and the cup backing bore, and information about the time limits for the nose gear assembly replacement.
B.
A longitudinal nose gear fairing extends aft that covers the shimmy dampener, upper part of the shock strut and the drag link spring. The shock strut trunnion is attached to the lower forward engine mount at two pivot lugs. The drag link spring is attached at the upper part of the wheel fork and to the lower side of the fuselage using two bearing blocks, in tandem. This arrangement allows for easy removal of the complete nose gear assembly when replacement is required, or when installing floats. For more information about the shimmy dampener, nose gear shock strut and nose gear fairing, refer to Shimmy Dampener - Maintenance Practices, Nose Gear Shock Strut - Maintenance Practices and Nose Gear Fairing - Maintenance Practices.
C.
Vertical loads encountered when you land and taxi are absorbed by the drag link spring and the nose gear shock strut. Minor loads, such as when you taxi, are absorbed primarily by the drag link spring, but as the rate of application of loads increases, such as when you land, a larger and larger proportion of the total load is absorbed by the shock strut. For more information about the nose gear shock strut, refer to Nose Gear Shock Strut - Maintenance Practices.
D.
The nose wheel is steerable through an arc of 15 degrees each side of center by use of the rudder pedals. By applying brakes, the angle may be increased up to 56 degrees either side of center.
Drag Link Spring Removal/Installation A.
Remove the Drag Link Spring (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7)
The nose gear drag link spring is formed to put 150 pounds, +25 or -25 pounds, preload on the extended nose gear strut. Shims are used to adjust the preload.
Lift the nose of airplane with jacks. Refer to Chapter 7, Jacking - Maintenance Practices. Remove the safety wire from the safety clips. Remove the bolts that attach the drag link spring fork to the bearing support. Pivot the nose gear strut forward to clear the drag link spring fork. Release the preload on the drag link spring. Keep the safety clip for installation. Hold the engine with a hoist and a sling. Refer to Powerplant - General. NOTE:
The hoist and sling will hold the front of the airplane while the jack is removed.
(8)
Remove the nose of the airplane from the jacks. Refer to Chapter 7, Jacking - Maintenance Practices. (9) Hold the drag link spring. (a) Remove the bolts, washers, and nuts that attach the forward drag link spring support to the drag link spring. (b) Remove the bolts and washers that attach the aft drag link spring support to the drag link spring. (c) Keep the shims, if installed, for installation. (10) Remove the drag link spring from airplane.
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Nose Landing Gear Installation Figure 201 (Sheet 1)
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Nose Landing Gear Installation Figure 201 (Sheet 2)
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Nose Landing Gear Installation Figure 201 (Sheet 3)
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Nose Landing Gear Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL B.
Install the Drag Link Spring (Refer to Figure 201). NOTE:
(1)
(2) (3) (4)
The nose gear drag link spring is formed to put 150 pounds, +25 or -25 pounds, preload on the extended nose gear strut. Shims are used to adjust the preload. The maximum thickness of the shims at either drag link spring support must not exceed 0.125 inch.
Put the drag link spring in position against the airplane. (a) Put the shims, if installed, in position. (b) Install the bolts, washers, and nuts that attach the forward drag link spring support to the drag link spring. (c) Install the bolts and washers that attach the aft drag link spring support to the drag link spring. Lift the nose of airplane with jacks. Refer to Chapter 7, Jacking - Maintenance Practices. Remove the hoist and a sling. Refer to Powerplant - General. Measure the distance between the centers of the attach holes in the bearing support and the drag link spring fork. The distance must measure 0.96 inch, +0.13 or -0.13 inch. (a) If the distance is greater, put shim(s), as necessary, between the aft support and the fuselage until the dimension is within the desired dimension. (b) If the distance is less, put shim(s), as necessary, between the forward support and the fuselage until the dimension is within the desired dimension. NOTE:
(5)
Put the safety clips in position on the bearing support. NOTE:
(6) (7) (8) 3.
On airplanes 2080001 thru 20800259 and 208B0001 thru 208B0581, CAB96952 must be included before the safety clip is installed.
The safety clips must fit tightly against the bearing support surface. The safety clips must not be able to turn.
Align the bolt attach holes in the drag link spring fork with the holes in the bearing support. (a) Install the bolts and washers that attach the drag link spring fork to the bearing support. (b) Torque the bolts from 60 to 85 inch-pounds. Install safety wire on the safety clips. Remove the airplane from the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
Drag Link Spring Support Liner Removal/Installation A.
Remove the Support Liner (Refer to Figure 201). (1) Remove the drag link spring. Refer to Drag Link Spring Removal and Installation (2) Remove the bolt, washer and nut that attach the aft drag link spring support to the drag link spring. (3) Remove the aft drag link spring support from the drag link spring. (4) Remove the forward drag link spring support from the drag link spring. (5) Remove the liner assembly from the forward drag link spring support.
B.
Install the Support Liner (Refer to Figure 201). (1) If the liner is not attached to the bushing, do the steps that follow: (a) Apply Loctite 609 or Loctite 680 Retaining Compound to the interior surface of the bushing. NOTE:
(2) (3) (4) (5) (6)
Do not use primer on the bushing.
(b) Immediately push the liner into the bushing. (c) Clean off the unwanted Loctite. Push the liner assembly into the forward drag link spring support. Use a staking tool to keep the liner assembly in place. Refer to Chapter 20, Bearings - Removal/ Installation. Install the forward drag link spring support on the drag link spring. Install the aft drag link spring support on the drag link spring. Install the bolt, washer and nut that attach the aft drag link spring support to the drag link spring.
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Install the drag link spring. Refer to Drag Link Spring Removal and Installation.
Nose Landing Gear Drag Link Spring Inspection/Repair
CAUTION: The damage must not be more than one square inch in area. This repair procedure is applicable for the repair of small areas of damage. Contact Cessna Propeller Aircraft Product Support for repair and replacement instructions for larger areas of damage. NOTE: A.
The repair procedure that follows may have lower life limits as a cause of the depth of damage that is repaired. Refer to Chapter 4, Replacement Time Limits - Description and Operation.
Do an Inspection of the Nose Landing Gear Drag Link Spring (Refer to Figure 202). (1) Remove the nose gear fairing to get access to the nose landing gear drag link spring. (2) Lift the airplane up on jacks. Refer to Chapter 7, Jacking - Maintenance Practices. (3) Remove the nose landing gear drag link spring. Refer to Drag Link Spring Removal/Installation. (4) Do an inspection of the nose landing gear drag link spring. Refer to Nose Landing Gear Inspection/Check. (a) If you do not find any damage, then go to step 4.B.(5) that gives instructions to install the nose landing gear drag link spring. (b) If damage is found, then go to the next step to measure the gouged areas of the landing gear spring. (5) Measure the gouged areas of the landing gear spring.
CAUTION: Do not remove paint from the landing gear spring with chemical stripper. If you use chemical stripper, the integrity of the shot peened surface will be compromised. (a)
Remove only enough paint and surface material to accurately measure the damaged area. NOTE: 1 2
Paint and scale or oxide are removed in the immediate area effected by the inspection, to get a correct measurement of the damaged area.
Use a 180 grit or finer abrasive where heavy layers of scale or oxide are to be removed. Use a steel brush or size 150 grit abrasive for any remaining paint, scale or oxide.
CAUTION: Use a pin-type depth micrometer with a tolerance of -0.001 or +0.001 when you measure the depth of the damage. When you measure the depth of the damage, accuracy is very important because of the high possibility of removal of the thin layer of shot peened material. (6)
Measure and record the difference between the undamaged spring surface and the most severe area of the damage with a pin-type depth micrometer. (a) If the measurement is equal to or less than 0.005 inch (0.127 mm), the damage does not need to be blended. 1 Prepare and touch-up the paint on the nose landing gear drag link spring. Refer to Nose Landing Gear - Cleaning and Painting. 2 Go to step 4.B.(5) that gives instructions to install the nose landing gear drag link spring.
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Nose Landing Gear Drag Link Spring Inspection/Repair Figure 202 (Sheet 1)
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Nose Landing Gear Drag Link Spring Inspection/Repair Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (b)
B.
If the measurement is more than 0.005 inch (0.127 mm), but is less than or equal to 0.075 inch (1.905 mm), repair or replace the nose landing gear spring link. 1 If damage is between 0.006 inch (0.152 mm) and 0.050 inch (1.270 mm), the damaged area must be blended out of the nose landing gear drag link spring within the next 200 landings. a If it is necessary for the nose gear drag link spring damage to be blended and an inspection done, then go to Repair the Nose Gear Drag Link Spring. b If it is necessary for the nose gear drag link spring to be blended and have an inspection done at a different time, then prepare and touch up the paint on the nose gear drag link spring as necessary and go to step 4.B.(5) that gives procedures to install the nose landing gear drag link spring. 2 If the damage is between 0.051 inch (1.295 mm) and 0.075 inch (1.905 mm), do an inspection of the nose landing gear drag link spring. a Do a magnetic particle inspection. Refer to the Model 208 Nondestructive Testing Manual, Part 8, Nose Landing Gear Spring. b Contact Cessna Propeller Aircraft Product Support at (316) 517-5800 or Fax (316) 942-9006, for more instructions. 3 Continued flight is permitted if no cracks are found, but you must obey operational limitations and restrictions while you wait for a Cessna disposition. a Use slow movements when you tow the airplane. b Use only paved runways. c Do not let the landing gear strut compress to the end of its travel, such as during a hard landing or a quick decrease in speed. d If a hard landing or quick decrease in speed that compress the landing gear strut to the end of its travel occur, then replace the nose gear drag link spring before flight. 4 If the measurement is more than 0.076 inch (1.930 mm), then replace the nose landing gear drag link spring. Refer to Drag Link Spring Removal/Installation.
Repair the Nose Gear Drag Link Spring (Refer to Figure 202). (1) Remove the paint and scale or oxide as necessary to clean the area to be blended. NOTE: (a) (b)
It is necessary to remove only the paint in the area you will repair.
Use a 180 grit or finer abrasive where heavy layers of scale or oxide are to be removed. Use a steel brush or size 150 grit abrasive for any remaining paint, scale or oxide.
CAUTION: Do not exceed 50 PSI when glass bead blasting the surface. (2)
Only do the bead blasting in the damaged area that needs to be blended. NOTE:
All of the surface will have the same satin finish if it is correctly bead blasted.
CAUTION: Remove only the amount of material necessary. If you remove too much material, you will increase the damaged area and replacement of the spring may be necessary. CAUTION: Use small hand-held type tools to do the blending procedures. Do not to do the blending procedures in one area for a long time because the spring material can get too hot in one area and cause damage. (3)
The damaged area must be blended until you get a smooth transition between the damage and the immediate adjacent areas. (a) Make sure only a minimum amount of material is removed. 1 Use a ratio of 20 to 1 along the vertically blended surface 2 Use a ratio of 5 to 1 across the circumference of the blended surface.
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(5) (6) (7) 5.
Measure the deepest part of the damaged area with a pin-type depth micrometer to find the depth of the damaged area to be blended. (a) Refer to the measurement you recorded of the difference between the undamaged spring surface and the deepest area of the damage in step 4.A.(6). 1 If the maximum depth of the damaged area after you blend the surface is 0.055 inch or less, go to step 4.B.(5) that gives instructions to install the nose landing gear drag link spring. 2 If the maximum depth of the damaged area after you blend the surface is greater than 0.055 inch (1.397 mm), contact Cessna Propeller Aircraft Product Support at (316) 942-9006, for more instructions. 3 Continued flight is permitted if no cracks are found, but you must obey operational limitations and restrictions while you wait for a Cessna disposition. a Use slow movements when you tow the airplane. b Use only paved runways. c Do not let the landing gear strut compress to the end of its travel, such as during a hard landing or a quick decrease in speed. d If a hard landing or quick decrease in speed that compress the landing gear strut to the end of its travel occur, then replace the nose landing gear drag link spring before flight. 4 If the maximum depth of the damaged area after you blend the surface is more than 0.075 inch (1.905 mm), then you must replace the nose landing gear drag link spring. Refer to step 4.B.(5) that gives instructions to install the nose landing gear drag link spring. Install the nose landing gear drag link spring. Refer to Drag Link Spring Removal/Installation. Install the nose landing gear drag link spring fairing. Refer to Nose Gear Fairing - Maintenance Practices. Remove the airplane from the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
Nose Wheel Grease Seal Bore And Cup Backing Bore Surface Rework NOTE: A.
The following procedure covers rework techniques for the nose wheel halves that have developed corrosion or pits on the grease seal bore or cup backing bore surfaces.
Removing Pits or Corrosion (Refer to Figure 203). (1) Pits or corrosion on the grease seal bore surface or cup backing bore surface can be removed by using 400 grit emery and hand polishing techniques. Care must be taken not to exceed the diameters indicated when working wheels. NOTE: (2)
B.
Do not use shop machinery (lathes, etc.) to clean up corrosion. Use hand polishing techniques.
If surface pits are removed and pits within the surface remain, they can be removed by glass beading at low pressure, being careful not to remove the parent material on the lip seal bore area. A good quality commercial stripper or acetone may also be used to remove embedded corrosion. Embedded pits must not extend deeper than 0.025 inch below the surface of the bore diameter indicated in Figure 203. If corrosion on surface A of Figure 203 is such that contaminants could enter the bearing bore area, wheel half must be replaced.
Returning Bearings to Service (Refer to Figure 203). (1) After removing pits or corrosion as best as possible, clean the area with mineral spirits or equivalent solution and dry with filtered compressed air. Because hand polishing or glass beading will have removed most of the anodize coating, it will be necessary to coat surfaces A and B with a protective coating. Chemically film-treat surfaces A and B using lridite 14-2 (or equivalent). Brush or swab a liberal coating over bare metal area being worked. Best results are achieved if coating is allowed to set for two to four minutes. Then, thoroughly rinse part with hot water (100°F to 110°F). Allow to dry. (2) Within 48 hours of chemical treatment, apply primer (Desoto 513X371 with 910X565 curing solution or equivalent) to surfaces by spraying or hand swabbing, using care not to cover the bearing cup surface C of Figure 203 with primer.
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Wheel Hub and Bearing Cutaway Figure 203 (Sheet 1)
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After adequate curing time (refer to primer manufacturers specifications), the wheels may be returned to service.
Nose Gear Assembly Replacement Information A.
The nose gear assemblies limits information is found in Chapter 4, Replacement Time Limits.
B.
This section shows the nose gear assemblies that you can replace (Refer to Figure 204).
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Nose Gear Replacement Assemblies Figure 204 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL NOSE LANDING GEAR - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the nose landing gear in a serviceable condition.
Task 32-20-00-220 2.
Nose Landing Gear Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the nose landing gear.
B.
Special Tools (1) None
C.
Access (1) Remove the nose wheel fender (if installed), drag link spring fairing, and the lower cowling to get access to nose gear attach points.
D.
Do a Nose Landing Gear Detailed Inspection. (1) Inspect the drag link spring fairing for cracks, wear, loose rivets, broken or missing attachment hardware. (2) Jack the airplane. Refer to Chapter 7, Jacking - Maintenance Practices. (3) Move the nose gear attach points manually at the engine mount and examine for looseness. (4) Examine the nose gear drag link spring attach structure for cracks, loose bolts, elongated holes, and corrosion. (a) Examine the nose gear spring yoke for corrosion, security of installation and freedom of rotation at the bearing. (b) Lubricate the nose gear spring yoke bearing. Refer to Chapter 12, Landing Gear Servicing. (5) Examine the nose gear shock strut for evidence of hydraulic leakage. (6) Examine the nose wheel fork for damage, corrosion and security of installation. (7) Examine the torque links for general condition, wear at the attach points, and security of installation. (a) Lubricate the torque links at the five lubrication points. Refer to Chapter 12, Landing Gear - Servicing. (8) Examine the nose gear trunnion bearings for looseness. (a) Lubricate the three lubrication points. Refer to Chapter 12, Landing Gear - Servicing. (9) Examine the nose gear shimmy damper for general condition, security, and freedom of movement through its full range of travel. (10) Examine the nose gear steering bungee attachment at the steering bellcrank. (11) Lower and remove the airplane from the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
E.
Restore Access (1) Install the nose wheel fender (if removed), drag link spring fairing, and the lower cowling. End of task 3.
Nose Wheel Grease Seal Bore and Cup Backing Bore Surfaces A.
Inspection Procedures. (1) Inspect grease seal bore and cup backing bore in nose gear wheel halves for corrosion and pitting due to water accumulation. If evidence of corrosion or pitting is discovered, rework in accordance with Nose Landing Gear - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL NOSE LANDING GEAR - CLEANING/PAINTING 1.
2.
General A.
This procedure provides cleaning and painting instructions for nose landing gear components. These components are constructed of aluminum and steel alloy. The following anticorrosion treatment processes are applied at time of manufacture: (1) Aluminum Alloy Parts: (a) Chemically degreased. (b) Chemically film-conversion-coated. (c) Epoxy primed. (d) Top coated with Polyurethane paint. (2) Steel Alloy Parts: (a) Chemically degreased. (b) Epoxy primed. (c) Top coated with Polyurethane paint or white polyester powder coat.
B.
For an illustration of nose landing gear construction materials, refer to Figure 701.
Tools, Equipment and Materials A.
Refer to Landing Gear - General for a list of required tools, equipment and materials. NOTE:
3.
Follow the directions of the manufacturer or supplier for storing, mixing and applying spray wash primers, brush chem-film primers, epoxy primers, and topcoats.
Refinishing High Stressed Steel Shot Peaned Surfaces
WARNING: The outer surface of the nose gear drag link spring is shot peened during final manufacture and prior to the application of protective coatings. The shot peened surface is thin and must not be disturbed or damaged. Do not use chemical strippers of any kind to remove paint from shot peened surfaces. Chemical strippers have acids that may cause hydrogen embrittlement. Also, do not sand or sand blast the nose gear drag link spring.
4.
A.
Refinishing minor nicks and scratches on the nose gear drag link spring. (1) Using a soft, hand wire brush or ScotchBrite pad (no power tools), gently remove loose paint, scale, and rust. Use good judgement to prevent damaging the shot peened surface. For large areas, the preferred method is media stripping. This includes glass bead, plastic bead, or wheat starch. (2) If the primer is damaged in an area larger than the size of a dime, the area should be hand solvent cleaned with an approved solvent and reprimed with epoxy primer. (3) Spot paint the damaged areas with Polyurethane paint.
B.
Complete refinishing of the nose gear drag link spring. (1) Remove the spring from the airplane. (2) Remove all paint by media stripping using glass bead, plastic bead, or wheat starch. (3) Remove all media and loose particles with compressed air, clean with an approved solvent and prime with epoxy primer. (4) Apply Polyurethane topcoat.
Refinishing Steel, and Aluminum Components A.
Refinishing of minor nicks and scratches. (1) Feather sand edges of finish around the effected area with 320 grit sandpaper, followed by 400 grit. Avoid sanding through the primer if possible. (2) For steel parts, if primer is damaged in an area larger than the size of a dime, the area should be hand solvent cleaned with an approved solvent then reprimed with a epoxy primer.
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Nose Landing Gear Construction Materials Figure 701 (Sheet 1)
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(4) B.
For aluminum parts, if primer is damaged in an area larger than the size of a dime, the area should be hand solvent cleaned with an approved solvent, then apply a spray wash primer or brush chem-film primer. (a) After spray wash primer or brush chem-film primer has dried for at least 30 minutes, apply epoxy primer. Spot paint damaged areas with Polyurethane paint.
Complete refinishing. (1) Degrease and remove sealants and heavy soil with approved solvents. (2) Strip original finish following recommendations of stripper manufacturer.
CAUTION: Do not allow stripper to come in contact with the nose gear drag link spring. Wear protective clothing and avoid contact with skin. Use of chemicals requires good ventilation and good fire safety practices. (3) (4) (5) (6) (7)
Use wire brush, ScotchBrite or fine aluminum oxide paper to remove any remaining lose paint, scale, and rust. Hand solvent clean with an approved solvent. Apply epoxy primer to steel parts. Apply spray wash primer, or brush chem-film to aluminum parts. Apply epoxy primer to aluminum parts. Apply Polyurethane topcoat.
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MODEL 208 MAINTENANCE MANUAL NOSE GEAR SHOCK STRUT - MAINTENANCE PRACTICES 1.
2.
General A.
This section gives procedures for the shock strut removal/installation and disassembly/assembly. For information about the time limits for the nose gear assembly replacement, and surface rework instructions for the grease seal bore and the cup backing bore, refer to Nose Landing Gear Maintenance Practices. For servicing of the nose gear shock strut, refer to Chapter 12, Nose Gear Shock Strut - Servicing.
B.
The nose gear consists of an oil snubber shock strut assembly mounted in a trunnion, a shimmy damper, nose wheel, tire and tube, a drag link spring assembly and a steering bungee linkage to the pilot's rudder pedals. An extended nose gear fork, which lets there be more propeller to ground clearance, is available as an option for Airplanes 20800095 thru 20800396 and 208B0150 thru 208B1173.
C.
A longitudinal nose gear fairing extends aft to cover the upper part of the shock strut and the drag link spring. The shock strut trunnion is attached to the lower forward engine mount at two pivot lugs. The drag link spring is attached at the upper part of the wheel fork and to the lower side of the fuselage using two bearing blocks, in tandem. This makes it easy to remove the complete nose gear assembly when replacement is necessary, or when you install floats. For more information about the nose gear fairing and drag link spring, refer to Nose Gear Fairing - Maintenance Practices and Nose Landing Gear - Maintenance Practices.
D.
Vertical loads that occur when you land and taxi are absorbed by the drag link spring and the nose gear shock strut. Minor loads, such as when you taxi, are absorbed primarily by the drag link spring, but as the rate of application of loads increases, such as when you land, a larger and larger proportion of the total load is absorbed by the shock strut. For more information about the nose gear fairing, refer to Nose Gear Fairing - Maintenance Practices.
E.
The nose wheel is steerable through an arc of 15 degrees each side of center by use of the rudder pedals, and by applying brakes, the angle may be increased up to 56 degrees either side of center.
Nose Gear Shock Strut Removal/Installation A.
Remove the Nose Gear Shock Strut (Refer to Figure 201). (1) Remove the left and right lower cowling sections and the nose gear fairing as necessary to get access to the nose gear shock strut. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices and Nose Gear Fairing - Maintenance Practices. (2) Lift the nose of the airplane with jacks. Refer to Chapter 7, Jacking - Maintenance Practices. (3) Disconnect the steering bungee from the bell crank.
CAUTION: The nose gear drag link spring has a preload of 150 pounds (667.23 N) applied to the extended nose gear fork. Support the drag link spring before removing the bolts to prevent damage to components. (4) (5) (6) (7) B.
Remove the bolt and washers to disconnect the drag link spring from the shock strut. (a) Remove the load on the extended shock strut. (b) Keep the safety clips. Pivot the shock strut slightly forward to clear the drag link spring fork. With the nose gear shock strut supported, remove the bolts from the trunnion assembly and the engine mount lugs. (a) Record the position of the washers. Slide the nose gear shock strut straight down to remove it from the airplane.
Install the Nose Gear Shock Strut (Refer to Figure 201). (1) Raise the nose gear shock strut assembly straight up to put the pivot holes in the trunnion in position with the holes in the engine mount lugs. (2) Put the washers in the position they were in before removal and installation of the bolts, then torque the nuts from 290 to 410 inch-pounds (32.7 to 46.3 N.m).
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Nose Gear Shock Strut Installation Figure 201 (Sheet 1)
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Nose Gear Shock Strut Installation Figure 201 (Sheet 2)
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Nose Gear Shock Strut Installation Figure 201 (Sheet 3)
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Apply approximately 150 pounds (680 N) upward force to the nose gear drag link spring. Align the holes in the fork with the holes in the nose wheel fork. Insert the safety clips between the gear spring fork and the trunnion with one leg resting against the boss of the trunnion bearing block. Make sure the leg of the safety clip is positioned tightly against the bearing block surface. NOTE:
The safety clip must not be able to turn.
Install bolts and torque from 60 to 85 inch-pounds (6.8 to 9.6 N.m). Safety the bolts with wire to the safety clip or washer. Connect the steering bungee to the bell crank. Install the lower cowling and the nose gear fairing as necessary. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices and Nose Gear Fairing - Maintenance Practices. (11) Remove the airplane from the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
(7) (8) (9) (10)
3.
Nose Gear Shock Strut Disassembly/Assembly (Airplanes 20800134 and On, 208B0099 and On, and Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Incorporating SK208-51) A.
Disassemble the Nose Gear Shock Strut (Refer to Figure 201). (1) Remove the plug from the outer barrel, then drain the hydraulic fluid. (2) Remove the cotter pin from the castellated nut. (3) Keep the castellated nut for assembly, but discard the cotter pin. (4) Remove the bolt assembly and washers, and keep the washers for assembly. (a) Using a micrometer, measure the outside diameter of the bolt assembly and the inside diameter of the lower bushing in the upper torque link assembly and the upper bushings in the lower torque link assembly. The difference in measurement in the bolt assembly and the bushing tolerances must not be more than 0.0031 inches (0.0787 mm). Refer to Table 201 for the necessary bolt assembly and bushing tolerances. (b) Replace the bolt assembly and/or bushing(s) as necessary. (5) Using an appropriate size drill bit, shank through the access hole in the lower forward surface of the outer barrel, dislodge the lock ring and engage the hook shaped tool to the lock ring, to remove the lock ring from the groove in the barrel. Remove the retainer ring. (6) Grasping the nose wheel fork, pull the inner barrel assembly free of the outer barrel assembly. Remove the inner backup rings, packing, and the outer packing from the support ring. Discard the inner backup rings, packing, and outer packing but keep the support ring.
CAUTION: Use care not to drop the bearing and races from the bearing block. Remove the nut, washers, spacer, and bolt from the nose wheel fork and separate the inner barrel and keep it for reassembly. (8) Remove the plug and metering pin assembly from the inner barrel and discard the packing. (9) Cut safety wire from the lower spacer, and using a suitable pin punch, drive out the roll pin and keep for reassembly. (10) Remove the pin and separate the lower torque link assembly from the nose wheel fork. (a) Using a micrometer, measure the outside diameter of the pin and the inside diameter of the lower bushings in the lower torque link assembly. The difference in measurement in pin and bushing tolerances must not exceed 0.002. Refer to Table 201 for required pin and bushing tolerances. (b) Replace the pin and/or bushing(s) as required. (11) Cut safety wire from the upper spacer, and using a suitable pin punch, drive out the roll pin and keep for reassembly. (12) Remove the pin assembly and separate the upper torque link from the outer barrel. (a) Using a micrometer, measure the outside diameter of the pin assembly and the inside diameter of the upper bushings in the upper torque link assembly. Refer to Table 201 for required pin assembly and bushing tolerances. The difference in measurement in the pin assembly and bushing tolerances must not exceed 0.002 inch (0.05 mm). (b) Replace the pin assembly and/or bushing(s) as required. (7)
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MODEL 208 MAINTENANCE MANUAL (13) Cut safety wire from the bolts and remove and keep the bolts, washers, spacers, shims and bumper. (14) Remove the shimmy damper (if installed). (15) Remove and keep the nut, washer, and bolt. Remove the strut tube. (16) Remove the cap plug from the steering ring. (17) Remove and keep the nut, washer, and bolt. Separate the steering ring from the top of the outer barrel. (18) Bend the locking tab(s) of the key washer flat, and using a suitable spanner wrench, remove and keep the spanner nut and seal. (19) Remove the upper bearing cone from the trunnion. NOTE:
The bearing cups are pressed into the trunnion and do not need to be removed unless damaged.
(20) Pull the outer barrel from the trunnion and remove the lower bearing cone. Table 201. Torque Link Bolt Assembly, Pin Assembly and Bushing Diameters COMPONENT
DIAMETER
MAXIMUM ALLOWABLE DIFFERENCE
Bolt Assembly
0.4364 Inch, +0.0005 or -0.0005 Inch (11.08 mm, +0.0127 or -0.0127 mm)
0.0031 Inch (0.0787 mm)
Lower Bushings, Upper Torque Link
0.4375 Inch , +0.0015 or -0.0015 Inch (11.1125 mm, +0.0381 or -0.0381 mm)
Upper Bushings, Lower Torque Link
0.4375 Inch , +0.0015 or -0,0015 Inch (11.1125 mm, +0.0381 or -0.0381 mm)
Pin Assembly
0.5300 Inch , +0.0004 or -0.0004 Inch (13.462 mm, +0.01016 or -0.01016 mm)
Upper Bushings, Upper Torque Link
0.5313 Inch , +0.0003 or -0.0003 Inch (13.495 mm, +0.00762 or -0.00762 mm)
Pin
0.5300 Inch , +0.0004 or -0.0004 Inch (13.462 mm, +0.01016 or -0.01016 mm)
Lower Bushings, Lower Torque Link
0.5313 Inch , +0.0003 or -0.0003 Inch (13.495 mm, +0.00762 or -0.00762 mm)
B.
0.002 Inch (0.0508 mm)
0.002 Inch (0.0508 mm)
Reassemble the Nose Gear Shock Strut (Refer to Figure 201). NOTE:
(1) (2)
Use all new packings, seals and backup rings when assembling the nose gear shock strut. Assemble these parts lubricated with a film of Petroleum VV-P-236, hydraulic fluid (MILPRF-5606), or Dow-Corning DC-7.
Clamp the trunnion in a suitable holding fixture, repack the lower bearing with MIL-G-21164 high and low temperature grease, and install on the lower bearing cup. Position the outer barrel assembly in the trunnion.
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(4) (5) (6)
Pack the upper bearing cone, using MIL-G-21164 high and low temperature grease and install in the upper trunnion bearing cup. (a) Install the seal, new key washer, and spanner nut. (b) Using a suitable spanner wrench, torque the nut until a slight drag is felt on the outer barrel turn action. Back off the nut to the first key tab and bend tab(s) 90 degrees to the nut keyway. Install the steering ring to the top of the outer barrel. Align the attach holes, and install the bolt, washer and nut. Lubricate all surfaces of the steering ring cavity with MIL-G-21164C high and low temperature grease before you install the cap plug. NOTE:
(7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21) (22) (23) (24) (25) (26) (27) (28) (29) (30) (31) (32) (33)
The steering ring cavity is the inside surface of the outer barrel that is above the bolt and cap plug in the outer barrel (Refer to Figure 201).
Clean the outside area of the cap plug and a 0.25 inch (6.35 mm) down the inner diameter of the outside barrel assembly with MEK or a suitable cleaning solvent. Apply a fillet of B-2 TYPE I sealer around the inner surface of the outer barrel cap plug flange before installation, refer to Nose Gear Shock Strut - Servicing. Install the cap plug into the barrel top. Apply a fillet of B-2 TYPE I sealer around the outside surface of the cap plug and barrel top, refer to Nose Gear Shock Strut - Servicing. Install the cap plug into the steering ring. Install the shimmy damper. Lubricate the new packing and install in the groove at the top of the strut tube. Insert the strut tube into the outer barrel. Align the attach holes, and install the bolt, nut and washers. Assemble the previously kept shims and bumper, and spacers to the lower torque link and install the washers and bolts. Safety wire the bolts. Install the new 2653085-200 lower bushings in the lower torque link assembly, as required. Position the spacer between bushings in the lower torque link assembly, and align the attach holes in the lower torque link assembly with holes in the nose wheel fork. Put the new or existing pin assembly through the holes. Make sure the roll pin hole in the pin assembly aligns with the roll pin hole in the spacer. Install the previously kept roll pin and safety wire. Install the new 2653085-200 upper bushings in the upper torque link assembly as required. Position the upper spacer beneath the over travel indicator cable and between the bushings in the upper torque link. Align the attach holes in the upper torque link assembly with holes in the outer barrel and put the new or existing pin assembly through the holes. Make sure the roll pin hole in the pin assembly aligns with the roll pin hole in the spacer. Install the previously kept roll pin and safety wire. Lubricate the new packing and install in the groove of the plug. Lubricate the new packing and install in the groove of the collar on the inner barrel. Pack the bearing with MIL-G-21164 high and low temperature grease. Assemble the races on each side of the bearing and insert it into the shallow (lower) recess of the bearing block. Slide the inner barrel through the bearing block, bearing, bearing races, and into the nose wheel fork. Position the plug and metering pin assembly through the bottom of the fork so that the attach holes through the plug, fork, and inner barrel are aligned. Install the previously kept bolt, spacers, washers and tighten nut. Lubricate the new packing and the new backup rings and assemble with backup rings on each side of the packing into the interior grooves of the support ring. Lubricate the new packing and install it into the exterior groove of the support ring. Install the lock ring, retainer ring, and the new wiper ring, lubricate with MIL-PRF-5606 hydraulic fluid, over the inner barrel. Position the support ring, with the packing installed, over the inner barrel. Install the bearing on the end of the inner barrel and secure it in place with the lock ring.
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MODEL 208 MAINTENANCE MANUAL (34) Slide the assembled inner barrel into the outer barrel assembly. Tap the lock ring into place in the groove at the bottom of the outer barrel. (35) Install the new bushing to the lower end of the upper torque link and the new bushing to the upper end of the lower torque link as required. Refer to the Model 208 Series Illustrated Parts Catalog for bushing part numbers. (36) Insert the new or existing bolt assembly and washer through the lower torque link and upper torque link and install the previously kept washer, castellated nut and the new cotter pin. Refer to the Model 208 Series Illustrated Parts Catalog for the cotter pin part number. (37) Service the shock strut with MIL-PRF-5606 hydraulic fluid, refer to Nose Gear Shock Strut Servicing. (38) Install the plug. 4.
Nose Gear Shock Strut Disassembly/Assembly (Airplanes 20800001 thru 20800133 and 208B0001 thru 208B0098 Except Airplanes Incorporating SK208-51) A.
Disassemble the Shock Strut (Refer to Figure 201). (1) Remove the plug from the outer barrel and drain hydraulic fluid. (2) Remove the cotter pin from the castellated nut and remove the castellated nut, the bolt, washers and bushing from the upper and lower torque link assemblies. Note the position of washers for reinstallation. Discard the cotter pin but keep the castellated nut, the bolt, and the washers and bushings. (3) Using the appropriate size drill bit shank through access hole in lower forward surface of the outer barrel, dislodge lock ring and engage hook shaped tool to the lock ring to remove the lock ring from the groove in the outer barrel. (4) Remove the retainer ring. (5) Grasping the nose wheel fork, pull the inner barrel free of the outer barrel assembly. (6) Remove and discard inner backup rings, packing, and the outer packing from the support. Keep the support.
CAUTION: Use care not to drop the bearing or races from bearing block. (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19)
Remove and keep the nut, bolt and washers from the nose wheel fork. Separate the inner barrel from the nose wheel fork (4). Remove the plug and metering pin from the inner barrel. Discard the packing. Cut safety wire from the spacer, and using suitable pin punch, drive out the roll pin and keep it for reassembly. Remove and keep the pin and separate the lower torque link assembly from the nose wheel fork. Cut safety wire from the spacer, and using suitable pin punch, drive out the roll pin from the spacer and keep for reassembly. Remove and keep pin and separate the upper torque link from the outer barrel. Cut safety wire from the bolts. Remove and keep the bolts and shims and the bumper block. Remove shimmy damper if installed. Remove and keep the nut, bolt and washers. Remove the strut tube. Remove the cap plug from the steering ring. Remove and keep the nut and bolt. Separate the steering ring from the top of the outer barrel. Bend the locking tab(s) of the key washer flat, and using a suitable spanner wrench, remove and keep the spanner nut, key washer and seal. Remove the upper bearing cone from the trunnion. NOTE:
Bearing cups are pressed into the trunnion and do not need to be removed unless they are damaged.
(20) Pull the outer barrel from trunnion, and remove and keep the lower bearing cone.
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MODEL 208 MAINTENANCE MANUAL B.
Reassemble the Shock Strut (Refer to Figure 201). NOTE:
(1) (2) (3) (4)
(5) (6)
Use all new packing, seals and backup rings when assembling the nose gear shock strut. Assemble these parts lubricated with a film of Petrolatum/VV- P-236, hydraulic fluid (MILPRF-5606), or Dow-Corning DC-7.
Clamp the trunnion in a suitable holding fixture, repack the lower bearing with MIL-G-21164 high and low temperature grease, and install on the lower bearing cup. Position the outer barrel assembly in the trunnion assembly. Pack the upper bearing cone, using MIL-G-21164 high and low temperature grease and install in the upper trunnion bearing cup. Install the seal, the new key washer and spanner nut. (a) Use the applicable spanner wrench and torque the nut until a slight drag is felt on the outer barrel turning action. (b) Back off the nut to the first key tab and bend the tab(s) of the key washer 90 degrees to the nut keyway. Install the steering ring to the top of the outer barrel. Align the attach holes and install the previously kept bolt, washer, and nut. Lubricate all surfaces of the steering ring cavity with MIL-G-21164C high and low temperature grease before you install the cap plug. NOTE:
(7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21) (22) (23) (24) (25) (26)
The steering ring cavity is the inside surface of the outer barrel that is above the bolt and plug in the outer barrel (Refer to Figure 201).
Clean the outside area of the cap plug and a 0.25 inch (6.35 mm) down the inner diameter of the outside barrel assembly with MEK or a suitable cleaning solvent. Apply a fillet of B-2 TYPE I sealer around the inner surface of the outer barrel cap plug flange before installation, refer to Nose Gear Shock Strut - Servicing. Install the cap plug into the barrel top. Apply a fillet of B-2 TYPE I sealer around the outside surface of the cap plug and barrel top, refer toNose Gear Shock Strut - Servicing. Install the cap plug into the steering ring. Install the shimmy damper. Lubricate new packing and install in the groove at the top of the strut tube. Insert the strut into the outer barrel. Align the attach holes, and install the previously kept bolt, washer, and nut. Assemble the previously kept shims and bumper block to the lower torque link assembly and install the kept bolts. Safety wire the bolts. Position the spacer between the lower bushings in the lower torque link assembly. Align the attach holes in the lower torque link assembly with holes in the nose wheel fork. Insert the previously kept pin through the holes. Make sure the role pin hole in pin aligns with the role pin hole in the spacer. Install the roll pin and the safety wire. Position the spacer between upper bushings in the upper torque link assembly. Align the attach holes in the upper torque link assembly with the holes in the outer barrel. Put the previously kept pin through the holes. Make sure the role pin hole in pin aligns with the role pin hole in the spacer. Install the previously kept roll pin and safety wire. Lubricate the new packing and install in the groove of the plug. Lubricate the new packing and install in the groove of the collar on the inner barrel. Pack the bearing with MIL-G-21164 high and low temperature grease. Assemble races on each side of the bearing and put into the shallow (lower) recess of the bearing block. Slide the inner barrel through the bearing block, bearing, bearing races, and into the nose wheel fork. Position the plug and metering pin assembly through the bottom of the fork so that the attach holes through the plug, fork, and inner barrel are aligned. Install the bolt and washer through the aligned attach holes and install the washer and nut.
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MODEL 208 MAINTENANCE MANUAL (27) Lubricate the new packing and new backup rings and assemble with the backup rings on each side of the packing into the interior grooves of the support ring. (28) Lubricate the new packing and install into the exterior groove of the support ring. (29) Install the lock ring, retainer ring, and the new wiper ring, lubricated with MIL-PRF-5606 hydraulic fluid, over the inner barrel. (30) Install the support ring, with the packings installed, over the inner barrel. (31) Install the bearing on the end of the inner barrel and secure it in place with the lock ring. (32) Slide the assembled inner barrel into the outer barrel assembly. Tap the lock ring into place in the groove at the bottom of the outer barrel. (33) Position the previously kept washers in the same relationship as removed at each side of the upper and lower torque links connection, and install the new or existing bushings. (34) Install the previously kept bolt, castellated nut, and the new cotter pin. Refer to Model 208 Series Illustrated Parts Catalog for cotter pin part number. (35) Service the shock strut with MIL-PRF-5606 hydraulic fluid, refer to Nose Gear Shock Strut Servicing. (36) Install the plug.
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MODEL 208 MAINTENANCE MANUAL SHIMMY DAMPENER - MAINTENANCE PRACTICES 1.
2.
General A.
This section gives procedures for the shimmy dampener removal/installation and disassembly/ assembly.
B.
A longitudinal nose gear fairing extends aft to cover the shimmy dampener, upper part of the shock strut and the drag link spring. For more information about the nose gear fairing, nose gear shock strut and drag link, refer to Nose Gear Fairing - Maintenance Practices, Nose Gear Shock Strut Maintenance Practices and Nose Landing Gear - Maintenance Practices.
C.
The nose wheel is steerable through an arc of 15 degrees each side of center by use of the rudder pedals. By applying brakes, the angle may be increased up to 56 degrees either side of center.
Shimmy Dampener Removal/Installation A.
Remove the Lord Shimmy Dampener (Airplanes 20800247 and On and 208B0502 and On and Airplanes 20800001 thru 20800246 and 208B0001 thru 208B0501 incorporating CAB96-3) (Refer to Figure 201). (1) Remove the nose gear fairings as necessary to get access to the shimmy dampener. Refer to Nose Gear Fairing - Maintenance Practices. (2) Open the upper and lower cowling doors and cowl panels as necessary to get access to the shimmy dampener. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices (3) Remove the cotter pin from the castellated nut. (4) Remove the castellated nut and washer from the bolt. (5) Remove the bolt and washer from the steering ring assembly. NOTE:
(6) (7)
Remove the cotter pin from the trunnion. Remove the eyebolt from the trunnion. NOTE:
(8) B.
Keep the shim washers that are found between the trunnion and the shimmy dampener attach lug. Make a record of the washer installation order, because the shim washer thickness can be different.
The head of the eyebolt is flat to let you get a good hold on to it for removal.
Remove the shimmy dampener from the airplane.
Install the Lord Shimmy Dampener (Airplanes 20800247 and On and 208B0502 and On and Airplanes 20800001 thru 20800246 and 208B0001 thru 208B0501incorporating CAB96-3) (Refer to Figure 201).
CAUTION: Do not exceed the 50 degree turn radius when you turn the nose gear. (1) (2)
Turn the nose gear to the right tow limit to let the steering ring assembly bolt clear the engine truss. Put the shimmy dampener in position so the mounting holes align in the steering ring assembly and the trunnion.
CAUTION: Do not put the cotter pin in position at this time. (3) (4) (5) (6)
Install the removed shim washers as they were installed before, between the shimmy dampener and the steering ring assembly. Lubricate the bolt with MIL-G-21164C, then put the bolt for the shimmy dampener in position in the steering ring assembly attach point. Make sure the shimmy dampener has a maximum vertical movement of 0.035 inches (0.889 mm). Make sure the shimmy dampener is free to turn in the steering ring assembly.
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Lord Shimmy Dampener Installation Figure 201 (Sheet 1)
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Put the shimmy dampener piston rod end into the trunnion attach point.
CAUTION: Do not install the cotter pin at this time. CAUTION: Do not exceed the 50 degree turn radius when you turn the nose gear. (8) (9) (10) (11) (12) (13) (14) (15) C.
Lubricate the eyebolt with MIL-G-21164C, then put the eyebolt for the shimmy dampener in position in the trunnion. Make sure the piston rod is free of any load or interference. Turn the nose gear left and right to the tow limits to make sure there is no interference or load after the initial movement of the nose gear. If there is any interference or load found after the initial movement of the nose gear, then install a different set of washers. Install the nut for the steering ring assembly, then torque the nut to approximately 50 to 70 inchpounds. (a) Align the cotter pin hole as necessary to install the cotter pin. Install the cotter pins. Install the cowling doors and cowl panels as necessary. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Install the nose gear fairings as necessary. Refer to Nose Gear Fairing - Maintenance Practices.
Remove the Shimmy Dampener (Airplanes 20800001 thru 20800246, and 208B00001 thru 208B0501, Not Incorporating CAB96-3) (Refer to Figure 202). (1) Remove the nose gear fairings as necessary to get access to the shimmy dampener. Refer to Nose Gear Fairing - Maintenance Practices. (2) Open the upper and lower cowling doors and cowl panels as necessary to get access to the shimmy dampener. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices (3) Remove the cotter pin out of the trunnion. (4) Remove the eyebolt from the trunnion. NOTE: (5) (6)
D.
The head of the eyebolt is flat to let you get a good hold on to it for removal.
Remove the cotter pin, nut, washers, and bolt that attach the shimmy dampener barrel to the steering bell crank. Remove the shimmy dampener from the airplane.
Install the Shimmy Dampener (Airplanes 20800001 thru 20800246, and 208B00001 thru 208B0501, Not Incorporating CAB96-3) (Refer to Figure 202).
CAUTION: Do not exceed the 50 degree turn radius when you turn the nose gear. (1) (2)
Turn the nose gear to the right tow limit to let the steering ring assembly bolt clear the engine truss. Put the shimmy dampener in position so the mounting holes align in the steering ring assembly and the trunnion.
CAUTION: Do not put the cotter pin in position at this time. (3) (4) (5) (6)
Install the removed shim washers as they were installed before, between the shimmy dampener and the steering ring assembly. Lubricate the bolt with MIL-G-21164C, then put the bolt for the shimmy dampener in position in the steering ring assembly attach point. Make sure the shimmy dampener has a maximum vertical movement of 0.035 inches (0.889 mm). Make sure the shimmy dampener is free to turn in the steering ring assembly.
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Shimmy Dampener Installation Figure 202 (Sheet 1)
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Put the shimmy dampener piston rod end into the trunnion attach point.
CAUTION: Do not install the cotter pin at this time. CAUTION: Do not exceed the 50 degree turn radius when you turn the nose gear. (8) Lubricate, then put the eyebolt for the shimmy dampener in position in the trunnion. (9) Make sure the piston rod is free of any load or interference. (10) Turn the nose gear left and right to the tow limits to make sure there is no interference or load after the initial movement of the nose gear. (11) If there is any interference or load found after the initial movement of the nose gear, then incorporate the trunnion modification procedures as necessary from Service Bulletin CAB96-3. (12) Install the nut for the steering ring assembly, then torque the nut to approximately 50 to 70 inchpounds. (a) Align the cotter pin hole as necessary to install the cotter pin. (13) Install the cotter pins. (14) Install the cowling doors and cowl panels as necessary. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (15) Install the nose gear fairings as necessary. Refer to Nose Gear Fairing - Maintenance Practices. 3.
Shimmy Dampener Disassembly/Assembly (Airplanes 20800394 and On and 208B1140 and On) A.
4.
The Lord Shimmy Dampener is sealed, and cannot be disassembled or assembled. Contact Cessna Propeller Aircraft Product Support at (316) 517-5800 or Fax (316) 942-9006, for more instructions about replacement.
Shimmy Dampener Disassembly/Assembly (Airplanes 20800001 thru 20800393 and 208B0001 thru 208B1139) NOTE:
These procedures include a check of the shimmy dampener components for a leaking shimmy dampener.
NOTE:
For procedures to make a piston insertion tool for the shimmy dampener servicing. Refer to Figure 203 for fabrication details.
NOTE:
The temperature compensating piston must be put in a position on the piston rod at a distance of 3.32 inches from the end of the piston rod (with ambient temperature 70°F). Mark a piece of welding rod or equivalent material with the 3.32 inch dimension from one end to be used as a positioning gauge in the following procedure.
A.
Leaking Shimmy Dampener. NOTE:
(1)
Leaking of the shimmy dampener at either end of the barrel, at the filler plugs, or at the end of the piston rod where the temperature compensating piston is installed, indicates a need for servicing of the shimmy dampener. Do a check of the position of the temperature compensating piston relative to the end of the piston rod to make sure there is a leaking condition. The piston is factory set at 3.32 inches, +0.05 or -0.05 inch, from the end of the piston rod at 70°F. As the temperature decreases below 70°F, the 3.32 inch dimension will increase to a maximum of 4.83 inches at 37°F. As the temperature increases above 70°F, the 3.32 inch dimension will decrease to a minimum of 2.74 inches at 98°F.
Use the Temperature versus Position Chart to service the shimmy dampener if the dimension from the end of the piston rod to the piston exceeds +0.15 inch from the normal line. Refer to Figure 202. If the shimmy dampener does not leak at the ends of the barrel or the piston, it may be possible to remove the filler plug and check the filler plug packing and service the dampener assembly without removal of the piston from the piston rod assembly.
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Shimmy Dampener Servicing Figure 203 (Sheet 1)
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Shimmy Dampener Servicing Figure 203 (Sheet 2)
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B.
Shimmy Dampener Servicing Figure 203 (Sheet 3) Disassembly of the Shimmy Dampener on Airplanes 20800001 thru 20800393, 208B0001 thru 208B1139 (Refer to Figure 203). NOTE:
(1) (2) (3)
Keep the shimmy dampener clean, especially the exposed portions of the piston rod, to prevent collection of dust and grit which could damage packings in the barrel. Use a clean lint free cloth saturated with MIL-PRF-5606 hydraulic fluid or ASTMD 3699-78 kerosene to keep the machined surfaces wiped free of dirt or dust. All surfaces must be wiped free of unwanted hydraulic fluid.
Remove the shimmy dampener from the airplane. Remove the safety wire from the filler plugs. Remove the filler plugs. NOTE:
(4) (5) (6)
There is packing found in front of the filler plugs.
Drain the hydraulic fluid from the barrel. Do a check of the condition of the O-rings and replace the O-rings if necessary. Remove the setscrew, spring and the temperature compensating piston from the piston rod. NOTE:
There is packing and a backup ring in front of the setscrew. Also, there is an O-ring in front of the temperature compensating piston.
NOTE:
Low pressure shop air applied to the filler port may be required to force the temperature compensating piston out of the piston rod.
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WARNING: Do not apply high air pressure to the filler port. Too much pressure will cause the temperature compensating piston to exit the piston rod at a high rate of speed. Make a way to stop the temperature compensating piston before complete egress from the piston rod occurs. NOTE: (7) (8) (9)
Do a check of the condition of the O-ring and replace if necessary. Do a check of the condition of the inside of the piston rod. Remove the snap ring and the end gland. NOTE:
(10) (11) (12) (13) (14) (15) C.
A piston extraction tool can be fabricated to remove the piston. Refer to Figure 203.
There is a backup ring and packing in front of the end gland.
Do a check of the backup ring and the packing to make sure they are serviceable. Remove the piston from the piston rod assembly. Do a check of the backup rings and the packing to make sure they are serviceable. Do a check of the condition of the roll pin and surface condition of the piston and piston rod. Do a check of the inside surface of the barrel. Do a check of the backup rings and the packing to make sure they are serviceable.
Assembly of the Shimmy Dampener (Airplanes 20800001 thru 20800393, 208B0001 thru 208B1139) (Refer to Figure 203). NOTE:
(1)
Before you assemble the shimmy dampener, make sure there are no sharp edges on any parts that could cause damage to the backup rings or packings during assembly. Lubricate the backup rings, packings, and metal parts with MIL-PRF-5606 hydraulic fluid or VSP Alba White Amojell Petroleum (Vaseline) to make installation easy.
Install a backup ring and packing in the barrel if previously removed.
CAUTION: Make sure that the hole for the temperature compensating piston and the piston rod are correctly aligned. (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13)
Install a roll pin through the piston and the piston rod. Install backup rings and packing on the piston. Install the piston rod in the barrel. Install the end gland in the end of the barrel. Position the piston rod all the way to the gland end of the barrel. Fill the chamber through the end port with MIL-PRF-5606 fluid. Install the packing and the filler plug. Fill the chamber through the end port with MIL-PRF-5606 fluid. Install the packing and the filler plug. Reposition the assembly so the piston rod is vertical with the end port down. Remove the side filler plug, setscrew, and piston. Push the piston rod down fully, to the opposite end of the barrel to force any remaining air through the orifice, then fill the rod assembly with fluid. NOTE:
Thumb pressure may be necessary to prevent fluid from coming out of the end of the rod assembly.
(14) Fill the piston rod with MIL-PRF-5606 fluid, then install the temperature compensating piston and the setscrew. (15) Finish filling the barrel with fluid and install the side filler plug. (16) Cycle the piston rod five or six times at the full stroke of travel. (17) Reposition the assembly horizontal with the side filler up, then remove the side plug. (18) Fill with additional fluid as necessary to bring fluid up to the top of the filler opening.
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MODEL 208 MAINTENANCE MANUAL (19) Reinstall the filler plug so it is not fully tightened, to let fluid be bleed. (20) Remove the setscrew. (21) Push the temperature compensating piston in to the 3.32 inch dimension (+0.05 inch tolerance), then tighten the side filler plug. (22) Use the spring insertion tool and slide it over the piston rod, then insert the spring and setscrew. NOTE:
The setscrew is put into position in the end of the piston rod.
(23) Operate the completed unit by hand at the full stroke of travel to check for air and correct operation. (24) If air is trapped in the shimmy dampener, then remove the temperature compensating piston and repeat the steps to remove air from the shimmy dampener. (25) If there is no air trapped in the shimmy dampener, then install the safety wires and reinstall the shimmy dampener on the airplane.
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MODEL 208 MAINTENANCE MANUAL SHIMMY DAMPENER - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the shimmy damper in a serviceable condition.
Task 32-20-02-720 2.
Shimmy Damper Functional Check A.
General (1) This task gives the procedures to do a functional check of the shimmy damper.
B.
Special Tools (1) None
C.
Access (1) Remove the nose gear fairings to get access to the shimmy damper. Refer to Nose Gear Fairing - Maintenance Practices. (2) Remove the left upper cowling door and the left lower cowl panel to get access to the shimmy damper. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a functional check of the shimmy damper.
CAUTION: Do not go more than the 50 degree turn radius limit when you turn the nose gear. (1) (2) (3)
With the nose gear turned to the limits (left and right), examine the damper for vertical preload at the piston rod attachment. Make sure that the shimmy damper has a maximum vertical movement of 0.035 inch (0.889 mm). If there is any interference or preload found after the initial movement of the nose gear, incorporate the trunnion modification procedures as necessary from service bulletin CAB96-3.
E.
Restore Access (1) Install the left upper cowling door and the left lower cowl panel. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (2) Install the nose gear fairings. Refer to Nose Gear Fairing - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL NOSE GEAR FAIRING - MAINTENANCE PRACTICES 1.
2.
General A.
This section gives procedures for the removal/installation of the drag link spring fairing.
B.
A longitudinal nose gear fairing extends aft that covers the shimmy dampener, upper part of the shock strut and the drag link spring. For more information about the shimmy dampener, nose gear shock strut and drag link, refer to Shimmy Dampener - Maintenance Practices, Nose Gear Shock Strut Maintenance Practices and Nose Landing Gear - Maintenance Practices.
Nose Gear Fairing Removal/Installation A.
Remove the Nose Gear Fairing (Refer to Figure 201). NOTE: (1)
B.
Removal procedures are typical for the left and right fairings.
Turn the quarter-turn fasteners counter clockwise on the nose gear fairing, then remove the nose gear fairing.
Install the Nose Gear Fairing (Refer to Figure 201). NOTE: (1)
Removal procedures are typical for the left and right fairings.
Put the nose gear fairing in position to the lower cowling, then turn the quarter-turn fasteners clockwise to attach the nose gear fairing to the airplane.
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Nose Gear Fairing Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL WHEELS AND BRAKES - DESCRIPTION AND OPERATION 1.
Description and Operation A.
Main wheels are fabricated of aluminum and are designed to be used with tires and tubes. Each main wheel consists of two wheel halves, two bearing cones, two bearing cups, two grease seals, brake disc assembly and snap rings. The wheel halves are secured together with bolts, washers and nuts. A hole in one wheel half is provided for installation of a valve stem. Standard on Models 208 and 208 Cargomaster are 6.50 X 10, 8-ply- rated tube tires; 8.50 X 10, 8-ply-rated tube tires are optional on Model 208, 208 Cargomaster and standard on Model 208B and 208B Super Cargomaster. The 29.11 X 10, 10-ply-rated tube tires are optional on the 208B Passenger. The wheel rotates on two bearing cones. Bearing cups are shrunk-fit into the wheel half hub. Bearings are protected against dirt, moisture, contamination, and loss of lubricant by a bearing seal. The wheel is secured to the axle with a washer, nut, and cotter pin.
B.
The nose wheel is fabricated of aluminum and is designed to be used with a tire and tube. The nose wheel consists of two wheel halves, two bearing cones, two bearing cups, two grease seals, and snap rings. Wheel halves are secured together with bolts, washers, and nuts. A hole in one wheel half is provided for installation of a valve stem. Standard on Models 208 and 208 Cargomaster are 6.50 X 8, 8-ply-rated tube tires; 22 X 8.00 X 8, 6-ply- rated tube tires are optional on Model 208, 208 Cargomaster and standard on Model 208B. The wheel is free-rolling on an independent axle and is used to steer the airplane on the ground by means of the nose wheel steering system.
C.
The brakes are hydraulically operated and are designed to use MIL- H-5606 hydraulic fluid. The brake consists of a magnesium housing containing four pistons, an inlet port, bleeder port, torque plate, backplates, pressure plate, shims and anchor bolt. The brake assembly is held together with bolts, washers, and nuts.
D.
Two brake master cylinders are installed, one for each brake. Master cylinders are located forward of the pilot's rudder pedals. Each brake master cylinder consists of a piston, ring, packing, spring and cylinder.
E.
The brake system reservoir is located in the engine compartment on the lower left corner of the firewall.
F.
A parking brake system is provided, consisting of a parking brake valve, located under the floor beneath the pilot's rudder pedals; lines from the valve to the master cylinders and brake cylinders; a flexible control wire and a parking brake control knob, located on the lower left instrument panel.
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MODEL 208 MAINTENANCE MANUAL WHEELS AND BRAKES - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Wheels and Brake Troubleshooting Chart Figure 101 (Sheet 1)
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Wheels and Brake Troubleshooting Chart Figure 101 (Sheet 2)
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Wheels and Brake Troubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL WHEELS AND BRAKES - MAINTENANCE PRACTICES 1.
General A.
2.
This section provides information on brake system bleeding, wheel removal/installation, tire mounting and dismounting, brake components removal/installation and brake inspection criteria.
Tire Mounting Precautions A.
Tire Mounting and Dismounting Criteria (Refer to Figure 201). (1) Prior to removing the wheel/tire assembly from the airplane, completely deßate the tire with a deßation cap. It is good practice to deßate the tire before removing the axle nut. When all pressure has been relieved, remove the valve core. Valve cores under pressure can be ejected like a bullet. If wheel or tire damage is suspected, approach the tire from the front or rear, not from the side (facing the wheel).
CAUTION: A tire/wheel assembly that has been damaged in service should be deßated by a remote means. If this is not possible, the tire/wheel assembly should be allowed to cool for a minimum of three hours before the tire is deßated. (2)
(3) (4) (5) (6)
Take special care when encountering difÞculty in freeing tire beads from wheel ßanges. Trying to pry beads free incorrectly may cause an accident. Even with tire tools, care must be taken to prevent damage to beads or wheel ßanges. On small tires, successive pressing with a two-foot length of wood close to the bead or tapping with a rubber mallet is generally sufÞcient. Ensure the mating tire and tube are speciÞed and correct for the wheel/tire assembly. Clean inside of tire, then lubricate lightly with talcum powder. Inßate tube to slightly round and insert in tire. This aids in mounting tire and tube to wheel half and helps prevent pinching tube. Align yellow stripe on tube with red balance dot on tire. Align red dot with valve if no stripe exists on tube. NOTE:
After inserting valve stem in hole on wheel half, connect a valve stem puller device to the valve stem to prevent the valve stem from receding from hole in wheel half.
(7)
When mounting tire and tube on wheel, ensure wheel bolts are torqued to wheel manufacturer's instructions before inßating. (8) Inßate tire in a safety cage to rated pressure. (9) Deßate to equalize stretch. (10) Reinßate to rated pressure. (11) After a 12-hour stretch period, reinßate to rated inßation pressure. 3.
Servicing Tires and Tubes A.
4.
Brake System Replenishing A.
5.
For servicing procedures related to tires and tubes, refer to Chapter 12, Tires - Servicing.
For replenishing of the brake master cylinders, refer to Chapter 12, Hydraulic Fluid - Servicing.
Brake System Bleeding NOTE:
A.
Anytime a brake line is disconnected or a spongy feel to the brake pedal is detected, there is likelihood air has entered the system. To assure proper braking action, all trapped air must be removed from the system by the following procedures.
Brake Bleeding Procedures. (1) Remove left and right brake fairings to gain access to brake bleeder valves.
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MODEL 208 MAINTENANCE MANUAL (2)
Ensure parking brake handle is off (fully in). NOTE:
(3) (4) (5) (6) (7) 6.
It is recommended that a pressure source of clean hydraulic ßuid (MIL-PRF-5606) be connected to the wheel cylinder bleeder valves.
Connect hydraulic pressure source, such as a hand pump or Hydro Fill unit, to right brake wheel cylinder bleeder valve. Open bleeder valve and begin pumping hydraulic ßuid into the system while observing ßuid level in brake system reservoir, located on lower left corner of Þrewall in engine compartment. When bubbles have ceased appearing in reservoir (approximately 1/2 full), close bleeder valve and remove pressure source. Repeat steps (3) through (5) for left brake system. Reinstall brake fairings.
Main Wheel and Tire Removal/Installation A.
Remove Main Wheel and Tire (Refer to Figure 201). (1) Jack airplane. Refer to Chapter 7, Jacking - Maintenance Practices.
WARNING: Make sure that you deßate the tires before you remove the wheel/tire assemblies from the airplane. When all pressure has been released, use an extraction tool to remove the valve core from the valve stem. Valve cores under pressure can be ejected from the valve stem and cause injury to personnel or damage to the airplane. If you think there is wheel or tire damage, get access to the tire from the front or rear, not from the side (facing the wheel). (2) (3) (4)
Deßate tire completely. Remove backplate bolts (12), washers and shims (3) attaching brake backplate to brake assembly. Remove backplate. Remove cotter pins and axle nut. NOTE:
(5) (6) (7) (8)
Bearings and bearing seals will be removed during disassembly.
Pull wheel from axle. Examine the axle Þttings for cracks, corrosion, pits, security, and any other damage. (a) Make sure that the attach bolts have sufÞcient thread engagement in the self-locking nuts. Examine the outer wheel axle surface and attachment bolts for condition, cracks, corrosion, signs of damage, and wear. Examine the inner main gear wheel axle surface for corrosion. NOTE:
The main gear axle is a time limited part. You should verify the current number of landings, inspection status and replacement time before you install the main wheel and tire. Refer to Chapter 4, Replacement Time Limits and Typical Inspection Time Limits for the applicable inspection criteria.
(9)
Examine the main gear axle attachment to the axle Þtting and the attachment bolts for corrosion, cracks, signs of damage, and security. (10) Examine the main gear axle Þtting where it is installed to the outer spring and the attachment bolts for corrosion, cracks, signs of damage, and security. (a) Make sure that the axle Þttings part number/serialization identiÞcation placards are securely attached. (11) Examine the Þllet seal around the main gear spring and at the axle Þtting. (a) If the seal is broken, loose, or deteriorated, replace it with a new Þllet seal using Type 1, Class B ½ sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing Maintenance Practices.
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Main Wheel, Tire and Tube Installation Figure 201 (Sheet 1)
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Main Wheel, Tire and Tube Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (12) Examine the axle spacer for wear at the grease seals contact area. If the grooves will not let the seal seat correctly, replace the axle spacer. NOTE:
Cessna Propeller Aircraft Product Support, 316-517-5800 or Fax 316-942-9006, is the source to get the information on damage criteria.
(13) Remove all wheel bearing cones. (a) Clean the bearing cones and cups and the wheel halves with stoddard solvent or an equivalent approved cleaning solvent.
WARNING: Use low pressure shop air to dry bearings. Do not spin bearing cones with compressed air. Dry running bearings without lubrication can explode. (14) Examine for, and replace the bearing cups if the cups are loose in the wheels, or there are scratches, pitting, corrosion, or signs of overheating. (15) Examine for, and replace the bearing cones if there are nicks, scratches, water staining, spalling, heat discoloration, roller wear, cage damage, cracks, or distortion. (16) Re-pack bearings with general purpose grease MIL-G-81322 or an equivalent grease. (17) Install the bearings in the wheels with new grease seals. B.
Install Main Wheel and Tire (Refer to Figure 201). (1) Place wheel assembly on axle. (2) Install axle nut and rotate wheel while torquing to 60 inch-pounds (6.8 N.m). Back off axle nut and torque to 30 inch-pounds (3.4 N.m) while rotating wheel. Tighten axle nut to next castellation but do not advance nut in excess of 15 degrees (one-half castellation). Install cotter pin. (a) If torque exceeds 40 inch-pounds (4.5 N.m) during Þnal tightening, remove nut and install washers as required to meet the torque requirements given above. Washers (Cleveland part number 153-00100) can be ordered from Cessna Parts Distribution. (3) Install brake backplate (1) and shim (3) using bolts (12) and washers. Torque bolts from 85 to 90 inch-pounds (9.6 to 10.2 N.m). NOTE:
(4) (5) 7.
Bolts (12) incorporate a special self-locking feature and are typically good for approximately four to six reuses. If bolt (12) can be fully engaged into the backplate by hand with no resistance, the self-locking feature of bolt (12) has been destroyed and the bolt should be rejected. Replacement bolts (part number 103-14300) can be ordered from Cessna Parts Distribution.
Inßate tire to proper pressure. Refer to Chapter 12, Tires - Servicing. Remove airplane from jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
Main Wheel, Tire and Tube Disassembly/Reassembly
WARNING: Injury can result from attempting to separate wheel halves with tube inßated. Take care to avoid damaging wheel halves when breaking tire beads loose. A.
Disassemble Main Wheel, Tire and Tube (Refer to Figure 201). NOTE: (1) (2) (3) (4)
Refer to Tire Mounting Precautions before disassembly of main wheel, tire and tube.
Ensure tube (11) is deßated completely then break beads of tire (10) loose. Remove thru-bolts (14) and separate wheel halves (6) and (12). Retain spacer (8). Remove tire (10), tube (11) and brake disk (13). Remove snap rings (3), grease seals (4) and bearing cone (5) from inboard wheel half (12).
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MODEL 208 MAINTENANCE MANUAL (5)
Remove snap ring (3), grease seal (4), bearing cone (5), from outboard wheel half (6). NOTE:
B.
Bearing cups (7) are a press-Þt in wheel halves (6) and (12) and should not be removed unless replacement is necessary. To remove bearing cups (7), heat wheel half in boiling water for 15 minutes. Using an arbor press, press out bearing cup and press in new bearing cup while wheel is still hot.
Assemble Main Wheel, Tire and Tube (Refer to Figure 201). NOTE: (1) (2)
Refer to Tire Mounting Precautions before reassembly of main wheel, tire and tube.
Clean the inside of both wheel halves with mineral spirits. Let it dry and apply Royco 103 or other MIL-C-16173 Type 1, Grade 1 protectant to all surfaces except the bearing outer race and the ßanges where the tire bead seats. NOTE:
(3) (4)
If replacing the bearing outer races, remove outer race and clean race cup bore surfaces of the wheel and apply a thin coating of protectant to the mating surface of the wheel. Install outer race while protectant is still wet. Brush coat the wheel surfaces around outer race after race installation to replace any protectant removed during the race installation. NOTE:
(5) (6)
A swab may be necessary to apply protectant to the wheel material inside the bolt holes.
Some protectant in the snap ring (3) groove is permitted, but do not Þll the groove. Too much protectant can be removed with mineral spirits.
Insert thru-bolts (14) through brake disk (13) and position in inner wheel half (12), using bolts to guide disc. Ensure disc is bottomed in wheel half. Position tire (10) and tube (11) on outboard wheel half (6). NOTE:
Lightweight point of tire is marked with a red dot on tire sidewall and heavyweight point of tube is marked with a contrasting color line (usually near valve stem). When installing tire, place these marks adjacent to each other.
(7) (8)
Valve stem (9) must protrude through hole in wheel half (6). With spacer (8) positioned into inboard wheel (12), place outboard wheel half (6) in position, applying a light force to keep wheel halves together. Do not pinch tube (11) between wheel halves. (9) Assemble washers (2) and nuts (1) on thru-bolts (14). (10) Torque nuts (1) evenly to 150 inch-pounds (16.9 N.m).
CAUTION: Uneven or improper torque of thru-bolt nuts (1) can cause bolt failure with resultant wheel failure. (11) Inßate tube to set tire beads, then adjust tire pressure. Refer to Chapter 12, Tires - Servicing. (12) Clean bearing cones (5) and repack with clean wheel bearing grease. (Refer to Chapter 12.) (13) Assemble bearing cone (5) and grease seal (4) into inboard wheel half (12) and secure with snap ring (3). (14) Assemble bearing cone (5) and grease seal (4) into outboard wheel half (6) and secure with snap ring (3). 8.
Brake Backplate and Pressure Plate Removal/Installation NOTE: A.
Brake linings can be removed without removing main wheels or disconnecting brake lines.
Remove Backplate and Pressure Plate (Refer to Figure 202). (1) Remove brake backplate bolts (12).
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Brake Assembly Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) B.
Remove backplates (1) and shim (3). Only one shim is allowed. Slide brake cylinder (13) clear of anchor bolts (8). Remove pressure plate (7) and linings (6) from anchor bolts (8). Retain brake piston insulators (9).
Install Backplate and Pressure Plate (Refer to Figure 202). (1) Slide new brake lining (6) and pressure plate (7) over anchor bolts (8). (2) Insert piston insulators (9) opposite pistons (11) and slide brake cylinder (13) over anchor bolts (8). (3) Position backplates (1) and shim (3) over brake disc (4) opposite backplate bolt holes in brake cylinder (13). (4) Install backplate bolts (12) and torque from 85 to 90 inch-pounds. NOTE:
9.
Brake Assembly Removal/Installation A.
Remove the Brake Assembly (Refer to Figure 202). (1) Without applying brakes, pull parking brake handle to the ON position (fully out). (2) Disconnect brake line at brake cylinder (13) and allow ßuid to drain from brake line. (3) Remove backplate bolts (12) and backplates (1). (4) Remove anchor bolt nuts (15) and slide brake cylinder assembly (13) off torque plate (5).
B.
Install the Brake Assembly (Refer to Figure 202 ). (1) Install pressure plate (7) over anchor bolts (8). (2) Slide brake cylinder assembly (13) onto torque plate (5). (3) With shim (3) positioned against backplates (1), install backplate bolts (12) and torque from 85 to 90 inch-pounds (9.6 to 10.2 N.m). NOTE:
(4) (5) (6) 10.
Backplate bolts (12) incorporate a special self-locking feature and are typically good for approximately four to six reuses. If backplate bolt (12) can be fully engaged into the backplate by hand with no resistance, the self- locking feature of backplate bolt (12) has been destroyed and the backplate bolt should be rejected. Replacement bolts (part number 103-14300) can be ordered from Cessna Parts Distribution.
Backplate bolts (12) incorporate a special self- locking feature and are typically good for approximately four to six reuses. If backplate bolt (12) can be fully engaged into the backplate by hand with no resistance, the self-locking feature of backplate bolt (12) has been destroyed and the backplate bolt should be rejected. Replacement bolts (part number 103- 14300) can be ordered from Cessna Parts Distribution.
Connect brake line at wheel cylinder Þtting. Push parking brake handle to the off position (fully in). Bleed brake system. Refer to Brake System Bleeding.
Brake Assembly Disassembly/Reassembly A.
Disassemble the Brake Assembly (Refer to Figure 202). (1) Remove pistons (11) and insulators (9) from brake cylinder (13). Remove O-rings (10) from piston (11) and discard. (2) Remove bleeder valve cap and bleeder valve (14).
B.
Assemble the Brake Assembly (Refer to Figure 202). (1) Install bleeder valve (14) in brake cylinder (13). (2) Using clean hydraulic ßuid (MIL-PRF-5606) as a lubricant, install new O-rings (10) on pistons (11). (3) Install pistons (11) into cylinder (13). (4) Install insulators (9) in pistons (11).
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11.
New Brake Burn-In A.
Airplanes 20800001 thru 20800135 and 208B0001 thru 208B0102. (1) Perform six consecutive light braking applications from 20 to 35 knots. Allow brake discs to cool substantially between stops.
CAUTION: Do not set the parking brakes while they are hot. This will help to prevent irregular friction surface mix transfer that can result in brake clatter, noise and vibration. B.
Airplanes 20800136 and On, 208B0103 and On, and All Spares. NOTE: (1)
The brake pads are of a metallic composition and require the following break-in procedure.
Perform two consecutive full stop braking applications from 30 to 35 knots.
CAUTION: Do not allow brake discs to cool substantially between stops. Use caution in performing this procedure, as higher speeds with successive stops could cause the brakes to overheat, resulting in warped discs and/or pressure plates. 12.
Nose Wheel Removal/Installation A.
Remove Nose Wheel (Refer to Figure 203). (1) Jack airplane. Refer to Chapter 7, Jacking - Maintenance Practices.
WARNING: Make sure that you deßate the tire before you remove the wheel/tire assembly from the airplane. When all pressure has been released, use an extraction tool to remove the valve core from the valve stem. Valve cores under pressure can be ejected from the valve stem and cause injury to personnel or damage to the airplane. If you think there is wheel or tire damage, get access to the tire from the front or rear, not from the side (facing the wheel). (2) (3) (4) (5) (6) (7)
Deßate tire completely. Remove cotter pin (1), nut (2) and washer (3) from one side of fork (9), withdraw axle stud (8). Using long punch through one axle bucket (4), tap out bucket at opposite side of fork (9). Remove both buckets (4) and pull tire and wheel (6) from fork (9). Remove spacers (5) or (10) and axle tube (7) before disassembling wheel. Examine the wheel axle tube for condition, cracks, corrosion, and wear. (a) Examine the axle spacer for wear at the grease seals contact area. If the grooves will not let the seal seat correctly, replace the axle spacer. (8) Examine the wheel spacers and the buckets for condition, cracks, and corrosion. (a) For airplanes 20800202 and On, 208B0256 and On, and airplanes that incorporate CAB9130, examine the seal for condition, cuts, and deterioration. (9) Examine the axle stud for condition, bends, damaged threads, cracks, and corrosion. (10) Remove all wheel bearing cones. (a) Clean the bearing cones and cups and the wheel halves with stoddard solvent or an equivalent approved cleaning solvent.
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Nose Wheel Installation Figure 203 (Sheet 1)
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WARNING: Use low pressure shop air to dry bearings. Do not spin bearing cones with compressed air. Dry running bearings without lubrication can explode. (11) Examine for, and replace the bearing cups if the cups are loose in the wheels, or there are scratches, pitting, corrosion, or signs of overheating. (12) Examine for, and replace the bearing cones if there are nicks, scratches, water staining, spalling, heat discoloration, roller wear, cage damage, cracks, or distortion. (13) Re-pack the bearings with general purpose grease MIL-G-81322 or an equivalent grease. (14) Install the bearings in the wheels with new grease seals. B.
Install Nose Wheel Airplanes 20800001 thru 20800201 and 208B0001 thru 208B0255 not incorporating CAB 91-30 (Refer to Figure 203). (1) Insert axle tube (7) into wheel and place a spacer (5) on each side of wheel. (2) Position wheel into fork (9) and install buckets (4) into fork recesses. Tap buckets with nonmetallic hammer until seated in fork. (3) Install axle stud (8), washer (3), and nut (2). (4) Tighten nut (2) until a slight drag can be felt when rotating the tire. (5) Rotate the tire, by hand, and measure the force at the outside diameter of tire. Force should be between 3 and 5 pounds (13.4 to 22.2 N). (6) Install new cotter pin (1) in nut (2). If cotter pin does not line up, back off nut to the next castellation. (7) Again, verify the measured force at the outside diameter of the tire is between 3 and 5 pounds (13.4 to 22.2 N). (8) Remove airplane from jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
C.
Install Nose Wheel Airplanes 20800202 and On and 208B0256 and On and Airplanes 20800001 thru 20800201 and 208B0001 thru 208B0255 incorporating CAB 91-30 (Refer to Figure 203). NOTE:
(1) (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13) 13.
Application of Permatex Anti-Seize Lubricant as indicated in this procedure will provide protection for nose wheel bearings and, if followed carefully, will allow 200 hours between bearing inspections and wheel bearing repacking.
Coat mating surfaces of seal (11) and spacer (10) with anti-seize lubricant, and then apply to seal surface which contacts wheel bearing seal. Insert axle tube (7) into wheel and place a seal (11) and spacer (10) on each side of wheel. Coat inside surface of the spacer (10) and outside surface of axle (7) where axle will slide inside the spacer (10) with anti-seize lubricant. Coat outer surface of buckets (4) with anti-seize lubricant. Position wheel into fork (9) and install buckets (4) into fork recesses. Tap buckets with nonmetallic hammer until seated in fork. Coat underside of axle stud (8) with anti-seize lubricant and install axle stud (8), washer (3), and nut (2). Tighten nut (2) until a slight drag can be felt when rotating the tire. Rotate the tire, by hand, and measure the force at the outside diameter of tire. Force should be between 3 and 5 pounds (13.4 to 22.2 N). Install new cotter pin (1) in nut (2). If cotter pin does not line up, back off nut to the next castellation. Again, verify the measured force at the outside diameter of the tire is between 3 and 5 pounds (13.4 to 22.2 N). Apply a small amount of anti-seize lubricant to threaded end of axle stud (8). Apply a small amount of anti-seize lubricant to both sides of wheel where seal (11) meets the snap ring of wheel bearing. Remove airplane from jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
Nose Wheel Tire and Tube Disassembly/Reassembly A.
Disassemble Nose Wheel Tire and Tube (Refer to Figure 204).
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WARNING: Injury can result from attempting to separate wheel halves with tube inßated. Take care to avoid damaging wheel halves when breaking tire beads loose. NOTE: (1) (2) (3) (4)
Refer to Tire Mounting Precautions before disassembly of nose wheel, tire and tube.
Ensure tube (8) is completely deßated then break beads of tire (7) loose. Remove thru-bolts (5) and separate wheel halves (6) and (9). Remove tire (7) and tube (8). Remove snap rings (1), grease seal (2) and bearing cones (3) from both wheel halves. NOTE:
B.
Bearing cups (4) are a press Þt in wheel halves (6) and (9) and should not be removed unless replacement is necessary. To remove bearing cups (4), heat each wheel half in boiling water for 15 minutes. Using an arbor press, press out bearing cup and press in new bearing cup while wheel is still hot.
Assemble Nose Wheel, Tire and Tube (Refer to Figure 204). NOTE: (1) (2)
Before reassembly of nose wheel, tire and tube, refer to Tire Mounting Precautions.
Clean the inside of both wheel halves with mineral spirits. Let it dry and apply Royco 103 or other MIL-C-16173 Type 1, Grade 1 protectant to all surfaces except the bearing outer race and the ßanges where the tire bead seats. NOTE:
(3) (4)
If replacing the bearing outer races, remove outer race and clean race cup bore surfaces of the wheel and apply a thin coating of protectant to the mating surface of the wheel. Install outer race while protectant is still wet. Brush coat the wheel surfaces around outer race after race installation to replace any protectant removed during the race installation. NOTE:
(5)
Some protectant in the snap ring (3) groove is permitted, but do not Þll the groove. Too much protectant can be removed with mineral spirits.
Place tire (7) and tube (8) on wheel half (9), aligning valve stem of tube with hole in wheel half. NOTE:
(6)
A swab may be necessary to apply protectant to the wheel material inside the bolt holes.
Lightweight point of tire is marked with a red dot on tire sidewall and heavyweight point of tube is marked with a contrasting color line (usually near valve stem). When installing tire, place these marks adjacent to each other.
Position wheel half (6) into position opposite assembled tire and wheel half (9), applying light force to keep wheel halves together. Do not pinch tube (8) between wheel halves.
CAUTION: Uneven or improper torque of thru-bolt nuts (10) can cause bolt failure with resultant wheel failure. (7)
Insert thru-bolts (5) through wheel halves and torque nuts (10) evenly to 150 inch-pounds (16.9 N.m). (8) Inßate tube to set tire beads, then adjust tire pressure. Refer to Chapter 12, Tires - Servicing. (9) Clean bearing cones (3) and repack with clean wheel bearing grease. (10) Assemble bearing cones (3) and grease seals (2) into both wheel halves, and secure with snap rings (1).
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Nose Wheel, Tire and Tube Installation Figure 204 (Sheet 1)
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14.
Inspection and Checks A.
Wheel Inspection. (1) Disassemble wheel. Refer to Main Wheel, Tire and Tube Disassembly/Reassembly or Nose Wheel, Tire and Tube Disassembly/Reassembly. (2) Visually inspect the wheel halves for cracks, corrosion, or other damage. Make sure to dye penetrant inspect any areas with suspected cracks. Cracked or badly corroded parts must be replaced. Small nicks, pits, and scratches may be polished out with Þne 400 grit wet or dry sandpaper and reÞnished. (3) Inspect bearing cups for looseness, scratches, pitting, corrosion or evidence of overheating. Replace cup if any defect exists. (4) Visually inspect bearing cones for nicks, scratches, water staining, spalling, heat discoloration, roller wear, cage damage, cracks or distortion. Replace if defective or worn. (5) Install new wheel bolts or the removed wheel bolts that have passed a visual and magnetic particle inspection. NOTE:
(6) (7)
If the wheel bolts are to be used again, they must pass an inspection for cracks, bending, thread damage, and excessive corrosion. A complete magnetic particle inspection must be done on each bolt, especially in the radius under the head and in the threads adjacent to the bolt shank.
Inspect thru-bolt nuts for self-locking feature. Replace nut if doubtful. Reassemble wheel. Refer to Main Wheel, Tire and Tube Disassembly/Reassembly or Nose Wheel, Tire and Tube Disassembly/Reassembly.
B.
Service Brake Disc. (Refer to Figure 202). (1) Discs are plated for special applications only; therefore, rust in varying degrees can occur. If a powder rust appears, one or two taxi/braking applications should wipe the disc clear. Rust allowed to progress beyond this point may require removal of the disc from the wheel assembly to properly clean both faces. (2) Wire brushing, followed by sanding with 220 grit sandpaper, can restore the braking surfaces for continued use.
C.
Inspect Brake Line For ChafÞng (Airplanes 20800001 thru 20800039). (1) Inspect brake line for chafÞng against lower main landing gear axle Þtting and fairing. If chafÞng is evident, brake line shall be re routed in accordance with Service Bulletin CAB85-4.
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MODEL 208 MAINTENANCE MANUAL WHEELS AND BRAKES - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the wheels and brakes in a serviceable condition.
Task 32-40-00-220 2.
Brakes Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the brakes.
B.
Special Tools (1) None.
C.
Access (1) None.
D.
Do the Brakes Detailed Inspection (Refer to Figure 601). (1) Remove the brake from the airplane. Refer to Brake Assembly Removal/Installation. (2) Disassemble the brake. Refer to Brake Assembly Disassembly/Reassembly. (3) Examine the brake linings for deterioration and maximum permissible wear. Replace the lining when worn to 0.100 inch (2.5 mm). (4) Examine the brake cylinder bores for evidence of scoring and deterioration. Replace scored cylinders. (5) Reassemble the brake. Refer to Brake Assembly Disassembly/Reassembly. (6) Before you install the brake, examine the disc for warpage, wear, grooves, deep scratches, and excessive general pitting or coning (refer to dimension A-A of Figure 601). (a) Coning beyond 0.015 inch (0.38 mm) in either direction is cause for replacement. (b) Single or isolated grooves up to 0.030 inch (0.76 mm) deep are not cause for replacement, although general grooving of the disc faces will reduce lining life. NOTE:
(7) (8)
Heat checks may develop on the wearing surface of the disc. Heat checks are considered to be superficial surface cracks and are not detrimental to braking performance, although brake disc replacement is necessary if any one crack has a length greater than 0.500 inch (12.7 mm), or a depth greater than 0.250 inch. (6.3 mm).
Replace the brake disc if more than three cracks are found in a disc, or if more than one crack per 90 degree quadrant is found in a disc. Install the brake. Refer to Brake Assembly Removal/Installation.
E.
Restore Access (1) None. End of task Task 32-40-00-610 3.
Brake System Servicing A.
General (1) This task gives the procedures to do the brake system servicing.
B.
Special Tools (1) MIL-PRF-5606 Hydraulic Fluid
C.
Access (1) Open the left engine cowling door.
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Brake Assembly Figure 601 (Sheet 1)
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Do the Brake System Servicing. (1) For the procedures necessary to do the brake system servicing, refer to Chapter 12, Hydraulic Fluid - Servicing.
E.
Restore Access (1) Close the left engine cowling door. End of task Task 32-40-00-222 4.
Main Landing Gear Wheels and Tires Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the main landing gear wheels and tires.
B.
Special Tools (1) Aircraft Jacks
C.
Access (1) None
D.
Do the Main Landing Gear Wheels and Tires Detailed Inspection. (1) If necessary, lift the main landing gear wheel and tire assemblies from the ground to turn the wheel assembly for this inspection. Refer to Chapter 7, Jacking - Maintenance Practices. (2) Use mild soap and water to remove oil, grease, and mud from the tires. (3) Examine the tires for wear, cuts, abrasion, flat spots, and correct pressure. Refer to Chapter 12, Tires-Servicing. (a) If the tire pressure is less than 85% of the recommended pressure, remove the tire from service for further inspection. (b) If the tires show signs of wear on the inside or outside edge, interchange the tires as necessary to the opposite axles to help get uniform tire wear. 1 If there is more than normal tire wear, examine the alignment. Refer to Main Landing Gear - Adjustment/Test. (4) Examine the wheel assemblies for condition, cracks (especially at the bolt holes), and corrosion. (a) If there is grease on the outside of the wheel, replace the grease seal on the applicable wheel half. (5) If airplane was jacked for the inspection, lower the airplane and remove the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
E.
Restore Access (1) None End of task Task 32-40-00-224 5.
Nose Landing Gear Wheel and Tire Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the nose landing gear wheel and tire
B.
Special Tools (1) Aircraft Jacks
C.
Access (1) None
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Do the Nose Landing Gear Wheel and Tire Detailed Inspection. (1) If necessary, lift the nose landing gear wheel and tire assembly from the ground to turn the wheel assembly for this inspection. Refer to Chapter 7, Jacking - Maintenance Practices. (2) Use mild soap and water to remove oil, grease, and mud from the tire. (3) Examine the tire for wear, cuts, abrasion, flat spots, and correct pressure. (a) If the tire pressure is less than 85% of the recommended pressure, remove the tire from service for further inspection. (4) Examine the wheel assembly for condition, cracks (especially at bolt holes), and corrosion. (a) If there is grease on the outside of the wheel, replace the grease seal on the applicable wheel half. (5) If airplane was jacked for the inspection, lower the airplane and remove the jacks. Refer to Chapter 7, Jacking - Maintenance Practices.
E.
Restore Access (1) None End of task Task 32-40-00-710 6.
Brakes Operational Check A.
General (1) This task gives the procedures to do an operational check of the brakes.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do an Operational Check of the Brakes. (1) Move the airplane to the parking ramp. (2) Set the parking brake. The parking brake must lock without more than necessary tension on the control and hold and release freely. NOTE: (3) (4) (5) (6) (7) (8)
Brake checks must include both pilot and copilot positions.
Start the engine and obey all operating limitations. Refer to Pilot's Operating Handbook and Approved Flight Manual. Release the parking brake. Taxi the airplane. Apply pressure to the pilot's brakes. (a) Make sure that the brakes do not drag, fade, or bypass fluid. (b) Make sure that the pedals do not oscillate from a warped or incorrectly aligned disc. Apply pressure to the copilot's brakes. (a) Make sure that the brakes do not drag, fade, or bypass fluid. (b) Make sure that the pedals do not oscillate from a warped or incorrectly aligned disc. Shut down the engine. Refer to Pilot's Operating Handbook and Approved Flight Manual.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL WHEELS AND BRAKES - CLEANING/PAINTING 1.
General A.
2.
Tools, Equipment and Materials A.
3.
This section provides information on cleaning, preparing and painting the wheels.
Refer to for a list of required tools, equipment and materials.
Wheel Preparation and Painting NOTE: A.
Wheels are made of aluminum and are painted with Vestal White baking enamel. Do not paint bearings or working surfaces.
Preparing Wheels For Painting. (1) Strip original finish of part following recommendations of stripper manufacturer. (2) Degrease and remove heavy soil by solvent wipe with clean cloth saturated in Methyl n-Propyl Ketone or equivalent. NOTE:
Cloth shall be folded each time the surface is wiped to present a clean area and avoid redepositing of grease.
B.
Applying Chemical Film to Unpainted Areas. (1) Refer to Main Landing Gear - Cleaning and Painting for generic instructions applicable to application of chemical film treatment to aluminum wheels.
C.
Priming Aluminum Components. (1) Refer to Main Landing Gear - Cleaning and Painting for generic instructions applicable to priming of aluminum wheels.
D.
Application of Baking Enamel. (1) Thin enamel to spray viscosity with toluene. (2) Spray one coat and air-dry for 10 minutes. (3) Apply a second coat and allow to flash off. (4) Bake for 5 to 10 minutes at 280°F to 300°F. (5) Dry film thickness should be 0.0015 to 0.0025 inch.
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MODEL 208 MAINTENANCE MANUAL WHEELS AND BRAKES - APPROVED REPAIRS 1.
General A.
2.
This section provides information on approved repairs to wheels, brakes and brake linings.
Approved Repair Methods A.
Wheels and Brakes. (1) Replace all worn or damaged parts. (2) Polish out minor nicks and scratches, using 400-grit wet or dry sandpaper.
B.
Brake Linings (Airplanes 20800001 Thru 20800135 and 208B0001 Thru 208B0102). (1) Remove brake backplates and pressure plate. Refer to Wheels and Brakes - Maintenance Practices. (2) Place a backplate on a table with backplate lining side down flat. Center a 9/64-inch (or slightly smaller) punch in rolled rivet and hit punch sharply with a hammer. Punch out all rivets securing linings to pressure plate in same manner. NOTE: (3) (4) (5) (6) (7) (8)
C.
A rivet setting kit, part number 199-00100, is available from the Cessna Supply Division. This kit consists of an anvil and a punch.
Clamp flat sides of anvil in a vise. Align new lining on backplate and place rivet in one hole with rivet head in lining. Place rivet head against anvil. Center rivet-setting punch on lips of rivet. While holding backplate down firmly against lining, hit punch with hammer to set rivet. Repeat blows on punch until lining is firmly against backplate. Realign lining on backplate and install remaining rivet. Install new linings on pressure plate in same manner. Install brake backplates and pressure plate. Refer to Wheels and Brakes - Maintenance Practices.
Brake Linings (Airplanes 20800136 and On, 208B0103 and On, and All Spares). (1) Remove brake backplates and pressure plate. Refer to Wheels and Brakes - Maintenance Practices. (2) If composition linings are being replace by metallic linings, drill out rivets using a 5/32 drill. (3) Install lining carrier pins (part number 177-00300) in backplate and pressure plate using a rivetsetting punch. Make sure pins are firmly installed. (4) Apply a light film of glue to the backside of each lining and allow to dry until tacky, and then snap new lining onto the carrier pins. NOTE: (5) (6) (7)
The glue is used only to retain the linings in place during reassembly of brake. It will burn off at first brake application.
If metallic linings are being replaced with metallic linings, snap old lining off carrier pins, then complete Step (4). Install brake backplate and pressure plate. Refer to Wheels and Brakes - Maintenance Practices. Burn in new brakes. Refer to Wheels and Brakes - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL BRAKE MASTER CYLINDER - MAINTENANCE PRACTICES 1.
General A.
2.
3.
This section provides information on brake master cylinder removal, installation and disassembly/ reassembly.
Master Cylinder Removal/Installation A.
Remove Master Cylinder (Refer to Figure 201). (1) Drain hydraulic fluid from brake system. (2) Remove floorboard cover immediately forward of pilot's control column mast. (3) Disconnect hoses (3), (4), (6), and (12) from master cylinders (5) and (9). Cap or plug ports and hoses. (4) Remove cotter pins (8) and clevis pins (7) from upper connection at rudder pedals of each master cylinder. (5) Remove cotter pins (11) and clevis pins (10) at floorboard mounting points and remove master cylinder.
B.
Install Master Cylinder (Refer to Figure 201). (1) Place master cylinders to floorboard mounting points, install clevis pins (10), and secure with cotter pins (11). (2) Connect piston rod clevis to rudder pedal bellcranks and secure with clevis pins (7) and cotter pins (8). (3) Connect brake hoses (3), (4), (6), and (12) to master cylinders (5) and (9). (Upper port is inlet, lower port is outlet.) (4) Bleed brake system. Refer to Wheels and Brakes - Maintenance Practices. (5) Install floorboard covers removed for access.
Master Cylinder Disassembly/Reassembly A.
Disassemble Master Cylinder (Refer to Figure 201). (1) Drain residual hydraulic fluid from open ports of body (9). (2) Loosen jam nut (16) and remove clevis (17) and nut (16) from piston rod (10). (3) Remove snap ring (15), and using piston rod (10), pull plug (11) from body. (9) (4) Remove packings (12) and (13) and backup ring (14) from plug (11) and discard. (5) Remove return spring (8) and washer (7) from piston rod (10). (6) Loosen and remove nut (6), allowing removal of spring (5), piston (3), Stat-O-Seal (2), and compensating sleeve (1). (7) Remove packing (4) from piston (3).
B.
Assemble Master Cylinder (Refer to Figure 201). (1) Using clean hydraulic fluid (MIL-PRF-5606) as lubricant, assemble new packing (4) in groove of piston (3). (2) Assemble Stat-O-Seal (2), piston (3), and spring (5) onto piston rod (10). Small diameter of spring (5) should rest against nut (6). Install nut (6) and tighten until clearance between piston (3) and Stat-O-Seal is 0.040 inch, +0.005 or -0.005 inch. (3) Install washer (7) and return spring (8) onto assembled piston rod and, with cylinder bore of body (9) lubricated with hydraulic fluid, insert piston rod assembly into body (9). (4) Place compensating sleeve (1), notched end toward piston, over piston rod (10). (5) Lubricate plug (11) with clean hydraulic fluid and insert new backup ring (14) and packing (13) in interior groove of plug (11). (6) Install new packing (12) onto exterior groove of plug (11). (7) Slide assembled plug (11) over piston rod (10) and into body (9). Install snap ring (15). (8) Screw jam nut (16) and clevis (17) onto piston rod end.
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Master Cylinder Installation Figure 201 (Sheet 1)
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Master Cylinder Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL PARKING BRAKE VALVE - MAINTENANCE PRACTICES 1.
General A.
2.
This section contains information on parking brake valve removal and installation. A faulty parking brake valve is not repairable, and must be replaced as an assembly.
Parking Brake Valve Removal/Installation A.
Remove Parking Brake Valve (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7)
B.
The parking brake valve is located on the aft side of the firewall (FS 100.00) between and just below the pilot's rudder pedals.
With parking brake handle (9) in off position, drain hydraulic fluid from brake system. Remove scuff plate and floorboard section forward of pilot's rudder pedals. Disconnect master cylinder hoses (12) and (13) from parking brake valve (11). Cap ports and hoses. Disconnect brake lines (1) and (2) from parking brake valve (11). Cap lines and ports. Loosen clamp bolt (3) on control lever (4) and remove control wire (5). Remove two bolts (15) attaching valve (11) to firewall. Remove valve (11) from airplane.
Install Parking Brake Valve (Refer to Figure 201). (1) Position valve (11) on aft side of firewall (10), aligning holes in valve with holes in firewall (10); install bolts and washers. (2) Remove caps from hoses (12) and (13) and lines (1) and (2) and connect to appropriate port of valve (11). (3) Install clamp bolt (3), washer, and nut on control lever (4) so bolt will swivel in control lever. (4) If knob (9) and control cable (5) were removed, install wire (5) and nut on forward side instrument panel, and Iock nut and knob (9) on aft side of panel. (5) Insert control wire (5) through clamp bolt (3) on control lever (4) and torque clamp bolt nut (16) to 15 inch pounds. (6) Bend control wire to 90 degrees to prevent disconnection in the event nut (16) should become loose. (7) Bleed brake system. Refer to Wheels and Brakes - Maintenance Practices. (8) Rig control knob (9) so that there is a 0.12-inch (minimum) cushion between locknut and knob (9) when control lever (4) is in OFF position. (9) Install floorboard section and scuff plate removed for access.
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Parking Brake Valve Installation Figure 201 (Sheet 1)
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Parking Brake Valve Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL NOSE GEAR STEERING - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Nose Gear Steering Figure 101 (Sheet 1)
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Nose Gear Steering Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL NOSE GEAR STEERING - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
This section provides description, operation, removal/installation and rigging procedures related to the nose gear steering system.
Nose gear steering is accomplished by the pilot’s rudder pedals. The steering system utilizes an adjustable steering tube/bungee assembly to connect torque arms on the rudder pedal shafts to a steering bellcrank mounted at the top of of the nose gear trunnion. A nylon boot seals off the cabin from the engine compartment where the steering tube passes through the firewall. The nose gear can be steered approximately 15 degrees each side of neutral using the steering system.
Nose Gear System Removal/Installation A.
Remove Nose Gear System (Refer to Figure 201). (1) Open upper cowling and remove scuff plate and floorboard cover plates around pilot’s rudder pedals to gain access to steering linkage. (2) With nose wheel set straight ahead and rudder pedals in neutral, note number of exposed threads on shank of rod end (12); loosen jam-nut (13). NOTE: (3) (4) (5) (6)
Remove bolt (11) and cut aft sta-tie (17). Back rod end (12) out of steering tube (7) and remove. Disconnect steering linkage at rod end (2) by removing bolt (4). Retain bushing (1) and washers. Withdraw linkage from cabin through bushing (15) into engine compartment. NOTE:
(7) (8) B.
Rudder pedal assembly can be pinned in neutral by utilizing alignment arms in rudder pedal torque tubes.
It may be necessary to remove bungee (5) from tube (7) to clear firewall.
Remove steering linkage from airplane. To remove remaining rudder system linkage, refer to Chapter 27, Rudder - Maintenance Practices .
Install Nose Gear System (Refer to Figure 201). (1) Set nose wheel straight ahead and block rudder pedals in neutral. NOTE:
Rudder pedal assembly can be pinned in neutral by utilizing alignment arms in rudder pedal torque tubes.
(2)
Ensure barrel nut (8) is threaded onto shaft (9A) with equal number of threads showing on each side. (3) From engine compartment side of firewall, insert steering tube (7) through bushing (15) into cabin. (4) Install bungee (5) if removed. (5) Connect rod end (2) to steering bellcrank (3) with bolt (4) and spacer (1). (6) From cabin side of firewall, thread jam-nut (13) onto bearing (12), and install as unit into steering tube (7) the same number of threads as noted during removal. (7) Rig steering linkage. (8) Install bolt (11) to connect bearing (12) to link (10). (9) Secure boot (18) to retainer (16) using new aft sta-tie (17). (10) Install floorboard covers and scuff plate removed to gain access.
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Nose Gear Stearing Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL C.
Rig Nose Gear Steering (Refer to Figure 201). (1) Block nose wheel straight ahead and rudder pedals in neutral. NOTE:
(2) (3) (4) (5) (6) (7) (8)
Rudder pedal assembly can be pinned in neutral by utilizing alignment arms in rudder pedal torque tubes. Verify that rudder gust lock is disengaged before rigging nose gear steering system.
Adjust rudder trim pointer so pointer is on the centerline of airplane mark. Gain access to rudder pedal torque arm (9) by removing scuff plate and floorboard covers as required. Remove bolt (11) and (23), if installed. Rotate forward end of shaft (9A) up and thread on link (10) until 0.50 inch of thread is showing on shaft (9A). Connect rod end (12) to link (10) with bolt (11). Install bolt (23), washers (22) and bearing and bushing (21) into guide (20). Ensure rod end (12) is sufficiently threaded into steering tube (7) by inserting safety wire into inspection hole in steering tube (7). Ensure rod end (2) has sufficient thread engagement with bungee shaft by inserting safety wire into inspection hole in rod end (2).
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33 CHAPTER
LIGHTS
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MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
33-00-00
Page 1
Aug 1/1995
33-10-00
Pages 1-2
Aug 1/1995
33-10-00
Pages 201-215
Apr 1/1996
33-20-00
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Aug 1/1995
33-20-00
Pages 201-212
Aug 1/1995
33-20-00
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Jun 1/2011
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Aug 1/1995
33-40-10
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Apr 1/2010
33-41-00
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Aug 1/1995
33-42-00
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Aug 1/1995
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Aug 1/1995
33-44-00
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Aug 1/1995
33-44-10
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Aug 2/2004
33-45-00
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Apr 1/1996
33-46-00
Pages 201-202
Aug 1/1995
33-Title 33-List of Effective Pages 33-Record of Temporary Revisions 33-Table of Contents 33-List of Tasks
33 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS LIGHTS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-00-00 33-00-00 33-00-00 33-00-00
FLIGHT COMPARTMENT LIGHTING - DESCRIPTION AND OPERATION . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-10-00 Page 1 33-10-00 Page 1 33-10-00 Page 1
FLIGHT COMPARTMENT LIGHTING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rheostat Control Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transistorized Dimming Assembly Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . Instrument Panel Post Light Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overhead Console Floodlight Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overhead Console Post Light Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Tank Selector and Oxygen Control Annunciator Light Removal/Installation . . . Lower Instrument Panel Floodlight Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . Left Sidewall Circuit Breaker Panel Floodlight Removal/Installation (Bottom Left Side of Instrument Panel Glareshield) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Left Sidewall Circuit Breaker Panel Floodlight Removal/Installation (Under Lower Left Side of Instrument Panel Glareshield) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Wheel Maplight Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo Dome Light Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outside Air Temperature Gage Post Light Removal/Installation . . . . . . . . . . . . . . . . . .
33-10-00 Page 201 33-10-00 Page 201 33-10-00 Page 201 33-10-00 Page 201 33-10-00 Page 201 33-10-00 Page 205 33-10-00 Page 205 33-10-00 Page 205 33-10-00 Page 205
PASSENGER/CARGO COMPARTMENT LIGHTING - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page 1 Page 1 Page 1 Page 1
33-10-00 Page 205 33-10-00 Page 211 33-10-00 Page 211 33-10-00 Page 211 33-10-00 Page 211 33-20-00 Page 1 33-20-00 Page 1 33-20-00 Page 1
PASSENGER/CARGO COMPARTMENT LIGHTING - MAINTENANCE PRACTICES . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Reading Light Removal/Installation (Model 208) . . . . . . . . . . . . . . . . . . . . . Passenger Reading Light Switch Removal/Installation (Model 208) . . . . . . . . . . . . . . Passenger Reading Light Removal/Installation (Model 208B) . . . . . . . . . . . . . . . . . . . Passenger Reading Light Switch Removal/Installation (Model 208B). . . . . . . . . . . . . No Smoking/Fasten Seat Belt Annunciator Warning Panel Removal/Installation. . . Cargo Entrance Door Light Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Entrance Door and Center Aisle Light Removal/Installation . . . . . . . . . . . Door Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Timer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lower Instrument Panel Passenger Compartment Light Switch Removal/ Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-20-00 Page 201 33-20-00 Page 201 33-20-00 Page 201 33-20-00 Page 201 33-20-00 Page 201 33-20-00 Page 201 33-20-00 Page 206 33-20-00 Page 206 33-20-00 Page 206 33-20-00 Page 206 33-20-00 Page 211
PASSENGER/CARGO COMPARTMENT LIGHTING - INSPECTION/CHECK. . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger/Cargo Compartment Lighting Operational Check . . . . . . . . . . . . . . . . . . . .
33-20-00 Page 601 33-20-00 Page 601 33-20-00 Page 601
EXTERIOR LIGHTING - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-40-00 Page 1 33-40-00 Page 1 33-40-00 Page 1
LANDING AND TAXI RECOGNITION LIGHTS - MAINTENANCE PRACTICES . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Landing and Taxi Recognition Lights Removal/Installation . . . . . . . . . . . . . . . . . . . . . . Landing and Taxi Recognition Lights Adjustment/Test . . . . . . . . . . . . . . . . . . . . . . . . . .
33-40-10 Page 201 33-40-10 Page 201 33-40-10 Page 201 33-40-10 Page 201
33-20-00 Page 211
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MODEL 208 MAINTENANCE MANUAL NAVIGATION LIGHTS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stinger Navigation Light Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Tip Navigation Light Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-41-00 Page 201 33-41-00 Page 201 33-41-00 Page 201 33-41-00 Page 201
ANTI-COLLISION STROBE LIGHTS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flash Tube Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Pack Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-42-00 Page 201 33-42-00 Page 201 33-42-00 Page 201 33-42-00 Page 201
FLASHING BEACON - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flashing Beacon Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flasher Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Resistor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-43-00 Page 201 33-43-00 Page 201 33-43-00 Page 201 33-43-00 Page 201 33-43-00 Page 201
ICE DETECTOR LIGHT - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ice Detector Light Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-44-00 Page 201 33-44-00 Page 201 33-44-00 Page 201
WINDSHIELD ICE INDICATOR LIGHT - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-44-10 Page 201 33-44-10 Page 201
COURTESY LIGHTS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Courtesy Light Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Timer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-45-00 Page 201 33-45-00 Page 201 33-45-00 Page 201 33-45-00 Page 201
EXTERNAL LIGHT SWITCHES - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . External Light Switches Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
33-46-00 Page 201 33-46-00 Page 201 33-46-00 Page 201
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LIST OF TASKS 33-20-00-710
Passenger/Cargo Compartment Lighting Operational Check
33-20-00 Page 601
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL LIGHTS - GENERAL 1.
Scope A.
2.
This chapter provides information on internal and external lighting.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Thermal Compound
Dow Corning #4
Dow Corning 3901 S. Saginaw Rd. Midland, MI 48640
To insulate transistors.
Locally Fabricate
To remove reading lights.
Schnee-Morehead Chemical Co. 111 N. Nursery Rd. Irving, TX 75017
To seal ice detector light housing.
Spanner Wrench Sealant
3.
Acryl-R-SS2S
Definition A.
This chapter is divided into sections and subsections to assist maintenance personnel in locating specific components and information. Consulting Table of Contents will further assist in locating a particular subject.
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MODEL 208 MAINTENANCE MANUAL FLIGHT COMPARTMENT LIGHTING - DESCRIPTION AND OPERATION 1.
General A.
2.
Flight compartment lighting consists of post lights for instrument panel lighting, outside air temperature gage, and overhead console lighting. Overhead flood lighting is provided for left sidewall switch and circuit breaker panel, instrument panel, and pilot's control pedestal. Overhead console incorporates two subminiature annunciator lights which illuminate labeling for ON/OFF oxygen control and the labeling for fuel selector. ln addition a maplight is internally mounted on the underside of pilot's control wheel to provide map lighting, and an instrument panel mounted annunciator panel.
Description and Operation A.
There are four concentric-type dual lighting controls installed on lower-left portion of instrument panel to left of control pedestal. These four controls vary intensity of instrument panel lighting, left sidewall switch and circuit breaker panel lighting, control pedestal lighting, and overhead console lighting. Following paragraphs describe controls that operate flight compartment lighting. (1) Large (outer) knob, labeled L FLT Panel, varies intensity of post lights illuminating left portion of instrument panel directly in front of pilot. Control also varies integral lighting intensity of digital clock, HSl, ADI, and radio instruments. Small (inner) knob, labeled FLOOD, varies brightness of right overhead floodlight which provides light for left map. Clockwise rotation of either knob increases lamp brightness and counterclockwise rotation decreases brightness. (2) Large (outer) knob labeled, ENG INST, varies intensity of post lights illuminating engine instruments located on top center of instrument panel. Small (inner) knob, labeled RADIO, controls integral lights and digital readouts of avionics equipment. Clockwise rotation of either knob increases lamp brightness and counterclockwise rotation decreases brightness. However, extreme counterclockwise rotation of RADIO knob turns digital readouts on bright for daylight viewing. (3) Large (outer) knob labeled, LWR PANEL/PED/OVHD, varies intensity of floodlights illuminating lower center portion of instrument panel, pedestal, overhead console, and outside air temperature gage. Small (inner) knob, labeled SW/CKT BKR, varies intensity of lights illuminating left sidewall switch and circuit breaker panel. Clockwise rotation of either knob increases lamp brightness and counterclockwise rotation decreases brightness. (4) Large (outer) knob, labeled R FLT PANEL, varies intensity of post lights illuminating right portion of instrument panel directly in front of right front passenger. Small (inner) knob, labeled R FLOOD, varies brightness of left overhead floodlight which provides light for right map.
B.
Flight compartment lighting circuits incorporate three transistorized light-dimming assemblies that are controlled by flight compartment lighting controls. Three transistorized dimming assemblies are mounted in back of left sidewall circuit breaker panel. (1) Top-Mounted Dimming Assembly is a three-transistorized unit and controls light dimming for: R FLT PANEL lights, R FLOOD light, and L FLT PANEL lights. (2) Center-Mounted Dimming Assembly is a three-transistorized unit and controls light dimming for: ENG INST lights, RADIO lights, and L FLOOD light. (3) Bottom-Mounted Dimming Assembly is a two-transistorized unit and controls light dimming for: LWR PANEL/PED/OVHD lights and SW/CKT BKR lights.
C.
Instrument panel postlighting is provided for left and right removable flight panels, left and right nonremovable flight panels, engine instruments, and lower left removable panel. Refer to Pilot's Operating Handbook for operation of post lights. Postlights are protected by circuit breakers mounted in left sidewall switch and circuit breaker panel, and light dimming is accomplished by a transistorized dimming assembly mounted on aft side of left sidewall switch and circuit breaker panel.
D.
Overhead console lighting provides for flight compartment floodlighting consisting of three lights mounted in overhead console. Two outside floodlights provide illumination for instrument panel, and a center floodlight provides illumination for control pedestal. Overhead console also incorporates a postlight to illuminate emergency flap switches, and also two subminiature-type lights provide illumination for FUEL TANK SELECTOR annunciator and OXYGEN ON/OFF annunciator. Concentric rheostats coupled to transistorized dimming assemblies vary light intensities. Protection for circuits is provided by circuit breakers mounted in left sidewall switch and circuit breaker panel.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL E.
There are two floodlights encased in a light-directing shield mounted on forward top side of pedestal which provides Iumination for lower center portion of instrument panel. Concentric rheostats mounted on lower instrument panel varies lighting intensity. Protection for circuit is provided by circuit breaker mounted in Ieft sidewall circuit breaker panel labeled RADIO/FLOODLlGHT.
F.
There are two floodlights which provide lighting for left sidewall circuit breaker panel. A floodlight is mounted on bottom left side of instrument panel glareshield to provide lighting for switches mounted on top of sidewall circuit breaker panel. A second floodlight is mounted under the lower left side instrument panel and provides lighting for the lower half of left sidewall circuit breaker panel. It is controlled by a rheostat labeled S/W CKT BKR. This rheostat is installed on lower left side of instrument panel and controls lamp intensity for all of left sidewall circuit breaker panel floodlights. Protection for floodlights is provided by a 10-amp circuit breaker, labeled CABIN LIGHT, which is installed on left sidewall circuit breaker panel.
G.
A control wheel maplight is internally mounted in control wheel. A rheostat, located on lower right side of wheel, controls lamp intensity. Protection for circuit is provided by a circuit breaker, labeled MAPLIGHT, mounted in left circuit breaker panel.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FLIGHT COMPARTMENT LIGHTING - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Flight compartment lights include instrument lights, three overhead floodlights, post lights for the instruments and overhead console, left sidewall circuit breaker panel, lower instrument panel, and a maplight mounted internally in control wheel.
Rheostat Control Removal/Installation A.
Remove Rheostat Control (Refer to Figure 201). (1) Align setscrew holes in both knobs (8) and (10), then loosen setscrew (9) and remove knobs (10) and (8). (2) Remove nut (7) securing rheostat to instrument panel. (3) Remove rheostats (1), (2), (3) or (4) from back of panel and retain lockwasher (6) and nut (5) for reinstallation. (4) Remove shrink tubing (12). (5) Unsolder electrical lead (11) and tag for reinstallation.
B.
Install Rheostat Control (Refer to Figure 201). (1) Place a one- inch section of shrink tubing (12) over each electrical lead (11). (2) Solder electrical lead to rheostat terminal, slide shrink tubing over terminals, and shrink tubing using a heating gun. (3) Remove tags from electrical leads. (4) Assemble nut (5) lockwasher (6) to rheostats (1), (2), (3) or (4). (5) Insert rheostat through instrument panel and secure using nut (7). (6) Install knobs (8) and (10), align setscrew holes in knobs, insert setscrew (9), and tighten.
Transistorized Dimming Assembly Removal/Installation A.
Remove Transistorized Dimming Assemby (Refer to Figure 202). (1) Remove left sidewall circuit breaker panel by removing screws securing panel to side of airplane. (2) Disconnect electrical connector (4). (3) Remove screws (1) and washers (2) securing dimming assembly to back of panel. (4) To remove transistors for circuit checking, remove screws (5) and washers (6), and retain insulators (15) for reinstallation.
B.
Install Transistorized Dimming Assemby (Refer to Figure 202). (1) Install transistors and insulators (15) on dimming assembly using screws (5) and washers (6). NOTE: (2) (3) (4)
4.
Apply Thermal Compound (Dow Corning #4 or equivalent) to all transistors.
Install dimming assembly using screws (1) and washers (2). Connect electrical connector (4). Install left sidewall circuit breaket panel to side of airplane using screws.
Instrument Panel Post Light Removal/Installation A.
Remove Instrument Panel Post Light (Refer to Figure 203). (1) Remove cap assembly (1) from socket assembly (3). (2) Remove bulb (2) from socket assembly (3). (3) Unscrew connector (8) from socket assembly (3), and retain nut (7), lockwasher (6), washer (5), and lockwasher (4) for reinstallation. (4) Remove socket assembly (3) from instrument panel.
B.
Install Instrument Panel Post Light (Refer to Figure 203). (1) Assemble lockwasher (4) and insert socket assembly (3) through instrument panel. (2) Assemble washer (5), lockwasher (6), and nut (7), and screw connector (8) onto socket assembly (3). (3) Insert bulb (2) into socket assembly (3). (4) Push cap assembly (1) over socket assembly (3).
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MODEL 208 MAINTENANCE MANUAL
Flight Compartment Lighting Controls Figure 201 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Transistorized Light Dimming Installations Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Instrument Panel Post Light Installation Figure 203 (Sheet 1)
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5.
6.
7.
8.
9.
Overhead Console Floodlight Removal/Installation A.
Remove Overhead Console Floodlight (Refer to Figure 204). (1) Unscrew cap (7) from socket assembly (6). (2) Remove two screws (8) securing socket assembly (6) to overhead console. (3) Remove socket assembly (6) from overhead console. (4) Slide shield (4) off of socket assembly. (5) Press in slighty on bulb (5) and turn counterclockwise to remove from socket assembly (6). (6) Remove electrical leads (9) and tag for reinstallation.
B.
Install Overhead Console Floodlight (Refer to Figure 204 ). (1) Connect electrical leads (9) and remove tags. (2) Insert bulb (5) into socket assembly (6), press in slightly, and turn clockwise. (3) Slide shield (4) over socket assembly (6). (4) Insert socket assembly (6) into overhead console and secure using two screws (8). (5) Screw cap (7) onto socket assembly (6).
Overhead Console Post Light Removal/Installation A.
Remove Overhead Console Post Light (Refer to Figure 204). (1) Remove cap assembly (3) from socket assembly (1). (2) Remove bulb (2) from cap assembly (3).
B.
Install Overhead Console Post Light (Refer to Figure 204). (1) Insert bulb (2) into cap assembly (3). (2) Push cap assembly (3) over socket assembly (1).
Fuel Tank Selector and Oxygen Control Annunciator Light Removal/Installation A.
Remove Fuel Tank Selector and Oxygen Control Annunciator Light (Refer to Figure 204). (1) Remove socket (12) by unscrewing from overhead console. (2) Remove bulb (10) from socket and retain shield (11) for reinstallation.
B.
Install Fuel Tank Selector and Oxygen Control Annunciator Light (Refer to Figure 204). (1) Assemble shield (11) to socket (12). (2) Insert bulb (10) into socket (12) and screw socket into overhead console.
Lower Instrument Panel Floodlight Removal/Installation A.
Remove Lower Instrument Panel Floodlight (Refer to Figure 205 ). (1) Remove screws (1) securing shield (2) to pedestal. (2) Press in slightly on bulb (3) and turn counterclockwise to remove from socket (4).
B.
Install Lower Instrument Panel Floodlight (Refer to Figure 205). (1) Insert bulb (3) into socket (4), press in slightly and turn clockwise. (2) Install screws (1) securing shield (2) to pedestal.
Left Sidewall Circuit Breaker Panel Floodlight Removal/Installation (Bottom Left Side of Instrument Panel Glareshield) A.
Remove Floodlight (Refer to Figure 206). (1) Grasp hood (13) and gently pull to remove hood. (2) Push in slightly on bulb (12) and turn counterclockwise to remove bulb from socket (11). (3) Remove two screws (10) securing bracket (9) to glareshield. (4) Unscrew housing (7) from socket (11) and keep retainer clip (8) for reinstallation.
B.
Install Floodlight (Refer to Figure 206). (1) Place socket (11) through bracket (9) and retainer clip (8), and screw into housing (7). (2) Install two screws (10) securing bracket (9) to glareshield. (3) Insert bulb (12) into socket (11), push in slightly and turn clockwise. (4) Push hood (13) over socket (11) after ensuring alignment pin matches alignment groove in hood.
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MODEL 208 MAINTENANCE MANUAL
Overhead Console Lighting Installation Figure 204 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Overhead Console Lighting Installation Figure 204 (Sheet 2)
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Lower Instrument Panel Floodlighting Installation Figure 205 (Sheet 1)
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Circuit Breaker Floodlights Installation Figure 206 (Sheet 1)
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Circuit Breaker Floodlights Installation Figure 206 (Sheet 2)
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10.
11.
12.
13.
Left Sidewall Circuit Breaker Panel Floodlight Removal/Installation (Under Lower Left Side of Instrument Panel Glareshield) A.
Remove Left Sidewall Circuit Breaker Panel Floodlight (Refer to Figure 206). (1) Grasp hood (15) and gently pull to remove hood. (2) Push in slightly on bulb (16) and turn counterclockwise to remove bulb from socket (17). (3) Remove two screws (19) and two nuts (23) securing bracket (18) to plate (22). (4) Unscrew housing (21) from socket (17) and retain retainer clip (20) for reinstallation.
B.
Install Left Sidewall Circuit Breaker Panel Floodlight (Refer to Figure 206). (1) Place socket (17) through bracket (18) and retainer clip (20) and screw into housing (21). (2) Install two screws (19) and two nuts (23) securing bracket (18) to plate (22). (3) Insert bulb (16) into socket (17), push in slightly and turn clockwise. (4) Push hood (15) over socket (17) after ensuring alignment pin matches alignment groove in hood.
Control Wheel Maplight Removal/Installation A.
Remove Control Wheel Maplight (Refer to Figure 207). (1) Rotate pilot's control wheel approximately 90 degrees clockwise for access to maplight. (2) Press in on bulb (1) and turn counterclockwise to remove bulb. (3) Remove nut (3), washer (4), and socket (2). (4) Loosen setscrew (6) in knob (5) and remove knob. (5) Remove nut (7) and rheostat (8) from lower left side of control wheel.
B.
Install Control Wheel Maplight (Refer to Figure 207). (1) Rotate pilot's control wheel approximately 90 degrees clockwise for access to maplight. (2) Place rheostat (8) through hole in lower side of control wheel. (3) Install nut (7) on rheostat (8) and tighten. (4) Install knob (5) on rheostat (8) and tighten setscrew (6). (5) Install socket (2), washer (4), and nut (3) and tighten. (6) Press bulb (1) into socket (2) and turn clockwise.
Cargo Dome Light Removal/Installation A.
Remove Cargo Dome Light (Refer to Figure 208). (1) Remove lens assembly (6) by removing screws. (2) Press in slightly on bulb (4) and turn counterclockwise to remove bulb from socket assembly (2). (3) Remove screws (5) and remove socket assembly (2) from splice plate (1).
B.
Install Cargo Dome Light (Refer to Figure 208). (1) Install screws (5) securing socket assembly (2) to splice plate (1). (2) Insert bulb (4) into socket assembly (2), press in slightly, and turn clockwise. (3) Ensure gasket (3) is in place. (4) Secure lens assembly (6) using screws.
Outside Air Temperature Gage Post Light Removal/Installation A.
Remove Outside Air Temperature Gage Post Light (Refer to Figure 209 ). (1) Remove outside air temperature gage to gain access to post light connector. (2) Remove cap assembly (1) from socket assembly (3). (3) Remove bulb (2) from socket assembly (3). (4) Unscrew electrical connector (8) from socket assembly (3). (5) Unscrew nut (7) from socket assembly (3) and retain nut (7), Iockwasher (6), washer (5), and lockwasher (4) for reinstallation.
B.
Install Outside Air Temperature Gage Post Light (Refer to Figure 209). (1) Assemble lockwasher (4) on socket assembly (3) and insert socket assembly (3) through trim panel. (2) Tighten nut (7) and screw electrical connector (8) onto socket assembly (3). (3) Install bulb (2) to socket assembly (3). (4) Install cap assembly (1) to socket assembly (3).
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Control Wheel Map Light Installation Figure 207 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (5)
Install outside air temperature gage.
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Dome Light Installation Figure 208 (Sheet 1)
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Outside Air Temperature Gage Post Light Installation Figure 209 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PASSENGER/CARGO COMPARTMENT LIGHTING - DESCRIPTION AND OPERATION 1.
General A.
2.
Model 208 passenger compartment lights consist of eight passenger reading lights, eleven reading lights on Model 208B Passenger, and NO SMOKING/FASTEN SEAT BELT annunciator warning panel on both models. Models 208 and 208B are equipped with three cabin lights (one light is located above center aisle, one above aft cargo door, and one above passenger entry door).
Description and Operation A.
Reading lights may be installed at each of aft passengers positions. Lights are located above window line in small convenience panels above each seat. A pushbutton type on-off switch mounted in each panel controls the lights. Lights can be pivoted in their mounting sockets to provide the most comfortable angle of illumination for each passenger. Reading lights are protected by a circuit breaker mounted in left sidewall circuit breaker panel, labeled RDNG LIGHT
B.
Panel may be installed in airplane to facilitate warning passengers of impending flight operations requiring fastening of seat belts and/or extinguishing of all smoking materials. Installation consist of an annunciator warning panel incorporating two lamps which illuminates, displaying international graphic symbolism for fasten seat belts and no smoking to rear cabin passengers. Panel is controlled by two switches, labeled NO SMOKE, SEAT BELT, located on lower left side of instrument panel. Protection for circuit is provided by a circuit breaker mounted in left sidewall circuit breaker panel, labeled SEAT BELT SIGN.
C.
There are three cargo/passenger lights installed in interior cabin headliner. One is above aft cargo door, another is directly opposite cargo door light, and a third is overhead and forward in the center aisle. Power is supplied from battery bus through a switch located on lower left instrument panel, labeled CABIN. Independent ON/OFF switching is controlled by switch(es) located just forward of cargo door and (Model 208 and 208B Passenger) passenger entry door. Protection is provided by clock circuit breaker, located in power box mounted in the upper left portion of engine compartment on firewall. Airplanes with courtesy light timer option use a solid-state timer which allows the lights to remain illuminated for a period of thirty minutes after airplane has been deplaned and secured.
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MODEL 208 MAINTENANCE MANUAL PASSENGER/CARGO COMPARTMENT LIGHTING - MAINTENANCE PRACTICES 1.
General A.
2.
3.
4.
5.
Passenger reading light maintenance practices consist of reading light removal/installation and reading light switch removal/installation.
Passenger Reading Light Removal/Installation (Model 208) A.
Remove Passenger Reading Light (Refer to Figure 201 ). (1) Pull lens assembly (13) from socket assembly (11). (2) Press in on bulb (12) and turn counterclockwise to remove. (3) Using a spanner wrench, unscrew socket assembly (11) counterclockwise and pull out of opening. (4) Disconnect electrical leads (9) and tag for reinstallation.
B.
Install Passenger Reading Light (Refer to Figure 201 ). (1) Connect electrical leads (9) to socket (11) and remove tags. (2) Using a spanner wrench, rotate socket assembly (11) approximately six turns counterclockwise, then screw clockwise into retainer (4). Electrical leads (9) should not be twisted. (3) Insert bulb (12) into socket (11), press in slightly, and turn clockwise. (4) Install lens assembly (13).
Passenger Reading Light Switch Removal/Installation (Model 208) A.
Remove Passenger Reading Light Switch (Refer to Figure 201 ). (1) Remove light. (2) Remove pushbutton (14) by gently pulling. (3) Remove nut (8). (4) Remove switch (3) by pushing it out of its mounting hole, and then pull it through light hole in cover (10). (5) Disconnect electrical wires (1) and (2) from switch (3), and tag for reinstallation. (6) Remove switch (3).
B.
Install Passenger Reading Light Switch (Refer to Figure 201 ). (1) Connect electrical leads (1) and (2) to switch (3), and push switch back through plenum (4) and insert switch (3) through its mounting holes in plenum (4), retainer fillet (5), cover (6), and cover (10). (2) Install nut (8) on switch (3) and tighten on shaft. (3) Install pushbutton on switch (3) by gently pushing. (4) Reinstall light.
Passenger Reading Light Removal/Installation (Model 208B) A.
Remove Passenger Reading Light (Refer to Figure 201 ). (1) Pull lens (12) from socket assembly (14). (2) Push in on bulb (13) and turn counterclockwise to remove. (3) Unscrew socket assembly (14) by turning counterclockwise. (4) Remove screws and washers securing electrical leads to socket assembly.
B.
Install Passenger Reading Light (Refer to Figure 201 ). (1) Connected electrical leads to socket assembly (14). (2) Rotate socket assembly (14) counterclockwise approximately eight turns then screw clockwise into bracket assembly (5). Electrical leads should not be twisted after installation. (3) Insert bulb (13) into socket assembly (14), press in slightly and turn clockwise. (4) Insert lens assembly (12) into socket assembly (14).
Passenger Reading Light Switch Removal/Installation (Model 208B). A.
Remove Passenger Reading Light Switch (Refer to Figure 201 ). (1) Remove reading light.
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Passenger Reading Lights Installation Figure 201 (Sheet 1)
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Passenger Reading Lights Installation Figure 201 (Sheet 2)
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Passenger Reading Lights Installation Figure 201 (Sheet 3)
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Passenger Reading Lights Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) (6) B.
6.
7.
8.
9.
Remove Wemac (11) by turning counterclockwise. Remove decorative nut (10) and dress ring (9) and remove cover (8). Open headliner as necessary for access to switch (3) electrical leads. Disconnect electrical leads from switch. Remove nut (7) and washer (6) and remove switch (3).
Install Passenger Reading Light Switch (Refer to Figure 201 ). (1) Position switch (3) through bracket(s) and install washer (6) and nut (7). (2) Connect electrical leads to switch (3) and close headliner. (3) Position cover (8) and install dress ring (9) and decorative nut (10). (4) Install wemac (11) by turning clockwise. (5) Install reading light.
No Smoking/Fasten Seat Belt Annunciator Warning Panel Removal/Installation A.
Remove No Smoking/Fasten Seat Belt Annunciator Warning Panel (Refer to Figure 202). (1) Remove screws (8) from escutcheon. (2) Disconnect plug (6) and remove sign assembly (7).
B.
Install No Smoking/Fasten Seat Belt Annunciator Warning Panel (Refer to Figure 202). (1) Connect plug (6). (2) Slide sign assembly (7) into position in escutcheon. (3) Install screws (8).
Cargo Entrance Door Light Removal/Installation A.
Remove Cargo Entrance Door Light (Refer to Figure 203). (1) Remove screws (10) and remove lens assembly (9). (2) Press in slightly on bulb (15) and turn counterclockwise to remove bulb from socket (6). (3) Remove screws (8) and remove escutcheon (7). (4) Remove screws (2) and washers (3) and (4) from socket (6) and remove terminals (5).
B.
Install Cargo Entrance Door Light (Refer to Figure 203). (1) Secure terminals (5) to socket (6) using screw (2) and washers (3) and (4). (2) Secure escutcheon (7) to bracket (1) using screws (8). (3) Insert bulb (15) into socket (6), press in slightly, and turn clockwise. (4) Secure lens assembly (9) using screws (10).
Passenger Entrance Door and Center Aisle Light Removal/Installation A.
Remove Passenger Entrance Door and Center Aisle Light (Refer to Figure 204 ). (1) Remove lens assembly (6) by removing screws. (2) Press in slightly on bulb (4) and turn counterclockwise to remove bulb from socket assembly (2). (3) Remove screws (5) and remove socket assembly (2) from bracket (1).
B.
Install Passenger Entrance Door and Center Aisle Light (Refer to Figure 204). (1) Install screws (5) securing socket assembly to bracket (1). (2) Insert bulb (4) into socket assembly (2), press in slightly, and turn clockwise. (3) Ensure gasket (3) is in place. (4) Secure lens assembly (6) using screws.
Door Switch Removal/Installation A.
Remove Door Switch (Refer to Figure 203). (1) Remove screws securing door panel (11). (2) Remove screws (12) securing switch (13) to bracket (14). (3) Lift switch up and out of hole and disconnect electrical leads by removing screws.
B.
Install Door Switch (Refer to Figure 203). (1) Connect electrical leads and position switch (13) in bracket (14). (2) Install screws (12) securing switch to bracket (14). (3) Install screws securing door panel (11).
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No Smoking/Fasten Seat Belts Lights Installation Figure 202 (Sheet 1)
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Cargo Entrance Door Light Installation Figure 203 (Sheet 1)
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Cargo Entrance Door Light Installation Figure 203 (Sheet 2)
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Passenger Entrance Door and Center Aisle Lights Installation Figure 204 (Sheet 1)
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10.
11.
Timer Removal/Installation A.
Remove Timer (Refer to Figure 203 ). (1) Disconnect electrical plug (18). (2) Remove four screws (17) securing timer (16). (3) Remove timer (16) from brackets (19).
B.
Install Timer (Refer to Figure 203). (1) Position timer (16) to brackets (19). (2) Install four screws (17) securing timer (16). (3) Connect electrical plug (18).
Lower Instrument Panel Passenger Compartment Light Switch Removal/Installation A.
Remove Lower Instrument Panel Passenger Compartment Light Switch (Refer to Figure 205 ). (1) Remove nut (10) securing switch (6) to instrument panel, push switch through instrument panel, and retain washers (8) and (9) for reinstallation. (2) Disconnect electrical leads and tag for reinstallation.
B.
Install Lower Instrument Panel Passenger Compartment Light Switch (Refer to Figure 205). (1) Connect electrical leads and remove tags. (2) Assemble washers (8) and (9), push switch (6) up and through instrument panel, and secure using nut (10).
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Passenger Compartment Light Switch Installation Figure 205 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PASSENGER/CARGO COMPARTMENT LIGHTING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the passenger/cargo compartment lighting system in a serviceable condition.
Task 33-20-00-710 2.
Passenger/Cargo Compartment Lighting Operational Check A.
General (1) This task gives the necessary information to do the operational check of the passenger/cargo compartment lighting system.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do a Detailed Inspection of the Passenger/Cargo Compartment Lighting. (1) Examine all passenger/cargo compartment lighting for condition and correct operation. (2) Examine the overhead reading light assemblies/cargo lighting assemblies, and electrical components for condition, accessible wiring for chafing, routing, and security. (3) Examine all lighting rheostats for condition and security.
E.
Do an Operational Check of the Passenger Cargo Compartment Lighting. (1) Cabin Lights (without timer) (a) Apply external electrical power to the airplane. (b) Make sure that the CABIN LTS circuit breaker on the J-Box panel in the engine bay is engaged. (c) One at a time, operate the switches that follow: • CABIN switch on the LIGHTS control panel in the cockpit • Passenger Door Switch (Forward of the Aft Passenger Door) • Cargo Door Switch (Forward of the Cargo Door). 1 Make sure that the lights that follow come on and go off when you operate each switch regardless of the other switch positions. • Forward Dome Light (Model 208B only) • Passenger Cabin Light • Passenger Door Light • Cargo Door Light (2 light bulbs) • Left Wing Courtesy Light • Right Wing Courtesy Light. NOTE:
(2)
(d) Seat (a) (b) (c) (d) (e) (f) (g)
One courtesy light is installed under the left wing, and one is installed under the right wing. The lights illuminate the area outside of the airplane adjacent to the crew entry doors. The lights operate in conjunction with the cabin lights and are controlled by the cabin light switch.
Set the CABIN switch on the LIGHTS control panel to the OFF position. Belt Sign and No Smoking Sign Make sure that the SEAT BELT SIGN circuit breaker is engaged. Set the SEAT BELT switch on the LIGHTS control panel to the ON position. Make sure that the Seat Belt sign comes on. Set the SEAT BELT switch on the LIGHTS control panel to the OFF position. Set the NO SMOKE switch on the LIGHTS control panel to the ON position. Make sure that the No Smoking sign comes on. Set the NO SMOKE switch on the LIGHTS control panel to the OFF position.
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MODEL 208 MAINTENANCE MANUAL (3)
Cabin Lights with timer (all cargo airplanes, and passenger airplanes if a timer is installed). (a) Make sure that the CABIN LTS circuit breaker on the J-Box panel in the engine bay is engaged. (b) Set the CABIN switch on the LIGHTS control panel to the ON position and make sure that the following lights operate correctly: NOTE:
(c)
(d)
(e)
One courtesy light is installed under the left wing, and one is installed under the right wing. The lights illuminate the area outside of the airplane adjacent to the crew entry doors. The lights operate in conjunction with the cabin lights and are controlled by the cabin light switch.
1 Forward Cabin Light (Model 208B only) 2 Main Cabin Light 3 Cargo Door Light (2 light bulbs) 4 Cabin Light opposite cargo door 5 Left Wing Courtesy Light 6 Right Wing Courtesy Light Do the following checks: 1 The momentary CABIN lights switch controls all lights regardless of the position of the cargo door light switch and the passenger door light switch, if installed. 2 On passenger airplanes, the passenger door light switch controls all lights except the cargo door light, and the cargo door light switch controls the cargo door light only. 3 On cargo airplanes, the cargo door light switch controls all lights. 4 Keep any of the lights on, and make sure that the timer shuts off those lights after approximately 30 minutes. Reading Lights (passenger airplanes) 1 Make sure that the RDNG LIGHT circuit breaker is engaged. 2 Make sure that each of the eight (Model 208) or the 14 (Model 208B) reading lights can be switched on and off with their own switch. Disconnect the external electrical power from airplane.
F.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL EXTERIOR LIGHTING - DESCRIPTION AND OPERATION 1.
General A.
2.
Exterior lights consist of landing lights, taxi/recognition, navigation, anti-collision strobe, flashing beacon, ice detector, and courtesy. Courtesy light circuit incorporates a solid-state timer which allows the lights to remain illuminated for a period of thirty minutes after airplane has been deplaned and secured.
Description and Operation A.
There are two landing lights installed on the airplane, one in each outboard leading edge adjacent to two taxi/recognition lights. The lights are controlled by switches, labeled LEFT LDG, RIGHT LDG, TAXI/RECOG, located on the lower left side of the instrument panel. Protection for circuit is provided by a circuit breaker mounted in left sidewall circuit breaker panel, labeled LEFT LDG, RIGHT LDG, TAXI LIGHT.
B.
There are three navigation lights installed on the airplane. One on each wing tip and one on stinger. Lights are controlled by a switch, labeled NAV, located on lower left instrument panel. Protection for circuit is provided by a circuit breaker, labeled NAV LIGHT, which is mounted in left sidewall circuit breaker panel
C.
Anti-collision strobe lights are mounted on each wing tip adjacent to navigation lights. Strobe lights are controlled by a switch, labeled STROBE, located on lower left instrument panel. Protection for circuit is provided by a circuit breaker, labeled STROBE LIGHT, mounted in left sidewall circuit breaker panel.
D.
A red flashing beacon is installed on tip of vertical fin. Beacon is controlled by a switch, labeled BCN, located on lower left instrument panel. A flasher is mounted on canted bulkhead at FS 388.68 on Model 208, and at FS 436.68 on the Model 208B. Since flasher is designed to accommodate two flashing beacon units, a (95 watt/60 ohm) resistor is installed just below flasher to eliminate radio noise feed-back. If an additional flasher is installed by customer, resistor can be removed from circuit without causing radio noise feedback. Protection for circuit is provided by a circuit breaker labeled BEACON LIGHT, mounted in left sidewall circuit breaker panel. If output driver portion of flasher assembly that is driving light assembly fails, useful life of the assembly may be extended by using other half of unit to drive light. This may be accomplished by crossing yellow and green wire within flasher assembly connector.
E.
Optional ice detector light is located on forward cabin top, forward wing root rib. Lamp is controlled by a switch, labeled WING LIGHT, located on the lower left instrument panel. Protection for circuit is provided by a circuit breaker, labeled WING ICE DET LIGHT, mounted in left sidewall circuit breaker panel. This light is standard equipment when flight into known-icing option is installed.
F.
Courtesy lights are located between wing strut fairing and fuselage on lower side of wings. Lights are controlled by a switch, labeled CABIN, mounted in lower left instrument panel. Lights use a solid-state timer (except the 208 Cargomaster) which allows lights to remain illuminated for a period of 30 minutes after airplane has been deplaned and secured.
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MODEL 208 MAINTENANCE MANUAL LANDING AND TAXI RECOGNITION LIGHTS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Landing and taxi recognition lights maintenance practices consist of landing and taxi light removal/ installation and adjustment/test.
Landing and Taxi Recognition Lights Removal/Installation A.
Remove Landing and Taxi Recognition Lights (Refer to Figure 201 ). (1) Remove screws (3) securing lens retainer (1) and remove lens (2). (2) Remove screws (4) and springs (9), pull lamp assembly forward and disconnect electrical leads. (3) Remove screws (6) securing bracket (5) and baffle assembly (5A) to plate (8) and remove lamp (7).
B.
Install Landing and Taxi Recognition Lights (Refer to Figure 201). (1) Place lamp (7) between bracket (5), baffle assembly (5A), and plate (8). Install screws (6). (2) Connect electrical leads to lamp (7). (3) Place screws (4) through lamp assembly and springs (9) and install lamp assembly. (4) Install lens (2) and lens retainer (1) with screws (3).
Landing and Taxi Recognition Lights Adjustment/Test A.
Adjust Landing and Taxi Recognition Lights (Refer to Figure 201). (1) Park airplane 50 feet from a wall or any suitable light reflecting surface. (2) Remove screws (3) securing lens retainer (1) and lens (2). (3) Level airplane. Refer to Chapter 8, Leveling - Maintenance Practices. Ensure pilot seat rail is on a horizontal plane. (4) Measure distance from ground surface up to center of lamps. (5) Apply external power and place master switch in the ON position. (6) Turn landing/taxi light switch ON. (7) Align lamps focal point straight forward and at a height 63 inches below center of lamps as measured in step 4, using alignment screws (1), (2), and (3). (8) Turn landing/taxi light switch OFF. (9) Disconnect external power. Turn master switch OFF. (10) Reinstall lens which was removed in step 2.
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Landing and Taxi Lights Installation Figure 201 (Sheet 1)
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Landing and Taxi Lights Installation Figure 201 (Sheet 2)
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Landing and Taxi Lights Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL NAVIGATION LIGHTS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Navigation lights maintenance practices consist of stinger navigation light removal/installation and wing tip navigation light removal/installation.
Stinger Navigation Light Removal/Installation A.
Remove Stinger Navigation Light (Refer to Figure 201). (1) Remove screws (8), lens retainer (7), lens (6), and gasket (4). (2) Press in on bulb (5) and turn counterclockwise to remove bulb. (3) Pull socket (3) from stinger and unscrew cap (2) to remove socket (3).
B.
Install Stinger Navigation Light (Refer to Figure 201). (1) Screw cap (2) on socket (3). (2) Press bulb (5) into socket (3) and turn clockwise. (3) Place socket (3) in position in stinger, then install gasket (4), lens (6), and lens retainer (7) with screws (8).
Wing Tip Navigation Light Removal/Installation A.
Remove Wing Tip Navigation Light (Refer to Figure 201). (1) Remove screws (7), lens retainer (6), and lens (8). (2) Press in on bulb (9) and turn counterclockwise to remove bulb.
B.
Install Wing Tip Navigation Light (Refer to Figure 201). (1) Press bulb (9) into lamp socket (11) and turn clockwise. (2) Install lens (8) and lens retainer (6) with screw (7).
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Navigation Lights Installation Figure 201 (Sheet 1)
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Navigation Lights Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL ANTI-COLLISION STROBE LIGHTS - MAINTENANCE PRACTICES 1.
General A.
2.
Anti-collision lights maintenance practices consist of flash tube removal/installation and power pack removal/installaion.
Flash Tube Removal/Installation
WARNING: Anti-collision strobe light system is a high-voltage device. Do not remove or touch flash tube assembly while in operation. Wait at least five minutes after turning off power before starting work. CAUTION: Extreme care should be taken when exchanging flash tube. Tube is fragile and can be easily cracked in a place where it will not be obvious visually. Make sure tube is seated properly on base of navigation light assembly and is centered in dome. A.
Remove Flash Tube (Refer to Figure 201 ). NOTE: (1) (2)
B.
3.
When checking defective power supply and flash tube, units from opposite wing may be used.
Remove screws (10) and remove lens retainer (9), taking care not to drop lens (8) or navigation light lens. Pull flash tube (6) from receptacle (4).
Install Flash Tube (Refer to Figure 201). (1) Install flash tube (6) in receptacle. (2) Position lens (8) through lens retainer (9), making sure gasket (7) is installed on lens (8). (3) Position navigation light lens over lamp, then install lens retainer (9) using screws (10).
Power Pack Removal/Installation A.
Remove Power Pack (Refer to Figure 201). (1) Remove screws (11). (2) Pull power pack (14) from wing tip and disconnect plug (15).
B.
Install Power Pack (Refer to Figure 201). (1) Connect plug (15) to power pack (14). (2) Slide power pack into wing tip making sure gasket (13) is installed, and install screws (11).
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MODEL 208 MAINTENANCE MANUAL
Anti-Collision Strobe Lights Installation Figure 201 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Anti-Collision Strobe Lights Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL FLASHING BEACON - MAINTENANCE PRACTICES 1.
General A.
2.
Flashing beacon maintenance practices consist of beacon removal/installation and flasher removal/ installation.
Flashing Beacon Removal/Installation
CAUTION: When installing or removing lamp (3) use a handkerchief or a tissue to prevent getting fingerprints on lamp. Fingerprints on lamp may shorten life of lamp.
3.
4.
A.
Remove Flashing Beacon (Refer to Figure 201). (1) Loosen screw securing clamp (5) then remove dome (1) and gasket (2). (2) Press in slightly on bulb (3) and turn counterclockwise to remove bulb from socket (6). (3) Remove screws securing fin tip (7) to fin. (4) Disconnect receptacle (10) from plug (9). (5) Remove screws (4) and remove socket (6) and baffle (8).
B.
Install Flashing Beacon (Refer to Figure 201 ). (1) Install socket (6) and baffle (8) on fin tip (7) with screws (4). (2) Connect receptacle (10) to plug (9). (3) Install fin tip (7) on fin. (4) Insert bulb (3) into socket (6), press in slightly, and turn clockwise.
Flasher Removal/Installation A.
Remove Flasher (Refer to Figure 201). (1) Remove lower empennage access plate for access to flasher (13). (2) Disconnect receptacle (15) from plug (14) and receptacle (16) from plug (17). (3) Remove screws (18) and remove flasher (13).
B.
Install Flasher (Refer to Figure 201). (1) Install flasher (13) with screws (18), be sure ground wire (19) is installed. (2) Connect receptacle (15) to plug (14) and receptacle (16) to plug (17). (3) Install lower empennage access plate.
Resistor Removal/Installation A.
Remove Resistor (Refer to Figure 201). (1) Remove lower empennage access plate for access to resistor. (2) Remove terminal (10). (3) Remove screw (12) and remove resistor (11).
B.
Install Resistor (Refer to Figure 201). (1) Install resistor (11) and secure with screw (12). (2) Install terminal (10). (3) Replace lower empennage access plate.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Flashing Beacon Installation Figure 201 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Flashing Beacon Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Flashing Beacon Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL ICE DETECTOR LIGHT - MAINTENANCE PRACTICES 1.
General A.
2.
Ice detector light maintenance practices consist of ice detector light removal/installation.
Ice Detector Light Removal/Installation A.
Remove Ice Detector Light (Refer to Figure 201). (1) Remove three screws (8), lens retainer (7), and lens (6). (2) Press in slightly on bulb (9) and turn counterclockwise to remove bulb from socket (3). (3) Remove old sealer from housing assembly (10) and lens (6), taking care not to scratch lens.
B.
Install Ice Detector Light (Refer to Figure 201). (1) Insert bulb (9) into socket (6), press in slightly, and turn clockwise. (2) Apply Acryl-R-SS2S sealant, or equivalent, around opening in housing assembly (10). Take care to apply sealant sparingly. (3) Press lens (6) into place and secure lens retainer (7) using three screws (8).
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MODEL 208 MAINTENANCE MANUAL
Wing Ice Detector Light Installation Figure 201 (Sheet 1)
33-44-00 © Cessna Aircraft Company
Page 202 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Wing Ice Detector Light Installation Figure 201 (Sheet 2)
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Page 203 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL WINDSHIELD ICE INDICATOR LIGHT - MAINTENANCE PRACTICES 1.
General A.
For removal and installation procedures for the windshield ice indicator light, refer to Chapter 30, Windshield Ice Indicator Light - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL COURTESY LIGHTS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Courtesy lights maintenance practices consist of courtesy light removal/installation and timer removal/ installation.
Courtesy Light Removal/Installation A.
Remove Courtesy Light (Refer to Figure 201). (1) Remove screws (14) securing plate (12) and remove plate. (2) Pull pin (3) from socket (2), and remove entire assembly from wing. (3) Remove screws (13), nuts (5 and 6) and remove plate (12) from shield assembly (7). (4) Press in on bulb (11) and turn counterclockwise to remove bulb from socket (8).
B.
Install Courtesy Light (Refer to Figure 201). (1) Insert bulb (11) into socket (8), press bulb into socket slightly, and turn clockwise. (2) Attach shield assembly (7) to plate (12) using screws (13), nuts (5 and 6), making sure ground wire (4) is connected as shown. (3) Insert pin (3) into socket (2). (4) Secure plate (12) by using screws (14).
Timer Removal/Installation A.
Remove Timer (Refer to Figure 201). (1) Disconnect electrical plug (17). (2) Remove four screws (16) securing timer (15). (3) Remove timer (15) from brackets (18).
B.
Install Timer (Refer to Figure 201). (1) Position timer (15) to brackets (18). (2) Install four screws (16) securing timer (15). (3) Connect electrical plug (17).
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Courtesy Lights Installation Figure 201 (Sheet 1)
33-45-00 © Cessna Aircraft Company
Page 202 Apr 1/1996
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL EXTERNAL LIGHT SWITCHES - MAINTENANCE PRACTICES 1.
General A.
2.
External light switches maintenance practices consist of light switch removal/installation.
External Light Switches Removal/Installation A.
Remove External Light Switches (Refer to Figure 201). (1) Remove nut (14) securing switch (10) to instrument panel, and push switch through instrument panel and retain washers (12) and (13) for reinstallation. (2) Disconnect electrical leads and tag for reinstallation.
B.
Install External Light Switches (Refer to Figure 201). (1) Connect electrical leads and remove tags. (2) Assemble washers (12) and (13), push switch (10) up and through instrument panel, and secure using nut (14).
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MODEL 208 MAINTENANCE MANUAL
External Light Switches Installation Figure 201 (Sheet 1)
33-46-00 © Cessna Aircraft Company
Page 202 Aug 1/1995
34 CHAPTER
NAVIGATION
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT 34-Title 34-List of Effective Pages 34-Record of Temporary Revisions 34-Table of Contents 34-List of Tasks 34-00-00 34-10-00 34-11-00 34-11-00 34-11-00 34-11-00 34-11-00 34-11-01 34-11-01 34-11-01 34-11-01 34-11-02 34-11-02 34-11-10 34-12-00 34-12-01 34-13-00 34-13-00 34-14-00 34-14-00 34-15-00 34-16-00 34-16-10 34-16-10 34-16-20 34-16-30 34-17-00 34-18-00 34-19-00 34-19-01 34-20-00 34-21-00 34-21-00 34-22-00 34-22-00 34-23-00 34-23-00 34-24-00 34-25-00
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DATE
Page 1 Page 1 Page 1 Pages 101-103 Pages 201-211 Page 501 Pages 601-609 Page 1 Pages 101-108 Pages 201-207 Pages 501-503 Pages 201-202 Pages 501-502 Pages 201-203 Pages 201-202 Pages 201-202 Pages 101-105 Pages 201-202 Pages 101-103 Pages 201-202 Pages 201-202 Page 1 Page 1 Page 201 Page 1 Page 1 Page 1 Page 1 Page 1 Page 1 Page 1 Pages 201-203 Pages 601-602 Pages 101-104 Pages 201-202 Pages 1-3 Pages 201-204 Pages 201-202 Pages 201-203
Aug 1/1995 Aug 1/1995 Apr 1/2010 Aug 1/1995 Apr 1/2010 Jun 1/2011 Jun 1/2011 Apr 1/2010 Apr 1/2010 Apr 1/2010 Apr 1/2010 Apr 1/2010 Apr 1/2010 Jun 1/2011 Aug 1/1995 Apr 1/2010 Aug 1/1995 Apr 1/1996 Aug 1/1995 Aug 1/1995 Mar 1/2008 Aug 1/1995 Aug 2/2004 Aug 2/2004 Mar 1/2008 Mar 1/2008 Aug 1/1995 Aug 1/1995 Aug 1/1995 Aug 2/2004 Aug 1/1995 Aug 1/1995 Jun 1/2011 Aug 1/1995 Aug 1/1995 Aug 1/1995 Mar 1/2008 Apr 1/2010 Mar 1/2008
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Oct 15/1999
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Aug 1/1995
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Oct 15/1999
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Oct 15/1999
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Jun 1/2011
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Oct 15/1999
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Apr 1/2010
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Jun 1/2011
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Jun 1/2011
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Jun 1/2011
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Jun 1/2011
34-54-20
Pages 201-205
Mar 1/2008
34-55-00
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Mar 1/2008
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Mar 1/2008
34-56-00
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Mar 1/2008
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Mar 1/2008
34-60-10
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Apr 1/2010
34-60-10
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Apr 1/2010
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CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS NAVIGATION - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-00-00 Page 1 34-00-00 Page 1 34-00-00 Page 1
FLIGHT ENVIRONMENTAL DATA - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-10-00 Page 1 34-10-00 Page 1
PITOT/STATIC SYSTEM - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-00 Page 1 34-11-00 Page 1
PITOT/STATIC SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-00 Page 101 34-11-00 Page 101
PITOT/STATIC SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot System Purging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Static System Purging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot Tube Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Alternate Static Source Valve Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . Static Source Drain Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overspeed Pressure Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Airspeed Warning Horn Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-00 Page 201 34-11-00 Page 201 34-11-00 Page 201 34-11-00 Page 201 34-11-00 Page 209 34-11-00 Page 210 34-11-00 Page 210 34-11-00 Page 210 34-11-00 Page 211 34-11-00 Page 211
PITOT/STATIC SYSTEM - ADJUSTMENT/TEST. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overspeed Pressure Switch Operational Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot System Inspection and Leak Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-00 Page 501 34-11-00 Page 501 34-11-00 Page 501 34-11-00 Page 501 34-11-00 Page 501
PITOT/STATIC SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot/Static System Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot Tube Heaters Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-00 Page 601 34-11-00 Page 601 34-11-00 Page 601 34-11-00 Page 608
LOW AIRSPEED AWARENESS SYSTEM - DESCRIPTION AND OPERATION With TKS Anti-Ice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-01 34-11-01 34-11-01 34-11-01
Page 1 Page 1 Page 1 Page 1
LOW AIRSPEED AWARENESS SYSTEM - TROUBLESHOOTING (With TKS Anti-Ice System) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-01 Page 101 34-11-01 Page 101
LOW AIRSPEED AWARENESS SYSTEM - MAINTENANCE PRACTICES With TKS Anti-Ice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Airspeed Awareness (LAA) Annunciator/Switch Removal/Installation . . . . . . . . Low Airspeed Awareness (LAA) Logic Module Removal/Installation. . . . . . . . . . . . . . Low Airspeed Awareness (LAA) System Airspeed Switch Removal/Installation . . . Airspeed Warning Horn Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-01 Page 201 34-11-01 Page 201 34-11-01 Page 201 34-11-01 Page 206 34-11-01 Page 206 34-11-01 Page 207
LOW AIRSPEED AWARENESS SYSTEM - ADJUSTMENT/TEST With TKS Anti-Ice System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Airspeed Awareness System Functional Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-01 Page 501 34-11-01 Page 501 34-11-01 Page 501 34-11-01 Page 501
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MODEL 208 MAINTENANCE MANUAL LOW AIRSPEED AWARENESS SYSTEM - MAINTENANCE PRACTICES With Pneumatic De-Icing Systems. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Airspeed Awareness Pressure System Switch Removal/Installation . . . . . . . . .
34-11-02 Page 201 34-11-02 Page 201 34-11-02 Page 201
LOW AIRSPEED AWARENESS SYSTEM - ADJUSTMENT/TEST With Pneumatic DeIcing Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . De-Icing Low Airspeed Awareness System Functional Test . . . . . . . . . . . . . . . . . . . . .
34-11-02 Page 501 34-11-02 Page 501 34-11-02 Page 501 34-11-02 Page 501
GARMIN GDC-74A AIR DATA COMPUTER - MAINTENANCE PRACTICES . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GDC-74A Air Data Computer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-11-10 Page 201 34-11-10 Page 201 34-11-10 Page 201
OUTSIDE AIR TEMPERATURE GAGE - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outside Air Temperature Gage Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-12-00 Page 201 34-12-00 Page 201 34-12-00 Page 201
GTP 59 OUTSIDE AIR TEMPERATURE (OAT) PROBE - MAINTENANCE PRACTICES General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outside Air Temperature (OAT) Sensor Removal/Installation . . . . . . . . . . . . . . . . . . . .
34-12-01 Page 201 34-12-01 Page 201 34-12-01 Page 201
VERTICAL SPEED INDICATOR - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-13-00 Page 101 34-13-00 Page 101
VERTICAL SPEED INDICATOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vertical Speed Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vertical Speed Indicator Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-13-00 Page 201 34-13-00 Page 201 34-13-00 Page 201 34-13-00 Page 201
TRUE AIRSPEED INDICATOR - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-14-00 Page 101 34-14-00 Page 101
TRUE AIRSPEED INDICATOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . True Airspeed Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . True Airspeed Indicator Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-14-00 Page 201 34-14-00 Page 201 34-14-00 Page 201 34-14-00 Page 201
ALTIMETER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Altimeter Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-15-00 Page 201 34-15-00 Page 201 34-15-00 Page 201
ENCODING ALTIMETER (TYPE 5035 SERIES) - DESCRIPTION AND OPERATION . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-16-00 Page 34-16-00 Page 34-16-00 Page 34-16-00 Page
ENCODING ALTIMETER (TYPE KEA-130A) - DESCRIPTION AND OPERATION . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-16-10 Page 1 34-16-10 Page 1 34-16-10 Page 1
ENCODING ALTIMETER (TYPE KEA-130A) - MAINTENANCE PRACTICES. . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . KEA-130A Radio Altimeter Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-16-10 Page 201 34-16-10 Page 201 34-16-10 Page 201
ENCODING ALTIMETER (TYPE EA-401A) - DESCRIPTION AND OPERATION . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-16-20 Page 34-16-20 Page 34-16-20 Page 34-16-20 Page
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1 1 1 1
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Page 2 of 6 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE EA-801A) AND ALTITUDE ALERTING WITH PRESELECT - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-16-30 34-16-30 34-16-30 34-16-30
Page 1 Page 1 Page 1 Page 1
ENCODING ALTIMETER (TYPE EA-401A) - DESCRIPTION AND OPERATION . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-17-00 34-17-00 34-17-00 34-17-00
Page 1 Page 1 Page 1 Page 1
ENCODING ALTIMETER (TYPE EA-801A) AND ALTITUDE ALERTING WITH PRESELECT - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-18-00 34-18-00 34-18-00 34-18-00
Page 1 Page 1 Page 1 Page 1
RADAR ALTIMETER (TYPE KRA-10A) - DESCRIPTION AND OPERATION . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-19-00 34-19-00 34-19-00 34-19-00
Page 1 Page 1 Page 1 Page 1
RADAR ALTIMETER (TYPE KRA-405B) - DESCRIPTION AND OPERATION . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-19-01 34-19-01 34-19-01 34-19-01
Page 1 Page 1 Page 1 Page 1
ATTITUDE AND DIRECTION - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-20-00 Page 1 34-20-00 Page 1
MAGNETIC COMPASS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Magnetic Compass Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-21-00 Page 201 34-21-00 Page 201 34-21-00 Page 201
MAGNETIC COMPASS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Magnetic Compass Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-21-00 Page 601 34-21-00 Page 601 34-21-00 Page 601
TURN AND BANK INDICATOR - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-22-00 Page 101 34-22-00 Page 101
TURN AND BANK INDICATOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turn and Bank Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-22-00 Page 201 34-22-00 Page 201 34-22-00 Page 201
HORIZON AND DIRECTIONAL GYROS - DESCRIPTION AND OPERATION . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-23-00 Page 1 34-23-00 Page 1
HORIZON AND DIRECTIONAL GYROS - MAINTENANCE PRACTICES . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Horizon Gyro Removal/Installation (Model 208 Only) . . . . . . . . . . . . . . . . . . Electrical Directional Gyro Removal/Installation (Model 208 Only) . . . . . . . . . . . . . . . Electrical Horizon Gyro Removal/Installation (Model 208B and 208B Passenger) . HSI Indicator Removal/Installation (Model 208B and 208B Passenger) . . . . . . . . . . .
34-23-00 Page 201 34-23-00 Page 201 34-23-00 Page 201 34-23-00 Page 201 34-23-00 Page 201 34-23-00 Page 204
GMU 44 MAGNETOMETER - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Troubleshooting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GMU 44 Magnetometer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-24-00 Page 201 34-24-00 Page 201 34-24-00 Page 201 34-24-00 Page 201
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MODEL 208 MAINTENANCE MANUAL ATTITUDE HEADING REFERENCE SYSTEM (AHRS) - MAINTENANCE PRACTICES General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Troubleshooting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GRS 77 AHRS #1 Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GRS 77 AHRS #2 Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-25-00 Page 201 34-25-00 Page 201 34-25-00 Page 201 34-25-00 Page 201 34-25-00 Page 203
LANDING AIDS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-30-00 Page 1 34-30-00 Page 1
INDEPENDENT POSITION DETERMINING - DESCRIPTION AND OPERATION . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-40-00 Page 1 34-40-00 Page 1
BF GOODRICH WX-1000+/E STORMSCOPE WEATHER MAPPING SYSTEM DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-40-01 Page 1 34-40-01 Page 1 34-40-01 Page 1
BF GOODRICH WX-1000+/E STORMSCOPE WEATHER MAPPING SYSTEM REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stormscope Computer Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WX1000/SKY497 Display Unit Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . Stormscope Antenna Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-40-01 Page 401 34-40-01 Page 401 34-40-01 Page 401 34-40-01 Page 401 34-40-01 Page 401
BF GOODRICH SKYWATCH (SKY497) TRAFFIC ADVISORY SYSTEM DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-40-02 Page 1 34-40-02 Page 1 34-40-02 Page 1
BF GOODRICH SKYWATCH (SKY497) TRAFFIC ADVISORY SYSTEM - REMOVAL/ INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TRC497 Transmitter/Receiver/Computer Removal/Installation. . . . . . . . . . . . . . . . . . . Stormscope/Skywatch Maintenance Switch Removal/Installation . . . . . . . . . . . . . . . . WX-1000/SKY497 Display Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Skywatch NY-164 Antenna Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-40-02 Page 401 34-40-02 Page 401 34-40-02 Page 401 34-40-02 Page 401 34-40-02 Page 401 34-40-02 Page 403
GARMIN GWX-68 WEATHER RADAR SYSTEM - DESCRIPTION AND OPERATION. . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-41-00 Page 34-41-00 Page 34-41-00 Page 34-41-00 Page
GARMIN GWX-68 WEATHER RADAR SYSTEM - MAINTENANCE PRACTICES . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GWX-68 Weather Radar Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-41-00 Page 201 34-41-00 Page 201 34-41-00 Page 201
GARMIN GWX-68 WEATHER RADAR SYSTEM - ADJUSTMENT/TEST. . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Weather Radar Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-41-00 Page 501 34-41-00 Page 501 34-41-00 Page 501
DEPENDENT POSITION DETERMINING - DESCRIPTION AND OPERATION . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-50-00 Page 1 34-50-00 Page 1
DEPENDENT POSITION DETERMINING - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transponder Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-50-00 Page 601 34-50-00 Page 601 34-50-00 Page 601
KN-53 VHF NAVIGATION SYSTEM - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-51-00 Page 34-51-00 Page 34-51-00 Page 34-51-00 Page
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MODEL 208 MAINTENANCE MANUAL KN-53 VHF NAVIGATION SYSTEM - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . KN-53 Transceiver Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . KI-204 Nav Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-51-00 Page 401 34-51-00 Page 401 34-51-00 Page 401 34-51-00 Page 401
KING KT-70 TRANSPONDER - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . KT-70 Transponder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-52-00 Page 401 34-52-00 Page 401 34-52-00 Page 401
BENDIX/KING KT-73 MODE-S TRANSPONDER - MAINTENANCE PRACTICES. . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . KT-73 Transponder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transponder Antenna Without Cargo Pod Removal/Installation. . . . . . . . . . . . . . . . . . Transponder Antenna With Cargo Pod Removal/Installation. . . . . . . . . . . . . . . . . . . . .
34-52-10 Page 201 34-52-10 Page 201 34-52-10 Page 201 34-52-10 Page 201 34-52-10 Page 204
BENDIX/KING KT-73 MODE-S TRANSPONDER - ADJUSTMENT/TEST . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . KT-73 Self Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
34-52-10 Page 501 34-52-10 Page 501 34-52-10 Page 501
GARMIN GTX 327 TRANSPONDER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin Transponder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transponder Antenna without Cargo Pod Removal/Installation . . . . . . . . . . . . . . . . . . Transponder Antenna with Cargo Pod Removal/Installation . . . . . . . . . . . . . . . . . . . . .
34-54-00 Page 201 34-54-00 Page 201 34-54-00 Page 201 34-54-00 Page 201 34-54-00 Page 204
GARMIN GTX 330 TRANSPONDER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin Transponder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transponder Antenna without Cargo Pod Removal/Installation . . . . . . . . . . . . . . . . . . Transponder Antenna with Cargo Pod Removal/Installation . . . . . . . . . . . . . . . . . . . . . Transponder Antenna (Fairing Installation) Removal/Installation . . . . . . . . . . . . . . . . .
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GTX-33 TRANSPONDER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GTX-33 Transponder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GTX-33 Diversity Transponder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transponder Antenna without Cargo Pod Removal/Installation . . . . . . . . . . . . . . . . . . Transponder Antenna with Cargo Pod Removal/Installation . . . . . . . . . . . . . . . . . . . . .
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GDL-69A FIS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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GDL-69A FIS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GDL-69A XM Weather Data Link Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . .
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GARMIN INTEGRATED AVIONICS UNIT (GIA 63) - MAINTENANCE PRACTICES . . . . GIA 63 Integrated Avionics Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . .
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GARMIN DISPLAY UNIT (GDU) - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Troubleshooting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin Display Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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GARMIN SYNTHETIC VISION TECHNOLOGY SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL GARMIN SYNTHETIC VISION TECHNOLOGY SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Garmin Synthetic Vision Technology System Software Configuration. . . . . . . . . . . . . Garmin Synthetic Vision Technology System - Operational Check . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 34-11-00-720
Pitot/Static System Functional Check
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34-11-00-710
Pitot Tube Heaters Operational Check
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34-21-00-720
Magnetic Compass Functional Check
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34-50-00-720
Transponder Functional Check
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MODEL 208 MAINTENANCE MANUAL NAVIGATION - GENERAL 1.
Scope A.
This chapter describes the navigation systems, units, and components which provide airplane navigational information. Included are pitot/static, gyros, compasses, VOR, and indicators. For Sperry 4008 Autopilot, Sperry 4008 IFCS and King Flight Control System information, refer to Chapter 22. NOTE:
2.
This chapter does not deal with specific instrument repairs. Federal Aviation Regulations require malfunctioning instruments be sent to an approved instrument overhaul and repair station or returned to the manufacturer for servicing. These subjects will be covered in this chapter as individual sections:
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the table of Contents will further assist in locating a particular subject. A brief definition of the sections incorporated in this chapter is as follows: (1) The Flight Environmental Data Section describes systems that sense environment conditions, and use data to influence navigation of the airplane. This includes systems that depend on pitot and static information. (2) The Attitude and Direction Section describes systems that use magnetic gyroscopic and inertia forces. This includes items like gyros, compasses, magnetic heading, and turn and bank. (3) The Landing Aids Section describes systems that provide guidance during approach, landing, and taxiing. This includes items such as Iocalizer, glide slope, and marker beacon. (4) The Independent Position Determining Section describes systems that provide information to determine position, and are mainly independent of ground installation. The weather radar system is described in this section. (5) The Dependent Position Determining Section describes systems that provide information to determine position, and are mainly dependent on ground installation. This includes systems like VOR, DIVIE, ADF, RNAV, and transponders.
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MODEL 208 MAINTENANCE MANUAL FLIGHT ENVIRONMENTAL DATA - DESCRIPTION AND OPERATION 1.
Description A.
This section covers that portion of the system which senses environmental conditions and uses the data to influence navigation. It includes items such as pitot/static, air temperature, vertical speed indicator (rate-of-climb), airspeed indicator, altimeter, etc.
B.
The airplane may utilize a second, optional, right pitot/static system which is independent from the left system. In addition, an optional right flight panel is offered, which incorporates a second airspeed indicator, altimeter, vertical speed indicator (rate-of-climb), and turn-and-bank indicator. Even though the mounting locations on the right panel are different from those on the left panel, the removal, installation, troubleshooting, and maintenance practices and procedures will be the same for both left and right installations. Only the maintenance procedures for the left instruments will be covered in the chapter, since the maintenance procedures for the right instruments would be the same. The right panel may also incorporate an optional electric gyro system, which is independent of the left vacuum driven gyros. Maintenance procedures for the electrical gyros are covered in this chapter. Refer to Chapter 37 for servicing information on vacuum driven gyros.
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MODEL 208 MAINTENANCE MANUAL PITOT/STATIC SYSTEM - DESCRIPTION AND OPERATION 1.
Description A.
The pitot system utilizes a pitot assembly and tube assembly, located in leading edge of left wing (right wing for right system), just inboard of landing lights (WS 185.00). Ram air passes through the pitot assembly, is then routed through lines to a pressure switch (located forward of instrument panel on instrument panel support), and then to airspeed indicator. The pitot assembly incorporates two electrical heating elements to prevent ice from obstructing passage of ram air through pitot assembly. The heating elements, powered by the electrical system, are controlled by a switch located at the lower left corner of instrument panel. One switch controls both left and optional right (if installed) pitot tube heating elements. Refer to Pitot/Static System - Maintenance Practices, Figure 201.
B.
The static system utilizes a static port, a static source drain valve, an alternate static source selector valve assembly, and necessary plumbing to operate airspeed indicator; vertical speed indicator, and altimeter. The static port is located in pitot assembly. A line runs from static port, through tube assembly along leading edge of wing to WS 33.50, and down forward door post at FS 154.00 to static source drain valve, which is located below and to the left of alternate static source selector valve. The static source drain valve is located at the lowest point in system and is utilized for draining any moisture in system. Refer to placard adjacent to drain valve for drain valve operation instructions. Refer to Chapter 5, Time Limits/Maintenance Checks for time limit intervals for draining moisture from system.
C.
The alternate static source valve is located on left lower instrument panel, allowing for an alternate source of static air pressure to be obtained from inside cabin. The right pitot/static system does not incorporate an alternate static source selector valve. The static ports in pitot tube are the only source of static air for the right system. NOTE:
The alternate static source is to be used only in emergency situations, when normal system is inoperative. When alternate static source valve is used, instrument readings may vary from normal readings due to static air source being obtained from inside cabin. Refer to Pilot's Operating Handbook for flight operation using alternate static source.
D.
The airplane has airspeed warning horns installed. The horns are installed behind the headliner above the pilot. Airplanes that have the G1000 system installed, the G1000 the activates airspeed warning horns. On airplanes that do not have the G1000 system installed, the overspeed pressure switch activates airspeed warning horns. The horns will operate when the airspeed is more than 175 KlAS (VMO).
E.
Airplanes that have a pneumatic or TKS anti-ice system installed also have a Low Speed Awareness (LAA) system installed as follows: (1) The 97.5 KIAS LAA system is installed in airplanes that have the optional TKS system installed. The LAA system operates from inputs given from the pitot-static system. The LAA system is installed with airplanes that have the G1000 system installed and airplanes that do not have the G1000 system installed. The warning horns operate when airspeed is less than 97.5 KIAS, +2 or -2 KIAS. A warning signal can also be heard in the pilot's headset. For more data on the 97.5 KIAS LAA system, refer to Low Airspeed Awareness System - Description and Operation (With TKS System). (2) The 110 KIAS LAA system is installed on airplanes that have pneumatic anti-ice system installed. The LAA system operates from inputs given from the pitot-static system. The warning horns operate when airspeed is less than 110 KIAS, +5 or -5 KIAS. A warning signal can also be heard in the pilot's headset. For more data on the 110 KIAS LAA system, refer to Low Airspeed Awareness System - Maintenance Practices (With Pneumatic System).
F.
The Altair Engine Trend Monitoring (ETM) System is attached to the pitot-static system behind the copilot's instrument panel at FS 117.55. It includes a pitot transducer and a static transducer. The transducers supply analysis to the ETM. For more information on the ETM system, refer to Altair ADAS+ Engine Trend Monitoring System - Description and Operation or Altair ADASd Engine Trend Monitoring System - Description and Operation.
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MODEL 208 MAINTENANCE MANUAL PITOT/STATIC SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 1)
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL PITOT/STATIC SYSTEM - MAINTENANCE PRACTICES 1.
General A.
Correct maintenance of the pitot-static system is essential for correct operation of the altimeter, vertical speed indicator, and airspeed indicator. Leaks, moisture, and obstructions in the pitot system will result in false airspeed indications. Static system malfunctions will affect indications of all three instruments. Cleanliness and correct installation are the principal rules for maintenance of the pitot-static system. The pitot tube and static port MUST be kept clean and clear of obstructions. When you replace pitot-static system components, use the minimum amount of antiseize compound on the male threads of both metal and plastic connections. Always avoid excess compound which might be able to enter the pitot-static system lines. Tighten connections firmly, but be very careful not to over tighten and distort fittings.
CAUTION: Except for the use of the system drains and alternate static source pressure valves, make sure to do a leak test after the static pressure system is opened or closed. Refer to Pitot System Inspection and Leak Test. CAUTION: If an autopilot or integrated flight control system is installed, make sure that any portion of the these systems that are interconnected with the static system are disconnected before you purge the static system. If a 400B autopilot or 400B integrated flight control system is installed, refer to the appropriate autopilot figure in the 208 Avionic Installations Service/Parts Manual for the location of the components and static line attachment. 2.
Pitot System Purging NOTE: A.
Moisture may collect at various points in the pitot system and can produce a partial obstruction.
Purge the Pitot System (Refer to Figure 201). (1) Identify and disconnect the pitot line from the pressure switch.
CAUTION: Never blow through the pitot or static lines toward the instruments or pressure switch. (2) (3) (4) (5) (6) 3.
Loosen the clamp and disconnect the pressure switch line from the true airspeed indicator, and put a cap on the airspeed indicator. Use clean, dry, low-pressure air, and blow out the pitot system from the pressure switch end of pitot line toward the pitot assembly. Connect the pitot line to the pressure switch. Connect the pressure switch line to the true airspeed indicator and tighten the clamp. Do the leak test. Refer to Pitot System Inspection and Leakage Test.
Static System Purging NOTE:
A.
Static pressure lines must also be kept clear and connections tight. A drain valve is used to drain moisture from the system. The static source drain valve is found at the low point in the system, which is at the lower left corner of the instrument panel, directly below the alternate static source valve. If the hoses are not correctly installed, moisture can also collect at other points in the system and cause a partial obstruction.
Purge the Static System. (1) Loosen the lower clamp and disconnect the alternate static source line at the lower tee fitting forward of the true airspeed indicator and vertical speed indicator. (2) Put a cap on the tee fitting.
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Pitot-Static System Installation Figure 201 (Sheet 1)
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Pitot-Static System Installation Figure 201 (Sheet 2)
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Pitot-Static System Installation Figure 201 (Sheet 3)
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Pitot-Static System Installation Figure 201 (Sheet 4)
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Pitot-Static System Installation Figure 201 (Sheet 5)
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Pitot-Static System Installation Figure 201 (Sheet 6)
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Pitot-Static System Installation Figure 201 (Sheet 7)
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MODEL 208 MAINTENANCE MANUAL (3)
Make sure the alternate static source valve is closed. NOTE:
(4) (5) (6)
The vent hole in the main body of the alternative static source valve must be clear and free from any blockages.
Disconnect the static source line from the tee fitting found above the static source drain valve. (a) Put a cap on the top of the tee fitting. Disconnect the static line at the pressure switch. Close the static source drain valve.
CAUTION: Never blow through the pitot or static lines toward the instruments or pressure switch. (7) (8) (9) (10) (11) (12) (13) (14) 4.
Use clean, dry, low pressure air, and blow out the static lines from the indicator end of the alternate static source line toward the static source drain valve. Put a cap on the pressure switch end of the static line. Open the static source drain valve and blow clean, dry, low pressure air from the indicator end of the alternate static source line toward the static source drain valve. Remove the cap from the lower tee fitting. Connect the alternate static source line to the lower tee fitting and tighten the lower clamp. Connect the static line to the pressure switch. Use clean, dry, low pressure air, and blow from the static source drain valve end of static source line toward the static port in the pitot assembly. Connect the static source line to the tee fitting found above the static source drain valve.
Pitot Assembly Removal/Installation A.
Remove the Pitot Assembly (Refer to Figure 201).
WARNING: Before you try to do a removal or installation of the pitot assembly, make sure the airplane battery switch is in the OFF position. (1) (2) (3)
B.
Remove the screws that attach the access plate (503BB left, 603BB right) and remove from the lower surface of the wing. (a) Identify and disconnect the pitot line and static source line found in the leading edge. Remove the four screws and washers that attach the pitot assembly to the pitot tube assembly. Pull the pitot assembly forward, identify and disconnect the components that follow: (a) Pitot heat electrical connector (b) Pitot line (c) Static line.
Install the Pitot Assembly (Refer to Figure 201). (1) Put the pitot assembly in position forward of the pitot tube assembly and connect the components that follow: (a) Pitot heat electrical connector (b) Pitot line (c) Static line. (2) Put the pitot assembly in its position on the pitot tube assembly. (3) Install the four attach screws and washers that attach the pitot assembly to the pitot tube assembly. (4) Connect the pitot line and static source line found in leading edge of wing. (5) Do the operational test. Refer to Chapter 34, Pitot System Inspection and Leak Test. (6) Carefully, make sure that the pitot tube becomes warm when the PITOT-STATIC HEAT switch is in the ON position. (7) Put the access plate (503BB left, 603BB right) in its position on the lower surface of the wing and attach with the screws.
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5.
Pitot Tube Assembly Removal/Installation A.
Remove the Pitot Tube Assembly (Refer to Figure 201). (1) Remove the pitot assembly. Refer to Pitot Assembly Removal/Installation. (2) Remove the landing light lens attach screws and the landing light lens. (3) Remove the pitot tube assembly attach screws and the pitot tube assembly. (4) Remove the electrical line, pitot line, and static source line from the pitot tube assembly. (5) Do an inspection to make sure the pitot tube assembly is serviceable.
B.
Install the Pitot Tube Assembly (Refer to Figure 201). (1) Install the serviceable electrical line, pitot line, and static source line in the pitot tube assembly. NOTE: (2) (3) (4)
6.
Install the pitot tube assembly into the leading edge of the wing with the pitot tube assembly attach screws. Install the landing light lens with the landing light lens attach screws. Install the pitot assembly. Refer to Pitot Assembly Removal/Installation.
Alternate Static Source Valve Assembly Removal/Installation A.
Remove the Alternate Static Source Valve Assembly (Refer to Figure 201). (1) Identify, disconnect, and put a cap on the alternate static source line from the alternate static source valve. (2) Disconnect the static source drain valve line from the alternate static source valve. (3) Remove the setscrew from the knob. (4) Pull the knob and remove it from the alternate static source valve. (5) Remove the screws and nuts that attach the alternate static source valve to the instrument panel. (6) Remove the alternate static source valve from the instrument panel.
B.
Install the Alternate Static Source Valve Assembly (Refer to Figure 201). (1) Install the alternate static source valve to the instrument panel with the screws and nuts. NOTE: (2) (3) (4)
7.
A replacement pitot tube must be installed if the removed pitot tube is not serviceable.
The vent hole in the main body of the alternative static source valve must be clear and free from any blockages.
Connect the static source drain valve line to the alternate static source valve. Identify, remove the cap, and connect the alternate static source line to the alternate static source valve. Push the knob onto the alternate static source valve and install the setscrew.
Static Source Drain Valve Removal/Installation A.
Remove the Static Source Drain Valve (Refer to Figure 201). (1) Remove the two clamps from the tee fitting. (a) Remove the screws that attach the clamps to the substrate forward wall panel for the left installation. (b) Remove the screws and nuts that attach the clamps to the avionic shelf sidewall for the right installation. (2) Pull the fitting out of the clamps. (3) Hold the fitting firmly to unscrew and remove the static source drain valve.
B.
Install the Static Source Drain Valve (Refer to Figure 201). (1) Hold the fitting firmly and screw the static source drain valve into the fitting. (2) Attach the fitting and the static source drain valve. (a) Put the fitting in the clamps, with the static source drain valve installed, and attach it to the substrate forward wall panel with one screw for the left installation. (b) Put the fitting in the clamps, with the static source drain valve installed, and attach it to the avionic shelf sidewall with one screw and nut for the right installation.
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8.
9.
Overspeed Pressure Switch Removal/Installation A.
Remove the Overspeed Pressure Switch (Refer to Figure 201). (1) Disconnect the electrical connector. (2) Identify, disconnect, and put a cap on the pitot line and the static line. (3) Remove two nuts and washers that attach the pressure switch to mounting bracket. (4) Remove the pressure switch.
B.
Install the Overspeed Pressure Switch (Refer to Figure 201). (1) Put the pressure switch in the mounting bracket with the ports marked P and S correctly oriented, and install it with two washers and nuts. (2) Identify the pitot line and static line, remove the caps, and connect the pitot line and static line to the ports marked P and S, respectively. (3) Connect the electrical connector to the pressure switch. (4) Do the operational test. Refer to Overspeed Pressure Switch Operational Test.
Airspeed Warning Horn Removal/Installation A.
Remove the Airspeed Warning Horn (Refer to Figure 201). (1) Remove four screws that attach the cover assembly to the trim assembly, found in the headliner above the pilot's seat. (2) Remove the cover assembly. (3) Remove the screw from the hinged mounted plate assembly. (4) Pull the aft end of the hinged mounting plate assembly down through the trim assembly. (5) Identify and disconnect the electrical connector. (6) Remove two nuts, washers, and screws from the airspeed warning horn. (7) Remove the airspeed warning horn.
B.
Install the Airspeed Warning Horn (Refer to Figure 201). (1) Put the airspeed warning horn against the hinged mounting plate assembly. (a) Install two screws, washers, and nuts to attach the airspeed warning horn to the hinged mounting plate assembly. (2) Identify and connect the electrical connector. (3) Push the hinged mounting plate assembly up through the trim assembly. (4) Install the attach screw for the hinged mounting plate assembly. (5) Put the cover assembly against the trim assembly and install the four attach screws.
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MODEL 208 MAINTENANCE MANUAL PITOT/STATIC SYSTEM - ADJUSTMENT/TEST 1.
General A.
2.
This section has procedures to do a test of the pitot and static systems.
Equipment NOTE:
Equivalent equipment may be substituted for that listed below.
NAME
NUMBER
MANUFACTURER
USE
Air Data Tester
101-00184
Barfield 4101 NW 29th Street Miami, FL 33142-5617
To supply pressure or vacuum for the pitot and static system tests.
Pitot Static Test Adaptor
PS4769
Nav-Aids Ltd. 2955 Diab Street Montreal, Quebec H4S 1M1
To attach portable air data tester to pitot system.
3.
Overspeed Pressure Switch Operational Test A.
4.
Do an Overspeed Pressure Switch Operational Test. (1) Engage the AIR SPEED WARN circuit breaker. (2) Put a cover on the drain hole on the pitot tube. (3) Install a very low pressure air source on the end of the pitot tube. (4) Slowly increase the pressure and make sure that the airspeed warning horn gives off an audible sound at 178 KIAS, +3 or -3 KIAS. (5) Slowly decrease the pressure and make sure that the airspeed warning horn sound stops at 178 KIAS, +3 or -3 KIAS. (6) Return the pitot/static system to field elevation and disconnect the pitot/static tester. (7) Remove the cover from the drain hole on the pitot tube.
Pitot System Inspection and Leak Test A.
Do a Leak Test of the Pitot System. (1) Put a piece of rubber or plastic tubing over the pitot tube. (2) Close the opposite end of the rubber or plastic tubing and slowly roll the tubing up until the airspeed indicator shows within the cruise range. (3) Close the tubing and, after a few minutes, examine the airspeed indicator. (a) If there is a leak present, the system pressure and airspeed indication will be decreased. (b) Examine all the connections and tighten them as necessary. (4) Slowly unroll the tubing, gradually decrease the pressure.
CAUTION: Make sure that the pressure is gradually decreased to prevent damage to the instrument. (5)
Remove the tubing from the pitot tube.
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MODEL 208 MAINTENANCE MANUAL PITOT/STATIC SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the pitot/static system in a serviceable condition.
Task 34-11-00-720 2.
Pitot/Static System Functional Check A.
General (1) This task provides procedures to perform a functional check of the pitot/static system. Airplanes equipped with Garmin G1000 proceed to Paragraph F. For an alternate method of compliance, (without air data test set) proceed to Paragraph G.
B.
Special Tools NOTE: (1) (2) (3) (4)
C.
Equivalent tools and equipment can be used.
Air Data Test Set - (LAVERSAB Model 65000); (Barfield 101-00184) Pitot/Static Test Adaptor - ( Nav-Aids Ltd. PS4769) External Electrical Power Unit, 28 VDC. Air Bulb (Optional)
Safety Precautions and Preparations
CAUTION: Do not disconnect pitot-static tubes, hoses, or test equipment while test pressures are applied. Connections that are not correct can cause damage to the instruments. Make sure that all of the plumbing connections have been installed correctly. CAUTION: Make sure that the static system pressure is not more than pitot system pressure, or instrument damage can occur. Do not apply pitot pressure to the static system or a vacuum to the pitot system. Do not do a leak test of the pitot and static system with soap and water or other liquids. CAUTION: Do not apply anti-ice power to pitot probes or static ports when adapters are installed. D.
Access (1) None
E.
Do a Functional Check of the Pitot/Static System. (1) Examine the pitot tube(s) and the static port(s) for condition, corrosion, and obstructions. (2) Examine the mast(s) for condition, bends, and damage. (a) Make sure that the sealant at the mast-to-wing joint is in good condition. (3) Examine all pitot/static system plumbing for condition and security. (a) Make sure that there are no low spots in the tubing that would cause water to collect. (4) Make sure that there is no moisture and/or restrictions caught in the static system. (5) Make sure that there are no alterations or deformations of the airframe surface that would affect the relationship between the air pressure in the static pressure system and the true ambient static air pressure for any flight configuration. (6) Examine the drain valve(s) for condition, water in static system, and security of installation.
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CAUTION: Do not open the autopilot drain plugs unless moisture is found in the left static system drain valve. If the autopilot static drain plug is removed to drain moisture, you must do a static system check after you install the plug. (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17) (18)
Examine the drain valve(s) tubing connections for condition and security. Connect external electrical power to the airplane. Set the External Power Switch to BUS. Set the Battery Switch to ON. Set the Avionics Switches 1 and 2 to ON. Do a self-test on the air data test set and record the leak rate for future use. Connect the air data test set to the left pilot's pitot/static probe in accordance with the manufacturer's instructions. Make sure that the altimeter(s) pressure display reads 29.92 IN (1013 HPA). Use the air data test set, to increase the pressure to the left pitot systems to generate airspeeds of 100, 125, 150, and 175 knots. (a) Make sure that the airspeed displayed is the same as the input +/-5 knots. Slowly increase the pressure and make sure that the airspeed warning horn gives an audible sound at 178 KIAS, +3 or -3 KIAS. Slowly decrease the pressure and make sure that the airspeed warning horn sound stops at 178 KIAS, +3 or -3 KIAS. With the test set input set at 175 knots, do the leak check on the system. (a) After 1 minute the maximum allowable loss must not be more than 5 knots. NOTE:
The airplane's leak rate is determined by subtracting the recorded test set's internal leakage.
(19) Set the altitude on the air data test set to 5,000, 7,500, 10,000, 12,500, and 15,000 feet. (a) Make sure that the altitude displayed on the altimeter(s) are the same +/-250 feet. (20) Slowly return the pitot/static system to the field elevation. (21) If installed, do the test again for the right copilot's pitot/static system. (22) Remove the air data test set in accordance with the manufacturer's instructions. (23) Set the PITOT-STATIC HEAT switch to ON for 30 seconds, then OFF. NOTE:
The pitot tubes have two heating elements, one in the front of and one behind the static port compensating ring. Make sure that both elements are operating.
WARNING: Use extreme caution when you touch the pitot tube surface with you bare hands. The pitot tube will cause severe burns to skin if it is left on too long. (24) Carefully make sure that the pitot tube becomes warm when the PITOT-STATIC HEAT switch is at the ON position. (25) Set the Avionics Switches 1 and 2 to OFF. (26) Set the Battery Switch to OFF. (27) Set the External Power Switch to OFF. (28) Remove the external electrical power from the airplane. (29) Do the Restore Access. F.
Do a Functional Check of the Pitot/Static Systems (Garmin G1000 Equipped). (1) Examine the pitot tube(s) and the static port(s) for condition, corrosion, and obstructions. (2) Examine the mast(s) for condition, bends, and damage. (a) Make sure that the sealant at the mast-to-wing joint is in good condition. (3) Examine all pitot/static system plumbing for condition and security. (a) Make sure there are no low spots in the tubing that would cause water to collect. (4) Make sure that there is no moisture and/or restrictions caught in the static system.
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MODEL 208 MAINTENANCE MANUAL (5) (6)
Make sure that there are no alterations or deformations of the airframe surface that would affect the relationship between the air pressure in the static pressure system and the true ambient static air pressure for any flight configuration. Examine the drain valve(s) for condition, water in static system, and security of installation.
CAUTION: Do not open the autopilot drain plugs unless moisture is found in the left static system drain valve. If the autopilot static drain plug is removed to drain moisture, you must do a static system check after you install the plug. (7) (8) (9) (10) (11) (12) (13)
Examine the drain valve(s) tubing connections for condition and security. Connect external electrical power to the airplane. Set the External Power Switch to BUS. Set the Battery Switch to ON. Set the Avionics Switches 1 and 2 to ON. Do a self-test on the air data test set and record the leak rate for future use. Connect the air data test set to the left pilot's pitot/static probe in accordance with the manufacturer's instructions.
CAUTION: Make sure that the static system pressure is not more than pitot system pressure, or instrument damage can occur. Do not apply pitot pressure to the static system or a vacuum to the pitot system. Do not do a leak test of the pitot and static system with soap and water or other liquids. CAUTION: Do not apply power to pitot probes when the test adapters are installed. NOTE:
The pressure sensors inside of the GDC 74 are internally heated and must stabilize before the test. The G1000 / Air Data System must be powered on for a minimum of 15 minutes before you take calibration readings.
(14) Push in on the BARO correction knob on PFD1 and make sure that the pressure display reads 29.92 IN (1013 HPA). (a) Set the barometric setting on the standby altimeter to 29.92 IN. (15) Use the air data test set to increase the pressure to the left pitot systems to generate the airspeeds (A/S) shown in Table 601. Record the airspeed displayed on the PFD and the standby airspeed indicator. Make sure that it is the same as the input +/- 5 knots. Table 601. Airspeed Display Check Input Airspeed
Airspeed on PFD1
Airspeed on Standby Airspeed Indicator
100 125 150 175 (16) Slowly increase the pressure and make sure that the airspeed warning horn gives an audible sound at 178 KIAS, +3 or -3 KIAS. (17) Slowly decrease the pressure and make sure that the airspeed warning horn sound stops at 178 KIAS, +3 or -3 KIAS.
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MODEL 208 MAINTENANCE MANUAL (18) With the test set input set at 175 knots do the leak check on the system. (a) After 1 minute the maximum allowable loss must not be more than 5 knots. NOTE:
The aircraft's leak rate is determined by subtracting the recorded test set's internal leakage.
(19) Set the altitude on the test set to the values shown in Table 602 and make sure that the altitude shown on PFD1 and the standby altimeter are the same +/- 250 feet. Table 602. Altitude Display Check Test Set Altitude
PFD1 Altitude
Standby Altimeter
5,000 7,500 10,000 12,500 15,000 (20) Do the test again for the right pitot/static system using PFD2. Record the data in Table 603 and Table 604. NOTE:
The standby altimeter and airspeed indicators are not connected to the right system.
Table 603. Airspeed Display Check Input Airspeed
Airspeed on PFD2
100 125 150 175 Overspeed Warning Speed Table 604. Altitude Display Check Test Set Altitude
PFD2 Displayed Altitude
5,000 7,500 10,000 12,500 15,000 (21) Slowly return the pitot/static system to the field elevation. (22) Remove the air data test set in accordance with the manufacturer's instructions. (23) Set the PITOT-STATIC HEAT switch to ON for 30 seconds, then OFF. NOTE:
The pitot tubes have two heating elements, one in the front of and one behind the static port compensating ring. Make sure that both elements are operating.
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WARNING: Use extreme caution when you touch the pitot tube surface with you bare hands. The pitot tube will cause severe burns to skin if it is left on too long. (24) Carefully make sure that the pitot tube becomes warm when the PITOT-STATIC HEAT switch is at the ON position. (25) Set the Avionics Switches 1 and 2 to OFF. (26) Set the Battery Switch to OFF. (27) Set the External Power Switch to OFF. (28) Remove the external electrical power from the airplane. (29) Do the Restore Access. G.
Do a Functional Check of the Pitot/Static Systems (Alternate Method). Refer to Figure 601. (1) Examine the pitot tube(s) and the static port(s) for condition, corrosion, and obstructions. (2) Examine the mast(s) for condition, bends, and damage. (a) Make sure that the sealant at the mast-to-wing joint is in good condition. (3) Examine all pitot/static system plumbing for condition and security. (a) Make sure there are no low spots in the tubing that would cause water to collect. (4) Make sure that there is no moisture and/or restrictions caught in the static system. (5) Make sure that there are no alterations or deformations of the airframe surface that would affect the relationship between the air pressure in the static pressure system and the true ambient static air pressure for any flight configuration. (6) Examine the drain valve(s) for condition, water in static system, and security of installation.
CAUTION: Do not open the autopilot drain plugs unless moisture is found in the left static system drain valve. If the autopilot static drain plug is removed to drain moisture, you must do a static system check after you install the plug. (7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17)
Examine the drain valve(s) tubing connections for condition and security. Connect external electrical power to the airplane. Set the External Power Switch to BUS. Set the Battery Switch to ON. Set the Avionics Switches 1 and 2 to ON. Connect a piece of rubber or plastic tubing over the left pilot's pitot/static probe. Close the opposite end of the rubber or plastic tubing and slowly roll the tubing up to generate airspeeds of 100, 125, 150, and 175 knots. (a) Make sure that the airspeed displayed is the same as the input +/-5 knots. Slowly increase the pressure and make sure that the airspeed warning horn gives an audible sound at 178 KIAS, +3 or -3 KIAS. Slowly decrease the pressure and make sure that the airspeed warning horn sound stops at 178 KIAS, +3 or -3 KIAS. With the airspeed set at 175 knots do the leak check on the system. (a) After 1 minute the maximum allowable loss must not be more than 5 knots. Slowly unroll the tubing and gradually decrease the pressure.
CAUTION: Make sure that the pressure is gradually decreased to prevent damage to the instrument. (18) (19) (20) (21)
Remove the tubing from the pitot tube. Close the static pressure alternate source valve. Set the altimeter to 29.92. Apply a source of suction to the remaining static pressure source opening. Refer to Figure 601 for one method.
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Static System Test Equipment Fabrication Figure 601 (Sheet 1)
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CAUTION: When you apply or release the suction, you must stay less than the range of the vertical speed or airspeed indicator. (22) Slowly apply suction until the altimeter shows a 1000-foot increase in altitude. (a) Air Bulb Method Squeeze the air bulb to remove as much air as possible. 1 2 Hold the suction hose firmly against the static pressure source opening. 3 Slowly release the air bulb to get the necessary suction. 4 Tightly close the hose to trap the suction in the system. (b) Close the suction source to keep the system closed for one minute. (c) Make sure that the decrease in altitude is not more than 100 feet as shown on the altimeter. (d) If the leakage rate is within tolerance, slowly release the suction source. (e) If the leakage rate is more than the maximum permitted rate, tighten all the connections and do a leakage test. (f) If the leakage rate is still more than the maximum permitted rate, do the steps that follow: Disconnect the static pressure lines from the airspeed indicator and the vertical speed 1 indicator. Use the correct fittings and connect the pressure lines together so the altimeter is the 2 only instrument connected to the static pressure system. 3 Do a leak test to find whether the static pressure system or the bypassed instruments are the cause of the leakage. If the leakage is the result of an instrument failure, the instrument must be a repaired by an approved repair station, or it must be replaced. If the leakage is the result of the static pressure system, find the leakage as b follows. Apply a source of positive pressure to the static source opening. Refer to Figure c 601 for one method to get a positive pressure.
CAUTION: Make sure that you do not apply a positive pressure when the airspeed indicator or the vertical speed indicator is connected to the static pressure system. d
Slowly apply a positive pressure until the altimeter shows a 500-foot decrease in altitude. NOTE:
e f g
Put leak detector solution or a mixture of mild soap and water on the line connections and the static source flange. Apply a positive pressure to keep the altimeter indication and look for bubbles which show the leaks. Slowly release the pressure. NOTE:
h i j k
For the air bulb method you must hold the pressure hose firmly against the static pressure source opening. To apply the desired pressure to the static system you slowly squeeze the air bulb. This will replace any air that is released through the leaks.
For the air bulb method you must slowly open the pressure bleed-off screw.
Remove the test equipment. Tighten all the connections that leak. Repair or replace the defective parts. Do the leak test again.
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MODEL 208 MAINTENANCE MANUAL (g)
After the leak test, release the suction slowly. Intermittently let a small amount of air to go into the static system. Air Bulb Method 1 Tilt the end of the suction hose away from the opening, then immediately tilt it a back against the opening. Continue to release a small amount of air intermittently until all of the suction is b released. (23) If installed, do the test again for the right copilot's pitot/static system. (24) Remove the test equipment. (25) Set the PITOT-STATIC HEAT switch to ON for 30 seconds, then OFF. NOTE:
The pitot tubes have two heating elements, one in the front of and one behind the static port compensating ring. Make sure that both elements are operating.
WARNING: Use extreme caution when you touch the pitot tube surface with you bare hands. The pitot tube will cause severe burns to skin if it is left on too long. (26) Carefully, make sure that the pitot/static tube(s) became warm when the PITOT-STATIC HEAT switch was placed in the ON position. (27) Set the Avionics Switches 1 and 2 to OFF. (28) Set the Battery Switch to OFF. (29) Set the External Power Switch to OFF. (30) Remove the external electrical power from the airplane. H.
Restore Access (1) None End of task Task 34-11-00-710 3.
Pitot Tube Heaters Operational Check A.
General (1) This task gives the information needed to operational check of the pitot tube heaters.
B.
Special Tools (1) External Electrical Power Unit
C.
Access (1) None
D.
Do the Pitot Tube Heater Operational Check. NOTE: (1) (2) (3) (4)
The pitot tubes have two heating elements, one in the front and one behind the static port compensating ring. Make sure that both elements operate.
Make sure that the covers are not installed on the pitot tubes. Connect the external electrical power unit to the airplane. Set the BATT switch to the ON position. Set the PITOT-STATIC HEAT switch to the ON position for 30 seconds, then to the OFF position.
WARNING: Use extreme caution when you touch the pitot tube surface with you bare hands. The pitot tube will cause severe burns to skin if it is left on too long. (5)
Carefully make sure that the pitot tube becomes warm when the PITOT-STATIC HEAT switch is at the ON position.
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Set the BATT switch to the OFF position. Remove the electrical power from the airplane.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL LOW AIRSPEED AWARENESS SYSTEM - DESCRIPTION AND OPERATION With TKS Anti-Ice System 1.
2.
General A.
The Low Speed Awareness (LAA) system (97.5 KIAS) is installed on airplanes that also have the TKS system installed. The LAA system is installed on airplanes that have the G1000 system installed and airplanes that do not have the G1000 system installed. The LAA system is installed on Airplanes 20800518 and On, Airplanes 208B2067 and On, and Airplanes 208B0001 Thru 208B2066 that incorporate CAB08-7. For airplanes with G1000, refer to Low Airspeed Awareness - Maintenance Practices, Figure 201, Figure 202 . For airplanes without G1000 refer to Low Airspeed Awareness Maintenance Practices, Figure 203, and Figure 204 as applicable.
B.
The LAA system gives a warning to the pilot if the airspeed goes below 97.5 KIAS, +2 or - 2 knots. The LAA system has an airspeed switch, (UI028), an annunciator/switch (SI033), and a logic module (UI027). The P/S HEAT/LOW A/S AWARE switch controls the electrical power supplied by the LEFT PITOT HEAT and RIGHT PITOT HEAT circuit breakers. The circuit breakers are on the left circuit breaker panel. When the P/S HEAT/LOW A/S AWARE switch is in the ON position, the LAA system has electrical power.
Description A.
3.
The annunciator switch is installed in the instrument panel near the front of the pilot position. The annunciator light has two colors, white and amber. The logic module is installed behind the left circuit breaker panel. The module gives the logic to operate the warning horn and cause the annunciator to come on. The airspeed switch is installed below the dash, approximately behind the middle of the instrument panel. When the airspeed goes below 97.5 KIAS , +2 -2 knots, the airspeed switch sends a discrete signal to the logic module. The module causes the warning horn to operate and the annunciator to come on.
Operation A.
The P/S HEAT/LOW A/S AWARE switch gives power to the LAA system when in the ON position. When the switch is put in the ON position before takeoff, the annunciator comes on white and shows BELOW ICING MIN SPD. On airplanes that have the GFC 700 auto pilot system, the system disengages. When the airspeed is more than 97.5 KIAS, +2 or -2 knots the annunciator goes off. If the airspeed goes below 97.5 KIAS +2 or -2, the annunciator shows amber and then white each second the stall warning horn operates on and off each second and the autopilot disengages. When the airspeed goes to or is more than 97.5 KIAS +2 or -2 knots, the annunciator goes off and the horn stops operating. When the low airspeed warning stops operating, the crew can engage the auto pilot system.
B.
Push the annunciator switch to set the annunciator to solid white to stop the horn operation, and to allow the autopilot to be engaged. The annunciator stays white until the airspeed is more than 97.5 KIAS, +2 or -2. If the P/S HEAT/LOW A/S AWARE switch is set to the ON position when the airspeed is more than 97.5 KIAS, +2 or -2 knots, the annunciator shows white for approximately one half a second. This shows that the LAA system has electrical power.
C.
The SW/CB PNLS/ANNUN knob controls the BELOW ICING MIN SPD annunciator intensity level. There are intensity levels for the day and for the night operation. Put the knob in the correct position necessary for the operation condition.
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MODEL 208 MAINTENANCE MANUAL LOW AIRSPEED AWARENESS SYSTEM - TROUBLESHOOTING (With TKS Anti-Ice System) 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 1)
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 2)
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 3)
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 4)
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 5)
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 6)
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Pitot-Static System Troubleshooting Chart Figure 101 (Sheet 7)
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MODEL 208 MAINTENANCE MANUAL LOW AIRSPEED AWARENESS SYSTEM - MAINTENANCE PRACTICES With TKS Anti-Ice System 1.
General A.
This section gives the maintenance procedures for the Low Airspeed Awareness (LAA) system. The LAA system is installed on airplanes that have the TKS system installed. The LAA system is installed on Airplanes 20800518 and On, Airplanes 208B2067 and On, and Airplanes 208B0001 Thru 208B2066 that incorporate CAB08-7.
B.
The LAA system is installed on airplanes that have the G1000 system installed and airplanes that do not have the G1000 system installed. The LAA system gives the crew warning indications if airspeed goes below 97.5 KIAS +2 or -2 KIAS. The maintenance procedures for the LAA systems on airplanes with G1000 installed and airplanes without G1000 installed are similar.
C.
Correct maintenance of the pitot-static system is necessary for correct operation of the altimeter, vertical speed indicator, airspeed indicator, and Low Speed Awareness (LAA) airspeed switch (UI028). Leaks, moisture, and blockages in the pitot system will cause incorrect airspeed indications. Static system malfunctions will affect indications of all three instruments and the airspeed switch. Clean components and correct installation are necessary for maintenance of the pitot-static system. When you replace pitot-static system components, use the minimum quantity of antiseize compound on the threads of the metal and the plastic connections. Always prevent too much compound that can go into the pitot-static system lines. Tighten connections tightly, but be very careful not to tighten too much and cause a distortion of the fittings.
CAUTION: Except for the use of the system drains and alternate static source pressure valves, make sure that you do a leak test after the static pressure system is opened or closed. Refer to Pitot System Inspection and Leak Test. 2.
Low Airspeed Awareness (LAA) Annunciator/Switch Removal/Installation A.
Remove the LAA Annunciator/Switch (SI033) (Refer to Figure 201, or Figure 203 as Applicable). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel, to the OFF position. (3) Make sure that the P/S HEAT/LOW A/S AWARE switch is in the OFF position. (4) Disengage the LEFT PITOT HEAT and RIGHT PITOT HEAT circuit breakers on the left circuit breaker panel. (5) Remove the face of the switch from the annunciator. NOTE: (6) (7) (8)
B.
The switch face stays connected to the switch because of its wiring.
Loosen the retainer tab screws. Disconnect the electrical connector (PI033) from the annunciator. Remove the annunciator from the instrument panel.
Install the LAA Annunciator/Switch (SI033) (Refer to Figure 201, or Figure 203 as Applicable). (1) Connect the electrical connector (PI033) to the annunciator. (2) Put the annunciator far enough in the instrument panel that the retainer tabs will engage the instrument panel structure. (3) Slowly tighten the retainer tabs. (4) Install the switch face to the switch. (a) Make sure that the face snaps in place. (5) Do a test of the LAA system. Refer to Low Airspeed Warning System - Adjustment/Test (With TKS Anti-Ice System).
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LAA Switch and Circuit Breaker Installation (G1000) Figure 201 (Sheet 1)
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LAA Components Installation (G1000) Figure 202 (Sheet 1)
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LAA Switch and Circuit Breaker Installation (Non-G1000) Figure 203 (Sheet 1)
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LAA Components Installation (Non-G1000) Figure 204 (Sheet 1)
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3.
Low Airspeed Awareness (LAA) Logic Module Removal/Installation A.
Remove the Logic Module (UI027) (Refer to Figure 201, and Figure 202, and Figure 203, and Figure 204 as Applicable). (1) Set the BATTERY switch on the circuit breaker switch panel, to the OFF position. NOTE:
(2) (3) (4)
For airplanes that have G1000 the reference designator for the BATTERY switch (SC005) and for airplanes that do not have G1000 the reference designator for the BATTERY switch is (S219).
Remove external electrical power from the airplane. Disconnect the battery terminals. Attach a warning tag to the battery and external power receptacle that have the statement that follows:
WARNING: DO NOT CONNECT ELECTRICAL POWER - MAINTENANCE IN PROGRESS (5) (6)
Make sure that the P/S HEAT/LOW A/S AWARE switch is in the OFF position. Disengage the LEFT PITOT HEAT and RIGHT PITOT HEAT circuit breakers on the left circuit breaker panel. (7) Remove the left circuit beaker panel. (8) Remove the electrical connector (PI027) from the LAA logic module. (9) Remover the screws that attach the logic module to the logic module bracket. (10) Remove the LAA logic module from the structure of the circuit breaker panel.
B.
4.
Install the LAA Logic Module (UI027) (Refer to Figure 201, and Figure 202, or Figure 203, and Figure 204 as Applicable). (1) Put the LAA logic module in its position on the module bracket. (2) Install the screws that attach the LAA logic module to the module bracket. (3) Connect the electrical connector (PI027) to the LAA logic module. (4) Install the circuit breaker panel. (5) Remove the warning tags from the external power receptacle and the battery. (6) Connect the battery terminals. (7) Do a test of the LAA system. Refer to Low Airspeed Warning System - Adjustment/Test (With TKS Anti-Ice System).
Low Airspeed Awareness (LAA) System Airspeed Switch Removal/Installation A.
Remove the LAA System Airspeed Switch (UI028) (For airplanes with the G1000, refer to Figure 201, and Figure 202. For airplanes that do not have G1000, refer to Figure 203, and Figure 204 as applicable). (1) Remove external electrical power from the airplane. (2) Set the BATTERY switch on the circuit breaker switch panel, to the OFF position. NOTE:
(3) (4) (5) (6) (7) (8)
For airplanes that have G1000 the reference designator for the BATTERY switch (SC005) and for airplanes that do not have G1000 the reference designator for the BATTERY switch is (S219).
Disconnect the electrical connector (PI028) from the LAA airspeed switch. Identify, and disconnect the lines that follow: • The pitot line • The static line. Put a cap on each of the two lines. Remove the two nuts and washers that attach the LAA pressure switch to the mounting bracket. Remove the pressure switch. Remove the two unions from the switch.
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MODEL 208 MAINTENANCE MANUAL B.
5.
Install the LAA System Airspeed Switch (UI028) (Refer to Figure 20,1 and Figure 202, or Figure 203, and Figure 204 as Applicable). (1) Install the two unions to the switch with new packing. (2) Put the airspeed switch in its position the mounting bracket. (3) Install the two washers and nuts that attach the airspeed switch to the mounting bracket. (4) Make sure that the ports marked P and S correctly oriented for the pitot and static line installation. (5) Remove the caps from the pitot and static lines. (6) Identify the pitot line and static line and connect them to the LAA airspeed switch as follows: (a) Connect the pitot line to the port marked P. (b) Connect the static line to the port marked S. (7) Connect the electrical connector (PI028) to the pressure switch. (8) Do a test of the LAA system. Refer to Low Airspeed Warning System - Adjustment/Test (With TKS Anti-Ice System).
Airspeed Warning Horn Removal/Installation A.
To remove and replace the airspeed warning horns (Refer to Pitot/Static - Maintenance Practices, Airspeed Warning Horn Removal/Installation).
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MODEL 208 MAINTENANCE MANUAL LOW AIRSPEED AWARENESS SYSTEM - ADJUSTMENT/TEST With TKS Anti-Ice System 1.
General A.
2.
This section gives the procedures to do a test of the Low Airspeed Awareness (LAA) system (97.5 KIAS). The LAA system is installed on airplanes that have the G1000 system installed and airplanes that do not have the G1000 installed. Do the steps applicable to the airplane systems installation.
Equipment NOTE: A.
The use of equivalent equipment is permitted for the equipment shown on Table 501.
Table 501 NAME
NUMBER
MANUFACTURER
USE
Air Data Tester
101-00184
Barfield 4101 NW 29th Street Miami, FL 33142-5617
To supply pressure or vacuum for the pitot and static system tests.
Pitot Static Test Adaptor
PS4769
Nav-Aids Ltd. 2955 Diab Street Montreal, Quebec H4S 1M1
To attach portable air data tester to pitot system.
3.
Low Airspeed Awareness System Functional Test A.
Do a Low Airspeed Awareness System Functional Test (Refer to Low Airspeed Awareness Maintenance Practices, Figure 201, Figure 202, Figure 203, and Figure 204 as applicable). (1) For airplanes with the G1000, disconnect the left and right pitot/static heater electrical connectors (PL004 left, PR004 right) at each wing. NOTE: (2) (3) (4) (5)
For airplanes that do not have G1000, disconnect the left and right pitot/static heater electrical connectors (P92 left, P152 right) at each wing. Connect an air data tester to the pitot/static system. Refer to Table 501. Connect external electrical power to the airplane. Set the BATTERY switch to the ON position. NOTE:
(6)
(7) (8)
When the pitot/static heat system is in operation the P/S HEAT L/R CAS message with show on the primary flight display.
For airplanes that have G1000 the reference designator for the BATTERY switch is (SC005) and for airplanes that do not have the G1000 the reference designator for the BATTERY switch is (S219)
Engage the circuit breakers on the left circuit breaker panel that follow: • LEFT PITOT HEAT • RIGHT PITOT HEAT • STALL WARN Set the P/S HEAT/ LOW A/S AWARE switch to the ON position. (a) Make sure that the BELOW ICING MIN SPD annunciator/switch (SI033) comes on (white). For airplanes with G1000 do the steps that follow: (a) Move the SW/CB PNLS/ANNUN dimmer (RI010) knob to the DAY position and then out of the DAY position. 1 Make sure that the intensity of the BELOW ICING MIN SPD annunciator/switch changes from day to night when the knob is moved in and out of the DAY position.
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MODEL 208 MAINTENANCE MANUAL (b)
Push the BELOW ICING MIN SPD annunciator/switch and at the same time, move the SW/CB PNLS/ANNUN dimmer knob to the DAY position, and then out of the DAY position. 1 Make sure that the intensity of the BELOW ICING MIN SPD annunciator/switch (amber) changes when the knob is moved. (9) For airplanes that do not have G1000, do the steps that follow: (a) Set the DAY/NIGHT switch to the NIGHT position. (b) Adjust the ENG INST dimmer. 1 Make sure that the intensity of the BELOW ICING MIN SPD annunciator/switch changes when the dimmer knob is moved in and out of the DAY position. (c) Push the BELOW ICING MIN SPD annunciator/switch and at the same time, move the SW/CB PNLS/ANNUN dimmer knob. 1 Make sure that the intensity of the BELOW ICING MIN SPD annunciator/switch (amber) changes when the knob is moved. (10) Use the air data tester to set the airspeed above 97.5 KIAS. (a) Make sure that the BELOW ICING MIN SPD annunciator/switch goes off when the airspeed exceeds 97.5 KIAS +2 or -2 KIAS. (11) Pull the control yoke back. NOTE:
This lets the airspeed warning horn operate.
(12) Use the air data tester to reduce the airspeed below 97.5 KIAS. (13) When the airspeed goes below 97.5 KIAS +2 or -2 KIAS make sure that: (a) The BELOW ICING MIN SPD annunciator/switch shows amber and then white. (b) The airspeed warning horn operates when the light is amber. (14) Move the control yoke forward. (a) Make sure that the airspeed warning horn stops operation. (15) Move the control yoke back. (a) Make sure that the airspeed warning horn operates. (16) Use the air data tester to increase the airspeed above 97.5 KIAS. (17) When the airspeed goes above 97.5 KIAS +2 or -2 KIAS. (a) The BELOW ICING MIN SPD annunciator/switch goes out. (b) The airspeed warning horn stops operation. (18) Use the air data tester to reduce the airspeed below 97.5 KIAS. (19) When the airspeed goes below 97.5 KIAS +2 or -2 KIAS make sure that: (a) The BELOW ICING MIN SPD annunciator/switch shows amber and then white. (b) The airspeed warning horn operates when the light is amber. (20) Push the BELOW ICING MIN SPD annunciator/switch. (a) Make sure that the annunciator/switch goes solid white. (b) Make sure that the horn stops operation. (21) Use the air data tester to set the airspeed above 97.5 KIAS. (a) Make sure that when the airspeed is more than 97.5 KIAS +2 or -2 KIAS, the BELOW ICING MIN SPD annunciator/switch goes off. (22) Set the P/S HEAT/ LOW A/S AWARE switch to the off position. (a) Make sure that the BELOW ICING MIN SPD annunciator/switch does not flash . (23) Set the P/S HEAT/ LOW A/S AWARE switch to the on position. (a) Make sure that the BELOW ICING MIN SPD annunciator/switch flashes for approximately one half a second. NOTE: (24) (25) (26) (27)
The airspeed must still be set above 97.5 KIAS.
Set the P/S HEAT/ LOW A/S AWARE switch to the OFF position. Set the airspeed to zero KIAS. Remove the air data tester from the airplane. For airplanes that have G1000 installed do the steps that follow: (a) Connect the left and right pitot/static heater electrical connectors (PL004 left, PR004 right) at each wing. (b) Disengage the LEFT PITOT HEAT circuit breaker on the left circuit breaker panel.
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MODEL 208 MAINTENANCE MANUAL (c)
Set the P/S HEAT/ LOW A/S AWARE switch to the ON position. 1 Make sure that the P/S HEAT L CAS message shows on the primary flight display. (d) Set the P/S HEAT/ LOW A/S AWARE switch to the OFF position. (e) Engage the LEFT PITOT HEAT circuit breaker on the left circuit breaker panel. (f) Disengage the RIGHT PITOT HEAT circuit breaker on the left circuit breaker panel. (g) Set the P/S HEAT/ LOW A/S AWARE switch to the ON position. 1 Make sure that the P/S HEAT R CAS message shows on the primary flight display. (h) Set the P/S HEAT/ LOW A/S AWARE switch to the OFF position. (i) Engage the RIGHT PITOT HEAT circuit breaker on the left circuit breaker panel. (28) For airplanes that do not have G1000 do the steps that follow: (a) Connect the left and right pitot/static heater electrical connectors (P92 left, P162 right) at each wing. (b) Disengage the LEFT PITOT HEAT circuit breaker on the left circuit breaker panel. (c) Set the P/S HEAT/ LOW A/S AWARE switch to the ON position. 1 Carefully make sure that the right pitot/static heater operates. 2 Carefully make sure that the left pitot/static heater does not operates. 3 Wait two minutes and make sure that the RIGHT PITOT HEATER circuit breaker is engaged. (d) Engage the LEFT PITOT HEAT circuit breaker on the left circuit breaker panel. (e) Disengage the RIGHT PITOT HEAT circuit breaker on the left circuit breaker panel. (f) Set the P/S HEAT/ LOW A/S AWARE switch to the ON position. 1 Carefully make sure that the left pitot/static heater operates. 2 Carefully make sure that the right pitot/static heater does not operates. 3 Wait two minutes and make sure that the LEFT PITOT HEATER circuit breaker is engaged. (g) Set the P/S HEAT/ LOW A/S AWARE switch to the OFF position. (29) Set the BATTERY switch to the OFF position. NOTE:
For airplanes that have G1000 the reference designator for the BATTERY switch is (SC005) and for airplanes that do not have the G1000 the reference designator for the BATTERY switch is (S219)
(30) Remove external electrical power from the airplane.
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MODEL 208 MAINTENANCE MANUAL LOW AIRSPEED AWARENESS SYSTEM - MAINTENANCE PRACTICES With Pneumatic De-Icing Systems 1.
General A.
This section gives the maintenance procedures for the Low Airspeed Awareness (LAA) system that is installed on airplanes with pneumatic anti-ice systems. The LAA system gives the crew warning indications if the airspeed goes below 110 KIAS, +5 or -5 KIAS.
B.
Correct maintenance of the pitot-static system is essential for correct operation of the altimeter, vertical speed indicator, and airspeed indicator. Leaks, moisture, and obstructions in the pitot system will result in false airspeed indications. Static system malfunctions will affect indications of all three instruments. Cleanliness and correct installation are the principal rules for maintenance of the pitot-static system. The pitot tube and static port MUST be kept clean and clear of obstructions. When you replace pitot-static system components, use the minimum amount of antiseize compound on the male threads of both metal and plastic connections. Always avoid excess compound which might be able to enter the pitot-static system lines. Tighten connections firmly, but be very careful not to over tighten and distort fittings.
CAUTION: Except for the use of the system drains and alternate static source pressure valves, make sure to do a leak test after the static pressure system is opened or closed. Refer to Pitot System Inspection and Leak Test. 2.
Low Airspeed Awareness Pressure System Switch Removal/Installation A.
Remove the Icing Low Airspeed Awareness System Pressure Switch (Refer to Figure 201). (1) Disconnect the electrical connector. (2) Identify, disconnect, and put a cap on the pitot line and the static line. (3) Remove the two nuts and washers that attach the pressure switch to the mounting bracket. (4) Remove the pressure switch. (5) Remove the two unions.
B.
Install the Icing Low Airspeed Awareness System (110 KIAS) Pressure Switch (Refer to Figure 201). (1) Install the two unions on the switch with new packings. (2) Put the pressure switch in the mounting bracket with the ports marked P and S correctly oriented, and install it with two washers and nuts. (3) Identify the pitot line and static line, remove the caps, and connect the pitot line and static line to the ports marked P and S, respectively. (4) Connect the electrical connector to the pressure switch. (5) Do the operational test. Refer to De-Icing Low Airspeed Awareness System Functional Test.
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LAA System Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL LOW AIRSPEED AWARENESS SYSTEM - ADJUSTMENT/TEST With Pneumatic De-Icing Systems 1.
General A.
2.
This section has the procedures to do a test of the Low Airspeed Awareness (LAA) system. The LAA system is installed on airplanes that have the pneumatic anti-ice system installed.
Equipment NOTE:
Equivalent equipment can be substituted for that listed below.
NAME
NUMBER
MANUFACTURER
USE
Air Data Tester
101-00184
Barfield 4101 NW 29th Street Miami, FL 33142-5617
To supply pressure or vacuum for the pitot and static system tests.
Pitot Static Test Adaptor
PS4769
Nav-Aids Ltd. 2955 Diab Street Montreal, Quebec H4S 1M1
To attach portable air data tester to pitot system.
3.
De-Icing Low Airspeed Awareness System Functional Test A.
Do an Icing Low Airspeed Awareness System Functional Test. (1) Put the BATTERY switch in the ON position. (2) Engage the PROP ANTI-ICE CONT and STALL WARN circuit breakers. (3) Push the annunciator panel lamp test button and make sure that the BELOW ICING MIN SPD annunciator light comes on. (4) Put the prop deice switch in the AUTO position and make sure that the BELOW ICING MIN SPD annunciator light comes on. (5) Put the DAY/NIGHT switch in the NIGHT position. (6) Adjust the engine instrument dimmer. (a) Make sure that the BELOW ICING MIN SPD annunciator becomes dimmer. (7) Put a cover on the drain hole on the pitot tube. (8) Install a pitot/static tester on the end of the pitot tube. (9) Use the pitot/static tester to bring the airspeed to more than 115 knots on the pilot's airspeed indicator. (a) Make sure that the BELOW ICING MIN SPD annunciator light goes off. (10) Use the pitot/static tester to decrease the airspeed to less than 105 knots on the pilot's airspeed indicator. (a) Make sure that the BELOW ICING MIN SPD annunciator light flashes between white and amber in color between 115 and 105 knots. (b) Make sure that the stall horn operates when the annunciator is amber in color. NOTE:
The control yoke must be pulled back to the aft position for the stall horn to operate.
(11) Use the pitot/static tester to increase the airspeed to more than 115 knots on the pilot's airspeed indicator. (a) Make sure that the BELOW ICING MIN SPD annunciator light is not on and the stall horn is not in operation. NOTE:
The control yoke must be pulled back to the aft position for the stall horn to operate.
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MODEL 208 MAINTENANCE MANUAL (12) Use the pitot/static tester to decrease the airspeed to less than 105 knots on the pilot's airspeed indicator. (a) Make sure that the BELOW ICING MIN SPD annunciator light flashes between white and amber in color between 115 and 105 knots. (b) Make sure that the stall horn operates when the annunciator is amber in color. NOTE:
The control yoke must be pulled back to the aft position for the stall horn to operate.
(13) Push the BELOW ICING MIN SPD annunciator switch. (a) Make sure that the BELOW ICING MIN SPD annunciator light stays solid white in color. (b) Make sure that the stall horn is not in operation. (14) Use the pitot/static tester to increase the airspeed to more than 115 knots on the pilot's airspeed indicator. (a) Make sure that the BELOW ICING MIN SPD annunciator light goes off between 105 and 115 knots on the pilot's airspeed indicator. (b) Return the pitot/static system to field elevation and disconnect the pitot/static tester. (15) Put the BATTERY switch in the OFF position.
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MODEL 208 MAINTENANCE MANUAL GARMIN GDC-74A AIR DATA COMPUTER - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives maintenance practices for the Garmin GDC-74A Air Data Computer and configuration module. The air data computer has electrical and pitot-static connections that must be disconnected before the unit can be removed. There are two air data computers installed, one for the pilot and one for the copilot. For a general description of the pitot-static system, refer to Pitot-Static System - Description and Operation.
GDC-74A Air Data Computer Removal/Installation A.
Remove the Air Data Computer (Refer to Figure 201). NOTE: (1)
(2) (3) (4) (5) (6) B.
Removal of the pilot's and copilot's air data computers is typical.
Remove electrical power from the airplane. (a) For the pilot's air data computer, disengage the ADC 1 circuit breaker on the circuit breaker panel. (b) For the copilot's air data computer, disengage the ADC 2 circuit breaker on the circuit breaker panel. Remove the pilot's or copilot's primary flight display (PFD). Refer to Garmin Display Unit Maintenance Practices. Disconnect the pitot and static hose adapters from the air data computer. (a) Put caps on the pitot and static hose adapters to prevent contamination in the pitot-static system. Disconnect the electrical connector from the air data computer. Loosen the thumbscrews on the mounting rack and plate that hold the air data computer in the G1000 system rack. Carefully remove the air data computer through the instrument panel.
Install the Air Data Computer (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6)
(7)
(8) (9)
Installation of the pilot's and copilot's air data computers is typical.
Put the air data computer in position on the G1000 system rack through the instrument panel. Install the plate and thumbscrews to the mounting rack to hold the air data computer in position on the G1000 system rack. Connect the electrical connector to the air data computer. Remove the caps from the pitot and static hose adapters and connect the hoses to the air data computer. Install the pilot's or copilot's primary flight display (PFD). Refer to Garmin Display Unit Maintenance Practices. Engage circuit breakers, as necessary. (a) For the pilot's air data computer, engage the ADC 1 circuit breaker on the circuit breaker panel. (b) For the copilot's air data computer, engage the ADC 2 circuit breaker on the circuit breaker panel. Make sure the air data computer operates correctly. (a) If a new unit is installed, load the software and configuration. Refer to the Garmin G1000 Line Maintenance Manual. (b) Do a check to make sure the air data computer operates correctly. Refer to the Garmin G1000 Line Maintenance Manual. Do the Pitot System Inspection and Leak Test. Refer to Pitot/Static System - Adjustment/Test. Do the Pitot/Static System Functional Check. Refer to Pitot/Static System - Inspection/Check.
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GDC-74A Air Data Computer Installation Figure 201 (Sheet 1)
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GDC-74A Air Data Computer Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL OUTSIDE AIR TEMPERATURE GAGE - MAINTENANCE PRACTICES 1.
General A.
2.
The airplane is equipped with a mechanical outside air temperature gage (OAT Gage). The gage is calibrated in both degrees Fahrenheit and Centigrade. It is located at upper left corner of windshield and extends up through fuselage.
Outside Air Temperature Gage Removal/Installation A.
Remove Outside Air Temperature Gage (Refer to Figure 201). (1) Unscrew sunshield tube (1) from OAT Gage (2), holding jamnut (3) in place. NOTE: (2) (3) (4)
B.
OAT Gage (2) must be held while removing or installing sunshield tube.
Slip sunshield tube (1) off of OAT Gage (2). Remove washer (4) and grommet (5) from surface of airplane skin (6). Remove OAT Gage (2), washer (7), and grommet (8) from airplane.
Install Outside Air Temperature Gage (Refer to Figure 201). (1) Position OAT Gage (2) through airplane trim panel (9), assuring jamnut (3), washer (7), and grommet (8) are in place on the OAT Gage. NOTE:
(2) (3)
Screw jamnut (3) out on OAT Gage (2) threads so that washer (7) and grommet (8) are pressed against fuselage (6) and sufficient clearance is allowed between OAT GAGE (2) dial and fuselage (6) for trim panel (9).
Slip grommet (5) and washer (4) onto OAT Gage (2). Screw sunshield tube (1) onto OAT Gage (2), while holding OAT Gage (2) in place. NOTE:
Top end of sunshield tube (1) should extend 0.50 inch beyond the top of OAT Gage (2) stem.
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Outside Air Temperature Gage Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GTP 59 OUTSIDE AIR TEMPERATURE (OAT) PROBE - MAINTENANCE PRACTICES 1.
2.
General A.
Each air data computer uses an outside air temperature (OAT) probe to calculate the outside ambient air data environment.
B.
The OAT probes are found on top of the fuselage at FS 157.00.
Outside Air Temperature (OAT) Sensor Removal/Installation A.
Remove the OAT Sensor (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5)
B.
Installation is typical for left and right probes.
Disconnect electrical power from the airplane. Remove the headliner above the crew seats. Refer to Cabin Upholstery-Maintenance Practices. Remove the jam nut and washer from the OAT sensor. Disconnect the electrical connector. Remove the OAT sensor from the airplane.
Install the OAT Sensor (Refer to Figure 201). (1) Clean and electrically bond (Type I) the installation surfaces of the airplanes skin and the OAT probe. Refer to Chapter 20, Electrical Bonding – Maintenance Practices. (2) Put the OAT sensor into the airplane. (a) Make sure the bonding jumper is installed between the probe and the airplanes skin. (3) Connect the electrical connector. (4) Install the washer and jam nut on the OAT sensor. (a) Use a Type I, Class B sealer to apply a fay seal between the washer and skin. Refer to Chapter 20, Fuel, Pressure, Weather and High-Temperature Sealing - Maintenance Practices. (b) Use a Type I, Class B sealer to apply a fay seal between the nut and washer. Refer to Chapter 20, Fuel, Pressure, Weather and High-Temperature Sealing - Maintenance Practices. (c) Use a Type I, Class B sealer to apply a shank seal between the OAT probe and the nut. Refer to Chapter 20, Fuel, Pressure, Weather and High-Temperature Sealing Maintenance Practices. (5) Make sure that the OAT probe is electrically bonded (Type I) to the airplane skin. (6) Install the headliner above the crew seats. Refer to Cabin Upholstery-Maintenance Practices. (7) Connect electrical power to the airplane. (8) Make sure that the OAT probe functions properly. (a) Make sure there are no red X's on the OAT and TAS indicators on PFD 1 or PFD 2. (b) Make sure the OAT indications on PFD 1 and PFD 2 are less than 5°F (3°C) different. (9) Disconnect electrical power from the airplane.
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GTP 59 Outside Air Temperature (OAT) Probe Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL VERTICAL SPEED INDICATOR - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system understanding. Refer to Figure 101.
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MODEL 208 MAINTENANCE MANUAL
Vertical Speed Indicator Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Vertical Speed Indicator Troubleshooting Chart Figure 101 (Sheet 2)
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Vertical Speed Indicator Troubleshooting Chart Figure 101 (Sheet 3)
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Vertical Speed Indicator Troubleshooting Chart Figure 101 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL VERTICAL SPEED INDICATOR - MAINTENANCE PRACTICES 1.
General A.
2.
3.
The vertical speed indicator, located in the pilot's instrument panel, measures the rate of change in static pressure when the airplane is climbing or descending. It indicates, by means of a pointer located on the face of the instrument, the rate of ascent or descent of the airplane in feet per minute. A zero adjust screw is located on the front of the VSl in the lower left corner to allow for pointer adjustment. An optional, dual, right located vertical speed indicator is also available. Maintenance practices for the right vertical speed indicator are the same as for the left indicator, which is covered in this chapter.
Vertical Speed Indicator Removal/Installation A.
Remove Vertical Speed Indicator (Refer to Figure 201). (1) Remove screws and washers securing removable flight panel (6) and slide flight panel (6) aft to gain access to back of vertical speed indicator (3). (2) Loosen clamp (1). (3) Disconnect, cap off, and identify line (2) from fitting (7), located on the back of the vertical speed indicator (3). (4) Remove screws (4) securing vertical speed indicator (3) and faceplate (5) to flight panel (6). (5) Remove vertical speed indicator (3) and faceplate (5). (6) Remove fitting (7) and plug opening in back of vertical speed indicator (3) to avoid possible contamination of the instrument.
B.
Install Vertical Speed Indicator (Refer to Figure 201). (1) Remove plug in back of vertical speed indicator (3) and install fitting (7). (2) Position vertical speed indicator (3) and faceplate (5) in appropriate opening in flight panel (6). (3) Install screws (4) securing faceplate (5) and vertical speed indicator (3) to flight panel (6). (4) Connect line (2) to fitting (7) and tighten clamp (1), located on the back of the vertical speed indicator (3). (5) Slide flight panel (6) forward and install screws and washers securing flight panel (6) to instrument panel.
Vertical Speed Indicator Adjustment A.
A mechanical zero adjust screw (8) is located on the front of the vertical speed indicator (3) in the lower left corner, allowing for pointer adjustment. Turning the zero adjust screw clockwise will make the pointer de flect downward, and turning the zero adjust screw counterclockwise will make the pointer deflect upward. Refer to Figure 201.
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Vertical Speed Indicator Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL TRUE AIRSPEED INDICATOR - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Airspeed Indicator Troubleshooting Chart Figure 101 (Sheet 1)
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Airspeed Indicator Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL TRUE AIRSPEED INDICATOR - MAINTENANCE PRACTICES 1.
General A.
2.
The true airspeed indicator, located in the upper left corner of the pilot's instrument panel, is equipped with a conversion ring, which may be rotated until pressure altitude is aligned with outside air temperature. Then, airspeed indicated on the instrument is read as true airspeed on the adjustable ring. An optional right located airspeed (not true airspeed) indicator is also available. Maintenance practices for the right airspeed indicator are the same as for the true airspeed indicator, except the retainer (9) and ring (10) on the true airspeed indicator are replaced by a faceplate on the right airspeed indicator installation.
True Airspeed Indicator Removal/Installation A.
Remove True Airspeed Indicator (Refer to Figure 201). (1) Remove screws and washers securing removable flight panel (7) and slide flight panel (7) aft to gain access to back of true airspeed indicator (6). (2) Loosen clamps (1) and (4). (3) Disconnect, cap off, and identify lines (2) and (3) from fittings (5), located on back of true airspeed indicator (6). (4) Remove screws (8) securing retainer (9), true airspeed ring (10), spacer (11), and true airspeed indicator (6) to flight panel (7). (5) Remove true airspeed indicator (6), retainer (9), true airspeed ring (10), and spacer (11). (6) Remove fittings (5) and plug openings in back of true airspeed indicator (6) to avoid possible contamination of instrument.
B.
Install True Airspeed Indicator (Refer to Figure 201). (1) Remove plugs in back of true airspeed indicator (6) and install fittings (5). (2) Position spacer (11) of adequate thickness required on true airspeed indicator (6). NOTE: (3) (4) (5)
Position true airspeed indicator (6) in its appropriate opening in flight panel (7) by inserting it up through forward side of flight panel. Position true airspeed ring (10) and retainer (9) on aft side of flight panel (7). Install screws (8) securing true airspeed indicator (6), true airspeed ring (10), and retainer (9) and spacer (11) to flight panel (7). NOTE:
(6) (7) (8) 3.
Use spacers (11) as required to provide adequate friction on true airspeed ring (10).
Do not overtighten screws (8) to the extent that true airspeed ring (10) and retainer (11) cannot be adjusted.
Connect lines (2) and (3) to fittings (5), located on back of true airspeed indicator (6). Tighten clamps (1) and (4). Adjust true airspeed indicator in accordance with procedures in True Airspeed Indicator Adjustment.
True Airspeed Indicator Adjustment A.
Adjust True Airspeed Indicator (Refer to Figure 201). (1) Loosen screws (8) securing true airspeed indicator (6), true airspeed ring (10), and (2) retainer (11) to flight panel (7). Rotate true airspeed ring (10) until 110 knots on true airspeed ring aligns with 110 knots on face of true airspeed indicator (6). (3) Holding this setting, move retainer (9) until 15°C aligns with zero pressure altitude; tighten screws (8).
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True Airspeed Indicator Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ALTIMETER - MAINTENANCE PRACTICES 1.
2.
General A.
The altimeter, located on pilot's instrument panel, converts static pressure into a visual indication of aircraft altitude above sea level. Pointers on instrument dial indicate altitude in increments of 100, 1000 and 10,000 feet, with a range of -1000 to 135,000 feet.
B.
A barometric scale is incorporated in the instrument. The scale is calibrated in inches of mercury and is set manually by a knob on lower left corner of altimeter case. Adjustment for local barometric conditions is made by manually rotating setting knob.
C.
An optional right located altimeter is also available. Maintenance practices for the right altimeter are the same as for the left altimeter covered in this chapter.
Altimeter Removal/Installation A.
Remove Altimeter (Refer to Figure 201). (1) Remove screws and washers securing removable ßight panel (6) and slide ßight panel aft to gain access to back of altimeter indicator (3). (2) Loosen clamp (1). (3) Disconnect and cap off line (2) from Þtting (7), located on back of altimeter (3). (4) Remove screws (4) securing altimeter (3) and faceplate (5) to ßight panel (6). (5) Remove altimeter (3) and faceplate (5) from ßight panel (6). (6) Remove Þtting (7) and plug opening in back of altimeter (3) to avoid possible contamination of instrument.
B.
Install Altimeter (Refer to Figure 201). (1) Remove plug in back of altimeter (3) and install Þtting (7). (2) Position altimeter (3) and faceplate (5) in appropriate opening in ßight panel (6). (3) Install screws (4) securing altimeter (3) and faceplate (5) to ßight panel (6). (4) Connect line (2) to Þtting (7), located on back of altimeter (3). (5) Tighten clamp (1). (6) Slide ßight panel (6) forward and install screws and washers securing ßight panel (6) to instrument panel. (7) (For airplanes equipped with KFC-225 autopilot.) If a new unit is installed or the unit is calibrated, do a system alignment. Refer to Introduction, the List of Manufacturers Technical Publications for the manufacturer's installation manual.
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Altimeter Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE 5035 SERIES) - DESCRIPTION AND OPERATION 1.
Description A.
2.
Operation A.
3.
Optional equipment (Model 208 Only) is mounted in upper portion of left instrument panel, and provides dual function of visual altitude indication and encoding altitude data, for further transmission to an Air Traffic Control Radar Beacon System ground facility through the airplane's transponder.
Refer to Section 9 in Pilot's Operating Handbook.
Maintenance Practices A.
Refer to Model 208 Avionic Installations Service/Parts Manual for removal, installation, and wiring diagrams. Refer to List of Publications in front of this manual for appropriate Service/Parts Manual.
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MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE KEA-130A) - DESCRIPTION AND OPERATION 1.
Description A.
2.
Optional equipment (Model 208 Only) is installed in the upper portion of left instrument panel, and gives dual function of visual altitude indication and encoding altitude data for transmission to an Air Traffic Control Radar Beacon System ground facility through the airplane's transponder.
Operation A.
Refer to Section 9 in Pilot's Operating Handbook.
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MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE KEA-130A) - MAINTENANCE PRACTICES 1.
General A.
2.
The airplane has an encoding altimeter installed on the left instrument panel. The encoding altimeter functions with the transponder to automatically report the altitude to air traffic control.
KEA-130A Radio Altimeter Maintenance Practices A.
Refer to the Model 208 Avionic Installations Service/Parts Manual for removal, installation, and wiring diagrams. Refer to the List of Publications in front of this manual for the appropriate Service/Parts Manual.
B.
If the airplane has a KFC-225 autopilot, refer to the autopilot configuration drawings supplied with the airplane for configuration and adjustment.
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MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE EA-401A) - DESCRIPTION AND OPERATION 1.
Description A.
2.
Operation A.
3.
Optional equipment (Model 208 Only) is mounted in upper portion of left instrument panel, and provides dual function of visual altitude indication and encoding altitude data for transmission to an Air TrafÞc Control Radar Beacon System ground facility through the airplane's transponder.
Refer to Section 9 in Pilot's Operating Handbook.
Maintenance Practices A.
Refer to Model 208 Avionic Installations Service/Parts Manual for removal, installation, and wiring diagrams. Refer to List of Publications in front of this manual for appropriate Service/Parts Manual.
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MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE EA-801A) AND ALTITUDE ALERTING WITH PRESELECT - DESCRIPTION AND OPERATION 1.
Description A.
2.
Operation A.
3.
The encoding altimeter is optional equipment (208 model only) and is mounted in upper portion of left instrument panel, and provides dual function of visual altitude indication and encoding altitude data for transmission to an Air TrafÞc Control Radar Beacon System ground facility, through the airplane’s transponder. The optional altitude alerter is mounted in left portion of instrument panel, and is used with encoding altimeter to supply preselected altitude signal to Integrated Flight Control System when installed. It also provides visual and aural warnings as airplane approaches, then deviates from selected altitude.
Refer to Section 9 in Pilot’s Operating Handbook.
Maintenance Practices A.
Refer to Model 208 Avionic Installations Service/Parts Manual for removal, installation, and wiring diagrams. Refer to List of Publications in front of this manual for appropriate SIGMATEC, INC./ARC Avionics Service/Parts Manual.
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MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE EA-401A) - DESCRIPTION AND OPERATION 1.
Description A.
2.
Operation A.
3.
Optional equipment (Model 208 Only) is mounted in upper portion of left instrument panel, and provides dual function of visual altitude indication and encoding altitude data for transmission to an Air Traffic Control Radar Beacon System ground facility through the airplane's transponder.
Refer to Section 9 in Pilot's Operating Handbook.
Maintenance Practices A.
Refer to Model 208 Avionic Installations Service/Parts Manual for removal, installation, and wiring diagrams. Refer to List of Publications in front of this manual for appropriate Service/Parts Manual.
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MODEL 208 MAINTENANCE MANUAL ENCODING ALTIMETER (TYPE EA-801A) AND ALTITUDE ALERTING WITH PRESELECT - DESCRIPTION AND OPERATION 1.
Description A.
2.
Operation A.
3.
The encoding altimeter is optional equipment (208 model only) and is mounted in upper portion of left instrument panel, and provides dual function of visual altitude indication and encoding altitude data for transmission to an Air Traffic Control Radar Beacon System ground facility, through the airplane’s transponder. The optional altitude alerter is mounted in left portion of instrument panel, and is used with encoding altimeter to supply preselected altitude signal to Integrated Flight Control System when installed. It also provides visual and aural warnings as airplane approaches, then deviates from selected altitude.
Refer to Section 9 in Pilot’s Operating Handbook.
Maintenance Practices A.
Refer to Model 208 Avionic Installations Service/Parts Manual for removal, installation, and wiring diagrams. Refer to List of Publications in front of this manual for appropriate SIGMATEC, INC./ARC Avionics Service/Parts Manual.
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MODEL 208 MAINTENANCE MANUAL RADAR ALTIMETER (TYPE KRA-10A) - DESCRIPTION AND OPERATION 1.
Description A.
2.
Operation A.
3.
The optional radar altimeter (Model 208 and Passenger) is mounted in left portion of instrument panel, and is designed to give absolute (depending on terrain re flectivity and airplane bank angle) altitude indication from 2500 to 35 feet AGL, +15 or -15 feet AGL.
Refer to Section 9 in Pilot’s Operating Handbook.
Maintenance Practices A.
Refer to King Avionic Installations Service/Parts Manual for removal and Installation. Refer to Cessna 208 Avionics Installations Service/Parts Manual for wiring diagrams. Refer to List of Publications in front of this manual for appriopriate King Service/Parts Manual.
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MODEL 208 MAINTENANCE MANUAL RADAR ALTIMETER (TYPE KRA-405B) - DESCRIPTION AND OPERATION 1.
Description A.
2.
Operation A.
3.
The optional radar altimeter (Model 208 and 208B) is installed in the aft fuselage of the airplane with an indicator installed in the left portion of the instrument panel. The altimeter gives an altitude indication range of 2000ft to -20ft, which can change with terrain reflectivity and airplane bank angle. The altitude indicator has two scales marked on the face. Between 0ft to 500ft, the scale is divided into 10ft increments, and from 500ft to 2000ft, the scale is divided into 100ft increments.
Refer to Section 9 in the Pilot’s Operating Handbook.
Maintenance A.
Refer to the King Avionic Installations Service/Parts Manual for the removal and Installation of the KRA-405B Radar Altimeter. Refer to the Cessna 208 Avionics Installations Service/Parts Manual for wiring diagrams. Refer to List of Publications in front of this manual for the applicable King Service/ Parts Manual.
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MODEL 208 MAINTENANCE MANUAL ATTITUDE AND DIRECTION - DESCRIPTION AND OPERATION 1.
Description A.
This section contains information pertaining to such items as magnetic compass, turn and bank indicator, and stall warning system. For information pertaining to standard gyros, refer to Chapter 37 of this manual, Vacuum Indicating - Maintenance Practices. For information pertaining to optional gyros, refer to Model 208 Avionics Installations Manual.
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MODEL 208 MAINTENANCE MANUAL MAGNETIC COMPASS - MAINTENANCE PRACTICES 1.
General A.
2.
The airplane is equipped with a magnetic compass, located on top left center of glareshield. The compass is liquid filled, containing a circular, calibrated compass card, visible through a window in compass case, with expansion provisions to compensate for temperature changes. It is equipped with compensating magnets and has two adjusting setscrews, one for North/South headings and one for East/West headings. These setscrews are located on face of compass, behind compass bezel. Lighting is integral and controlled by lower panel lights rheostat, located on left lower instrument panel.
Magnetic Compass Removal/Installation A.
Remove Magnetic Compass Assembly (Refer to Figure 201). (1) Remove screws securing cowl deck cover (11) and remove cowl deck cover (11). (2) Cut wires (1) at both ends of permanent splice. Identify wires and discard permanent splice. (3) Remove placard (2). (4) Remove screws (3). (5) Remove bezel (4) and compass assembly (7) from mounting bracket (5). (6) Remove screws (6) and mounting bracket (5).
B.
Install Magnetic Compass Assembly (Refer to Figure 201). (1) Position mounting bracket (5) to glareshield. (2) Install screws (6). (3) Position compass assembly (7) and bezel (4) to mounting bracket (5). (4) Install screws (3). (5) Install placard (7) on compass bezel (4). (6) Position wires (1) through grommet (12) in cowl deck. (7) Splice wires (1). (8) Position cowl deck cover (11) and secure with attaching screws. NOTE:
Anytime compass assembly (2) has been installed, refer to Align Compass, and perform a compass alignment check.
C.
Remove Compass (Refer to Figure 201). (1) Remove compass assembly (7) from mounting bracket (5). Refer to Remove Magnetic Compass Assembly. (2) Remove screw (8) and pull compass (10) out of compass cup (9).
D.
Install Compass (Refer to Figure 201). (1) Position compass (10) inside compass cup (9). (2) Install screw (8) securing compass (10) to compass cup (9). (3) Install compass assembly (7) as outlined in Install Magnetic Compass Assembly. NOTE:
E.
For calibration of compass (10), compass must be installed in airplane in compass cup (9). Refer to Align Compass, for compass alignment procedures, and Adjust Compass Calibration, for compass calibration procedures.
Align Compass (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5)
Compass alignment shall be performed on a calibrated compass rose.
Using a hand held magnetic compass, check all ferrous material parts for magnetism near magnetic compass. Degauss any parts within two feet which cause greater than a 10 degree deflection, and any part within four feet which causes greater than 90 degree deflection of magnetic compass. Ensure all electrical instruments for aircraft are properly installed and operating correctly. Ensure other aircraft and vehicles are of a safe distance away. Position aircraft on 270 degree heading of compass rose.
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Magnetic Compass Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (6) (7)
With engine running and power lever in idle position, turn on the following: all circuit breakers, all lights except landing lights and reading lights, all avionics systems, and all electrical systems except pitot heat and stall warning heat. Record compass error in degrees. NOTE:
High readings are positive errors, low readings are negative errors.
(8) (9) (10) (11)
Position aircraft on 360 degree heading of compass rose, and repeat steps (6) and (7). Position aircraft on 90 degree heading of compass rose, and repeat steps (6) and (7). Position aircraft on 180 degree heading of compass rose, and repeat steps (6) and (7). Algebraically sum North and South errors, divide this sum by two, and change sign of result. Resulting number is amount and direction of North/South calibration adjustment. (12) Repeat step (11) for east/west calibration adjustment using East/West errors. F.
Adjust Compass Calibration. (1) Errors obtained in compass alignment procedure, steps (6) through (12), will be used to determine required amount and degree of calibration for compass.
G.
Calibrate Compass. (1) At one cardinal heading, adjust appropriate calibration screw amount calculated in compass alignment procedure, steps (6) through (12). (2) Rotate aircraft 90 degrees and adjust appropriate calibration screw and amount calculated. (3) Rotate aircraft to next two cardinal headings and ensure that no error greater than 5 degrees is present. (4) With normal aircraft power on and all electrical systems on, rotate aircraft to 30 degree headings (including cardinals). Stop on each heading long enough to allow compass to stabilize. (5) Record headings indicated by compass at 30-degree positions. No error greater than +5 or -5 degrees is to be allowed.
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MODEL 208 MAINTENANCE MANUAL MAGNETIC COMPASS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the magnetic compass in a serviceable condition.
Task 34-21-00-720 2.
Magnetic Compass Functional Check A.
General (1) This section gives the information needed to do a functional check of the magnetic compass.
B.
Special Tools (1) Non Ferrous Screw Driver
C.
Access (1) None
D.
Do a Functional Check (Alignment) of the Magnetic Compass. NOTE: (1) (2) (3)
The following procedures are to be done on a calibrated compass rose.
Start and taxi the airplane to an approved calibrated compass rose. Refer to the Model 208 FAA Approved Airplane Flight Manual. Put the airplane on a north heading line, 0/360 degrees +0.5 or -0.5 degree. Put the airplane in a test configuration with the engine running and the power lever in idle position, all circuit breakers engaged, and the following turn on: all lights except landing lights and reading lights, all avionics systems, and all electrical systems except pitot heat and stall warning heat. NOTE:
(4)
Record the compass error in degrees. NOTE:
(5)
(6)
This configuration is used to record the compass errors at the different headings on compass rose.
High readings are positive errors, low readings are negative errors.
Put the airplane in the test configuration and record the compass errors with the airplane in the headings that follow: (a) 90 degree heading (b) 180 degree heading (c) 270 degree heading (d) 0/360 degree heading. Add the errors for the north and south heading, then divide by 2. (a) If the number is negative, adjust the magnetic compass in a positive direction. (b) If the number is positive, adjust the magnetic compass in a negative direction. NOTE:
(c)
Example: -7° error + 4° error = -3° error. -3° divided by 2 = -1.5° error correction factor. The magnetic compass would be adjusted in a positive direction 1.5 degrees.
Do the steps again for the east and west errors.
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MODEL 208 MAINTENANCE MANUAL E.
Do a Functional Check (Calibration) of the Magnetic Compass. NOTE: (1) (2) (3) (4) (5) (6) (7) (8)
The recorded error adjustments from the functional check (calibration) of the magnetic compass are used to find the necessary amount and degree of calibration for the compass.
At one cardinal heading, adjust applicable calibration screw the necessary amount calculated in the compass alignment procedure. Turn the airplane 90 degrees and adjust the applicable calibration screw and amount calculated. Turn the airplane to the next two cardinal headings and make sure that there are not errors more than 5 degrees. With normal electrical power on the airplane, and all of electrical systems on, turn the airplane to the different 30 degree headings (including cardinals). (a) Stop at each heading for a sufficient amount of time to let the compass to stabilize. Record the headings shown on the compass at each of the 30-degree positions. (a) Errors that are more than +5 or -5 degrees are not permitted. Taxi the airplane back to the necessary area. Stop the engine. Refer to the Model 208 FAA Approved Airplane Flight Manual. Remove electrical power from the airplane.
F.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL TURN AND BANK INDICATOR - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technicain in system troubleshooting. Refer to Figure 101.
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Turn and Bank Indicator Troubleshooting Chart Figure 101 (Sheet 1)
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Turn and Bank Indicator Troubleshooting Chart Figure 101 (Sheet 2)
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Turn and Bank Indicator Troubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL TURN AND BANK INDICATOR - MAINTENANCE PRACTICES 1.
General A. (1)
2.
The turn and bank indicator is a combination instrument located in the pilot's instrument panel to the left of control column. The indicator consists of an electrically driven, gyroscopic rate of turn indicator that operates only when aircraft battery switch is in the ON position. A fluid dampened inclinometer is also located in the instrument. The inclinometer consists of a curved, liquid filled glass tube in which a ball, moving with dampened motion, changes positions according to gravitational and centrifugal force acting upon airplane. An optional right located turn and bank indicator is also available. Trouble shooting and maintenance practices for the right turn and bank indicator are the same as for the left turn and bank indicator covered in this chapter.
Turn and Bank Indicator Removal/Installation A.
Remove Turn and Bank Indicator (Refer to Figure 201). (1) Remove screws and washers securing removable flight panel (8) and slide flight panel (8) aft to gain access to back of turn and bank indicator (3). (2) Remove two postlight assemblies (1) in accordance with procedures in Chapter 33, Lighting. (3) Disconnect electrical connector (2), located on back of turn and bank indicator (3). (4) Remove screws (4) securing indicator (3) to faceplate (7). (5) Remove screws (5) and nuts (6) securing faceplate (7) to flight panel (8).
B.
Install Turn and Bank Indicator (Refer to Figure 201). (1) Position faceplate (7) to flight panel (8) and secure with screws (5) and nuts (6). (2) Position indicator (3) to faceplate (7) and secure with screws (4). (3) Connect electrical connector (2). (4) Install two postlight assemblies (1) in accordance with procedures in Chapter 33, Lighting. (5) Slide flight panel (8) forward and install screws and washers securing flight panel (8) to instrument panel.
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Turn and Bank Indicator Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL HORIZON AND DIRECTIONAL GYROS - DESCRIPTION AND OPERATION 1.
Description A.
For maintenance information for standard vacuum driven horizon and directional gyros, left and right, not connected to an autopilot or IFCS, refer to Chapter 37, Vacuum Indicating - Maintenance Practices. For location, Refer to Figure 1.
B.
For electrical wiring information for optional vacuum driven horizon and directional gyros, installed with a 400B Autopilot or 400B IFCS, refer to 208 Avionic Installations Service/Parts Manual. Refer to List of Publications in the front of this manual for the appropriate King Maintenance Manual.
C.
For maintenance information for standard electronically driven horizon and directional gyros, refer to List of Publications in the front of this manual for the appropriate King Flight Control and Avionics System Installation Maintenance Manual. For location, Refer to Figure 1.
D.
Optional electronically driven horizon and directional gyros are available on Model 208 only. For location, Refer to Figure 1.
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Gyro Locations Figure 1 (Sheet 1)
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Gyro Locations Figure 1 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL HORIZON AND DIRECTIONAL GYROS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
4.
This section gives information to remove and install the components of the electronically driven horizon and directional gyros.
Electrical Horizon Gyro Removal/Installation (Model 208 Only) A.
Remove Electrical Horizon Gyro (Refer to Figure 201). (1) Remove screws and washers securing right removable ßight panel, and slide panel aft to gain access to back of gyro (1). (2) Disconnect electrical plug (2). (3) Cut sta-strap (10). (4) Remove screws (3), securing gyro (1) and faceplate (4) to ßight panel, and lift gyro (1) from ßight panel. (5) Remove hose (11).
B.
Install Electrical Horizon Gyro (Refer to Figure 201). (1) Position hose (11) and sta-strap (10). (2) Position gyro (1) and faceplate (4) to removable ßight panel. (3) Install screws (3). (4) Connect sta-strap (10). (5) Connect electrical plug (2). (6) Slide removable ßight panel back against instrument panel, and install screws and washers. (7) (For airplanes equipped with KFC-225 autopilot.) If a new unit is installed or the unit is calibrated, do a system alignment. Refer to Introduction, the List of Manufacturers Technical Publications for the manufacturer's installation manual.
Electrical Directional Gyro Removal/Installation (Model 208 Only) A.
Remove Electrical Directional Gyro (Refer to Figure 202). (1) Remove screws and washers, securing right removable ßight panel, and slide panel aft to gain access to back of gyro (5). (2) Disconnect electrical plug (6). (3) Cut sta-strap (10). (4) Remove screws (7), securing gyro (5) and faceplate (8) to ßight panel, and lift gyro (5) from ßight panel. (5) Remove hose (11).
B.
Install Electrical Directional Gyro (Refer to Figure 202). (1) Position hose (11) and sta- strap (10). (2) Position gyro (5) and faceplate (8) to removable ßight panel. (3) Install screws (7). (4) Connect sta-strap (10). (5) Connect electrical plug (6). (6) Slide ßight panel back against instrument panel, and install screws and washers. (7) (For airplanes equipped with KFC-225 autopilot.) If a new unit is installed or the unit is calibrated, do a system alignment. Refer to Introduction, the List of Manufacturers Technical Publications for the manufacturer's installation manual.
Electrical Horizon Gyro Removal/Installation (Model 208B and 208B Passenger) A.
Remove Electrical Horizon Gyro (Refer to Figure 202). (1) Remove screws and washers securing right removable ßight panel, and slide panel aft to gain access to back of gyro (2). (2) Disconnect electrical plug (1). (3) Remove screws (5) securing faceplate (4) and gyro (2), and lift gyro (2) from ßight panel.
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Optional Electric Gyros Installation Figure 201 (Sheet 1)
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Electric Gyros Installation Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
5.
(Refer to Figure 202). (1) Position gyro (2) in removable ßight panel, and install faceplate (4) and gyro to panel using screws (5). (2) Connect electrical plug (1) to gyro(2). (3) Slide removable ßight panel back against instrument panel, and install screws and washers. (4) (For airplanes equipped with KFC-225 autopilot.) If a new unit is installed or the unit is calibrated, do a system alignment. Refer to Introduction, the List of Manufacturers Technical Publications for the manufacturer's installation manual.
HSI Indicator Removal/Installation (Model 208B and 208B Passenger) A.
Remove HSI Indicator (Refer to Figure 202). (1) Remove screws and washers securing left removable ßight panel, and slide panel aft to gain access to back of HSI indicator (7). (2) Disconnect electrical plugs (6) and (11). (3) Remove screws (10) securing faceplate (9) and HSl (7), and lift HSl (7) from ßight panel.
B.
Install HSI Indicator (Refer to Figure 202). (1) Position HSI (7) in removable ßight panel, and install faceplate (9) and HSl (7) to panel using screws (10). (2) Connect electrical plugs (6) and (11) to HSl (7). (3) Slide panel back against instrument panel, and install screws and washers.
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MODEL 208 MAINTENANCE MANUAL GMU 44 MAGNETOMETER - MAINTENANCE PRACTICES 1.
2.
General A.
On airplanes with Garmin G1000, the GMU 44 magnetometers sense magnetic field information. The data is used by the GRS 77 AHRSs to find aircraft magnetic heading.
B.
Maintenance practices give procedures for the removal and installation of the GMU 44 magnetometer. The unit is removed and installed through an access panels on the bottom side of both the left and right wing.
Troubleshooting A.
3.
For troubleshooting procedures, refer to the Garmin G1000 Line Maintenance Manual.
GMU 44 Magnetometer Removal/Installation A.
Remove the Magnetometer (Refer to Figure 201). (1) Put the MASTER switch in the off position. (2) Put the AVIONICS switch in the off position.
CAUTION: Do not use magnetized tools or screws around the magnetometer. Use of magnetized tools or screws can cause an incorrect heading indication. (3) (4) (5) B.
Remove the access plates 523AB for left wing, 623AB for right wing) to get to the magnetometer. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. Remove the screws that attach the magnetometer to the flux detector bracket. Disconnect the electrical connector.
Install the Magnetometer (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7)
If a new unit is installed, the software must be loaded.
Make sure the electrical connector and connector pins have no damage. (a) Replace the electrical connector or connector pins if applicable. Refer to the Model 208 Wiring Diagram Manual and the Garmin G1000 Line Maintenance Manual. Connect the electrical connector. Attach the magnetometer to the flux detector bracket with the screws. (a) Put the magnetometer in position on the flux detector bracket, temporarily aligned parallel to the longitudinal axis of the airplane. If a new unit is installed, load the software. Refer to the Garmin G1000 Line Maintenance Manual. Do the calibration procedure. Refer to the Garmin G1000 Line Maintenance Manual. Install access plates. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. Do a check to make sure the magnetometer operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
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Magnetometer Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ATTITUDE HEADING REFERENCE SYSTEM (AHRS) - MAINTENANCE PRACTICES 1.
2.
General A.
On airplanes with Garmin G1000, the GRS 77 AHRS is an attitude, heading, and reference unit that gives airplane attitude and ßight characteristics information to the Primary Flight Display (PFD) and Multi-Function Display (MFD) and the GIA 63 Integrated Avionics Units. The unit has advanced tilt sensors, accelerometers, and rate sensors. In addition, the GRS 77 AHRS interfaces with both the GDC 74A Air Data computer and the GMU 44 Magnetometer. The GRS 77 AHRS also uses GPS signals sent from the GIA 63.
B.
Maintenance practices give procedures for the removal and installation of the GRS 77 AHRS.
Troubleshooting A.
3.
For troubleshooting procedures, refer to the Garmin G1000 Line Maintenance Manual.
GRS 77 AHRS #1 Removal/Installation NOTE: A.
If the mounting screws that attach the mounting rack to the airplane structure are loosened after post-calibration has been completed, the GRS 77 AHRS must be calibrated.
Remove the AHRS unit (Refer to Figure 201). (1) Disconnect electrical power from the aircraft. (a) Disengage AHRS 1 circuit breaker. (2) Remove the end cover from the copilot's avionics rack. NOTE: (3) (4) (5)
B.
The copilot's avionics rack is found in front of the copilot's door under the instrument panel.
Disconnect the electrical connector. Loosen the thumbscrews that attach the AHRS unit to the mounting tray. Lift the unit out of the mounting tray.
Install the AHRS unit (Refer to Figure 201). NOTE:
If a new AHRS unit is installed, the software must be loaded.
NOTE:
If the mounting screws that attach the mounting rack to the airplane structure are loosened after post-calibration has been completed, the GRS 77 AHRS must be calibrated.
(1) (2) (3) (4) (5) (6) (7)
(8)
Make sure the electrical connector and connector pins are not damaged. (a) Replace the electrical connector or connector pins if applicable. Refer to the Model 208 Wiring Diagram Manual and the Garmin G1000 Line Maintenance Manual. Put the AHRS unit in position in the mounting tray. Attach the AHRS unit with the thumbscrews. Connect the electrical connector. Engage the AHRS 1 circuit breaker. Connect electrical power to the aircraft. Make sure the AHRS unit operates correctly. (a) If the mounting screws that attach the mounting rack to the airplane structure have been loosened after post-calibration has been completed, calibrate the AHRS unit. Refer to the Garmin G1000 Line Maintenance Manual. (b) If a new unit is installed, load the software. Refer to the Garmin G1000 Line Maintenance Manual. (c) Do a check to make sure the AHRS operates correctly. Refer to the Garmin G1000 Line Maintenance Manual. Install the end cover on the copilot's avionics rack.
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Attitude Heading Reference System (AHRS) Installation Figure 201 (Sheet 1)
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4.
GRS 77 AHRS #2 Removal/Installation NOTE:
If the mounting screws that attach the mounting rack to the airplane structure are loosened after post-calibration has been completed, the GRS 77 AHRS must be calibrated.
A.
Remove the AHRS unit (Refer to Figure 201). (1) Disconnect electrical power from the aircraft. (a) Disengage AHRS 2 circuit breaker. (2) Remove the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (3) Remove access panel 232BR. Refer to Chapter 6, Access Plates and Panels IdentiÞcation. (4) Disconnect the electrical connector. (5) Loosen the thumbscrews that attach the AHRS unit to the mounting tray. (6) Lift the unit out of the mounting tray.
B.
Install the AHRS unit (Refer to Figure 201). NOTE:
If a new AHRS unit is installed, the software must be loaded.
NOTE:
If the mounting screws that attach the mounting rack to the airplane structure are loosened after post-calibration has been completed, the GRS 77 AHRS must be calibrated.
(1) (2) (3) (4) (5) (6) (7)
(8) (9)
Make sure the electrical connector and connector pins are not damaged. (a) Replace the electrical connector or connector pins if applicable. Refer to the Model 208 Wiring Diagram Manual and the Garmin G1000 Line Maintenance Manual. Put the AHRS unit in position in the mounting tray. Attach the AHRS unit with the thumbscrews. Connect the electrical connector. Engage the AHRS 2 circuit breaker. Connect electrical power to the aircraft. Make sure the AHRS unit operates correctly. (a) If the mounting screws that attach the mounting rack to the airplane structure have been loosened after post-calibration has been completed, calibrate the AHRS unit. Refer to the Garmin G1000 Line Maintenance Manual. (b) If a new unit is installed, load the software. Refer to the Garmin G1000 Line Maintenance Manual. (c) Do a check to make sure the AHRS operates correctly. Refer to the Garmin G1000 Line Maintenance Manual. Install the access panel. Install the copilot's seat.
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MODEL 208 MAINTENANCE MANUAL LANDING AIDS - DESCRIPTION AND OPERATION 1.
Description A.
The landing and taxiing systems provide guidance during approach, landing, and taxiing. localizer, glide scope, and marker beacon systems are utilized for approach and landing.
B.
Sperry Marker Beacon System (Model 208 Only). (1) This system is internally incorporated in the Sperry (Type SMA-90) Audio Control Panel with an antenna mounted on the lower surface of the fuselage.
C.
King Marker Beacon System. (1) This optional system includes a crystal controlled superheterodyne marker beacon receiver with 3-light presentation incorporated in KMA-24 Audio Control Panel. The system also incorporates an external bottom mounted antenna. Refer to 208 Avionic Installations Service/Parts Manual for electrical wiring diagrams. Refer to KMA-24 Audio Control Panel with Marker Beacon Service/ Parts Manual, listed in List of Publications in front of this manual, for maintenance information on this marker beacon system.
D.
Allied-Signal Avionics (KR21) Beacon System. (1) This system includes a 3-light presentation incorporated into the KR21 marker receiver. The system also incorporates an external bottom mounted antenna. Refer to 208 Avionics Installations Service/Parts Manual for electrical wiring diagrams.
E.
Sperry 400 (Type R-443B) Glide Slope System (Model 208 Only). (1) This optional system can be installed as a single or dual installation. The system consists of a remote mounted receiver coupled to an existing 300 or 400 navigation system, a panel mounted indicator, and an externally mounted antenna. Refer to 208 Avionic Installations Service/Parts Manual for installation and electrical wiring information. Refer to Sperry 400 Glide Slope Manual, listed in List of Publications in front of this manual, for maintenance information on this glide slope system.
F.
King Nav/Com (Type KX-165) With Integral Glide Slope Receiver. (1) The King Glide Slope System can be installed as a single or dual installation. The system consists of a remote mounted receiver coupled to an existing KX-165 navigation system, a panel mounted HSI, and an externally mounted antenna. Refer to 208 Avionics Installations Service/Parts Manual for wiring diagrams. Refer to King KX-165 Nav/Com With Integral Glide Slope Receiver Service Parts Manual, listed in List of Publications in front of this manual, for maintenance information on this glide slope system.
G.
Sperry and King Localizers for Glide Scope and Marker Beacon Options. (1) The Sperry and King receivers for localizer, glide scope, and marker beacon and VHF omnidirectional range (VOR) are all combined into one navigational receiver (NAV 1); also, NAV 1, when dual glidescopes are installed. Optional Sperry NAV receivers will be either 300 Nav/Com (RT-385A) or 400 Nav/Com (RT-485B) on 208 Models only. The optional King Nav receivers (KX-165 Nav/Coms) are available on 208 and 208B Passenger Models. (2) For installation and wiring information on Sperry and King Nav/Coms, refer to 208 Avionic Installations Service/Parts Manual. Refer to appropriate Sperry or King Nav/Com manual, listed in List of Publications in front of this manual, for maintenance information on Nav/Com radios.
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MODEL 208 MAINTENANCE MANUAL INDEPENDENT POSITION DETERMINING - DESCRIPTION AND OPERATION 1.
Description A.
The independent position determining systems provide information to aid in determining airplane position, and operate mainly independent of ground installation, including weather radar.
B.
King KWX-56 Color Weather Radar (Model 208, 20800061 and On). (1) This optional system consists of a wing mounted receiver/transmitter pod, and a stabilized Xband radar antenna and a panel mounted radar indicator. Refer to 208 Avionic Installations Service/Parts Manual for wiring diagrams. Refer to List of Publications in front of this manual for appropriate King Maintenance Manual on this system.
C.
Bendix RDS-82 Color Weather Radar (Model 208, 20800061 and On). (1) This optional system consists of a wing mounted receiver/transmitter pod, and a stabilized Xband radar antenna and a panel mounted radar indicator. Refer to 208 Avionic Installations Service/Parts Manual for installation and electrical wiring information. Refer to List of Publications in front of this manual for appropriate Bendix Maintenance Manual on this system.
D.
Bendix RDS-81 Color Weather Radar (Model 208B and Model 208B Passenger). (1) This optional system consists of a wing mounted receiver/transmitter pod, and a stabilized X-band radar antenna and a panel mounted radar indicator. Refer to 208 Avionic Installations Service/Parts Manual for installation and electrical wiring information. Refer to List of Publications in front of this manual for appropriate Bendix Maintenance Manual on this system.
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MODEL 208 MAINTENANCE MANUAL BF GOODRICH WX-1000+/E STORMSCOPE WEATHER MAPPING SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
WX-1000+/E Stormscope Weather Mapping System is a lightning detection system which detects, locates, and maps areas of vertical electrical charge activity (lightning) 360 degrees around airplane, to a distance of 200 nautical miles.
B.
WX-1000+/E Stormscope System contains an electronic checklist which may be customized to meet specific requirements. Refer to manufacturer’s installation manual listed in the Introduction, List of Manufacturers Technical Publications .
C.
WX-1000+/E Stormscope Weather Mapping System is capable of connecting to an external analog synchro (gyro) heading source to enable airplane heading to be displayed, as well as maintaining proper lightning display orientation while turning.
D.
Additionally, WX- 1000+/E can be interfaced with the BF Goodrich Skywatch System, to provide dual use with one indicator. A switch located near the indicator selects the desired mode. If Indicator switch is selected to display, Stormscope and a traffic advisory is detected by Skywatch System, display will automatically switch to Skywatch mode.
E.
WX-1000+/E Stormscope Weather Mapping System consists of three major components - a receiver/ transmitter antenna, processor, and display unit. NOTE:
F.
2.
Display unit is also utilized by the BF Goodrich Skywatch Sky-497 system, and is referred to as Stormscope/Skywatch display
A self-test is automatically initiated each time system is turned on. Following the self-test, a main menu presents the operator with a menu of system operating modes. Refer to the manufacturer’s installation manual listed in the Introduction, List of Manufacturers Technical Publications for the operation of available modes.
Description and Operation A.
An aerodynamically designed, flat pack, combined crossed loop-and-sense antenna receives both electric and magnetic fields emitting from lightning generated by thunderstorms. Operating parameters permit the antenna to sense electrical and magnetic fields around the airplane, then forwards this information to a processor for analysis. Differential drive amplifiers are used for both X and Y loop signals and for sense antenna electric field. Test generators are built into the antenna to continuously verify proper antenna performance of each antenna section.
B.
Analog processing is utilized by the processor to detect, discriminate, sample, hold, and convert signals to digital format. Software algorithms are used to process digital signal components to simultaneously analyze the field data to determine and validate the range, and magnetic field data to determine azimuth. Analysis is completed in less than two milliseconds and results forwarded to display unit. The processor has built-in auto-test and diagnostic functions including antenna, analog processing, program, video, and data memories, heading flag, clock battery, and stuck mic key (over one minute). Tests are conducted during power up or may be operator initiated. Six primary functions are continuously tested several times per minute during system operation. The processor also contains the system power supply.
C.
The display unit has a high resolution (256 x 256), raster-scan three inch diameter cathode ray tube (CRT) with precise, real-time lightning activity information within a 200 nautical mile range instantly plotted on the CRT. It receives its video, horizontal sync 16,390 Hz, and vertical syncs the 60 Hz from the processor. The basic circuit functions similarly to a conventional monochrome monitor, except using balanced input drives. Four buttons control the processor functions for various displays. The power switch is coupled to brightness control.
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MODEL 208 MAINTENANCE MANUAL BF GOODRICH WX-1000+/E STORMSCOPE WEATHER MAPPING SYSTEM - REMOVAL/INSTALLATION 1.
General A.
This section describes the removal and installation of the WX-1000+/E Stormscope Weather Mapping System. (1) The WX-1000+/E Stormscope is located on the aft avionics shelf. (2) The Stormscope antenna is located on the upper fuselage at FS 263.96 and RBL 5.00
B.
Components consist of an antenna, computer, and display unit. Maintenance practices for these components consists of removal and installation. NOTE:
C. 2.
3.
4.
The Stormscope display is known as a WX- 1000/SKY497 Display and is shared with the Skywatch (SKY497) Traffic Advisory System. It is located in the instrument panel. A switch mounted to the lower left of the display allows the crew to select either option. The display will automatically switch to Skywatch if traffic is detected.
For additional maintenance procedures, tests, and troubleshooting charts, refer to manufacturers publications listed in Introduction, List of Manufacturers Technical Publications.
Stormscope Computer Removal and Installation A.
Remove Stormscope Computer (Refer to Figure 401) (1) Disengage STM SCP circuit breaker on left circuit breaker panel. (2) Locate the Stromscope computer on the aft avionics shelf. (3) Loosen knurl nut securing Stormscope computer to tray assembly. (4) Gently pull computer forward and carefully remove Stormscope computer from mounting tray. (5) Remove Stormscope computer from airplane.
B.
Install Stormscope Computer (Refer to Figure 401). (1) Guide Stormscope computer aft into mounting tray ensuring proper alignment of electrical connectors. (2) Tighten knurl nut securing Stormscope computer to tray assembly. (3) Engage STM SCP circuit breaker on circuit breaker panel.
WX1000/SKY497 Display Unit Removal and Installation A.
Remove WX-1000/SKY497 Display (Refer to Figure 401). (1) Disengage circuit breakers STM SCP and SKY WATCH on lower left circuit breaker panel. (2) On instrument panel, remove WX-1000/SKY497 display by removing screws. (3) Move display outward from instrument panel and disconnect electrical connector. (4) Remove display from airplane.
B.
Install WX1000/SKY497 Display Unit (Refer to Figure 401) (1) Position display near mounting hole and connect electrical connector. (2) Place display into instrument panel and secure with screws. (3) Engage circuit breakers STM SCP and SKY WATCH on left circuit breaker panel.
Stormscope Antenna Removal and Installation A.
Remove Stormscope Antenna (Refer to Figure 401) (1) Disengage circuit breaker STM SCP on left circuit breaker panel. (2) Remove hard shell headliner panel inside cabin just below the Stromscope antenna located at FS 263.96 and RBL 5.00 by removing fasteners. (a) Allow panel to drop enough to reach coax connector. (b) Disconnect coax connector. (3) Remove old sealant from top of antenna screws. (4) Remove screws securing antenna and spacers to airplane. (5) Carefully part sealant from antenna and airplane skin. (6) Remove antenna and antenna spacers from airplane.
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WX1000+ Stormscope Components Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
Install Stormscope Antenna (Refer to Figure 401) (1) Clean antenna and airplane mounting surface using an approved solvent. (2) Place antenna and spacers in position. Secure antenna and spacers to airplane with screws. (3) Fillet seal around antenna base, spacers, and doubler with Type I, Class B sealant. Refer to Chapter 20, Fuel, Weather, Pressure and High-Temperature Sealing - Maintenance Practices. (4) Fill in top of antenna screws with Class B Type 1 sealant. (5) From cabin, connect coax connector to antenna. (6) Install hard shell headliner panel using fasteners. (7) Engage STM SCP circuit breaker on left circuit breaker panel.
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MODEL 208 MAINTENANCE MANUAL BF GOODRICH SKYWATCH (SKY497) TRAFFIC ADVISORY SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
The BF Goodrich Avionics Systems SKYWATCH (SKY497) is an airborne traffic advisory system that advises the flight crew where to look for airplanes posing a collision threat. SKYWATCH (SKY497) alerts the crew to nearby transponder equipped airplanes that may present a danger. NOTE:
Display unit is also utilized by the BF Goodrich WX1000+/E Stormscope system and is referred to as an WX-1000/SKY497 display.
NOTE:
The BF Goodrich WX1000+ indicates an enhanced version of the Stormscope. The added E indicates the WX- 1000+ Stormscope with NAVAID.
B.
Traffic information is graphically displayed on the display unit. A range of either 2 or 6 nautical miles may be selected.
C.
SKYWATCH (SKY497) components include a TRC497 Transmitter Receiver Computer (TRC), a TRC mounting tray, a shared WX-1000/SKY497 display and a NY-164 antenna located on the forward upper fuselage. A selector switch mounted on the instrument panel next to the indicator allows the crew to choose between either STORMSCOPE OR SKYWATCH operation.
Operation A.
With SKYWATCH mode selected, the system can track up to 30 intruder airplanes with only the 8 highest priority showing on the display. Horizontal surveillance tracking radius is 11 nautical miles maximum with a relative altitude tracking range of plus or minus 10,000 feet maximum. An ON/OFF control and display brightness is controlled through the OFF/BRT switch located on the WX- 1000/ SKY497 display.
B.
SKYWATCH/STORMSCOPE mode switch is required if the WX-1000+/E Stormscope Weather Mapping System is installed. This switch permits the flight crew to switch the display between the SKY497 and WX- 1000+/E. If a traffic advisory is detected in the Stormscope display mode, the display will switch to the SKYWATCH mode automatically.
C.
The WX-1000+/Emaintenance switch is located on the aft tailcone avionics shelf. This switch, when moved to SKYWATCH position, allows the SKYWATCH system to be powered up if the WX-1000+/E has been removed. If the SKYWATCH TRC is removed, then a shorting plug is used to allow the WX-1000+/E to continue operating.
D.
The NY-164 directional antenna is a teardrop shaped antenna. Connections are made through two TNC and one BNC connector. The antenna is sealed by using an O-ring to ensure a tight seal between antenna and skin of the airplane.
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MODEL 208 MAINTENANCE MANUAL BF GOODRICH SKYWATCH (SKY497) TRAFFIC ADVISORY SYSTEM - REMOVAL/INSTALLATION 1.
2.
General A.
This section describes the removal and installation procedures for the BF Goodrich SKYWATCH (SKY497) Traffic Advisory System components.
B.
The system consists of a transmitter/receiver/computer, a maintenance switch, display, and antenna. (1) The TRC497 Transmitter/Receiver/Computer (TRC), and maintenance switch is located on the aft avionics shelf in the tailcone. (2) The NY-164 antenna is mounted on the forward upper fuselage at FS 161.25 and LBL 5.00. (3) The WX-1000/SKY497 display is located on the instrument panel.
TRC497 Transmitter/Receiver/Computer Removal/Installation A.
Remove TRC497 (Refer to Figure 401) (1) Disengage circuit breaker SKY WATCH on left circuit breaker panel. (2) Locate the TRC497 on the aft avionics shelf. NOTE: (3) (4) (5) (6) (7)
B.
Disconnect interconnect cable from TRC497. Disconnect power cable from TRC497. Disconnect all coax antenna cables from TRC497. Loosen mounting tray hold- down knobs and remove TRC497 from mounting tray. Remove TRC497 from airplane.
Install TRC497 (Refer to Figure 401) (1) Position TRC497 into mounting tray on aft avionics shelf and push into place. Securely (2) Tighten mounting tray hold-down knobs. NOTE: (3) (4) (5) (6)
3.
4.
All connections are on the front of the TRC497.
All connections are on the front of the TRC497.
Connect all coax antenna cables to TRC497. Connect power cable to TRC497. Connect interconnect cable to TRC497. Engage circuit breaker SKY WATCH on left circuit breaker panel.
Stormscope/Skywatch Maintenance Switch Removal/Installation A.
Remove Maintenance Switch (Refer to Figure 401) (1) Disengage circuit breakers STM SCP and SKY WATCH on the lower left circuit breaker panel. (2) Locate maintenance switch on aft avionics shelf. (3) Remove hex nut holding switch in place. (4) Remove switch from angle bracket and remove screws attaching wires to switch terminals. Note color and position of each wire. (5) Remove switch from airplane.
B.
Install Maintenance Switch (Refer to Figure 401) (1) Attach wires to switch terminals in same order they were removed. (2) Position switch in angle bracket. (3) Install hex nut and tighten. (4) Engage circuit breakers STM SCP and SKY WATCH on the lower left circuit breaker panel.
WX-1000/SKY497 Display Removal/Installation A.
Remove WX-1000/SKY497 Display (Refer to Figure 401) (1) Disengage circuit breakers STM SCP and SKY WATCH on lower left circuit breaker panel. (2) On instrument panel, remove screws from WX-1000/SKY497 display. (3) Move display outward from instrument panel and disconnect electrical connector. (4) Remove display from airplane.
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Skywatch (SKY497) Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
5.
Install WX-1000/SKY497 Display(Refer to Figure 401) (1) Position display at instrument panel and connect electrical connector. (2) Position display into instrument panel and install screws. (3) Engage circuit breakers STM SCP and SKY WATCH on the lower left circuit breaker panel.
Skywatch NY-164 Antenna Removal/Installation A.
Remove Skywatch Antenna (Refer to Figure 401) (1) Disengage circuit breaker STM SCP and SKY WATCH on left circuit breaker panel. (2) Remove hard shell headliner panel inside the cabin just below the Skywatch antenna located at FS 161.25 and LBL 5.00 by removing fasteners. (3) Allow panel to drop enough to disconnect coax connectors. (4) Identify and disconnect all coax connectors. (5) From outside, remove screws securing antenna to airplane. (6) Carefully pry antenna from airplane skin. Clean old sealant from antenna and airplane skin. (7) Remove antenna from airplane.
B.
Install Skywatch Antenna (Refer to Figure 401) (1) Clean antenna and airplane mounting surface using an approved solvent. Refer to Chapter 20, General Solvents/Cleaners - Maintenance Practices. (2) Position antenna and secure with screws. (3) Fillet seal around antenna base with Type I, Class B sealant. Refer to Chapter 20, Fuel, Weather, Pressure and High-Temperature Sealing - Maintenance Practices. (4) From inside cabin, connect all coax connectors. (5) Install hard shell headliner panel using fasteners. (6) Engage circuit breaker STM SCP and SKY WATCH on left circuit breaker panel.
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MODEL 208 MAINTENANCE MANUAL GARMIN GWX-68 WEATHER RADAR SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
Description A.
3.
The Garmin GWX-68 is a 6.5 kilowatt weather radar installed to help the pilot monitor areas of precipitation in the ßight path of the airplane. This section gives a description and operation of the Garmin GWX-68 Weather Radar System.
Garmin GWX-68 Weather Radar. (1) The weather radar is installed behind a radome on the right wing. The weather radar assembly includes an antenna, receiver, and transmitter in one assembly. (2) The weather radar is adjustable to many scan proÞles (20 to 90 degrees) and gives a highdeÞnition target display. The system also includes a vertical scan function to help the pilot look at thunderstorm tops, gradients and cell buildup activity at many altitudes. The GWX-68 Weather Radar has extended Sensitivity Time Control (or STC) logic that digitally integrates weather attenuation and distance compensation. This component prevents a display change in the size of severe weather cells as distance to the cells changes. (3) Garmin’s WATCH™ (Weather Attenuated Color Highlight) feature identiÞes shadow effects of short-range cell activity. This system identiÞes the areas behind intense weather cells, or large areas of less intense precipitation, where the radar display can be less accurate. (4) The GWX-68 Weather Radar also has an Automatic Target Alert feature that looks ahead for intense cell activity in the 80 to 320 nautical mile range. This component will give a warning, even if the pilot does not actively monitor the displays.
Operation A.
Garmin GWX-68 Weather Radar. (1) The weather radar on this airplane is a typical weather radar installation. It uses pulsed microwave signals, transmitted by the phased array antenna to look for reßections (echoes) of precipitation. The reßected signal is received by the same phased array antenna. Detection is a two-way process that needs 12.36 ms for a signal to travel out to the target and come back to the antenna. The center of the phased array antenna has a higher signal energy, which decreases toward the edge of the antenna. (2) The weather radar gives current precipitation data and is shown on the inset map of the pilot's and copilot's primary ßight display (PFD) and on the multifunction display (MFD). The system uses a four-color display to show intensity and location of precipitation. (3) The weather radar with the Garmin G1000 avionics system gives the operational features that follow. • Range modes of 2.5, 5, 10, 20, 40, 60, 80, 100, 120, 160, and 320 nautical miles • Vertical scan angle of 60 degrees • Horizontal scan mode (20 to 90 degrees) • Weather and ground mapping modes. (4) The weather radar communicates to the Garmin G1000 system through the HSDB bus and Garmin GDL-69A Datalink system. The weather radar also receives power from the same electrical connector in the radome. The weather radar receives power through the RADAR R/T circuit breaker on the avionics circuit breaker panel.
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MODEL 208 MAINTENANCE MANUAL GARMIN GWX-68 WEATHER RADAR SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives maintenance practices for the Garmin GWX-68 Weather Radar System. The GWX-68 is the only component of the weather radar system included in this section. For a general description of the GWX-68 weather radar, refer to Garmin GWX-68 Weather Radar - Description and Operation.
GWX-68 Weather Radar Removal/Installation A.
Remove the GWX-68 Weather Radar (Refer to Figure 201). (1) Remove electrical power from the airplane. (a) Disengage the RADAR R/T circuit breaker on the avionics circuit breaker panel. (2) Remove the radome to get access to the weather radar. (3) Disconnect the electrical connector from the weather radar. (4) Carefully hold the weather radar and remove the bolts that hold the weather radar on the bulkhead assembly. (5) Remove the weather radar from the bulkhead assembly.
B.
Install the GWX-68 Weather Radar (Refer to Figure 201). (1) Put the weather radar in position on the bulkhead assembly. (2) Install the bolts that hold the weather radar on the bulkhead assembly. (a) Make sure the bonding straps are installed under the bottom bolts. (3) Make sure that the weather radar is Type I electrically bonded to the bulkhead assembly. Refer to Chapter 20, Electrical Bonding - Maintenance Practices. (4) Connect the electrical connector to the weather radar. (5) Engage the RADAR R/T circuit breaker on the copilot's circuit breaker panel. (6) Install the radome. (7) If a new weather radar is installed, do the operational check. Refer to Garmin G1000 Line Maintenance Manual. (8) Do the Weather Radar Test. Refer to Garmin GWX-68 Weather Radar - Adjustment/Test.
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GWX-68 Weather Radar Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GARMIN GWX-68 WEATHER RADAR SYSTEM - ADJUSTMENT/TEST 1.
General A.
This section gives adjustment and test procedures for the weather radar system installed on the airplane. The weather radar control indications and mode data is shown on the Multifunction Display (MFD) and primary flight displays (PFD's). It is not permitted to do these weather radar test procedures inside a hangar.
WARNING: Be very careful when you use the WEATHER or GROUND mode on the radar during tests. When the weather radar is operated in the typical flight mode (WEATHER), or in GROUND it will radiate dangerous RF energy. 2.
Weather Radar Test A.
Do the Weather Radar Test. (1) Connect external power to the airplane. (2) Put the EXTERNAL POWER switch in the BUS position. (3) Put the BATTERY switch in the ON position. (4) Put the AVIONICS 1 and AVIONICS 2 switches in the ON position. (5) Use the outer FMS knob to select the MAP page group (highlighted). (6) Use the inner FMS knob to select the MAP-WEATHER RADAR page. Immediately make sure that the weather radar is in the OFF mode. (a) If the weather radar is in the WEATHER or GROUND mode, immediately push the MODE softkey, then the OFF softkey. (7) Make sure that the radar is in the OFF mode by the indications that follow: (a) OFF (cyan) shows in the upper left corner of the radar map on the MFD (b) OFF (white) shows in the center of the radar map on the MFD (c) The OFF softkey shows as a black letters on a gray background. (8) Set the weather radar mode to STANDBY. (a) Push the MODE softkey. (b) Push the STANDBY softkey. (9) Make sure that the radar is in the STANDBY mode by the indications that follow: (a) STANDBY (cyan) shows in the upper left corner of the radar map on the MFD (b) For about 60 seconds, WARM-UP (white) shows BELOW the center of the second range ring. A count down timer (white) shows below the WARM-UP indication. (c) At the end of the warm-up period, STANDBY (white) shows in place of the WARM-up indication. (d) The standby softkey shows as a black letters on a gray background. (10) Make sure that a white WARM-UP message is displayed for approximately 60 seconds below the center of the second range ring on the MFD. NOTE:
A white timer will be shown below the WARM-UP indication. At the end of the warm-up period, the WARM-UP indication will be replaced with the white STANDBY indication.
(11) Push the inner FMS knob to start a cursor that highlights the field to the right of the TILT indication. (12) Turn the outer FMS knob and make sure that the cursor does not move. (13) Make sure that the MODE softkey is available. (a) If the MODE softkey is not available, push the BACK softkey. (14) Put the system in test mode. (a) To enter test mode, push the 7 and 9 softkey on the MFD in the following order: • Softkey 7 • Softkey 9 • Softkey 9 • Softkey 7
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MODEL 208 MAINTENANCE MANUAL (15) Push the MODE softkey. (a) When asked if you are sure, select YES. NOTE:
The antenna will move to the horizontal and vertical stops, then start the azimuth scan. When the antenna reaches a stop, it will ratchet against the stop. This is known as synchronization. It will cause significant vibration and noise, and will stop in less than 10 seconds.
(16) Push the TEST softkey. (17) Push the BACK softkey. NOTE:
The antenna will move to the horizontal and vertical stops, then start the azimuth scan. When the antenna reaches a stop, it will ratchet against the stop. This is known as synchronization ratcheting. It will cause significant vibration and noise, and will stop in less than 10 seconds.
(18) Push the BRG softkey. NOTE:
A cyan bearing selector will come on the radar azimuth scan area, and a cursor will come on in the field to the left of the BEARING indication in the bottom right quadrant of the MFD display.
(19) Change the bearing angle to L 45° and push the VERTICAL softkey. (a) While the antenna sweeps up and down, make sure that the antenna does not touch the radome or structure. NOTE:
The antenna will continue its horizontal sweep until it reaches the left 45-degree azimuth and start a vertical sweep.
(20) Change the bearing angle to L 30° and push the VERTICAL softkey. (a) While the antenna sweeps up and down, make sure that the antenna does not touch the radome or structure. NOTE:
The antenna will continue its horizontal sweep until it reaches the left 30-degree azimuth and start a vertical sweep.
(21) Change the bearing angle to L 15° and push the VERTICAL softkey. (a) While the antenna sweeps up and down, make sure that the antenna does not touch the radome or structure. NOTE:
The antenna will continue its horizontal sweep until it reaches the left 15-degree azimuth and start a vertical sweep.
(22) Change the bearing angle to L 0° and push the VERTICAL softkey. (a) While the antenna sweeps up and down, make sure that the antenna does not touch the radome or structure. NOTE:
The antenna will start a vertical sweep.
(23) Change the bearing angle to R 15° and push the VERTICAL softkey. (a) While the antenna sweeps up and down, make sure that the antenna does not touch the radome or structure. NOTE:
The antenna will continue its horizontal sweep until it reaches the right 15-degree azimuth and start a vertical sweep.
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MODEL 208 MAINTENANCE MANUAL (24) Change the bearing angle to R 30° and push the VERTICAL softkey. (a) While the antenna sweeps up and down, make sure that the antenna does not touch the radome or structure. NOTE:
The antenna will continue its horizontal sweep until it reaches the right 30-degree azimuth and start a vertical sweep.
(25) Change the bearing angle to R 45° and push the VERTICAL softkey. (a) While the antenna sweeps up and down, make sure that the antenna does not touch the radome or structure. NOTE:
The antenna will continue its horizontal sweep until it reaches the right 45-degree azimuth and start a vertical sweep.
(26) Push the HORIZON softkey. NOTE:
This will bring the radar back to the azimuth scan.
(27) Push the inner FMS knob to start a cursor that highlights the field to the right of the TILT indication. (28) Push the TILT softkey. NOTE:
A cursor will come on in the field to the left of the TILT indication, in the bottom right quadrant of the MFD display.
(29) Change the tilt angle to UP 15.00°. (a) While the radar sweeps left and right, make sure that the antenna does not touch the radome and structure. NOTE:
The antenna will move to the horizontal and vertical stops, then start the azimuth scan. When the antenna reaches a stop, it will ratchet against the stop. This is known as synchronization. It will cause significant vibration and noise, and will stop in less than 10 seconds.
(30) Change the tilt angle to UP 10.00°. (a) While the radar sweeps left and right, make sure that the antenna does not touch the radome and structure. (31) Change the tilt angle to UP 5.00°. (a) While the radar sweeps left and right, make sure that the antenna does not touch the radome and structure. (32) Change the tilt angle to UP 0.00°. (a) While the radar sweeps left and right, make sure that the antenna does not touch the radome and structure. (33) Change the tilt angle to DOWN 5.00°. (a) While the radar sweeps left and right, make sure that the antenna does not touch the radome and structure. (34) Change the tilt angle to DOWN 10.00°. (a) While the radar sweeps left and right, make sure that the antenna does not touch the radome and structure. (35) Change the tilt angle to DOWN 15.00°. (a) While the radar sweeps left and right, make sure that the antenna does not touch the radome and structure. (36) Disengage the RADAR R/T circuit breaker on the avionics circuit breaker panel. (a) Make sure the RADAR FAILED message appears on the weather radar page on the MFD. (37) Engage the RADAR R/T circuit breaker on the avionics circuit breaker panel. (38) Put the AVIONICS 1 and AVIONICS 2 switches in the OFF position. (39) Put the BATTERY switch in the OFF position. (40) Put the EXTERNAL POWER switch in the OFF position.
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CAUTION: Disconnect external power when not in use to prevent an accidental discharge of the battery. (41) Disconnect external power from the airplane.
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MODEL 208 MAINTENANCE MANUAL DEPENDENT POSITION DETERMINING - DESCRIPTION AND OPERATION 1.
Description A.
The dependent position determining systems include navigation receiver(s), DME, ADF(s), transponder(s), and area navigation and course indicators.
B.
Sperry Dependent Position Determining Equipment (Model 208 Only). (1) Sperry standard equipment consists of a 300 Nav/Com(s) (RT-385A), 300 ADF (R-546E), SDM77A DME (RT-377A), and 400 Transponder (RT-459A). (2) Sperry optional equipment consists of a 400 Nav/Com(s) (RT-485B), 400 ADF (R-446A), 400 RNAV (RN-479A), and RMl (IN-404A). NOTE:
C.
Refer to 208 Avionic Installations Service/Parts Manual for installation and electrical wiring information on Sperry equipment listed in items (1) and (2) above. Refer to List of Publications in front of this manual for appropriate Sperry Maintenance Manual, covering equipment listed in items (1) and (2) above.
King Dependent Position Determining Equipment (1) The King exchange equipment consists of a KX-165 Nav/Com with KI-206 Indicator, KR-87 ADF(s), KN- 63 or KN-63(01) DME, KI-229 RMI, (Model 208 and 208B Passenger), KNI-582 RMI (Model 208 and 208B Passenger), KT-79 Transponder with KEA-130 Encoder, and KNS81 Integrated Nav System. NOTE:
Refer to List of Publications in front of this manual for appropriate King Maintenance Manual covering the above equipment. Refer to 208 Avionic Installations Service/ Parts Manual for wiring diagrams.
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MODEL 208 MAINTENANCE MANUAL DEPENDENT POSITION DETERMINING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the dependent position determining components in a serviceable condition.
Task 34-50-00-720 2.
Transponder Functional Check
CAUTION: When you set a transponder code, make sure that you do not radiate any of the following codes unintentionally: Code 7777 - Military Intercept Code, Code 7500 - Hijack Code, Code 7600 - VHF Com Receiver Failure Code, or Code 7700 - Emergency Code. NOTE:
It is necessary to calibrate the transponders at intervals in accordance with 14 CFR Part 91.413. You must calibrate the transponders in an approved altimeter repair facility by approved personnel with approved equipment and documentation established in 14 CFR Part 91.413. It is necessary to do an altitude reporting check after the transponder is certified and installed in the airplane.
A.
General (1) This task gives the procedures to do a functional check of the transponder.
B.
Special Tools NOTE: (1) (2) (3) (4)
Equivalent tools and equipment can be used.
ATC-601 Ramp Test Set Air Data Test set - LAVERSAB Model 65000 or Barfield 101-00184 Pitot Static Test Adaptor - (Nav-Aids Ltd. PS4769) External Electrical Power Unit
C.
Access (1) None
D.
ATC-601 Ramp Test Set Setup. (1) Connect the coaxial cable supplied with the ATC-601 Ramp Test Set (antenna connector) to the flat plate antenna, which is also supplied with the test set. (2) Put the antenna in a position so that no large metal objects are between the tester antenna and the aircraft transponder antenna. (3) Apply power to the ATC-601 Ramp Test Set. (a) Push the Setup button. (4) Measure the horizontal distance in feet along the shop floor or ramp between the antennas. (5) Use the Select and Slew buttons to put that value in the Range field. (6) Measure the vertical distance between the aircraft transponder antenna and the ATC-601 Ramp Test Set antenna. (7) Use the Select and Slew buttons to put that value in the Height field. (8) Use the gains from the ATC-601 Ramp Test Set flat plate antenna label and put the values in the Gain 1030= and Gain 1090= fields. (9) Use the value from the sleeve on the coax cable and put in the value in the Loss= field. (10) Make the bottom antenna the selection for the test. (a) If an upper transponder antenna is installed, it is necessary to do the test again for that antenna. This test will apply only to the Mode S configured airplane.
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MODEL 208 MAINTENANCE MANUAL E.
Do the Transponder Functional Check. NOTE:
This functional check is applicable to all Model 208 Mode A , C, & S transponders to include Honeywell KT-70 series transponder systems, Garmin GTX 327, 330, & 33 series transponder systems, and the Sperry RT 459 transponder system.
NOTE:
If airplane is Mode A & C only, it will have a failure for all Mode S functions. The display will show Modes that are satisfactory. Disregard Mode S failures.
(1) (2) (3) (4) (5)
Apply external electrical power to the airplane. Set the external power switch to BUS. Set the battery switch to ON. Set the avionics switches 1 and 2 to ON. Make sure that the transponder circuit breakers XPDR 1 and XPDR 2 on the lower left circuit breaker panel are engaged. (6) If applicable, make sure that the XPDR 1 or XPDR 2 switch is set to the correct position for the test. (7) Tune and make sure that the transponder code is 0600 on the transponder to be tested. (8) Select the transponder being tested to the ON position and let it warm-up for fifteen minutes. (9) Select the ALT position on the transponder being tested. (10) Push the Auto Test button on the ATC-601 Ramp Test Set. (a) Push the Run/Stop button to start the test. 1 Make sure that the transponder passes all of the tests. 2 During the test, make sure that the reply annunciator on the front panel shows R. (11) Use the Select button on the ATC-601 to display the Squitter test selection. (a) Make sure that the United States tail number or foreign registration code is shown correctly. (12) Push the Select button on the ATC-601 Ramp Test Set.
CAUTION: Make sure that the ATC-601 Ramp Test Set altitude is shown in 100's to agree with the transponder. (a) (b)
(13) (14) (15)
(16) (17)
Make sure that the ALT= field agrees with the pilot’s encoding altimeter. Make sure that the CODE= field agrees with the squawk code on the face of the transponder. (c) The transmitting power specification (Spec) is from 125 watts minimum to 500 watts maximum. (d) The receiver MTL Spec for Mode S transponders is -74 dbm +3 or -3 dbm. 1 The receiver MTL Spec for all other transponders is -73 dbm +4 or -4 dbm. (e) The frequency Spec for Mode S transponders is 1090 MHz +1 or -1 MHz. 1 The frequency MTL Spec for all other transponders is 1090 MHz +3 or -3 MHz. Connect the air data tester to the airplane pitot/static system. Adjust the Baro on the encoding altimeter to show a barometric setting of 29.92 inches of Mercury. Use the air data test set and increase the altitude to 25,000 feet. (a) Continuously make sure that the altitude displays on the encoding altimeter, transponder, and the ATC-601 Ramp Test Set show the same altitude, +125 or -125 feet, as the altitude increases to 25,000 feet. Use the air data test set and slowly decrease the altitude to the field elevation. Push the XPDR IDENT switch on the pilot’s control wheel. (a) Make sure that the R (Reply) annunciator shows to the right of the digits of the transponder code in the window. NOTE:
The IDT annunciator comes on for 18 seconds, +1 or - 1 second, after the start of an IDENT.
(18) If applicable, do this test again for the upper antenna. (19) If applicable, do this test again for the second transponder system.
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MODEL 208 MAINTENANCE MANUAL (20) (21) (22) (23) (24) (25) (26) End of task
After the tests are complete, select the OFF position on the transponder. Set the avionics switches 1 and 2 to OFF. Set the battery switch to OFF. Set the external power switch to OFF. Remove the air data tester from the airplane. Remove the external electrical power from the airplane. Remove the power from the ATC-601 Ramp Test Set.
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MODEL 208 MAINTENANCE MANUAL KN-53 VHF NAVIGATION SYSTEM - DESCRIPTION AND OPERATION 1.
2.
General A.
The KN-53 is an instrument panel mounted VHF VOR/LOC Transceiver with glideslope capabilities. This system provides the pilot with means to navigate using an analog indicator. The system is FAA approved for both navigation and landing operations.
B.
The KN-53 system consists of a panel mounted transceiver with digital style readout, and a panel mounted indicator. The system utilizes existing antennas. Other components such localizer, glideslope, marker beacon and DME may be interfaced with the KN-53 system.
Description A.
3.
VOR (1) The VOR/LOC transceiver operates 200 channels in the frequency range of 108.00 to 117.95 MHz. The localizer operates on 40 channels between the frequency range of 108.10 to 111.95 MHz. When selecting a localizer frequency/channel, a corresponding glideslope frequency/ channel is automatically selected. (2) VOR/LOC signals are received by a pair of existing navigation antennas. Coaxial cable connect the antennas to the KN-53 and run the length of the fuselage.
Operation A.
The KN-53 transceiver contains all operating controls and displays on the front panel. A brief description of the controls and their functions follows. (Refer to Figure 1)
B.
Frequency select knobs - Clockwise rotation increases and counterclockwise rotation decreases the particular displays in the various modes of operation. (1) Frequency Select Knobs provide frequency selection from 108.00 to 117.95 MHz (VOR/LOC) and 329.15 to 335.00 MHz (Glideslope) in either 1 MHz or 50 kHz increments. (2) The larger knob (outer concentric) will increase or decrease the MHz portion of the display in 1 MHz steps with rollover at each band range. (3) The small tuning knob (inner concentric) will increase or decrease the display in 50 kHz increments.
C.
ON/OFF/VOLUME - Clockwise rotation of the ON/OFF/VOLUME control knob applies power to the system. Pulling the knob out of detent to the "ID" position allows the Ident Tone signal plus NAV voice information to be heard. Pushing the knob into detent allows NAV voice information to be heard with no Ident Tone and also adjusts the volume level. Rotating the knob fully counterclockwise removes power from the NAV transceiver.
D.
Frequency Transfer button - Depressing the Frequency Transfer Button transfers the selected frequency displayed in the "STBY" window to the "USE" window, and the displayed "USE" frequency to the "STBY" window.
E.
Photocell - The brightness for the displays is controlled automatically by the photocell, which reacts to ambient light.
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KN-53 Navigation Transceiver Figure 1 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL KN-53 VHF NAVIGATION SYSTEM - REMOVAL/INSTALLATION 1.
2.
General A.
This section provides removal and installation procedures for components of the KN-53 VHF Navigation system.
B.
System components include a KN-53 VHF transceiver and KI-204 indicator.
C.
For system maintenance practices not covered in this chapter, refer to manufacturer’s maintenance manuals.
KN-53 Transceiver Removal/Installation A.
Remove KN-53 Transceiver (Refer to Figure 401) NOTE:
(1) (2) (3) (4) B.
The KN-53 transceiver is mounted in the instrument panel. Removal/Installation is accomplished by releasing the unit and sliding from/into the rack. A 3/32 inch hex wrench is used to lock/unlock the unit to/from its mounting tray.
Disengage NAV 2 circuit breaker on the left side circuit breaker panel. Insert the hex wrench into the small hole located on the right side of the front panel of the transceiver. Rotate the hex wrench counterclockwise until the locking mechanism becomes noticeably loose. Continue counterclockwise rotation until mechanism just begins to become snug again. Do not rotate locking mechanism until it binds and becomes tight. Pull the transceiver unit outward from mounting tray by pulling on the sides of the front panel. Do not pull unit from tray using knobs. Remove transceiver from tray.
Install KN-53 Transceiver. (Refer to Figure 401) (1) Ensure front lug on mounting tray locking mechanism is in up position. Insert transceiver into mounting tray and slide into tray completely. (a) Insert hex wrench into hole in transceiver front panel to engage set screw. Rotate hex wrench clockwise until snug. Remove hex wrench, then pull gently on front panel to ensure unit is properly secured in mounting tray.
CAUTION: Hand tighten only. Avoid over torquing or damage may result. (2) 3.
Engage NAV 2 circuit breaker on left circuit breaker panel.
KI-204 Nav Indicator Removal/Installation A.
Remove KI-204 Nav Indicator. (Refer to Figure 401 ) (1) Remove electrical power from the airplane. Disengage NAV 2 circuit breaker from the left hand circuit breaker panel. (2) Remove screws and adapter plate from Nav Indicator. (3) Withdraw indicator from instrument panel and disconnect electrical connector. (4) Remove indicator from airplane.
B.
Install KI-204 Nav Indicator. (Refer to Figure 401) (1) Connect electrical connector to indicator. (2) Place indicator in instrument panel and secure with adapter plate and screws. (3) Engage NAV 2 circuit breaker on the left circuit breaker panel.
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KN-53 Navigation System Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL KING KT-70 TRANSPONDER - REMOVAL/INSTALLATION 1.
2.
General A.
This section covers the removal and installation of the King KT-70 transponder.
B.
The transponder is located in the center console.
KT-70 Transponder Removal/Installation A.
Remove Transponder. (1) Disengage circuit breaker XPDR 1 on lower left circuit breaker panel. (2) Loosen set screw securing transponder. NOTE: (3) (4)
B.
Transponder set screw is accessed from the transponder front panel using a long hex wrench.
Lift transponder from rack. Remove transponder from airplane.
Install Transponder. (1) Put the transponder in the mounting rack. Engage the set screw in the front panel with a long hex wrench. (2) Fasten the transponder in rack.
CAUTION: Hand tighten set screw only. Avoid over torquing or damage may result. (3)
Engage circuit breaker XPDR 1 on left circuit breaker panel.
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MODEL 208 MAINTENANCE MANUAL BENDIX/KING KT-73 MODE-S TRANSPONDER - MAINTENANCE PRACTICES 1.
2.
3.
General A.
This section gives removal and installation procedures for the KT-73 Mode-S transponder.
B.
The KT-73 transponder is in the center radio panel.
KT-73 Transponder Removal/Installation A.
Remove the KT-73 transponder (Refer to Figure 201). (1) Disengage the XPDR 1 circuit beaker on the left circuit breaker panel. (2) Put a 3/32-inch hex wrench through the front panel of the transponder to get access to the lock mechanism. (3) Turn the hex wrench counterclockwise to loosen the transponder in the mounting rack. (4) Remove the hex wrench. (5) Pull the transponder from the mounting rack.
B.
Install the KT-73 Transponder (Refer to Figure 201). (1) Put the KT-73 transponder in position in the mounting rack. (2) Use the 3/32-inch hex wrench to lock the transponder. (3) Turn the hex wrench clockwise to lock the transponder in position. (4) Remove the hex wrench. (5) Engage the XPDR 1 circuit breaker on the left circuit breaker panel. (6) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check.
Transponder Antenna Without Cargo Pod Removal/Installation NOTE:
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. For the Model 208, the transponder antennas are located on the bottom of the airplane at FS 237.00, RBL 12.25 for transponder antenna 1, and FS 237.00, LBL 12.25 for transponder antenna 2. For the Model 208B, the transponder antennas are located on the bottom of the airplane at FS 285.00, RBL 12.25 for transponder antenna 1, and FS 285.00, LBL 12.25 for transponder antenna 2.
A.
Remove the Transponder Antenna (Refer to Figure 202). (1) Disengage the XPDR 1 transponder circuit breaker, on the lower left circuit breaker panel. (2) Remove the screws from the transponder antenna. (3) Carefully pull the transponder antenna away from the skin. (4) Disconnect the coax connector from the transponder antenna. (5) Remove the transponder antenna from the airplane.
B.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position near the skin. (3) Connect the coax connector to the transponder antenna. (4) Install the screws in the transponder antenna and attach the transponder antenna to the skin. (5) Do a check of the electrical bonding. Refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage XPDR 1 transponder circuit breaker on the lower left circuit breaker panel. (7) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices.
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Bendix/King KT-73 Transponder Installation Figure 201 (Sheet 1)
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Transponder Antenna without Cargo Pod Figure 202 (Sheet 1)
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4.
Transponder Antenna With Cargo Pod Removal/Installation NOTE:
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. For the Model 208 and the Model 208B, the transponder antennas are located on the bottom of the cargo pod at FS 154.30, LBL 12.25 for transponder antenna 1 and FS 154.30, RBL 12.25 for transponder antenna 2.
A.
Remove the Transponder Antenna (Refer to Figure 203). (1) Disengage XPDR 1 transponder circuit breaker on the lower left circuit breaker panel. (2) Open the cargo pod compartments to get access to the transponder antenna. (3) Disconnect the coax connector from the transponder antenna. (4) Remove the screws, washers, and nuts from the transponder antenna. (5) Carefully pull the transponder antenna away from the cargo pod. (6) Remove the transponder antenna from the airplane.
B.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position on the cargo pod. (3) Install the screws, washers, and nuts in the transponder antenna. (4) Connect the coax connector to the transponder antenna. (5) Do a check of the electrical bonding. Refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage XPDR 1 transponder circuit breaker on the lower left circuit breaker panel. (7) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (9) Close the cargo pod compartments.
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Transponder Antenna with Cargo Pod Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL BENDIX/KING KT-73 MODE-S TRANSPONDER - ADJUSTMENT/TEST 1.
2.
General A.
This section gives information necessary to do a self test of the Bendix/King KT-73 Mode-S transponder.
B.
The KT-73 Mode-S transponder is in the center radio panel.
KT-73 Self Test A.
Do a Self-Test of the KT-73 Transponder. (1) Make sure that there is power to the KT-73 transponder. (2) Turn the selection knob to the TST position. (3) Make sure that all display segments come on for a minimum of four (4) seconds after the TST position is selected. (4) Make sure that a TEST OK message comes on after four (4) seconds. (5) If any other message comes on after four (4) seconds, refer to the Model 208 Illustrated Parts Catalog - List of Vendor Publications. (6) Set the KT-73 transponder to the SBY position.
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MODEL 208 MAINTENANCE MANUAL GARMIN GTX 327 TRANSPONDER - MAINTENANCE PRACTICES 1.
2.
General A.
This section includes the removal and installation of the Garmin 327 transponders and associated antennas.
B.
The Garmin GTX 327, Mode A/Mode C transponders are located in the center radio panel.
Garmin Transponder Removal/Installation NOTE:
The removal and installation procedures for the transponders are typical.
A.
Remove Garmin Transponder (Refer to Figure 201). (1) Disengage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (2) Insert a 3/32 hex wrench through the front panel of the transponder and engage the lock mechanism. (3) Turn the lock mechanism counterclockwise to loosen and unlock the transponder from the mounting rack. (4) Remove the hex wrench from the transponder. (5) Pull the transponder from the mounting rack. (6) Remove the transponder from airplane.
B.
Install Garmin Transponder (Refer to Figure 201). (1) Put the transponder in position in the mounting rack. (2) Insert a 3/32 hex wrench through the front panel of the transponder and engage the lock mechanism. (3) Turn the lock mechanism counterclockwise and move the transponder forward into the mounting rack until it stops. (4) Turn the lock mechanism clockwise until the transponder bezel is flush with the radio panel. NOTE: (5) (6) (7) (8)
3.
This will lock the transponder in the mounting rack and engage the connectors.
Remove the hex wrench from the transponder. Engage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. Make sure that the transponder configuration agrees with the instructions in the GTX 327 Configuration Procedure drawing, which is supplied with the airplane. Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check.
Transponder Antenna without Cargo Pod Removal/Installation NOTE:
A.
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. For the Model 208 the transponder antennas are located on the bottom of the airplane at FS 237.00, RBL 12.25 for transponder antenna 1, and FS 237.00, LBL 12.25 for transponder antenna 2. For the Model 208B the transponder antennas are located on the bottom of the airplane at FS 285.00, RBL 12.25 for transponder antenna 1, and FS 285.00, LBL 12.25 for transponder antenna 2.
Remove the Transponder Antenna (Refer to Figure 202). (1) Disengage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (2) Remove the screws from the transponder antenna. (3) Carefully pull the transponder antenna away from the skin. (4) Disconnect the coax connector from the transponder antenna. (5) Remove the transponder antenna from the airplane.
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Garmin Transponder Figure 201 (Sheet 1)
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Transponder Antenna without Cargo Pod Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
4.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position near the skin. (3) Connect the coax connector to the transponder antenna. (4) Install the screws in the transponder antenna and attach the transponder antenna to the skin. (5) Check the electrical bonding, refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (7) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices.
Transponder Antenna with Cargo Pod Removal/Installation NOTE:
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. For the Model 208 and the Model 208B the transponder antennas are located on the bottom of the cargo pod at FS 154.30, LBL 12.25 for transponder antenna 1 and FS 154.30, RBL 12.25 for transponder antenna 2.
A.
Remove the Transponder Antenna (Refer to Figure 203). (1) Disengage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (2) Open the cargo pod compartments to get access to the transponder antenna. (3) Disconnect the coax connector from the transponder antenna. (4) Remove the screws, washers, and nuts from the transponder antenna. (5) Carefully pull the transponder antenna away from the cargo pod. (6) Remove the transponder antenna from the airplane.
B.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position on the cargo pod. (3) Install the screws, washers, and nuts in the transponder antenna. (4) Connect the coax connector to the transponder antenna. (5) Check the electrical bonding, refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (7) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (9) Close the cargo pod compartments.
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Transponder Antenna with Cargo Pod Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GARMIN GTX 330 TRANSPONDER - MAINTENANCE PRACTICES 1.
2.
General A.
This section includes the removal and installation of the Garmin 330 transponders and associated antennas.
B.
The Garmin GTX 330, Mode A/Mode C/Mode S transponders are located in the center radio panel.
Garmin Transponder Removal/Installation NOTE:
The removal and installation procedures for the transponders are typical.
A.
Remove Garmin Transponder (Refer to Figure 201). (1) Disengage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (2) Insert a 3/32 hex wrench through the front panel of the transponder and engage the lock mechanism. (3) Turn the lock mechanism counterclockwise to loosen and unlock the transponder from the mounting rack. (4) Remove the hex wrench from the transponder. (5) Pull the transponder from the mounting rack. (6) Remove the transponder from airplane.
B.
Install Garmin Transponder (Refer to Figure 201). (1) Put the transponder in position in the mounting rack. (2) Insert a 3/32 hex wrench through the front panel of the transponder and engage the lock mechanism. (3) Turn the lock mechanism counterclockwise and move the transponder forward into the mounting rack until it stops. (4) Turn the lock mechanism clockwise until the transponder bezel is flush with the radio panel. NOTE: (5) (6) (7)
Remove the hex wrench from the transponder. Engage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. Make sure that the transponder configuration and the Mode S Code agree with the instructions in the GTX 330 Configuration Procedure drawing, which is supplied with the airplane. NOTE:
(8) 3.
This will lock the transponder in the mounting rack and engage the connectors.
For the foreign-registered airplanes, the Mode S Code, for the airplane in which the transponder is installed, must be known in hexadecimal format.
Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check.
Transponder Antenna without Cargo Pod Removal/Installation NOTE:
A.
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. For the Model 208 the transponder antennas are located on the bottom of the airplane at FS 237.00, RBL 12.25 for transponder antenna 1, and FS 237.00, LBL 12.25 for transponder antenna 2. For the Model 208B the transponder antennas are located on the bottom of the airplane at FS 285.00, RBL 12.25 for transponder antenna 1, and FS 285.00, LBL 12.25 for transponder antenna 2.
Remove the Transponder Antenna (Refer to Figure 202). (1) Disengage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (2) Remove the screws from the transponder antenna. (3) Carefully pull the transponder antenna away from the skin.
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Garmin Transponder Figure 201 (Sheet 1)
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Transponder Antenna without Cargo Pod Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (4) (5) B.
4.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position near the skin. (3) Connect the coax connector to the transponder antenna. (4) Install the screws in the transponder antenna and attach the transponder antenna to the skin. (5) Check the electrical bonding, refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (7) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices.
Transponder Antenna with Cargo Pod Removal/Installation NOTE:
5.
Disconnect the coax connector from the transponder antenna. Remove the transponder antenna from the airplane.
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. For the Model 208 and the Model 208B the transponder antennas are located on the bottom of the cargo pod at FS 154.30, LBL 12.25 for transponder antenna 1, and FS 154.30, RBL 12.25 for transponder antenna 2.
A.
Remove the Transponder Antenna (Refer to Figure 203). (1) Disengage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (2) Open the cargo pod compartments to get access to the transponder antenna. (3) Disconnect the coax connector from the transponder antenna. (4) Remove the screws, washers, and nuts from the transponder antenna. (5) Carefully pull the transponder antenna away from the cargo pod. (6) Remove the transponder antenna from the airplane.
B.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position on the cargo pod. (3) Install the screws, washers, and nuts in the transponder antenna. (4) Connect the coax connector to the transponder antenna. (5) Check the electrical bonding, refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage the transponder circuit breaker (XPDR 1, CBU02 or XPDR 2, CBU18) on the lower left circuit breaker panel. (7) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (9) Close the cargo pod compartments.
Transponder Antenna (Fairing Installation) Removal/Installation NOTE:
A.
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. For the Model 208 the standard transponder antenna is found at FS 154.67, LBL 11.50 , the optional second transponder antenna is found at FS 154.67, RBL 11.50.
Remove the Transponder Antenna (Refer to Figure 204). (1) Disengage the applicable circuit breaker for the transponder.
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Transponder Antenna with Cargo Pod Figure 203 (Sheet 1)
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Transponder Antenna (Fairing Installation) Figure 204 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) (6) B.
Remove the aft fairing. Refer to Chapter 30, TKS Fluid Tank - Maintenance Practices, Remove the Aft Fairing. Disconnect the coax connector from the applicable antenna. Remove the screws, and washers from the applicable antenna. Carefully pull the applicable antenna away from the fairing. Remove the applicable antenna from the airplane.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position on the fairing. (3) Install the screws and washers that attaches the transponder antenna to the fairing. (4) Connect the coax connector to the transponder antenna. (5) Check the electrical bonding, refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage the applicable circuit breaker for the antenna you installed. (7) Do the Transponder Functional Check. Refer to Dependent Position Determining - Inspection/ Check. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (9) Install the aft fairing. Refer to Chapter 30, TKS FLUID TANK - Maintenance Practices, Remove the Aft Fairing.
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MODEL 208 MAINTENANCE MANUAL GTX-33 TRANSPONDER - MAINTENANCE PRACTICES 1.
2.
3.
General A.
The maintenance practices give the removal and installation procedures for the transponder, which is installed on the avionics shelf behind the instrument panel.
B.
The GTX-33 Transponder is an integrated component with the Garmin G1000 avionics system. The transponder is operated and monitored through the use of the Garmin Display Unit (GDU).
GTX-33 Transponder Removal/Installation A.
Remove the Transponder (Refer to Figure 201). (1) Disconnect electrical power to the airplane. (2) Disengage the XPDR 1 circuit breaker. (3) Remove the MFD from the instrument panel. Refer to Chapter 34, Garmin Display Unit Maintenance Practices. (4) Remove the screw from the lock lever. (5) Lift the lock lever to release the transponder from the avionics rack. (6) Remove the transponder from the airplane.
B.
Install the Transponder (Refer to Figure 201). (1) Install the transponder in the avionics rack. (2) Lower the lock lever. (3) Install the screw in the lock lever. (4) Install the MFD. Refer to Chapter 34, Garmin Display Unit - Maintenance Practices. (5) Engage the XPDR 1 circuit breaker. (6) Connect electrical power to the airplane. (7) If a new unit is installed, load the software and conÞguration. Refer to the Garmin G1000 Line Maintenance Manual. (8) Do a check to make sure the transponder operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
GTX-33 Diversity Transponder Removal/Installation A.
Remove the Transponder (Refer to Figure 201). (1) Disconnect electrical power to the airplane. (2) Disengage the XPDR 2 circuit breaker. (3) Remove the MFD from the instrument panel. Refer to Chapter 34, Garmin Display Unit Maintenance Practices. (4) Remove the screw from the lock lever. (5) Lift the lock lever to release the transponder from the avionics rack. (6) Remove the transponder from the airplane.
B.
Install the Transponder (Refer to Figure 201). (1) Install the transponder in the avionics rack. (2) Lower the lock lever. (3) Install the screw in the lock lever. (4) Install the MFD. Refer to Chapter 34, Garmin Display Unit - Maintenance Practices. (5) Engage the XPDR 2 circuit breaker. (6) Connect electrical power to the airplane. (7) If a new unit is installed, load the software and conÞguration. Refer to the Garmin G1000 Line Maintenance Manual. (8) Do a check to make sure the transponder operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
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GTX-33 Transponder Installation Figure 201 (Sheet 1)
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4.
Transponder Antenna without Cargo Pod Removal/Installation NOTE:
5.
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. The transponder antennas are found on the bottom of the airplane at FS 124.00, RBL 9.40 for transponder antenna 1, and FS 124.00, LBL 9.40 for transponder antenna 2.
A.
Remove the Transponder Antenna (Refer to Figure 202). (1) Disengage the circuit breaker on the lower left circuit breaker panel. (2) Remove the screws from the transponder antenna. (3) Carefully pull the transponder antenna away from the skin. (4) Disconnect the coax connector from the transponder antenna. (5) Remove the transponder antenna from the airplane.
B.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position near the skin. (3) Connect the coax connector to the transponder antenna. (4) Install the screws in the transponder antenna and attach the transponder antenna to the skin. (5) Check the electrical bonding, refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage the circuit breaker on the lower left circuit breaker panel. (7) Do a check to make sure the transponder operates correctly. Refer to the Garmin G1000 Line Maintenance Manual. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices.
Transponder Antenna with Cargo Pod Removal/Installation NOTE:
The airplane can have one or two transponder antennas. The removal and installation procedures for the transponder antennas are typical. The transponder antennas are found on the bottom of the cargo pod at FS 154.30, LBL 11.53 for transponder antenna 1, and FS 154.30, RBL 11.53 for transponder antenna 2.
A.
Remove the Transponder Antenna (Refer to Figure 203). (1) Disengage the circuit breaker on the lower left circuit breaker panel. (2) Open the cargo pod compartments to get access to the transponder antenna. (3) Disconnect the coax connector from the transponder antenna. (4) Remove the screws, washers, and nuts from the transponder antenna. (5) Carefully pull the transponder antenna away from the cargo pod. (6) Remove the transponder antenna from the airplane.
B.
Install the Transponder Antenna (Refer to Figure 202). (1) Make sure that the mounting surface area and the transponder antenna are clean and free of old sealant and contamination. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (2) Put the transponder antenna in position on the cargo pod. (3) Install the screws, washers, and nuts in the transponder antenna. (4) Connect the coax connector to the transponder antenna. (5) Check the electrical bonding, refer to Chapter 20, Electrical Bonding - Maintenance Practices. (6) Engage the transponder circuit breaker on the lower left circuit breaker panel. (7) Do a check to make sure the transponder operates correctly. Refer to the Garmin G1000 Line Maintenance Manual. (8) Fillet seal the exterior mating surfaces with Type 1 sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices. (9) Close the cargo pod compartments.
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Transponder Antenna without Cargo Pod Figure 202 (Sheet 1)
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Transponder Antenna with Cargo Pod Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GDL-69A FIS - DESCRIPTION AND OPERATION 1.
General A.
The GDL-69A Flight Information System (FIS) is a remote-mounted component of the Garmin G1000 avionics system. The GDL-69A gives weather and FIS information to the pilot. The information is controlled and seen through the Multi-Function Display (MFD). Information is sent from the data link receiver to the MFD through the high-speed data bus ethernet data path. With a current subscription, XM satellite radio service is available with the GDL-69A. The signals that the data link receives from satellites give better coverage than land-based transmissions. The XM radio is tuned through the MFD. Analog audio is sent to the audio panel and shares the AUX music input with the external audio entertainment input. GDL-69A capabilities include: • Graphical NEXRAD Data (NEXRAD) • Graphical METAR Data (METAR) • Textual METAR Data • Textual Terminal Aerodrome Forecasts (TAF) • City Forecast Data • Graphical Wind Data (WIND) • Graphical Echo Tops (ECHO TOP) • Graphical Cloud Tops (CLD TOP) • Graphical Lightning Strikes (XM LTNG) • Graphical Storm Cell Movement (CELL MOV) • NEXRAD Radar Coverage (displayed with NEXRAD data) • SIGMETs/AIRMETs (SIG/AIR) • Surface Analysis with City Forecasts (SFC) • County Warnings (COUNTY) • Freezing Levels (FRX LVL) • Hurricane Track (CYCLONE) • Temporary Flight Restrictions (TFR).
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MODEL 208 MAINTENANCE MANUAL GDL-69A FIS - MAINTENANCE PRACTICES 1.
General A.
2.
The maintenance practices give the removal and the installation procedures for the GDL-69A XM Weather Data Link.
GDL-69A XM Weather Data Link Removal/Installation A.
Remove the Data Link (Refer to Figure 201). (1) Disconnect electrical power from the airplane. (a) Disengage the XM-DATA LINK circuit breaker. (2) Remove the MFD from the instrument panel. Refer to Chapter 34, Garmin Display Unit Maintenance Practices. (3) Remove the screw from the lock lever. (4) Lift the lock lever to release the data link from the avionics rack. (5) Remove the data link from the airplane.
B.
Install the Data Link (Refer to Figure 201). (1) Install the data link in the avionics rack. (2) Lower the lock lever. (3) Install the screw in the lock lever. (4) Install the MFD. Refer to Chapter 34, Garmin Display Unit - Maintenance Practices. (5) Engage the XM-DATA LINK circuit breaker. (6) Connect electrical power to the airplane. (7) If a new unit is installed, load the software and conÞguration. Refer to the Garmin G1000 Line Maintenance Manual. (8) Do a check to make sure the data link operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
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GDL-69A XM Weather Data Link Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GARMIN INTEGRATED AVIONICS UNIT (GIA 63) - MAINTENANCE PRACTICES 1.
GIA 63 Integrated Avionics Unit Removal/Installation A.
Remove the Integrated Avionics Unit (Refer to Figure 201). (1) Disconnect electrical power to the airplane. (2) Remove the pilot's or copilot's primary ßight display (PFD). Refer to Garmin Display Unit Maintenance Practices. (3) Remove the screw from the lock lever. (4) Lift the lock lever to release the IAU from the avionics rack. (5) Remove the IAU from the airplane.
B.
Install the Integrated Avionics Unit (Refer to Figure 201). (1) Install the IAU in the avionics rack. (2) Lower the lock lever. (3) Install the screw in the lock lever. (4) Install the pilot's or copilot's primary ßight display (PFD). Refer to Garmin Display Unit Maintenance Practices. (5) Connect electrical power to the airplane. (6) If a new unit is installed, load the software and conÞguration. Refer to the Garmin G1000 Line Maintenance Manual. (7) Do a check to make sure the transponder operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
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GIA 63 Integrated Avionics Unit Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL GARMIN DISPLAY UNIT (GDU) - MAINTENANCE PRACTICES 1.
General A.
2.
Troubleshooting A.
3.
The GDU 1040 has a 10 inch LCD display with 1024x768 resolution. The cockpit has three GDU 1040s. Two are conÞgured as a Primary Flight Display (PFD) and the other is conÞgured as the Multi-Function Display (MFD). The MFD shows navigation, engine, and airframe information. The PFD shows primary ßight information, in place of gyro systems. All GDU 1040s connect and show all functions of the G1000 system during ßight. The displays communicate with each other and the GIA 63 Integrated Avionics Units (IAU) through a High-Speed Data Bus (HSDB) Ethernet connection.
For troubleshooting procedures, refer to the Garmin G1000 Line Maintenance Manual.
Garmin Display Unit Removal/Installation
CAUTION: If possible, do not touch the lens. The GDU 1040 lens has a layer of antireßective material which is very sensitive to skin oils, waxes and abrasive cleaners. CAUTION: Do not use cleaners that contain ammonia. Ammonia will cause damage to the anti-reßective material. It is very important to clean the lens with a clean, lint-free cloth and an eyeglass lens cleaner that is speciÞed as safe for anti-reßective material. A.
Remove the Garmin Display Unit (GDU) (Refer to Figure 201). (1) Disengage the applicable Primary Function Display (PFD) or Multi-Function Display (MFD) circuit breaker for the GDU. (2) Turn the quick release fasteners 1/4 turn counterclockwise with a 3/32" hex drive tool. (3) Carefully pull the GDU from the instrument panel. (4) Disconnect the electrical connector from the GDU.
B.
Install the Garmin Display Unit (GDU) (Refer to Figure 201).
CAUTION: If possible, do not touch the lens. The GDU 1040 lens has a layer of anti-reßective material which is very sensitive to skin oils, waxes and abrasive cleaners. CAUTION: Do not use cleaners that contain ammonia. Ammonia will cause damage to the anti-reßective material. It is very important to clean the lens with a clean, lint-free cloth and an eyeglass lens cleaner that is speciÞed as safe for anti-reßective material. NOTE:
If a new unit is installed, it is necessary to load the software and conÞguration.
NOTE:
If the initial unit is installed in the initial location or in the opposite location, it is not necessary to load the software and conÞguration.
(1) (2) (3) (4)
Make sure the connector and connector pins have no damage. (a) Replace the connector or connector pins if applicable. Refer to The Wiring Diagram Manual and the Garmin G1000 Line Maintenance Manual. Connect the electrical connector to the GDU. Put the GDU in position ßush with the instrument panel. Make sure the locking-stud alignment marks are in the vertical position. NOTE:
Light forward pressure can be required to engage the quick release fasteners.
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Garmin Display Unit (GDU) Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (5) (6)
Turn the quick release fasteners 1/4 turn clockwise with a 3/32" hex drive tool. Make sure the GDU operates correctly. (a) If a new unit is installed, load the software and conÞguration. Refer to the Garmin G1000 Line Maintenance Manual. (b) Do a check to make sure the GDU operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
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MODEL 208 MAINTENANCE MANUAL GARMIN SYNTHETIC VISION TECHNOLOGY SYSTEM - DESCRIPTION AND OPERATION 1.
Scope A.
2.
Description A.
3.
This chapter gives data about the optional Garmin Synthetic Vision Technology System.
Garmin Synthetic Vision Technology System. (1) The Synthetic Vision System (SVT) gives the pilot a three-dimensional view of the adjacent terrain, obstacles and other airplane positions. This gives the pilot more situational awareness to these items to improve flight safety. (2) The SVT terrain display shows a forward-looking view of land contours, large water features, towers, and other obstacles. Obstacles show that are more than 200 feet above ground level (AGL) from the obstacle database. Surface features, such as roads, highways, railroad tracks, cities, and state boundaries, do not show even if those features show on the MFD map. The terrain display also includes a compass grid to help in direction relative to the terrain. The colors used to show the terrain elevation contours are equivalent to those found on a topographic map view. (3) The operation of the SVT is controlled with Primary Flight Display (PFD) softkeys. (4) The Terrain Awareness and Warning System (TAWS) is integrated within SVT to give visual and auditory alerts to indicate the presence of terrain and obstacle threats relevant to the projected flight path. Terrain alerts are displayed in red and yellow shades on the PFD. (5) The SVT visual additions that show on the PFD are as follows: • Pathways • Flight path marker • Horizon heading marks • Traffic display • Airport signs • Runway display • Terrain alerting • Obstacle alerting.
Operation A.
The SVT display shows a perspective view of the terrain ahead of the airplane up to 35° to the left and 35° to the right of the airplane heading. The display shows land contours, large areas of water, and all obstacles in the G1000 data base that are over two hundred feet. Other ground objects such as roads, railroad tracks and cities are not shown on the PFD display even if they do show on the MFD map. The G1000 Terrain or TAWS system generates an alert or warning and if the obstacle is in the FPM field of view, the obstacle is colored yellow or red, respectively. The G1000 database covers the terrain as follows: (1) The terrain database covers north 75° latitude to south 60° for all longitudes. (2) The airport database covers the United States, Canada, Mexico, Latin America, and South America. (3) The obstacle database covers the United States.
B.
The SVT shows a green circular barbed symbol Flight Path Marker (FPM) that represents the current path of the airplane relative to the terrain display. The FPM is shown when the ground speed exceeds thirty knots and the SVT is shown on the PFD. The FPM shows only the current path and will change with a change in airspeed, crosswind, and power setting.
C.
A pathway shows on the PFD if PATHWAY is enabled on the SVT menu and a defined navigation path is entered on the G1000. The pathway representation of the programmed flight path. The pathway shows for GPS and ILS navigation paths but not for heading legs, VOR, LOC, BC, and ADF navigation aids. The pathway is not to be used as the primary navigation path reference.
D.
The SVT utilizes the TAS to present traffic symbols on both PDF displays. If traffic that is within the SVT field of view is detected by the TAS, white diamond symbols with show on the two PDF displays. If the traffic is close enough to generate an alert the traffic symbols will show as solid yellow circles.
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MODEL 208 MAINTENANCE MANUAL E.
If APTSIGNS is enabled an airport signpost shows on the SVT when the airplane is within 15 miles of airports that are in the G1000 database. There is also a runway representation shown on the PDF. The runway helps the pilot visually acquire the runway.
F.
The SVT continually shows a white horizontal line that represents the true horizon The horizon line shows the terrain above and below the airplane altitude.
G.
For more operational data, refer to the Garmin G1000 Integrated Flight Deck Pilot’s Guide. Refer to the Model 208 Illustrated Parts Catalog, Introduction, Supplier Publication List.
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MODEL 208 MAINTENANCE MANUAL GARMIN SYNTHETIC VISION TECHNOLOGY SYSTEM - MAINTENANCE PRACTICES 1.
General A.
The Model 208 airplane has an optional Synthetic Vision Technology (SVT ) available for installation. The SVT is a visual addition to the G1000 Integrated Avionics System. SVT shows a forward display of the terrain immediately in front of the airplane. The field of view is 30 degrees to the left and 35 degrees to the right. SVT data shows on the Primary Flight Display (PFD), or on the Multifunction Display (MFD) in Reversionary Mode.
B.
Installation of SVT requires a Garmin supplied Secure Digital (SD) card, 010-00330-55 to unlock the SVT on two PFD G1000 systems. NOTE:
2.
Once used the SD card is serialized by the first airplane system unlock, you will not be able to unlock another airplane SVT with the same SD card. The unlock card should be kept with the airplane for future use.
Garmin Synthetic Vision Technology System Software Configuration NOTE:
Garmin 0767.05 or later version software is required for PATHWAY function operation.
NOTE:
Garmin Supplemental Data SD cards, 010-00330-43 (one per display) with the 9 arc/second Terrain database are required.
NOTE:
When new system software and configuration files are loaded onto a PFD/MFD, the SVT option must be unlocked again for that PFD/MFD.
NOTE:
When the airframe configuration file is loaded onto the Garmin G1000 system, the SVT option for the PFD and MFD must be unlocked again.
A.
Enable the Garmin Synthetic Vision Technology System. (1) Set the EXTERNAL POWER switch (SC006) on the circuit breaker switch panel to the OFF position. (2) Set the BATTERY switch (SC005) on the circuit breaker switch panel to the OFF position. (3) Set the AVIONICS 1 switch (SC016) on the circuit breaker switch panel to the OFF position. (4) Set the AVIONICS 2 switch (SC018) on the circuit breaker switch panel to the OFF position. (5) Disengage the PFD 1, PFD 2, and MFD circuit breakers on the avionics circuit breaker panel. (6) Remove the GDU supplemental database SD cards from the bottom slot at each of the three flight displays. NOTE:
(7) (8) (9) (10) (11) (12) (13) (14) (15) (16) (17)
If the SD cards you removed are not 010-00330-43 data cards, make sure that you have 010-00330-43 data cards ready for installation at the completion of the SVT installation. If new cards are used all the latest Garmin database cycles must be installed. Also if Jeppesen Chartview is installed, the latest update must be installed.
Connect external electrical power to the airplane. Set the EXTERNAL POWER switch to the ON position. Set the AVIONICS 1 switch to the ON position. Set the AVIONICS 2 switch to the ON position. Push and hold the ENT key on the PFD 2 and engage the PFD 2 circuit breaker on the avionics circuit breaker panel. When INITIALIZING SYSTEM shows on PFD 2, release the ENT key. Push and hold the ENT key on the MFD and engage the MFD circuit breaker on the avionics circuit breaker panel. When INITIALIZING SYSTEM shows on the MFD, release the ENT key. Install the 010-00330-55 SVT unlock card in the PFD 1 top card slot. Push and hold the ENT key on the PFD 1 and engage the PFD 1 circuit breaker on the avionics circuit breaker panel. When INITIALIZING SYSTEM shows on PFD 1, release the ENT key.
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MODEL 208 MAINTENANCE MANUAL (18) Turn the small FMS knob on PFD 1 to go to the SYSTEM UPLOAD page. (19) Push the small FMS knob to activate the cursor. (20) Turn the small FMS knob on PFD 1 to highlight the CONFIGURATION FILE in the AIRFRAME field. NOTE:
Do not use the DISABLE SVT function. The pilot controls SVT by the soft keys on the PFDs.
(21) Push the PFD 1 ENT key. (22) Turn the small FMS knob on PFD 1 to expand the drop down menu. (23) Turn the small FMS knob on PFD 1 to highlight the ENABLE SVT DUAL PFD in the FILE field and push the PFD ENT key. (24) Push the load soft key. (25) Monitor the status of the upload. (26) When the UPDATE COMPLETED shows, push the ENT key. NOTE:
If the upload fails you can do the procedure again a maximum of four times. If the upload does fail four times there is a hardware problem that must be corrected. Do not cancel an upload before it is complete or fails.
(27) Push the UPDT CFG soft key, select YES, and then push the ENT key. NOTE: (28) (29) (30) (31) (32) (33) (34)
This updates the PFD 1 configuration module.
Monitor the status of the configuration module update. When the UPDATE CONFIG COMPLETE dialog box shows, push the ENT key. Set the EXTERNAL POWER switch on the circuit breaker switch panel to the OFF position. Set the AVIONICS 1 switch on the circuit breaker switch panel to the OFF position. Set the AVIONICS 2 switch on the circuit breaker switch panel to the OFF position. Remove external electrical power from the airplane. Remove the SVT unlock card from the PFD 1 top slot. NOTE:
Keep the SVT uplock card with the Pilot's Operating Handbook.
(35) Put the three GDU supplemental data cards in the bottom slot of the three flight displays. (a) Make sure that the data cards are 010-00330-43 data cards 3.
Garmin Synthetic Vision Technology System - Operational Check A.
Do a test of the Synthetic Vision Technology System Installation. (1) Make sure the airplane is out of a hanger with a clear view of the sky or in the hanger with a Global Positioning System (GPS) repeater system. NOTE: (2) (3) (4) (5) (6) (7) (8) (9)
The airplane must have a sufficient GPS positional fix to prevent a negative effect on the SVT display.
Put external electrical power on the airplane. Set the EXTERNAL POWER switch to the ON position. Set the AVIONICS 1 switch to the ON position. Set the AVIONICS 2 switch to the ON position. Push the ENT key on the MFD to clear the start screen. Push the PFD soft key on PFD 1. (a) Make sure that SYN VIS soft key shows on the PFD display. Push the SYN VIS soft key. Push the SYN TERR soft key. NOTE:
On the first initialization of the 010-00300-43 SD data card, it can take a maximum of five minutes before the SVT is initialized and the display shows.
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MODEL 208 MAINTENANCE MANUAL (10) Make sure that the soft keys that follow show on the SVT menu: • SYN TERR • HRZN HDG • APTSIGNS. • PATHWAY (for software 767.05 or later) B.
Do a test of the Synthetic Vision Technology System Field of View. (1) Turn the large FMS knob on the MFD to go to the MAP group. (2) Push the MENU key on the MFD. (3) Turn the small FMS knob to highlight the MAP SETUP option. (4) Push the ENT key. (5) Turn the small FMS knob to expand the drop down menu and select MAP in the group field. (6) Push the ENT key. (7) Turn the large FMS knob to scroll through the aviation group options to the FIELD OF VIEW option. (8) Turn the small FMS knob to select ON in the FIELD OF VIEW option. (9) Push the small FMS knob to return to the navigation map page on the MFD. (10) Set the EXTERNAL POWER switch on the circuit breaker switch panel to the OFF position. (11) Set the AVIONICS 1 switch on the circuit breaker switch panel to the OFF position. (12) Set the AVIONICS 2 switch on the circuit breaker switch panel to the OFF position. (13) Remove external electrical power from the airplane.
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35 CHAPTER
OXYGEN
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
35-00-00
Page 1
Dec 1/2006
35-01-00
Pages 1-9
Jan 2/2006
35-01-00
Pages 101-103
Aug 1/1995
35-01-00
Pages 201-205
Aug 1/1995
35-01-00
Pages 601-603
Jun 1/2011
35-11-00
Pages 201-203
Aug 1/1995
35-13-00
Pages 201-208
Mar 1/1999
35-14-00
Pages 201-203
Aug 1/1995
35-15-00
Pages 201-204
Aug 1/1995
35-15-00
Page 601
Jun 1/2011
35-Title 35-List of Effective Pages 35-Record of Temporary Revisions 35-Table of Contents 35-List of Tasks
35 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS OXYGEN - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-00-00 35-00-00 35-00-00 35-00-00
OXYGEN SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-01-00 Page 1 35-01-00 Page 1 35-01-00 Page 1
OXYGEN SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-01-00 Page 101 35-01-00 Page 101
OXYGEN SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Cylinder-Regulator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Cylinder-Regulator Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-01-00 Page 201 35-01-00 Page 201 35-01-00 Page 201 35-01-00 Page 202 35-01-00 Page 205
OXYGEN SYSTEM - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen System Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Bottle Restoration (Hydrostatic Test) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Bottle Discard. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-01-00 Page 601 35-01-00 Page 601 35-01-00 Page 601 35-01-00 Page 602 35-01-00 Page 603
OXYGEN FILLER VALVE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Filler Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-11-00 Page 201 35-11-00 Page 201 35-11-00 Page 201
HIGH PRESSURE OXYGEN LINE AND OUTLET VALVE ASSEMBLIES MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . High-Pressure Line Assembly Leak Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . High-Pressure Line Assembly Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Capillary Line Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection of High-Pressure Oxygen Lines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Outlet Valve Assembly Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . Crew Outlet Valve Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Outlet Valves Inspection/Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Test Oxygen System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-13-00 Page 201 35-13-00 Page 201 35-13-00 Page 201 35-13-00 Page 201 35-13-00 Page 201 35-13-00 Page 205 35-13-00 Page 205 35-13-00 Page 205 35-13-00 Page 205 35-13-00 Page 208
OXYGEN GAGE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Gage Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection of Gage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-14-00 Page 201 35-14-00 Page 201 35-14-00 Page 201 35-14-00 Page 201
OXYGEN MASKS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Mask Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Mask Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-15-00 Page 201 35-15-00 Page 201 35-15-00 Page 201 35-15-00 Page 201
OXYGEN MASKS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen Mask Restoration (Overhaul) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
35-15-00 Page 601 35-15-00 Page 601 35-15-00 Page 601
35 - CONTENTS © Cessna Aircraft Company
Page 1 Page 1 Page 1 Page 1
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 35-01-00-710
Oxygen System Operational Check
35-01-00 Page 601
35-01-00-780
Oxygen Bottle Restoration (Hydrostatic Test)
35-01-00 Page 602
35-01-00-960
Oxygen Bottle Discard
35-01-00 Page 603
35-15-00-960
Oxygen Mask Restoration (Overhaul)
35-15-00 Page 601
35 - LIST OF TASKS © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL OXYGEN - GENERAL 1.
Scope A.
2.
This chapter provides information on components associated with storage and distribution of oxygen to crew and passengers.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Teßon Lubricating Tape
S1465
Commercially Available
To lubricate threads and Þttings.
Trichlorethylene
ASTM D4080
Commercially Available
To clean oxygen lines.
Naptha
TT-N-95
Commercially Available
To ßush oxygen lines.
Sherlock Leak Detector
Type 1 MIL-PRF25567
Lub-O-Seal Co. Inc., 17519 Lewis Drive Cypress, Texas 77433
To leak test the oxygen system.
Flowrater (0 to 10 Liters Per Minute)
Commercially Available
To check pressure ßow to passenger mask.
Pressure gage (0 to 100 PSIG)
Commercially Available
To check oxygen ßow.
3.
DeÞnition A.
This chapter is divided into sections and subsections to assist maintenance personnel in locating speciÞc systems and information. For locating information within the chapter, refer to the Contents at the beginning of the chapter.
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MODEL 208 MAINTENANCE MANUAL OXYGEN SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
There are seven oxygen systems that are available for the 208 Caravan: A seventeen-port system with two pressure compensated regulators, a fourteen-port system with two pressure compensated regulators, a thirteen-port system with two compensated regulators, a ten-port system with two pressure compensated regulators, a ten-port system with a single pressure compensated regulator, an eight-port system with two pressure compensated regulators, and a two-port system with a non-pressure compensated regulator.
Description and Operation A.
The two-port system uses a 50.67 cubic-foot capacity oxygen cylinder. The eight-port, ten-port, thirteen-port, fourteen-port and seventeen-port systems use a 116.95 cubic-foot capacity oxygen cylinder. Both oxygen cylinders are composite construction and include a shutoff valve. All 116.95 cubic-foot capacity oxygen cylinders have an altitude compensated regulator which changes oxygen pressure with altitude. On airplanes equipped with the ten-port system, (Airplanes 208000208 thru 208000395), there is a single pressure compensated regulator to change the oxygen output with altitude. On airplanes that have the eight-port system (Airplanes 208000396 and On), ten-port system (Airplanes 208B000466 thru 208B1170) and seventeen-port system (Airplanes 208B000466 and On), there are two altitude compensated regulators to change the oxygen output with altitude. The 50.67 cubic-foot oxygen cylinders without an altitude compensated regulator keep an operating pressure of 70 psi and must have quick-don oxygen masks with a mounted diluter demand regulator. (1) On Models 208 and 208 Cargomaster the oxygen cylinder is attached to brackets that are installed in the upper part of the tailcone, aft of Fuselage Station 308.00. On Models 208B, and 208B Passenger the oxygen cylinder is attached to brackets that are installed aft of Fuselage Station 356.00. On Models 208 and 208 Cargomaster, there is an oxygen cylinder filler valve that is installed below a cover plate on the right side of the tailcone, aft of Fuselage Station 308.00. On Model 208B and 208B Passenger, there is an oxygen cylinder filler valve that is installed below a cover plate on the right side of the tailcone, aft of Fuselage Station 356.00. (2) A remote shutoff valve control with an ON/OFF label, is installed in the overhead console above pilot's and front passenger's seats. The remote shutoff valve control is used to turn the oxygen supply on or off as necessary. The shutoff valve control is mechanically connected to a cable that connects to the shutoff valve at the oxygen cylinder. (3) The oxygen outlets for the pilots and front passengers are attached in the cabin ceiling directly overhead and immediately outboard of each seat. On the Model 208 there are passenger oxygen outlets that are attached directly overhead and adjacent to the air vent outlets. The oxygen outlets on the Model 208 are attached in the same locations on the Model 208B Passenger. (4) One permanent microphone-equipped oxygen mask is provided for the pilot, and all other masks are partial rebreathing type, equipped with vinyl plastic flow indicators. (Refer to Figure 1). (5) All hoses provided for the pilot and passengers are the high-flow type and are color-coded with a blue band adjacent to the plug-in fitting. (a) An adapter cord is furnished with the pilot's microphone-equipped mask to mate the mask microphone lead to the auxiliary microphone jack located on the lower left outer portion of the instrument panel. (b) To connect the oxygen mask microphone, connect mask lead to the adapter cord and plug cord into the auxiliary microphone jack. (If an optional microphone-headset combination has been in use, the microphone lead from this equipment is already plugged into the auxiliary microphone jack. It will be necessary to disconnect this lead from the auxiliary microphone jack so that the adapter cord from the oxygen mask microphone can be plugged into the jack.) NOTE:
Airplanes equipped with only a partial oxygen system will incorporate the complete system less masks, oxygen cylinder, regulator, outlets, gage, control, filler, and some connecting plumbing
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MODEL 208 MAINTENANCE MANUAL B.
The oxygen flow to the outlet ports is provided when the oxygen control valve knob, located in the overhead console, is placed in the ON position and mask hoses are plugged into the overhead oxygen ports. NOTE:
C.
Each oxygen port contains a spring-loaded valve which prevents flow of oxygen until a mask hose is plugged in. Each mask hose contains an oxygen flow indicator for visual proof of oxygen flow.
The following information is permanently stamped on the shoulder, neck, or top head of the oxygen cylinder to aid in proper identification. (1) Cylinder specification followed by service pressure such as ICC or DOT-E8162. NOTE:
(2)
(3) (4) (5) (6) (7)
Effective 1 January 1970, all newly-manufactured cylinders are stamped DOT (Department of Transportation), rather than ICC (Interstate Commerce Commission). An example of the new designation would be: DOT-E8162.
Cylinder serial number is stamped below or directly following cylinder specification. The symbol of the purchaser, user, or maker, if registered with the Bureau of Explosives, may be located directly below or following the serial number. The cylinder serial number may be stamped in an alternate location on the cylinder top head. Inspectors official mark near serial number. Date of manufacture: This is the date of the first hydrostatic test (such as 6-84 for June 1984). The dash between the month and the year figures may be replaced with the mark of the testing or inspection agency (e.g., 6L84). Hydrostatic test date: Dates of subsequent hydrostatic tests shall be steel-stamped (month and year) directly below the original manufacturer date. the dash between month and year figures can be replaced with the mark of the testing agency. A Cessna identification placard is located near the center of cylinder body. Halogen test stamp: Halogen Tested, date of test (month, day, and year) inspector’s mark appears directly underneath the Cessna identification placard.
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Oxygen System Schematic Figure 1 (Sheet 1)
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Oxygen System Schematic Figure 1 (Sheet 2)
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Oxygen System Schematic Figure 1 (Sheet 3)
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Oxygen System Schematic Figure 1 (Sheet 4)
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Oxygen System Schematic Figure 1 (Sheet 5)
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Oxygen System Schematic Figure 1 (Sheet 6)
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Oxygen System Schematic Figure 1 (Sheet 7)
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MODEL 208 MAINTENANCE MANUAL OXYGEN SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Oxygen System Troubleshooting Chart Figure 101 (Sheet 1)
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Oxygen System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL OXYGEN SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Before any maintenance is performed on oxygen system, personnel should read and understand the following. Careful adherence to these instructions will aid in maintaining a trouble-free system.
Precautions
WARNING: Do not permit smoking or open flame near airplane while maintenance is being performed on the oxygen system. Ensure all electrical power is disconnected and that airplane is properly grounded. In addition, oils, grease, and solvents may burn or explode spontaneously when contacted by oxygen under pressure. A.
Use extreme care to ensure every port on system is kept thoroughly clean and free of water, oil, grease, and solvent contamination.
B.
Cap all openings immediately upon removal of any component. Do not use tape or caps which will induce moisture.
C.
Lines and fittings shall be clean and dry. One of the following methods may be used to clean lines.
CAUTION: Most air compressors are oil-lubricated and a minimum amount of oil may be carried by airstream into system. A water-lubricated compressor should be used to blow tubing clean only when nitrogen or argon are not available. The air must be clean, dry, and filtered. (1) (2) (3) (4) (5) (6) (7)
Wash with a vapor-degreasing solution of stabilized trichloroethylene conforming to MlL-T-7003, followed by blowing tubing clean with a jet of nitrogen gas (BB-N- 411) Type 1, Class 1, Grade A or Technical Argon (MIL-A-18455). Flush with naptha conforming to Specification TT-N-95, then blow clean and dry with clean, dry filtered air. Flush with anti-icing fluid conforming to MlL-F-5566 or anhydrous ethyl alcohol. Rinse thoroughly with fresh water and dry with a jet of nitrogen gas (BB-N-411) Type 1, Class 1, Grade A or Technical Argon (MIL-A-18455). Flush with hot inhibited alkaline cleaner until free from oil and grease. Rinse with fresh water and dry with a jet of nitrogen gas (BB-N-411) Type 1, Class 1, Grade A or Technical Argon (MIL-A-18455). Cap all lines immediately after drying. Fabrication of pressure lines is not recommended. Lines should be replaced with factory parts, by part number. Use only S1465 Teflon lubricating tape on threads of male fittings. No lubricating tape is used on coupling sleeves or outside of flares. Maintenance personnel must ensure that their hands are free of dirt and grease prior to installation of oxygen tubing or fittings.
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WARNING: Use nonsparking tools. CAUTION: With oxygen bottle charged, do not place control in the on position with outlet ports (low pressure) open to atmosphere. Damage to regulator metering poppet may occur. CAUTION: Whenever a component of the oxygen system has been removed, reinstalled, replaced, or system has been disassembled in any way, the oxygen system must be leak-checked and purged. D.
All tools used for installation of oxygen tubes or fittings must be free of dirt, grease and oils. NOTE:
3.
If a cylinder is recharged more than an average of once every other day, an accurate record of the number of recharges must be maintained by the owner or his agent.
Oxygen Cylinder-Regulator Removal/Installation A.
Remove Oxygen Cylinder-Regulator (Refer to Figure 201 and Figure 202). (1) Remove aft baggage partition to gain access to oxygen cylinder assembly.
CAUTION: Ensure that the oxygen shutoff valve arm is still in the off position after removing the cable end (17) from shutoff valve arm. (2) (3) (4) (5) (6) (7) (8) B.
Straighten cable end (17), remove bolt (21) from nut (20) and slip nut (20) off cable end (17). Loosen two screws securing control cable assembly (1) to cable housing clamp (16) and then remove cable assembly (1) from cable housing clamp (16) on regulator assembly (19). Remove and cap high-pressure line (9), and cap regulator port. Do not remove safetywired adapter from regulator. Remove and cap low-pressure line (10) Remove nipple (11) from compensated regulator (12) on 208 and regulator (13) on Model 208 Federal Express, and plug port. Loosen and remove washers (6) and bolts (7) attaching clamp (8) to mounting bracket (3). Remove safety-wire from clamp (8), loosen clamps and remove cylinder.
Install Oxygen Cylinder- Regulator (Refer to Figure 201 and Figure 202). (1) Slip clamps (8) over cylinder end, being certain that orientation is correct for attachment to mounting brackets (3). (2) Attach brackets (3) to support assemblies (4) using washers (6), bolts (7), and nutplate (5). (3) Tighten and safety-wire clamp (8). NOTE:
Observe all previously listed cautions and warnings when installing line fittings.
(4)
Install nipple (11) in pressure compensated regulator (12) on Model 208, or regulator (13) on Federal Express airplanes. (5) Attach low-pressure line (10) to nipple (11). (6) Attach high-pressure line (9) to regulator adaptor. (7) Insert cable (1) through cable housing clamp (16). (8) Insert cable end (17) through nut (20) and tighten bolt (21). (9) Test operation of control system to ensure that control will operate from the overhead console. (10) Bend cable end (1) 90 degrees. (11) Reinstall aft baggage partition.
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Two-Port Oxygen Bottle and Regulator Installation Figure 201 (Sheet 1)
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Ten-Port Oxygen Bottle and Regulator Installation Figure 202 (Sheet 1)
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4.
Oxygen Cylinder-Regulator Inspection A.
Inspect Oxygen Cylinder-Regulator. (1) A careful visual inspection of the oxygen cylinder should be performed during routine maintenance and periodic inspections. If the acceptability of the cylinder is questionable, return cylinder to manufacturer. Acceptable damage consists of such items as scratched paint or cuts and abrasions. (a) Scratches or Cuts. Cuts or scratches less than 0.005 inch deep are acceptable. (b) Abrasions. Minor abrasions such as scuffs, are acceptable unless the damage is deep enough to expose groups of fibers. Abrasions with isolated groups of fibers exposed or flat spots with depth less than 0.010 inch must be epoxy coated to avoid water entrapment. A group of fibers is defined as 0.010 inch thick and 0.125 inch wide. (c) Paint Removal. Paint removal is not recommended. In the event that paint removal for inspection or other reasons is required, the suitability of the paint removal procedure must be verified by the cylinder manufacturer. Some chemical paint removers may damage the composite. Abrasive or other mechanical means of paint removal, such as shot blast or wire brush are prohibited. (2) Regulator shall be checked to see that it functions properly during hydrostatic testing. (3) Actuate regulator controls and valve to check for ease of operation.
CAUTION: Damage to regulator will occur if the control of a charged oxygen cylinder is turned on with the low-pressure side of the regulator open to the atmosphere. (4)
Pressurize the system and check for leaks. NOTE:
For oxygen cylinder inspection, also refer to publication CGA C-612, Compressed Gas Association, Inc., Arlington, VA. 22202.
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MODEL 208 MAINTENANCE MANUAL OXYGEN SYSTEM - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the oxygen system in a serviceable condition.
Task 35-01-00-710 2.
Oxygen System Operational Check A.
General (1) This task gives the procedures to do an operational check of the oxygen system and the oxygen masks.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do a Detailed Inspection of the Oxygen System.
WARNING: When working on or around oxygen systems, always be alert for contaminants such as dirt or petroleum base materials. When mixed with contaminants, gaseous oxygen will explode. (1) (2) (3) (4) (5) (6) (7) (8) E.
On the cockpit overhead panel, examine the control handle for condition, security, and correct cable attachment. Examine the gauge installation for condition and security. Examine the mounting structure of the control handle and gauge for condition and security. Examine all mask assemblies for condition, security of components, and cleanliness. Examine the hoses for condition, security of attachment at the mask and security of the hose connector. Examine the hose end for condition, wear, and cleanliness. Examine the oxygen system outlets in the cockpit and cabin overhead for condition, security, and cleanliness. Make sure that the system servicing is correct. Refer to Chapter 12, Oxygen - Servicing.
Do an Oxygen System Operational Check.
WARNING: Do not smoke or let any open flame near the airplane while maintenance or other work is done on the oxygen system. Make sure that all electrical power is disconnected and that airplane is properly grounded. In addition, oils, grease, and solvents may burn or explode spontaneously when contacted by oxygen under pressure. NOTE: (1) (2) (3) (4) (5)
Perform this operational check on each oxygen mask and each outlet port installed in the airplane. Make sure that an adequate supply of oxygen is available for this test.
Connect an oxygen mask to an outlet port. Operate the control handle from OFF to ON and check for ease of operation. Make sure that the flowmeter is fully in the GREEN band. Operate the control handle from the ON to the OFF position, and examine for ease of operation and positive shutoff of the handle. If the operational check of the mask is satisfactory, disconnect the mask and correctly stow it in the airplane.
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If the operational check of the mask is unsatisfactory, replace the mask. Service the oxygen system. Refer to Chapter 12, Oxygen - Servicing.
F.
Restore Access (1) None End of task Task 35-01-00-780 3.
Oxygen Bottle Restoration (Hydrostatic Test) A.
General (1) This task includes the steps necessary to do a restoration (hydrostatic test) of the oxygen cylinder.
B.
Special Tools (1) None
C.
Access (1) Remove the aft baggage partition to get access to the oxygen cylinder assembly.
D.
Do the Oxygen Bottle Restoration (Hydrostatic Test).
WARNING: Make sure that personnel obey the safety precautions. Refer to Oxygen System - Maintenance Practices. CAUTION: Oxygen cylinders and pressure regulators are supplied as assemblies. Removal, repair, and installation of the oxygen pressure regulators in the field can cause contaminants to enter the oxygen system. CAUTION: Make sure that unserviceable pressure regulators or pressure regulators that need disassembly are interchanged for replacement oxygen cylinder and pressure regulator assemblies. CAUTION: Make sure that the oxygen cylinder and the pressure regulator assemblies are disassembled, repaired, inspected, cleaned, hydrostatically tested, reassembled, and serviced by the manufacturer or other FAA approved facility. NOTE: (1) (2) (3) (4) (5)
The pressure regulator is safety wired to the open position.
Remove the oxygen cylinder. Refer to Oxygen System - Maintenance Practices. Send the oxygen cylinder to an approved service facility for the hydrostatic test. Install the oxygen cylinder. Refer to Oxygen System - Maintenance Practices. Do the servicing of the oxygen cylinder. Refer to Chapter 12, Oxygen System - Servicing. Install safety wire on the shutoff valve knob while the knob is in the ON position. Refer to Safetying - Maintenance Practices.
E.
Restore Access (1) Install the aft baggage partition. End of task
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MODEL 208 MAINTENANCE MANUAL Task 35-01-00-960 4.
Oxygen Bottle Discard A.
General (1) This task has the procedures for the discard of the oxygen cylinder.
B.
Special Tools (1) None
C.
Access (1) Remove the aft baggage partition to get access to the oxygen cylinder assembly.
D.
Discard the Oxygen Bottle.
WARNING: Make sure that personnel obey the safety precautions. Refer to Oxygen System - Maintenance Practices. CAUTION: Oxygen cylinders and pressure regulators are supplied as assemblies. Removal, repair, and installation of the oxygen pressure regulators in the field can cause contaminants to enter the oxygen system. CAUTION: Make sure that unserviceable pressure regulators or pressure regulators that need disassembly are interchanged for replacement oxygen cylinder and pressure regulator assemblies. CAUTION: Make sure that the oxygen cylinder and the pressure regulator assemblies are disassembled, repaired, inspected, cleaned, hydrostatically tested, reassembled, and serviced by the manufacturer or other FAA approved facility. NOTE: (1) (2) (3) (4) (5)
The pressure regulator is safety wired to the open position.
Remove the oxygen cylinder. Refer to Oxygen System - Maintenance Practices. Send the bottle to an authorized discard facility. Install a new oxygen cylinder. Refer to Oxygen System - Maintenance Practices. Do the servicing of the oxygen cylinder. Refer to Chapter 12, Oxygen System - Servicing. Install safety wire on the shutoff valve knob while the knob is in the ON position. Refer to Safetying - Maintenance Practices.
E.
Restore Access (1) Install the aft baggage partition. End of task
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MODEL 208 MAINTENANCE MANUAL OXYGEN FILLER VALVE - MAINTENANCE PRACTICES 1.
General A.
2.
The oxygen filler valve is located behind a cover assembly located on the right side of the airplane at FS 309.90 on Model 208, and FS 357.90 on the 208B. The cover assembly is attached by 4-1/4 turn flush-head studs.
Oxygen Filler Valve Removal/Installation
WARNING: Disconnect the high-pressure line from the regulator before attempting any service procedure on the high-pressure side of the system (indicating or filler valve). A.
Remove Oxygen Filler Valve (Refer to Figure 201). (1) Remove aft baggage partition to gain access to filler valve assembly. (2) Loosen line assembly (11) from filler valve body (10). (3) Unscrew filler valve from body filler valve end (5). (4) Remove sleeve assembly (8) from filler valve end (5). (5) Remove O-rings (12) and (13) from sleeve assembly (8). NOTE: (6)
B.
The O-rings are the only field-repairable/replaceable parts of the filler valve assembly.
If filler valve end (5) is damaged, remove by loosening screws (4) from outside of airplane through access port.
Install Oxygen Filler Valve (Refer to Figure 201). (1) Install O-rings (12) and (13) on sleeve assembly (8). (2) Ensure that valve core (9) is properly positioned in filler body assembly and bronze filter (7) is positioned in sleeve assembly (8). (3) Screw filler valve assembly into filler valve end (5). (4) Attach line assembly (11) to filler valve body (10). (5) Test for leaks. (6) Reinstall aft baggage partition.
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Oxygen Filler Valve Installation Figure 201 (Sheet 1)
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Oxygen Filler Valve Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL HIGH PRESSURE OXYGEN LINE AND OUTLET VALVE ASSEMBLIES - MAINTENANCE PRACTICES 1.
2.
General A.
The high-pressure oxygen line assembly consists of the tubing and fittings connecting the oxygen filler valve to the cylinder-regulator and the capillary line to the pressure gage.
B.
The capillary line is routed from the back of the gage, then overhead through holes in the bulkheads to the oxygen cylinder regulator. The line is supported by grommets and ties at the various bulkheads.
C.
Overhead, outboard of each passenger station and centered between the pilot/copilot stations, are individual oxygen outlet valves. The valves are adjacent to the passenger lighting/ventilation ports and adjacent to the pilot/copilot floodlights.
High-Pressure Line Assembly Leak Check A.
Service High-Pressure Line Assembly (1) Charge the oxygen system in accordance with Chapter 12. (2) Allow 30 minutes for cylinder pressure to stabilize between 1800 and 1850 PSIG, indicated on pressure gage. (3) Record the cylinder pressure and ambient temperature. (4) After 24 hours, record cylinder pressure and ambient temperature. Maximum allowable pressure drop is 50 PSIG (correcting for temperature change, using formula of 618F 53.4 PSIG). NOTE:
(5) (6) 3.
4.
A shorter interval than 24 hours may be used. In this case, multiply the pressure change (which has been corrected for any temperature change) by 24/H where H is the number of hours between pressure readings. This gives how much the pressure drop would be in 24 hours.
If the pressure drop derived from the formula in the preceding step exceeds 50 PSIG, test the oxygen system for leakage by applying leak detector fluid Type CG-1 or equivalent to all fittings and connections, and observe for formation of bubbles. Remove all traces of solution. Repair or replace leaky fitting and repeat the preceding procedures.
High-Pressure Line Assembly Removal/Installation A.
Remove High-Pressure Line Assembly (Refer to Figure 201). (1) Ensure that oxygen control is in the OFF position. (2) Remove aft baggage partition. (3) Remove high-pressure line (9) from regulator adapter and tee (9A).
B.
Install High-Pressure Line Assembly (Refer to Figure 201). (1) Attach line assembly (9) to tee (9A) and regulator adapter (8). (2) Recharge system. Refer to Chapter 12, Oxygen System - Servicing. (3) Test system for leaks. Refer to High Pressure Line Assembly Leak Check, in this section.
Capillary Line Removal/Installation A.
Remove Capillary Line (Refer to Figure 201). (1) Remove overhead panels as necessary to expose capillary line. (2) Clip ties supporting capillary line. (3) Disconnect capillary line from tee (9A). (4) Disconnect capillary line from oxygen gage (11). (5) Remove and cap capillary line.
B.
Install Capillary Line (Refer to Figure 201). (1) Beginning at oxygen cylinder, route capped capillary line through holes in bulkheads overhead forward to gage position on overhead console panel. (2) Attach capillary line to gage (11). (3) Perform leak check. Refer to High-Pressure Line Assembly Leak Check.
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208/208B Outlet Valve Installation Figure 201 (Sheet 1)
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208/208B Outlet Valve Installation Figure 201 (Sheet 2)
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208/208B Outlet Valve Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (4) (5) 5.
Secure capillary line with ties and grommets as required. Replace overhead panels as necessary.
Inspection of High-Pressure Oxygen Lines A.
Visually inspect lines and fittings for cracks, nicks, corrosion, kinks, dents, rust, or visible damage. Presence of any of these will require replacement of affected area. NOTE:
6.
7.
Disassembly and reassembly of high-pressure lines should only be attempted by personnel familiar with high-pressure gases.
Passenger Outlet Valve Assembly Removal/Installation A.
Remove Passenger Outlet Valve Assembly (Refer to Figure 202). (1) Remove Wemac valve (31) by turning counterclockwise. (2) Remove light (34) by turning counterclockwise, then remove screws and washers securing electrical leads to light. (3) Remove dress ring (29) from switch (23). (4) Remove decorative nut (30) from outlet valve (21) and remove cover (28). (5) Open headliner as necessary to gain access to outlet valve (21). (6) Ensure oxygen control is in the OFF position then disconnect line (22) from outlet valve (21), cap line (22) and outlet valve (21). (7) Remove nut (35) and washer (36) then remove outlet valve (21).
B.
Install Passenger Outlet Valve Assembly (Refer to Figure 202). (1) Position outlet valve (21) through bracket (25) and install washer (36) and nut (35). (2) Remove cap from outlet valve (21) and line (22) and connect line (22) to outlet valve (21). Leak check connection. (3) Close headliner. (4) Position cover (28) and install dress ring (29) on switch (23). (5) Install decorative nut (30) on outlet valve (21). (6) Screw Wemac valve (31) into bracket (25). (7) Connect electrical leads to light (34). (8) Rotate light assembly approximately eight turns counterclockwise, then position in bracket (25) and screw in clockwise. Electrical leads should not be twisted after installation.
Crew Outlet Valve Assembly Removal/Installation A.
Remove Crew Outlet Valve Assembly (Refer to Figure 201 ). (1) Remove overhead console. (2) Remove jamnuts (5) holding valve assembly (7) to oxygen valve flange (6). (3) Loosen line assembly (9) from tee (9A) and cap line. (4) Remove valve (7) out of oxygen valve flange (6) and remove escutcheon (6A).
B.
Install Crew Outlet Valve Assembly (Refer to Figure 201). (1) Install escutcheon (6A) and insert valve (7) into hole in oyxgen valve flange (6) and install jamnuts (5) loosely. (2) Attach line assembly (9) to tee (9A) and attach tee (9A) to flared end of adapter (8). NOTE: (3) (4)
8.
No Teflon tape or sealant compound is to be used on flared connectors.
Test fittings for leaks. Reinstall overhead console after adjusting jamnut to ensure flush mounting of decorative ring (1).
Oxygen Outlet Valves Inspection/Test A.
Inspect Oxygen Outlet Valves. (1) Ensure that oxygen system is fully charged. (2) Insert an oxygen outlet adapter connected to a pressure gage into the oxygen outlet valve.
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208B Passenger Outlet Valve Installation Figure 202 (Sheet 1)
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208B Passenger Outlet Valve Installation Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (3) (4) 9.
Test retainer assembly adapter junction for leaks with fluid leak detector. No bubbles are permitted. After completion of leak tests, fully charge the oxygen system.
Functional Test Oxygen System NOTE:
Whenever the oxygen system regulator (cylinder- regulator assembly) has been replaced or overhauled, perform a flow test to determine that system functions properly.
A.
Functional Test Uncompensated Oxygen System. (1) Fully charge oxygen system. Refer to Chapter 12, Oxygen System - Servicing. (2) Install an oxygen outlet adapter, Part Number C166005-0506, into a pressure gage calibrated in one-pound increments from 0 to 100 PSIG. Insert adapter into pilot's oxygen outlet. (3) Place oxygen control in the ON position and verify pressure is 70 PSIG, +10 or -10 PSIG. (a) If pressure is not 70 PSIG, +10 or -10 PSIG, replace cylinder and regulator assembly. Repeat steps 9.A.(1) thru (3). (4) Recharge oxygen system as required. Refer to Chapter 12, Oxygen System - Servicing.
B.
Functional Test Compensated Oxygen System. (1) Fully charge oxygen system. Refer to Chapter 12, Oxygen System - Servicing. (2) Install an oxygen outlet adapter, Part Number C166005-0506, into a pressure gage, calibrated in one-pound increments from 0 to 100 PSIG. Insert adapter into pilot's oxygen outlet. (3) Place oxygen control in the ON position. (4) Insert adapters or mask line assemblies into all remaining outlets. (a) With oxygen flowing from all outlets, verify pressure conforms to Table 201. (b) If pressure at given altitude is different than shown per Table 201 check oxygen pressure at altitude compensating regulator inlet and verify pressure is 70 PSIG, +10 or -10 PSIG. (c) If pressure cannot be obtained per Table 201, and pressure is 70 PSIG, +10 or -10 PSIG at line to inlet port of compensating regulator, replace compensating regulator. (d) If 70 PSIG, +10 or -10 PSIG cannot be obtained at compensating regulator, replace cylinder regulator and repeat steps 9.B.(1) thru (4). (5) Position control to OFF and return all masks to mask storage. (6) Recharge oxygen system as required. Refer to Chapter 12, Oxygen System - Servicing.
Table 201. Altitude Pressure ALTITUDE ABOVE SEA LEVEL
PRESSURE GAGE
Sea Level
7.30 PSIG, +2.50 or -1.50 PSIG
1000
7.83 PSIG, +2.50 or -2.50 PSIG
1330
8.00 PSIG, +2.50 or -2.50 PSIG
2000
8.34 PSIG, +2.50 or -2.50 PSIG
3000
8.83 PSIG, +2.50 or -2.50 PSIG
4000
9.31 PSIG, +2.50 or -2.50 PSIG
5000
9.77 PSIG, +2.50 or -2.50 PSIG
6000
10.22 PSIG, +2.50 or -2.50 PSIG
8000
11.08 PSIG, +2.50 or -2.50 PSIG
10000
11.89 PSIG, +2.50 or -2.50 PSIG
14000
17.57 PSIG, +2.50 or -2.50 PSIG
17000
21.55 PSIG, +2.50 or -2.50 PSIG
20000
24.45 PSIG, +2.50 or -2.50 PSIG
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MODEL 208 MAINTENANCE MANUAL OXYGEN GAGE - MAINTENANCE PRACTICES 1.
General A.
2.
The oxygen gage is mounted in the overhead console.
Oxygen Gage Removal/Installation A.
Remove Oxygen Gage (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6) (7) (8)
B.
3.
The system does not have to be discharged before removing high-pressure lines as there is a check valve in the regulator to shut off the flow of oxygen when a connection is broken.
Remove overhead console to gain access to oxygen gage. Remove fuel selector valve handles. Remove knobs (10). Remove screws securing clear panel allowing access to the oxygen gage (1). Remove screws (12) securing retainer (3). Remove screw (4) securing retainer (3) to oxygen gage (1). Disconnect capillary line (2) from gage (1) Remove gage from overhead console area.
Install Oxygen Gage (Refer to Figure 201). (1) Position oxygen gage (1) in overhead console area. (2) Connect capillary line (2) to gage (1). (3) Install screw (4) securing retainer (3) to oxygen gage. (4) Install screws (12) securing retainer. (5) Position clear panel and install screws. (6) Install knobs (10) and secure using screw (11). (7) Install fuel selector valve handles. (8) Position and secure overhead console. (9) Install overhead console floodlights.
Inspection of Gage A.
The only inspection possible is to observe indicated pressure rise as the system is charged and decrease as oxygen is bled off.
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Oxygen Gage and Control Valve Installation Figure 201 (Sheet 1)
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Oxygen Gage and Control Valve Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL OXYGEN MASKS - MAINTENANCE PRACTICES 1.
General A.
2.
Oxygen Mask Inspection A.
3.
The Model 208 is equipped with one pilot's mask and nine passenger masks. The pilot's mask has a built-in microphone. On Models 208 and Federal Express 208B pilots and copilots quick-don masks with built-in microphones are provided. The Model 208 Super Cargomaster and 208B Passenger masks are of the constant-flow type with a metering orifice in the quick-connect adapter. The Model 208 and Federal Express 208B masks are the demand type. A flowmeter built into the line approximately 6 inches from the connector provides a visual indication of proper oxygen flow, showing red when no flow is taking place, red and green with a partial flow, and green with full flow. The masks are color- coded by a blue sleeve adjacent to the quick-connect adapter.
Inspect Oxygen Mask (Refer to Figure 201). (1) Check oxygen masks for cracks and rough face seals. (2) Flex the mask hose gently over its entirety and check for evidence of deterioration or dirt. (3) Examine mask and hose storage compartment for cleanliness and general condition. (4) Observe that each mask breathing tube end is free of nicks and that the tube end will slip into the cabin oxygen receptacle with ease and will not leak. (5) If a mask assembly is defective (leaks, does not allow breathing, or contains a defective microphone) it is advisable to return the mask assembly to the manufacturer or a repair station. (6) Replace hose if it shows evidence of deterioration.
Oxygen Mask Cleaning A.
Clean Oxygen Mask. (1) Clean and disinfect mask assemblies after use with rubbing alcohol, as appropriate. (2) If installed, remove microphone from mask.
CAUTION: DO NOT ALLOW RUBBING ALCOHOL TO MICROPHONE OR ELECTRICAL CONNECTIONS. (3) (4)
ENTER
Apply rubbing alcohol to mask with a cotton swab or the equivalent, as required, to remove contamination. If used, install microphone.
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Mask Assembly and Smoke Goggles Figure 201 (Sheet 1)
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Mask Assembly and Smoke Goggles Figure 201 (Sheet 2)
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Mask Assembly and Smoke Goggles Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL OXYGEN MASKS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the oxygen masks in a serviceable condition.
Task 35-15-00-960 2.
Oxygen Mask Restoration (Overhaul) A.
General (1) The crew masks are located on the crew doors. The passenger oxygen masks are located in different areas of the cabin depending on the seating configuration.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the Oxygen Mask Restoration (Overhaul). Refer to Table 601 for the correct Cessna and supplier part numbers. NOTE: (1) (2) (3)
A restoration (overhaul) on the mask is done at six years from the date of the manufacture or six years from the date of the last overhaul.
Remove the masks from the storage containers. Send the masks to an approved repair facility for an overhaul. Install the masks in the storage containers.
Table 601. Oxygen Mask Restoration (Overhaul) Cessna Part Number
Supplier Name and Part Number
Restoration (Overhaul) Interval
C166015-0101
B/E Aerospace 174406-01
6 years
C166015-0101
Avox 802065-03
6 years
C166015-0102
B/E Aerospace 174441-01
6 years
359A3552
Avox 359A3552
6 years
E.
Restore Access (1) None End of task
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CHAPTER
PNEUMATIC
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
36-00-00
Page 1
Aug 2/2004
36-10-00
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Apr 1/2010
36-Title 36-List of Effective Pages 36-Record of Temporary Revisions 36-Table of Contents
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
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Issue Date
By
Date Removed
By
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CONTENTS PNEUMATIC SYSTEM - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
36-00-00 Page 1 36-00-00 Page 1 36-00-00 Page 1 36-00-00 Page 1
PNEUMATIC DISTRIBUTION - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Regulator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pneumatic Distribution Line Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Connecting Shop Air to Pneumatic System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Component Cleaning/Servicing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Regulator Output Adjustment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Regulator Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL PNEUMATIC SYSTEM - GENERAL 1.
Scope A.
2.
Definition A.
3.
This section provides description and operation information on the components used to distribute pneumatic engine bleed air to using systems.
This chapter is divided into sections to aid maintenance technicians in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief description of the section follows: (1) The section on distribution describes that portion of the system used to regulate and distribute air to the vacuum and airfoil anti-ice system.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following listed items:
NAME
NUMBER
MANUFACTURER
USE
Adhesive (Gasoila)
26416351
Parker Hannifin Airborne Air & Fuel Products 711 Taylor Street Elyria, OH 44305
Used on pressure regulator.
Isopropyl Alcohol
Federal Specification TT-I-735
Commercially Available
Used for cleaning.
Naptha Solvent
PD-680 Type III
Commercially Available
Used for cleaning.
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MODEL 208 MAINTENANCE MANUAL PNEUMATIC DISTRIBUTION - MAINTENANCE PRACTICES 1.
General A.
This section gives procedures and data which apply to the pneumatic distribution system. This includes how to remove and install the components in the distribution system, connect the pneumatic system to shop air, clean, adjust, and test the pressure regulator.
CAUTION: When you replace a pneumatic system component, make sure all the connections are correct to prevent damage to the gyro system. CAUTION: Do not use teflon tape, pipe dope, or thread lubricants of any type on the fitting threads. Do not tighten the connections too much. (1)
2.
3.
When a component is removed, identify and use a cover on all open lines, hoses, and fittings to prevent dirt or foreign material from entering the system. Make sure installation is correct. When you replace a component, examine all the hoses carefully to make sure they are clean and free of debris, oil, solvent, collapsed inner liners, or external damage. Replace the hoses that are old, hard, cracked, or brittle.
Pressure Regulator Removal/Installation A.
Remove Pressure Regulator (Refer to Figure 201). (1) Open right engine cowling door to gain access to pressure regulator (6). (2) Disconnect line (2) from pressure regulator (6) and cap line. (3) Detach tube nut (3) from pressure regulator (6) and cap tee (5). (4) Disconnect pneumatic line (1) from pressure regulator (6). (5) Remove pressure regulator (6).
B.
Install Pressure Regulator (Refer to Figure 201). (1) Position pressure regulator (6) in engine compartment and connect pneumatic line (1) to regulator. (2) Remove cap from tee (5) and attach tube nut (3) to pressure regulator (6). (3) Connect line (2) to regulator. (4) Close engine cowling door.
Pneumatic Distribution Line Removal/Installation A.
Remove Pneumatic Distribution Line (Refer to Figure 201). (1) Open left and right engine cowling to gain access to pneumatic line (1). (2) Loosen clamps (4) securing pneumatic line (1) to engine mount and firewall. (3) Remove pneumatic line (1). (a) Airplanes 20800001 thru 20800143 and 208B0001 thru 208B0143, remove pneumatic line (1) from pressure regulator (6) and union (7) or deice line tee, if installed, at ejector (8). (b) Airplanes 20800144 and On and 208B0144 and On, and Airplanes 20800001 thru 20800143 and 208B0001 thru 208B0143 incorporating CAB90-14, remove line (1) from pressure regulator (6) and cross fitting (9) at ejector (8). (c) Airplanes 20800222 and On and 208B0317 and On, and Airplanes 20800001 thru 20800121 and 208B0001 thru 208B0316 incorporating CAB93-2, remove line (1) from pressure regulator (6) and cross assembly (9C) at ejector (8).
B.
Install Pneumatic Distribution Line (Refer to Figure 201). (1) Install pneumatic line (1). (a) Airplanes 20800001 thru 20800143 and 208B0001 thru 208B0143, position and install pneumatic line (1) to pressure regulator (6) and union (7) or deice line tee, if installed, at ejector (8). (b) Airplanes 20800144 and On and 208B0144 and On, and Airplanes 20800001 thru 20800143 and 208B0001 thru 208B0143 incorporating CAB90-14, position and install pneumatic line (1) to pressure regulator (6) and cross fitting (9) at ejector (8).
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Pneumatic System Installation Figure 201 (Sheet 1)
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Pneumatic System Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (c)
(2) (3) 4.
Connecting Shop Air to Pneumatic System NOTE:
5.
Connect Air to Pneumatic Distribution Line (Refer to Figure 201). (1) Remove pneumatic line (1) from pressure regulator (6). Cap open fitting on pressure regulator (6). (2) Connect regulated shop air to pneumatic line (1) and turn control valve on.
B.
Disconnect Air from Pneumatic Distribution Line (Refer to Figure 201). (1) Turn control valve off and disconnect pneumatic line (1) from regulated shop air. (2) Remove cap from pressure regulator (6) fitting and connect pneumatic line (1).
Component Cleaning/Servicing Pressure Regulator Cleaning Procedures. NOTE:
Use this procedure when pressure regulator output is too high, too low, erratic, diaphragm is intact, as shown by an absence of flow from vent hole in the dome and from around joint between dome and body, or no other obvious condition exists.
NOTE:
This procedure provides pressure regulator adjustment. However, pressure regulator adjustment should not be accomplished independently, as it may mask an internal problem, allowing condition to worsen.
(1)
The procedures to clean the pressure regulator are found in the vendor maintenance manual. Refer to the List Of Publications found in the introduction section of this manual for the applicable vendor maintenance manual.
Pressure Regulator Output Adjustment A.
7.
The pneumatic system may be operated without running the engine if a source of compressed air is available. Shop air must be filtered, regulated from 15 to 18 PSI, and equipped with a control valve.
A.
A.
6.
Airplanes 20800222 and On and 208B0317 and On, and Airplanes 20800001 thru 20800121 and 208B0001 thru 208B0316 incorporating CAB93-2, position and install pneumatic line (1) to pressure regulator (6) and cross assembly (9C) at ejector (8). Secure pneumatic line (1) to engine mount and firewall using clamps (4). Close left and right engine cowling doors.
Adjust pressure regulator output. (1) Loosen locknut and turn adjustment screw clockwise to increase pressure or counterclockwise to decrease pressure. Begin with one turn of adjustment screw, adjusting as required, until 18 PSIG, +1 or - 1 PSIG, is obtained. (2) Apply 26416351 Gasoila adhesive to adjustment screw and locknut when adjustments are completed.
Pressure Regulator Functional Test A.
Functional Test Procedures (Refer to Figure 202). (1) Remove relief valve or cap from union, tee, cross fitting (1A) or cross assembly (1), located near vacuum ejector (7). (2) Connect hose and pressure gage to uncapped fitting. Ensure hose length is sufficient to allow gage to reach cockpit. Refer to Figure 202. (3) Run engine and check regulator discharge pressure. Discharge pressure shall range from 17.0 to 20.0 PSIG over a power range of 70 percent Ng to momentary full power (as limited by the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual) with both increasing and decreasing power.
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Bleed Air Pressure Regulator Discharge Inspection Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (4)
(5)
Check for proper operation of pneumatic system, vacuum system, deice system and heater system. Refer to Chapter 37, Vacuum - General, Chapter 30, Ice and Rain Protection - General, and Chapter 21, Compressor Bleed Air Heater - Maintenance Practices. NOTE:
The 17.0 to 20.0 PSIG pressure range is greater than the normal 18.0 PSIG pressure regulator setting to allow for the lesser precision of airplane testing, compared to bench testing, and to allow for the pressure drop at union, tee, cross fitting (1A) or cross assembly (1) at vacuum ejector (7).
NOTE:
Other than a pressure regulator problem, low output may be caused by pressure leakage in plumbing downstream of pressure regulator, particularly deice plumbing; whereas, high output may be caused by blockage to the vacuum ejector nozzle (7), which could produce a low vacuum indication combined with the high regulator pressure.
Upon completion of test and adjustment, remove pressure gauge and recap union, tee, cross fitting (1A) or cross assembly (1).
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CHAPTER
VACUUM
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
37-00-00
Page 1
Aug 1/1995
37-01-00
Pages 1-3
Mar 3/1997
37-01-00
Pages 101-103
Aug 1/1995
37-10-00
Pages 201-206
Mar 3/1997
37-10-00
Pages 601-602
Jun 1/2011
37-20-00
Pages 201-205
Mar 1/2008
37-Title 37-List of Effective Pages 37-Record of Temporary Revisions 37-Table of Contents 37-List of Tasks
37 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
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MODEL 208 MAINTENANCE MANUAL
CONTENTS VACUUM - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
37-00-00 Page 1 37-00-00 Page 1 37-00-00 Page 1
VACUUM SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
37-01-00 Page 1 37-01-00 Page 1 37-01-00 Page 1
VACUUM SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
37-01-00 Page 101 37-01-00 Page 101
VACUUM DISTRIBUTION - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vacuum Ejector Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vacuum Relief Valve Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Connecting Shop Air to Vacuum System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vacuum Relief Valve Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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VACUUM DISTRIBUTION - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vacuum System Central Air Filter Discard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vacuum Relief Valve Filter Discard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
37-10-00 Page 601 37-10-00 Page 601 37-10-00 Page 601 37-10-00 Page 601
VACUUM INDICATING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vacuum-Low Warning Switch Removal/Installation (Model 208) . . . . . . . . . . . . . . . . . Vacuum-Low Warning Switch Removal/Installation (Model 208B) . . . . . . . . . . . . . . . . Suction Gage Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Horizon Gyro Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Directional Gyro Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 37-10-00-960
Vacuum System Central Air Filter Discard
37-10-00 Page 601
37-10-00-961
Vacuum Relief Valve Filter Discard
37-10-00 Page 601
37 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL VACUUM - GENERAL 1.
Scope A.
2.
This chapter describes those units and components used to provide suction (vacuum) necessary to operate the horizon and directional indicator gyros.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief de finition of the sections incorporated in this chapter is as follows: (1) The section on distribution describes those components used in the distribution of vacuum air. (2) The section on indicating describes those components used to indicate relative vacuum pressure in the system.
37-00-00 © Cessna Aircraft Company
Page 1 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL VACUUM SYSTEM - DESCRIPTION AND OPERATION 1.
2.
General A.
This section provides description and operation information on the components used to distribute and indicate vacuum air (suction) for use in the horizon and directional gyros.
B.
For vacuum system schematic, refer to Figure 1.
Description and Operation A.
The vacuum system consists of a vacuum ejector, vacuum relief valve, air filter, suction gage, lowvacuum warning switch and low-vacuum warning annunciator light. The vacuum system furnishes vacuum air (suction) for operation of horizon gyro and directional gyro. A brief description of the system components follows: (1) The bleed air pressure regulator provides regulated bleed air for the vacuum system, (Refer to Chapter 36, Pneumatic System - General). (2) Bleed air flowing through an orifice in the vacuum ejector located on left firewall creates the necessary suction to operate instruments. (3) The vacuum relief valve incorporates an adjustment to obtain correct vacuum for proper system operation and is located on left aft side of firewall. (4) The air filter provides continual filtering for proper operation of vacuum system and is located on left aft side of firewall. (5) The suction gage, located on left side of instrument panel is calibrated in inches of mercury and indicates suction available for operation of horizon and directional gyro indicators. (6) A red vacuum low warning light is installed on the annunciator panel to warn pilot of a possible low vacuum condition existing in the vacuum system. Illumination of light is caused when suction is less than approximately 3.0inches Hg. and activation of the warning switch occurs. (7) The horizon gyro indicator is mounted in the left removable flight panel and provides the pilot with a visual indication of the airplane's pitch and roll attitude with respect to the earth. Optional horizon gyro will also provide the autopilot with electrical roll and pitch signals. (8) The directional gyro indicator is mounted directly below the horizon gyro indicator and displays a stable indication of the airplane heading to the pilot when properly set to agree with the magnetic compass. Optional directional gyros also provide the autopilot with electrical heading information.
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Page 1 Mar 3/1997
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MODEL 208 MAINTENANCE MANUAL
Vacuum System Schematic Figure 1 (Sheet 1)
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Page 2 Mar 3/1997
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Vacuum System Schematic Figure 1 (Sheet 2)
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Page 3 Mar 3/1997
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL VACUUM SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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MODEL 208 MAINTENANCE MANUAL
Vacuum System Troubleshooting Figure 101 (Sheet 1)
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Vacuum System Troubleshooting Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL VACUUM DISTRIBUTION - MAINTENANCE PRACTICES 1.
General A.
This section provides information on removal and installation of components used in the vacuum distribution system, as well as procedures used to connect the vacuum system to shop air and adjustment of the vacuum relief valve. NOTE:
When replacing a vacuum system component, ensure all connections are made correctly to avoid damage to gyro system. When a component is removed, cap off and identify all open lines, hoses, and fittings to prevent dirt from entering system, and to ensure proper reinstallation. Upon component replacement, check all hoses carefully to be sure they are clean and free of debris, oil, solvent, collapsed inner liners, and external damage. Replace old, hard, cracked, or brittle hoses.
CAUTION: Do not use teflon tape, pipe dope, or thread lubricants of any type on fitting threads, and avoid over tightening connections B. 2.
3.
For replacement of the vacuum system air filter, refer to Chapter 12, Vacuum System Central Air Filter - Servicing.
Vacuum Ejector Removal/Installation A.
Remove Vacuum Ejector (Refer to Figure 201). (1) Open left engine cowling door. (2) Detach line (1) from regulator (6) and union (7) or deice line tee if installed (20800001 Thru 20800143 and 208B0001 Thru 208B0143); detach line (1) from regulator (6) and cross fitting (13A) (20800144 and On and 208B0001 Thru 208B0143 incorporating CAB90-14); or detach line (1) from regulator (6) and cross fitting (13D) (20800222 and On and 20800001 Thru 20800121 incorporating CAB93-2; and 208B0317 and On:, 208B0001 Thru 208B0316 incorporating CAB93-2). (3) Disconnect exhaust fitting (9) from ejector (8). (4) Disconnect tube nut (13) from fitting (10). (5) Unscrew nut (12) securing ejector to firewall and retain washers (11) for reinstallation. (6) Remove ejector (8) from firewall.
B.
Install Vacuum Ejector (Refer to Figure 201 ). (1) Assemble washer (11) and insert fitting (10) through firewall. (2) Assemble washer (11) and nut (12) securing ejector to firewall. (3) Connect tube nut (13) to fitting (10) and tighten. (4) Connect exhaust fitting (9) to ejector (8). (5) Connect line (1) from regulator (6) and union (7) or deice line tee if installed (20800001 Thru 20800143 and 208B0001 Thru 208B0143); connect line (1) from regulator (6) and cross fitting (13A) (20800144 and On and 208B0001 Thru 208B0143 incorporating CAB90-14); or connect line (1) from regulator (6) and cross fitting (13D) (20800222 and On and 20800001 Thru 20800121 incorporating CAB93-2; and 208B0317 and On:, 208B0001 Thru 208B0316 incorporating CAB93-2). (6) Close and secure left engine cowling door.
Vacuum Relief Valve Removal/Installation A.
Remove Vacuum Relief Valve (Refer to Figure 201 ). (1) Loosen two clamps (16) and slide hoses (15) from relief valve (14). (2) Detach tube nut (13) and remove relief valve.
B.
Install Vacuum Relief Valve (Refer to Figure 201 ). (1) Replace relief valve (14) attach and tighten tube nut (13). (2) Slip hoses (15) onto relief valve and tighten two clamps (16).
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4.
Connecting Shop Air to Vacuum System NOTE:
5.
Refer to Chapter 36, Pneumatic Distribution - Maintenance Practices, for procedures on connecting shop air to pneumatic system.
Vacuum Relief Valve Adjustment A.
Adjustment Procedures (Refer to Figure 201 ). (1) Start engine according to procedures outlined in the Pilot's Operating Handbook and FAA Approved Flight Manual. (2) With engine operating at 68% Ng, the suction gage should read 5.0 inches of mercury. (3) If not, adjust valve by straightening tabs on knurled lock nut (17) and making necessary adjustment to obtain desired reading. Clockwise rotation will increase vacuum, and counterclockwise rotation will decrease vacuum. (4) After adjusting, bend knurled locknut tabs up. (5) Shut down engine according to procedures outlined in the Pilot's Operating Handbook and FAA Approved Flight Manual.
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MODEL 208 MAINTENANCE MANUAL VACUUM DISTRIBUTION - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the vacuum distribution system in a serviceable condition.
Task 37-10-00-960 2.
Vacuum System Central Air Filter Discard
CAUTION: Do not operate the vacuum system with the filter removed or a vacuum line disconnected. Dust and other foreign objects can enter the system and damage the vacuum operated instruments. A.
General (1) This task gives the instructions to discard the vacuum system central air filter.
B.
Special Tools (1) None
C.
Access (1) None
D.
Discard the Vacuum System Central Air Filter. (1) Remove the vacuum system central air filter. Refer to Chapter 12, Vacuum System Central Air Filter - Servicing. (a) Discard the filter. (2) Install a new vacuum system central air filter. Refer to Chapter 12, Vacuum System Central Air Filter - Servicing.
E.
Restore Access (1) None End of task Task 37-10-00-961 3.
Vacuum Relief Valve Filter Discard
CAUTION: Do not operate the vacuum system with the filter removed or a vacuum line disconnected. Dust and other foreign objects can enter the system and damage the vacuum operated instruments. A.
General (1) This task gives the instructions to discard the vacuum relief valve filter.
B.
Special Tools (1) None
C.
Access (1) None
D.
Discard the Vacuum Relief Valve Filter. (1) Get access to the relief valve behind the attitude gyro. (2) Carefully stretch the foam element filter over the top of the retaining bezel. (3) Remove the filter from the relief valve and discard it. (4) Stretch a new relief valve filter over the top of the retaining bezel. (5) Make sure that the filter is secure on the relief valve.
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Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL VACUUM INDICATING - MAINTENANCE PRACTICES 1.
General A.
2.
This section provides a brief description of the components and instruments used to indicate vacuum; and removal/installation instructions for the various components.
Description and Operation A.
The required components that provide an indication for operation of the vacuum system include Vacuum-Low warning switch, suction gage, Vacuum-Low warning annunciator light, horizon gyro indicator and directional gyro indicator. (1) The Vacuum-Low warning switch provides electrical activation for the annunciator light when suction is less than approximately 2.5 to 3.5 in. Hg. (2) The suction gage indicates in inches of mercury for the vacuum system. (3) The Vacuum-Low warning annunciator light may illuminate, warning the pilot of a possible Vacuum-Low condition. For annunciator lamp replacement, refer to Chapter 31, Master Warning and Annunciator Panel - Maintenance Practices. (4) The horizon indicator gyro indicates the pitch and roll ßight attitude of the airplane in relationship to the earth. NOTE:
(5) 3.
4.
5.
The horizon and directional gyro indicators represented in this chapter are standard equipment. Refer to Model 208 Avionic Installations Manual for alternate gyro systems used with 400B Autopilot and 400B IFCS installations.
The directional gyro indicator displays the airplane heading when properly set to agree with the magnetic compass.
Vacuum-Low Warning Switch Removal/Installation (Model 208) A.
Remove Vacuum-Low Warning Switch (Refer to Figure 201). (1) Remove electrical leads (3) and tag for reinstallation. (2) Loosen clamp (2) and slide hose (1) from Vacuum-Low warning switch (4).
B.
Install Vacuum-Low Warning Switch (Refer to Figure 201). (1) Slip hose (1) over switch (4) and tighten clamp (2). (2) Install electrical leads (3) and remove tags.
Vacuum-Low Warning Switch Removal/Installation (Model 208B) A.
Remove Vacuum-Low Warning Switch (Refer to Figure 201 ). (1) Remove electrical leads (3) and tag for reinstallation. (2) Loosen clamp (2) and slide hose (1) from Vacuum-Low warning switch (4).
B.
Install Vacuum-Low Warning Switch (Refer to Figure 201). (1) Replace electrical leads (3) and remove tags. (2) Slip hose (1) over switch (4) and tighten clamp (2).
Suction Gage Removal/Installation A.
Remove Suction Gage (Refer to Figure 201). (1) Remove screws securing removable panel and slide the panel forward to gain access to the back of the suction gage. (2) Reach through opening and loosen clamps (6) and slide hoses (5) off gage (7). Loosen clamps (2) and remove Vacuum-Low warning switch (4). (3) Remove screws (8) and carefully remove suction gage through opening.
B.
Install Suction Gage (Refer to Figure 201 ). (1) Position suction gage through opening and secure to instrument panel using screws (8). (2) Slide hoses (5) over gage (7) and tighten clamps (6). Slide hose (1) over gage (7) and tighten clamps (2).
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MODEL 208 MAINTENANCE MANUAL (3) 6.
7.
Slide removable ßight panel back against instrument panel and tighten screws.
Horizon Gyro Removal/Installation A.
Remove Horizon Gyro (Refer to Figure 201). (1) Remove screws securing removable panel and slide panel forward to gain access to back of gyro (11). (2) Loosen clamps (10) and slide hoses (9) off gyro (11). (3) Remove screws (12) securing gyro to instrument panel and lift gyro from panel.
B.
Install Horizon Gyro (Refer to Figure 201). (1) Position gyro (11) in instrument panel and secure using screws (12). (2) Slide hoses (9) onto gyro (11) and tighten clamps (10). (3) Slide removable panel back against instrument panel and tighten screws. (4) (For airplanes equipped with KFC-225 autopilot.) If a new unit is installed or the unit is calibrated, do a system alignment. Refer to Introduction, the List of Manufacturers Technical Publications for the manufacturer's installation manual.
Directional Gyro Removal/Installation A.
Remove Directional Gyro (Refer to Figure 201). (1) Remove screws securing removable panel and slide panel forward to gain access to back of gyro (14). (2) Loosen clamps (15) and slide hoses (16) off gyro (14). (3) Remove screws (13) securing gyro to instrument panel and lift gyro from panel.
B.
Install Directional Gyro (Refer to Figure 201). (1) Position gyro (14) in instrument panel and secure using screws (13). (2) Slide hoses (16) onto gyro (14) and tighten clamps (15). (3) Slide removable panel back against instrument panel and tighten screws.
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38 CHAPTER
WATER/WASTE
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
38-00-00
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38-30-01
Pages 201-202
Aug 1/1995
38-Title 38-List of Effective Pages 38-Record of Temporary Revisions 38-Table of Contents
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CONTENTS WATER/WASTE - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RELIEF TUBE - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Relief Tube Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Relief Tube Servicing and Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
38-00-00 38-00-00 38-00-00 38-00-00
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38 - CONTENTS © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL WATER/WASTE - GENERAL 1.
Scope A.
2.
This chapter contains information on the systems used to dispose of waste.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Liquid Detergent
Commercially Available
To clean relief tube assembly after use.
Lysol
Commercially Available
To disinfect relief tube assembly after use.
3.
Definition A.
This chapter provides removal, installation and cleaning instructions for the optional relief tube mounted in the aft cargo area.
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MODEL 208 MAINTENANCE MANUAL RELIEF TUBE - MAINTENANCE PRACTICES 1.
General A.
2.
Tools, Equipment and Materials A.
3.
Airplanes may be equipped with an optional relief tube, located in the right side panel of the aft cargo area. The relief tube vents liquid waste overboard through a venturi tube in the outer skin. This section covers removal, installation and servicing of the relief tube.
For a list of required tools, equipment and materials, refer to Water/Waste - General.
Relief Tube Assembly Removal/Installation A.
Remove Relief Tube Assembly (Refer to Figure 201). (1) Remove access plate located in floorboard adjacent to relief tube storage compartment. (2) Remove clamp (4) that secures hose (3) to drain assembly (5) located in skin. (3) Open compartment door and lift cup (1) and hose (3) out. (4) Detach tie (2) securing hose (3) to cup (1). (5) Loosen clamp (4) on airplanes without cargo pod to remove complete hose. NOTE:
B.
Install Relief Tube Assembly (Refer to Figure 201). (1) Attach hose (3) to cup (1) using tie (2). (2) Route hose (3) through hole in relief tube storage compartment. (3) Attach hose (3) to drain assembly (5) using clamp (4). (4) Reinstall access plate in floorboard. NOTE:
4.
On airplane with a cargo pod, clamp (7) secured by screw (8), washer (9), and nut (10) must be loosened, and clamp (4) must be loosened by gaining access thru the adjacent cargo compartment door in order to completely remove the hose (3).
On airplane with a cargo pod, complete items (1) and (2) and then route hose thru clamp (7), grommet (6) and attach to drain assembly (5) with clamp (4). Secure clamp (4) by tightening screw (8), washer (9) and nut (10). Close cargo compartment door.
Relief Tube Servicing and Cleaning A.
After relief tube assembly has been used, rinse cup (1) and hose (3) thoroughly with a warm watermild detergent solution. Treat with a commercially available disinfectant (Lysol or equivalent). Allow cup to dry completely before using.
B.
Periodically (unconditional monitoring), the relief tube assembly and hose should be removed and thoroughly scoured with water-soap solution, then rinsed, and then treated with a commercially available disinfectant (Lysol or equivalent). NOTE:
If the airplane is to be stored or parked for an extended period of time, then the relief tube should be thoroughly cleaned and sterilized to prevent bacterial growth.
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CHAPTER
STANDARD PRACTICES STRUCTURES
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MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
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DATE
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51-01-00
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51-01-05
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CONTENTS STRUCTURES - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51-00-00 Page 1 51-00-00 Page 1 51-00-00 Page 1
CORROSION PREVENTION AND CONTROL PROGRAM - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Corrosion Prevention and Control Program Objective. . . . . . . . . . . . . . . . . . . . . . . . . . . Corrosion Prevention and Control Program Function . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Prevention and Control Program Application . . . . . . . . . . . . . . . . . . . . . . . . . . . Baseline Program . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Baseline Program Implementation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reporting System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Periodic Review . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Corrosion Related Airworthiness Directives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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CORROSION PREVENTION AND CONTROL PROGRAM (APPENDIX) DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Appendix A - Development Of The Baseline Program . . . . . . . . . . . . . . . . . . . . . . . . . . Appendix B - Procedures For Recording Inspection Results . . . . . . . . . . . . . . . . . . . . . Appendix C - Guidelines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Application Of The Corrosion Program Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Determination of the Corrosion Levels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Typical Actions That Follow the Determination of the Corrosion Level. . . . . . . . . . . . Factors Influencing Corrosion Occurrences . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reporting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Program Implementation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Page 1 Page 1 Page 1 Page 1 Page 2 Page 2 Page 3 Page 4 Page 5 Page 7 Page 7
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CORROSION - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Types of Corrosion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Typical Corrosion Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Corrosion Detection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Corrosion Repair. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51-11-00 51-11-00 51-11-00 51-11-00 51-11-00 51-11-00
CORROSION SEVERITY MAPS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL STRUCTURES - GENERAL 1.
Scope A.
2.
This chapter provides a description of general airplane structures and corrosion characteristics. For repair of structural members and repair techniques used throughout the airplane, refer to the Model 208 Series Structural Repair Manual.
Definition A.
This chapter is divided into two sections briefly described below. (1) The section on structures provides an overall description of the airplane structure and methods of construction used on the airplane. (2) The section on corrosion provides a general description of corrosion characteristics, types of corrosion and typical corrosion areas.
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MODEL 208 MAINTENANCE MANUAL CORROSION PREVENTION AND CONTROL PROGRAM - DESCRIPTION AND OPERATION 1.
2.
Introduction A.
As the airplane ages, corrosion occurs more often, while, at the same time, other types of damage such as fatigue cracks occur. Corrosion can cause damage to the airplane’s structural integrity, and if it is not controlled, the airframe will carry less load than what is necessary for continued airworthiness. (1) To help to prevent this, we started a Corrosion Prevention and Control Program (CPCP) for the Models 208 and 208B. A CPCP is a system to control the corrosion in the airplane’s primary structure. It is not the function of the CPCP to stop all of the corrosion conditions, but to control the corrosion to a level that the airplane's continued airworthiness is not put in risk.
B.
The initial and repeat inspection intervals are based on Model 208/208B service experience. It is recommended that you record the results of the inspection. After the second CPCP inspection, it is possible to shorten or lengthen the Repeat Interval (RI). Proposed changes to the RI should be submitted with supporting data to the regulatory authority.
Corrosion Prevention and Control Program Objective A.
3.
The objective of the CPCP is to help to prevent or control the corrosion so that it does not cause a risk to the continued airworthiness of the Models 208 and 208B airplanes.
Corrosion Prevention and Control Program Function A.
The function of this document is to give the minimum procedures necessary to control the corrosion so that the continued airworthiness is not put in risk. The CPCP consists of a Corrosion Program Inspection number, the area where the inspection will be done, specified corrosion levels and compliance times (Implementation Thresholds and Repeat Intervals). The CPCP also includes procedures to let Cessna Aircraft Company and the regulatory authorities know of the findings and the data associated with Level 2 and Level 3 corrosion. This includes the actions that were done to decrease possible corrosion in the future to Level 1.
B.
Maintenance or inspection programs need to include a good quality CPCP. The level of corrosion identified on the Principal Structural Elements (PSEs) and other structure listed in the Baseline Program will help make sure the CPCP provides good corrosion protection. NOTE:
C.
A good quality program is one that will control the corrosion of all of the structure at Level 1 or better.
Corrosion Program Levels. NOTE: (1)
(2)
(3)
In this manual the corrosion inspection task are referred to as the corrosion program inspection.
Level 1 Corrosion. (a) Corrosion damage occurring between successive inspection tasks, that is local and can be reworked or blended out with the allowable limit. (b) Local corrosion damage that exceeds the allowable limit but can be attributed to an event not typical of the operator’s usage or other airplanes in the same fleet (e.g., mercury spill). (c) Operator experience has demonstrated only light corrosion between each successive corrosion task inspection; the latest corrosion inspection task results in rework or blend out that exceeds the allowable limit. Level 2 Corrosion. (a) Level 2 corrosion occurs between two successive corrosion inspection tasks that requires a single rework or blend-out that exceeds the allowable limit. A finding of Level 2 corrosion requires repair, reinforcement or complete or partial replacement of the applicable structure. Level 3 Corrosion. (a) Level 3 corrosion occurs during the first or subsequent accomplishments of a corrosion inspection task that the operator determines to be an urgent airworthiness concern.
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4.
References A.
5.
This is a list of references for the Corrosion Prevention and Control Program. (1) FAA Advisory Circular AC120-CPCP, Development and Implementation of Corrosion Prevention and Control Program (2) FAA Advisory Circular AC43-4A, Corrosion Control for Aircraft (3) Cessna Model 208 Illustrated Parts Catalog (4) Cessna Model 208 Maintenance Manual (5) Cessna Model 208 Structural Repair Manual, Chapter 51-11-00, Corrosion Repair (6) Cessna Model 208 Nondestructive Testing Manual
Control Prevention and Control Program Application A.
The Models 208 and 208B Corrosion Prevention and Control Program gives the information required for each corrosion inspection. Maintenance personnel must fully know about corrosion control. The regulatory agency will give approval and monitor the CPCP for each airplane. (1) The CPCP procedures apply to all Models 208 and 208B airplanes that are at or have more than the Baseline Program Implementation Threshold for each location on the airplane. Refer to the Glossary and the Baseline Program. (a) Cessna Aircraft Company recommends that the CPCP be done first on older airplanes and areas that need greater changes to the maintenance procedures to meet the necessary corrosion prevention and control requirements. (2) Maintenance programs must include corrosion prevention and control procedures that limit corrosion to Level 1 or better on all Principal Structural Elements (PSEs) and other structure specified in the Baseline Program. If the current maintenance program includes corrosion control procedures in an inspection area and there is a report to show that corrosion is always controlled to Level 1 or better, the current inspection program can be used. (a) The Baseline Program is not always sufficient if the airplane is operated in high humidity environments, has a corrosive cargo leakage or has had an unsatisfactory maintenance or repair. When this occurs, make adjustments to the Baseline Program until the corrosion is controlled to Level 1 or better. (3) The CPCP consists of the corrosion inspection applied at a specified interval, and, at times, a corrosion inspection interval can be listed in a Service Bulletin. For the CPCP to be applied, remove all systems, equipment and interior furnishings that prevent sufficient inspection of the structure. A nondestructive test (NDI) or a visual inspection can be necessary after some items are removed if there is an indication of hidden corrosion such as skin deformation, corrosion under splices or corrosion under fittings. Refer to the Baseline Program. (4) The corrosion rate can change between different airplanes. This can be a result of different environments the airplane operates in, flight missions, payloads, maintenance practices (for example more than one owner), variation in rate of protective finish or coating wear. (a) Some airplanes that operate under equivalent environments and maintenance practices can be able to extend the Implementation Threshold (IT) or Repeat Interval (RI) if a sufficient number of inspections do not show indications of corrosion in that area. Refer to the Glossary. (5) Later design and/or production changes done as a result of corrosion conditions can delay the start of corrosion. Operators that have done corrosion-related service bulletins or the improved maintenance manual procedures listed in the Corrosion Program Inspection can use that specified implementation threshold or repeat inspection interval. Unless the instructions tell you differently, the requirements given in this document apply to all of the Models 208 and 208B airplanes. (6) Another system has been added to report all Level 2 and Level 3 corrosion conditions identified during the second and each subsequent CPCP inspection. This information will be reviewed by Cessna Aircraft Company to make sure the Baseline Program is sufficient and to change it as necessary.
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Baseline Program A.
The Baseline Program is part of the Models 208 and 208B Corrosion Prevention and Control Program (CPCP). It is divided into Basic Task, Implementation Threshold (IT) and Repeat Interval (RI). In this manual the Basic Tasks are referred to as the Corrosion Program Inspection. This program is to be used on all airplanes without an approved CPCP. Those who currently have a CPCP that does not control corrosion to Level 1 or better must make adjustments to the areas given in the Baseline Program.
B.
Typical Airplane Zone Corrosion Program Inspection Procedures. (1) Remove all the equipment and airplane interior (for example the insulation, upper upholstery panel, lower upholstery panel) as necessary to do the corrosion inspection. (2) Clean the areas given in the corrosion inspection before you inspect them. (3) Do a visual inspection of all of the Principal Structural Elements (PSEs) and other structure given in the corrosion inspection for corrosion, cracking, and deformation. (a) Carefully examine the areas that show that corrosion has occurred before. NOTE:
Areas that need a careful inspection are given in the corrosion inspection.
(b)
(4)
(5) (6)
(7)
(8)
Nondestructive testing inspections or visual inspections can be needed after some disassembly if the inspection shows a bulge in the skin, corrosion under the splices or corrosion under fittings. Remove all of the corrosion, examine the damage, and repair or replace the damaged structure. (a) Apply a protective finish where it is required. Refer to Chapter 20, Interior and Exterior Finish - Cleaning/Painting or Chapter 51, Corrosion - Description and Operation. (b) Clean or replace the ferrous metal fasteners with oxidation. Remove blockages of foreign object debris so that the holes and clearances between parts can drain. For bare metal on any surface of the airplane, apply fuel and corrosion resistant primer MILPRF-23377. (a) Apply a polyurethane topcoat paint to the exterior painted surface. Refer to the manufacturer's procedures. Cessna Aircraft Company recommends that you apply a corrosion preventive compound once every two years to areas with a high possibility for severe corrosion identified in the corrosion inspection. (a) On the Model 208, apply LPS-3 Heavy-Duty Rust Inhibitor, or equivalent, in the bilge area below the floorboards between FS 168.00 and FS 211.00. Refer to Chapter 51, Corrosion - Description and Operation. (b) On the Model 208B, apply LPS-3 Heavy-Duty Rust Inhibitor, or equivalent, in the bilge area below the floorboards between FS 168.00 and FS 356.00. Refer to Chapter 51, Corrosion - Description and Operation Apply compounds that will replace water and prevent corrosion. Refer to Chapter 51, Corrosion - Description and Operation. (a) Apply one layer of LPS-3 Heavy-Duty Rust Inhibitor, or equivalent, that will soak into the fayed surfaces to replace water and prevent corrosion.
Table 1. Areas and Items Not to Apply Compounds to Replace Water and Prevent Corrosion AREA or ITEM Oxygen System Lines and Components Cables, Pulleys, and Trim Tab Pushrod Plastics, Elastomers Lubricated and Teflon Surfaces (Greased Joints, Sealed Bearings, and Grommets) Adjacent to Tears and Holes in Insulation (Not Waterproof)
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MODEL 208 MAINTENANCE MANUAL Table 1. Areas and Items Not to Apply Compounds to Replace Water and Prevent Corrosion (continued) Areas with Electrical Arc Potential, Wiring Interior Upholstery Panels (Changes the Flammability Properties) Cargo Pod (Changes the Flammability Properties) Pitot Tubes Fuel Cap Tie Down Lugs Chrome Items (handles, locks) Standard Polished Spinner Stall Warning Detector (9) Install the dry insulation blankets. (10) Install the equipment and airplane interior (for example the upper upholstery panel and lower upholstery panel) that was removed to do the corrosion inspection. 7.
Baseline Program Implementation A.
The Baseline Program is divided into specific inspection areas and zone locations. Both of these items have an Implementation Threshold (IT) and Repeat Interval (RI) to do the corrosion program inspection. The inspection areas and zone locations apply to all 208/208B airplanes for which the given airplane calendar age is equal to or greater than the IT. For the structural parts that have been replaced, the IT can be calculated from the time of the installation of the new part. Refer to Table 2 and Table 3.
B.
Calculate the Implementation Threshold (IT). NOTE: (1) (2)
The corrosion inspections apply to airplanes with a calendar year age that is equal to or greater than the IT.
To calculate the IT on a new structural part, start at the time the structural part was replaced. Start the Baseline Program during scheduled maintenance when the maintenance check intervals are equal to or greater than the IT. NOTE:
The Baseline Program must be started before the airplane gets to the point of IT + RI for any corrosion inspection.
NOTE:
For airplanes near or greater than IT + RI, Cessna Aircraft Company recommends that you apply the Baseline Program immediately after the regulatory authority approves the schedule.
NOTE:
Early implementation of the Baseline Program is highly recommended.
Table 2. Corrosion Prevention and Control Program Example CORROSION INSPECTION NUMBER
INTERVAL (YEARS) IT
ZONE
ACCESS
RI
MAINTENANCE MANUAL REFERENCE SECTION
CORROSION INSPECTION DESCRIPTIONS
C32.701.01E
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Table 3. Corrosion Program Inspection Legend ITEM
DESCRIPTION
CORROSION INSPECTION NUMBER
The Corrosion Inspection Number identifies the area on the airplane where you will do the inspection. Each Corrosion Inspection Number consists of an ATA code, airplane zone, and the item number.
EXAMPLE: C32.701.01E C
The first letter identifies this as a Corrosion Prevention and Control Program inspection.
32
This is the ATA code (Chapter 32).
701
This is the airplane zone between 100 and 999.
01
This is the sequence number between 1 and 99.
E
This letter identifies this as an external inspection. An I would identify it as an internal inspection.
INTERVAL IMPLEMENTATION THRESHOLD (IT)
The specified Implementation Threshold for an airplane is the date that you must do the initial corrosion inspection in an area and zone. Use the age of the airplane to find the IT.
REPEAT INTERVAL (RI)
The Repeat Interval is the time interval between the successive corrosion inspections for an area and zone.
ZONE
This is the airplane zone number in which the inspection will be done. Refer to Chapter 6, Airplane Zoning - Description and Operation.
ACCESS
This is the list of access plates or panels that you need to remove to do a corrosion inspection. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
MAINTENANCE MANUAL REFERENCE SECTION
This gives the location of the removal and installation procedures in the maintenance manual.
CORROSION INSPECTION DESCRIPTION
This describes the inspection to be done.
8.
Reporting System A.
Corrosion Prevention and Control Program Reporting System (Refer to Figure 1). (1) The Corrosion Prevention and Control Program (CPCP) includes a system to report to Cessna Aircraft Company data that will show that the Baseline Program is sufficient and, if necessary, make changes. (2) At the start of the second Corrosion Program Inspection of each area, report all Level 2 and Level 3 Corrosion results that are listed in the Baseline Program to Cessna Aircraft Company. Send the Control Prevention and Control Program Damage Reporting Form to:
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Corrosion Prevention and Control Program Damage Report Form Figure 1 (Sheet 1)
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Cessna Aircraft Company Technical Support Services Department 751 P.O. Box 7706 Wichita, Kansas USA 67277 9.
Periodic Review A.
10.
Use the Service Difficulty Reporting System to report all Level 2 and Level 3 Corrosion results to the FAA and to Cessna Aircraft Company. All corrosion reports received by Cessna Aircraft Company will be reviewed to determine if the Baseline Program is adequate.
Corrosion Related Airworthiness Directives A.
Safety-related corrosion conditions transmitted by a service bulletin can be mandated by an Airworthiness Directive (AD). The service bulletins and ADs will be listed in this section. There are no corrosion-related ADs for the Models 208 and 208B.
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MODEL 208 MAINTENANCE MANUAL CORROSION PREVENTION AND CONTROL PROGRAM (APPENDIX) - DESCRIPTION AND OPERATION 1.
Appendix A - Development Of The Baseline Program A.
2.
Appendix B - Procedures For Recording Inspection Results A.
3.
The Corrosion Prevention and Control Program Baseline Program. (1) The function of the Corrosion Prevention and Control Program (CPCP) is to give the minimum procedures necessary to prevent and control corrosion so that continued airworthiness is not risked. The Principle Structural Elements (PSEs) are areas where the CPCP applies. (2) The CPCP Baseline Program consists of a Corrosion Program Inspection (CPI), Implementation Threshold (IT), and a Repeat Interval (RI). Each inspection is to be done in an airplane zone. (3) The corrosion reports that are sent to Cessna Aircraft Company and data from the FAA Service Difficulty Records from 1995 to 2002 were used to identify the inspection areas of the Baseline Program. When more than one incident of corrosion was identified at a specified location, an inspection was included for that location in the Baseline Program. (4) When corrosion was found once, the data was examined to find if the corrosion was caused by one specified occurrence or if other airplanes could have corrosion in the same location. If so, this inspection was added to the Baseline Program. (5) The inspection interval was specified by the duration and corrosion severity.
Record the Inspection Results. (1) It is not an FAA mandatory procedure to record the CPCP results, but Cessna Aircraft Company recommends that records be kept to assist in program adjustments when necessary. The inspection of records will make sure the identification, repeat, and level of corrosion is monitored. The data can identify whether there is more or less corrosion at repeat intervals. The data can also be used to approve increased or decreased inspection intervals.
Appendix C - Guidelines A.
Glossary (1) The following additional information clarifies the previous sections of this document.
B.
The Glossary of General Descriptions. Refer to Table 1, Figure 1, Figure 2, Figure 3, and Figure 4. WORD
GENERAL DESCRIPTION
Allowable Limit
The allowable limit is the maximum amount of material (usually expressed in material thickness) that may be removed or blended out without affecting the ultimate design strength capability of the structural member. Allowable limits may be established by the design approval holder. The FAA (or applicable regulatory authority) may also establish allowable limits. The design approval holder normally publishes allowable limits in the structural repair manual or in service bulletins.
Baseline Program
A Baseline Program is a CPCP developed for a specific model airplane. The design approval holder typically develops the Baseline Program. However, it may be developed by a group of operators who intend to use it in developing their individual CPCP. It contains the corrosion program inspection, an implementation threshold, and a repeat interval for the procedure accomplishment in each area or zone.
Basic Task
Refer to corrosion program inspection.
Corrosion Program Inspection (CPI)
The corrosion program inspection (CPI) is a specific and fundamental set of work elements that should be performed repetitively in all task areas or zones to successfully control corrosion. The contents of the CPI may vary depending upon the specific requirements in an airplane area or zone. The CPI is developed to protect the primary structure of the airplane.
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WORD
GENERAL DESCRIPTION
Corrosion (Metal)
The physical deterioration of metals caused by reaction to an adverse environment.
Corrosion Prevention and Control Program (CPCP)
A Corrosion Prevention and Control Program is a comprehensive and systematic approach to controlling corrosion such that the load carrying capability of an airplane structure is not degraded below a level necessary to maintain airworthiness. It contains the corrosion program inspections, a definition of corrosion levels, implementation thresholds, a repeat interval for task accomplishment in each area or zone, and specific procedures if corrosion damage exceeds Level 1 in any area or zone.
Design Approval Holder
The design approval holder is either the type certificate holder for the aircraft or the supplemental type certificate holder.
Implementation Threshold (IT)
The implementation threshold for a specific airplane is the date, based on that airplane’s age, by which the initial corrosion inspection task should be accomplished in an area or zone.
Inspection Area
The inspection area is a region of airplane structure to which one or more CPIs are assigned. The inspection area may also be referred to as a Zone.
Level 1 Corrosion
Level 1 Corrosion is one or more of the items that follow: 1. Corrosion damage occurring between successive inspections, that is local and can be reworked or blended out with the allowable limit. 2. Local corrosion damage that exceeds the allowable limit but can be attributed to an event not typical of the operator’s usage or other airplanes in the same fleet (e.g., mercury spill). 3. Operator experience has demonstrated only light corrosion between each successive corrosion task inspection; the latest corrosion inspection task results in rework or blend out that exceeds the allowable limit.
Level 2 Corrosion
Level 2 corrosion occurs between two successive corrosion inspection tasks that requires a single rework or blend-out that exceeds the allowable limit. A finding of Level 2 corrosion requires repair, reinforcement or complete or partial replacement of the applicable structure.
Level 3 Corrosion
Level 3 corrosion occurs during the first or subsequent accomplishments of a corrosion inspection task that the operator determines to be an urgent airworthiness concern. NOTE:
If Level 3 corrosion is determined at the implementation threshold or any repeat inspection, it should be reported. Any corrosion that is more than the maximum acceptable to the design approval holder or the FAA (or applicable regulatory authority) must be reported in accordance with current regulations. This determination should be conducted jointly with the design approval holder.
Light Corrosion
Light corrosion is corrosion damage so slight that removal and blendout over multiple repeat intervals (RI) may be accomplished before material loss exceeds the allowable limit.
Local Corrosion
Generally, local corrosion is corrosion of a skin or web (wing, fuselage, empennage, or strut) that does not exceed one frame, stringer, or stiffener bay. Local corrosion is typically limited to a single frame, chord, stringer, or stiffener, or the corrosion of more than one frame, chord, stringer, or stiffener where no corrosion exists on two adjacent members on each side of the corroded member.
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WORD
GENERAL DESCRIPTION
Principal Structural Element (PSE)
A PSE is an element that contributes significantly to carrying flight, ground or pressurization loads, and whose integrity is essential in maintaining the overall structural integrity of the airplane.
Repeat Interval (RI)
The repeat interval is the calendar time between the accomplishment of successive corrosion inspection tasks for a Task Area or Zone.
Task Area
Refer to Inspection Area.
Urgent Airworthiness Concern
An urgent airworthiness concern is damage that could jeopardize continued safe operation of any airplane. An urgent airworthiness concern typically requires correction before the next flight and expeditious action to inspect the other airplanes in the operator’s fleet.
Widespread Corrosion
Widespread corrosion is corrosion of two or more adjacent skin or web bays (a web bay is defined by frame, stringer, or stiffener spacing). Or, widespread corrosion is corrosion of two or more adjacent frames, chords, stringers, or stiffeners. Or, widespread corrosion is corrosion of a frame, chord, stringer, or stiffener and an adjacent skin or web bay.
Zone
Refer to Inspection Area.
4.
Application Of The Corrosion Program Inspection NOTE: A.
In this manual the Basic Tasks are referred to as the Corrosion Program Inspection (CPI).
Typical Airplane Zone Corrosion Program Inspection Procedures. (1) Remove all the equipment and airplane interior (for example the insulation, upper upholstery panel, lower upholstery panel) as necessary to do the corrosion inspection. (2) Clean the areas given in the corrosion inspection before you inspect them. (3) Do a visual inspection of all of the Principal Structural Elements (PSEs) and other structure given in the corrosion inspection for corrosion, cracking, and deformation. (a) Carefully examine the areas that show that corrosion has occurred before. NOTE:
Areas that need a careful inspection are given in the corrosion inspection.
(b)
(4)
(5) (6)
(7)
Nondestructive testing inspections or visual inspections can be needed after some disassembly if the inspection shows a bulge in the skin, corrosion under the splices or corrosion under fittings. Remove all of the corrosion, examine the damage, and repair or replace the damaged structure. (a) Apply a protective finish where it is required. Refer to Chapter 20, Interior and Exterior Finish - Cleaning/Painting or Chapter 51, Corrosion - Description and Operation. (b) Clean or replace the ferrous metal fasteners with oxidation. Remove blockages of foreign object debris so that the holes and clearances between parts can drain. For bare metal on any surface of the airplane, apply fuel and corrosion resistant primer MILPRF-23377. (a) Apply a polyurethane topcoat paint to the exterior painted surface. Refer to the manufacturer's procedures. Cessna Aircraft Company recommends that you apply a corrosion preventive compound once every two years to areas with a high possibility for severe corrosion identified in the corrosion inspection. (a) On the Model 208, apply LPS-3 Heavy-Duty Rust Inhibitor, or equivalent, in the bilge area below the floorboards between FS 168.00 and FS 211.00. Refer to Chapter 51, Corrosion - Description and Operation. (b) On the Model 208B, apply LPS-3 Heavy-Duty Rust Inhibitor, or equivalent, in the bilge area below the floorboards between FS 168.00 and FS 356.00. Refer to Chapter 51, Corrosion - Description and Operation
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Local Corrosion Found in Non-Adjacent Skin Panels Figure 1 (Sheet 1)
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Widespread Corrosion Found in Adjacent Skin Panels Figure 2 (Sheet 1)
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Local Corrosion Found in Non-Adjacent Frames Figure 3 (Sheet 1)
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Corrosion Found in Adjacent Frames Figure 4 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (8)
Apply compounds that will replace water and prevent corrosion. Refer to Chapter 51, Corrosion - Description and Operation. (a) Apply one layer of LPS-3 Heavy-Duty Rust Inhibitor, or equivalent, that will soak into the fayed surfaces to replace water and prevent corrosion.
Table 1. Areas and Items Not to Apply Compounds to Replace Water and Prevent Corrosion AREA or ITEM Oxygen System Lines and Components Cables, Pulleys, and Trim Tab Pushrod Plastics, Elastomers Lubricated and Teflon Surfaces (Greased Joints, Sealed Bearings, and Grommets) Adjacent to Tears and Holes in Insulation (Not Waterproof) Areas with Electrical Arc Potential, Wiring Interior Upholstery Panels (Changes the Flammability Properties) Cargo Pod (Changes the Flammability Properties) Pitot Tubes Fuel Cap Tie Down Lugs Chrome Items (handles, locks) Standard Polished Spinner Stall Warning Detector (9) Install the dry insulation blankets. (10) Install the equipment and airplane interior (upper upholstery panel, lower upholstery panel) that was removed to do the corrosion inspection. 5.
Determination of the Corrosion Levels A.
Find the Corrosion Levels (Refer to Figure 5). (1) Corrosion found on a structure when you use the Corrosion Program and Corrosion Prevention (CPCP) Baseline Program will help find the extent of the corrosion. (2) The second and subsequent inspections will find how well the CPCP program has been prepared, or if there is a need to make adjustments to the Baseline Program. (3) A good quality CPCP is one that controls corrosion to Level 1 or better. (4) If Level 2 corrosion is found during the second or subsequent inspection, you must do something to decrease the future corrosion to Level 1 or better. (5) If Level 3 corrosion is found, you must also do something to decrease the future corrosion to Level 1. Also, a plan to find or prevent Level 3 corrosion in the same area on other airplanes must be added to the CPCP. (6) All the corrosion that you can repair in the allowable damage limits, found in the Model 208 Structural Repair Manual, is Level 1 corrosion. (7) If all corrosion is Level 1, the CPCP is correctly prepared. (8) If you must reinforce or replace the part because of corrosion, the corrosion is Level 2. (9) If the part is not airworthy because of the corrosion, you must do an analysis to find out if the corrosion is Level 3. (10) The chart found in this section will help find the level of the corrosion. (11) The probability that the same problem will occur on another airplane is dependent on several factors such as: past maintenance history, operating environment, years in service, inspectability of the corroded area and the cause of the problem.
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Determination of Corrosion Level Figure 5 (Sheet 1)
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Determination of Corrosion Level Figure 5 (Sheet 2)
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Determination of Corrosion Level Figure 5 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL B.
6.
Level 2 Corrosion Findings. (1) All Level 2 corrosion that is more than the rework limits of the Structural Repair Manual must be reported to Cessna Aircraft Company. Cessna Aircraft Company engineering will do an analysis to make sure the corrosion is not a urgent airworthiness concern. (2) When doing the analysis, Cessna Aircraft Company will consider: (a) Can the cause of the corrosion be identified, such as a chemical spill or protective finish breakdown? (b) Has the same level of corrosion been found on other airplanes? (c) Are the corrosion protection procedures applied during manufacture the same for earlier and later models? (d) Age of the corroded airplane compared to others checked. (e) Is the maintenance history different than the other airplanes in the fleet?
Typical Actions That Follow the Determination of the Corrosion Level. A.
If corrosion is found, find the corrosion level, then do the necessary steps for a specific inspection.
B.
If Level 1 corrosion is found during the first CPCP inspection. (1) Repair the structure. Refer to the Model 208 Structural Repair Manual or a Cessna Aircraft Company approved repair procedure. (2) Continue with the Baseline Program. (a) Optional: Document the results of the inspection for use in validating program compliance.
C.
If Level 2 corrosion is found during the first CPCP inspection. (1) Repair the structure. Refer to the Model 208 Structural Repair Manual or Cessna Aircraft Company approved repair procedure. (2) Report the details of the corrosion you see to Cessna Aircraft Company and the FAA (or applicable regulatory authority). (3) Continue to use the Baseline Program but check the corroded area carefully when you do a subsequent CPCP inspection. (4) It is recommended that you record the results of the inspection to show compliance with the program.
D.
If Level 3 corrosion is found during the first CPCP inspection. (1) Immediately contact Cessna Aircraft Company and the FAA or regulatory authority of the corrosion you found. Refer to Reporting System. (2) Give sufficient information to make sure that the condition is a possible urgent airworthiness concern for your fleet. Get assistance from Cessna Propeller Aircraft Product Support to develop a plan of action. (3) Apply the corrosion program inspection, which includes the repair of the structure. Refer to the Model 208 Structural Repair Manual or a Cessna Aircraft Company approved repair procedure. (4) Do a report that has the information of the findings. Refer to Corrosion Prevention And Control Program Reporting System - Description And Operation. (5) Continue with the Baseline Program and other steps of procedure required by the FAA, or applicable regulatory authority. Examine this area carefully during future inspections.
E.
If no corrosion is found during the second or subsequent CPCP inspection: (1) Continue with the current Corrosion Prevention and Control Program. No adjustment of the current program is required. (2) It is recommended that you record the results of the inspection for a possible increase of the corrosion inspection Implementation Threshold (IT) and/or Repeat Interval (RI).
F.
If Level 1 corrosion is found on the second or subsequent CPCP inspection: (1) Do the corrosion program inspection, which includes the repair of the structure. Refer to the Model 208 Structural Repair Manual or a Cessna Aircraft Company approved repair procedure. (2) Continue with the Baseline Program. (3) No adjustment of the existing program is required. (4) It is recommended that you record the corrosion inspection number and the results of the inspection to show that the program was obeyed.
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MODEL 208 MAINTENANCE MANUAL G.
If Level 2 corrosion is found on the second or subsequent CPCP inspection: (1) Repair the structure. Refer to the Model 208 Structural Repair Manual or a Cessna Aircraft Company approved repair procedure. (2) Do a report that shows the information about the corrosion and send it to Cessna Aircraft Company and the FAA (or applicable regulatory authority). (3) If corrosion damage required the removal of material just beyond the allowable limits (within 10 percent), complete a check of the other airplanes in the fleet before you change the maintenance program. (a) If the corrosion is typical of Level 2, use the fleet data to find what changes are required to control corrosion to Level 1 or better. (b) If fleet damage is typically Level 1, examine the corroded area during subsequent inspections on all affected airplanes. (c) Make changes to the maintenance program if the typical corrosion becomes Level 2. (4) Further evaluation by Cessna Aircraft Company is recommended for Level 2 corrosion findings that are well beyond the allowable limits, and there is an airworthiness concern in which prompt action is required. NOTE:
(5)
(6) (7) H.
The airworthiness concern is because of the possibility to have similar but more severe corrosion on any other airplane in the operator's fleet prior to the next scheduled inspection of that area.
Find the action required to control the corrosion to a Level 1 or better, between future successive inspections. These can include the items that follow: (a) A structural modification, such as additional drainage. (b) Improvements to the corrosion prevention and control inspections, such as more care and attention to corrosion removal, reapplication of protective finish, drainage path clearance. (c) Decrease the Implementation Threshold (IT) for additional airplanes that go into the program. (d) Decrease the Repeat Interval (RI). Send a plan of corrective action to the FAA or applicable regulatory authority for approval and to Cessna Aircraft Company. Use the approved plan of action.
If Level 3 corrosion is found on the second or subsequent CPCP inspection: (1) contact Cessna Aircraft Company and the FAA (or applicable regulatory authority) about the corrosion that was found. (2) Send a plan to examine the same area on other affected airplanes in the operator's fleet. NOTE: (3)
Circumstances can dictate the need to examine airplanes younger than the corresponding Baseline Program Implementation Threshold.
Apply the corrosion program inspection, which includes the repair of the structure. Use the Model 208 Structural Repair Manual or a Cessna Aircraft Company approved repair procedure.
I.
Find the action needed to control the corrosion finding to Level 1 or better, between future successive inspections. These can include any or all of the following: (1) A structural modification, such as additional drainage. (2) Improvements to the corrosion prevention and control inspections, such as more care and attention to corrosion removal, reapplication of protective finish, drainage path clearance. (3) A decrease in the Implementation Threshold (IT) for additional airplanes entering the program. (4) A decrease in the Repeat Interval (RI).
J.
Send a plan of corrective action to the FAA (or applicable regulator authority for approval) as needed.
K.
Use the approved plan of action.
L.
It is recommended that you give the details of the findings to Cessna Aircraft Company.
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7.
Factors Influencing Corrosion Occurrences A.
8.
If you find Level 2 or Level 3 corrosion, when you think about how to change your CPCP, think about the list that follows: (1) Is there a presence of LPS-3 Heavy-Duty Rust Inhibitor? (2) Is there a presence or condition of protective finish? (3) What was the Implementation Threshold (IT) of the operator's Corrosion Control and Control Program (CPCP)? (4) What was the length of time since the last inspection and/or application of corrosion inhibiting compound Repeat Interval (RI)? (5) Was there inadequate clean-up/removal of corrosion prior to application of corrosion inhibiting compound, during previous maintenance of the area? (6) Are the moisture drains blocked or is there inadequate drainage? (7) What was the environment, the time of exposure to the environment and the use of the airplane? (a) Was the environment tropical, desert, salt water or industrial? (b) Are there electrolytes or water and moisture, salt water or battery fluid? (8) Was there a variation in past maintenance history and or use of the airplanes in the operator's fleet? (9) Were there variations in the production build standard in the operator's fleet?
Reporting A.
The minimum requirements to prevent or control the corrosion in the Corrosion Prevention and Control Program (CPCP) were made on the best information, knowledge and experience available at the time. As this experience and knowledge get better, the CPCPs will be changed at intervals as necessary. A reporting system for this is in Section 4.0. (1) You must contact the Cessna Aircraft Company about all Level 2 or 3 corrosion of the structure that is on the list in the Baseline Program that is found during the second and subsequent corrosion program inspections. Refer to Reporting System. NOTE:
9.
You do not have to contact the Cessna Aircraft Company about corrosion that is found on structure that is not on the list in the Baseline Program, for example the secondary structure.
Program Implementation A.
When a CPCP is started it is important to do the items that follow: (1) Start inspections where the airplane age is equal to or greater than the Baseline Program Implementation Threshold age (IT). (2) Once the corrosion program inspection (CPI) is started, the subsequent applications of the CPI are given by the Repeat Interval (RI) for each CPI. (3) You can start a CPCP on the basis of individual CPIs or groups of CPIs. (4) Cessna Aircraft Company highly recommends to start all of the CPIs as soon as possible. This is the most cost effective way to prevent or control corrosion.
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MODEL 208 MAINTENANCE MANUAL CORROSION - DESCRIPTION AND OPERATION 1.
General A.
Corrosion is a natural phenomenon which destroys metal by chemical or electrochemical action and converts it to a metallic compound such as an oxide, hydroxide, or sulfate. All metals used in airplane construction are subject to corrosion. Attack may take place over an entire metal surface or it may be penetrating in nature, forming deep pits. It may follow grain boundaries or it may penetrate a surface at random. Corrosion may be accentuated by stress from external loads or from lack of homogeneity in the metallic structure or from improper heat treatment. It is promoted by contact between dissimilar metals or with materials which absorb moisture, such as rubber, felt, dirt, salt, etc..
B.
Corrosion can take many different forms, and the corrosion resistance of materials used in the airplane can drastically change with only small environmental changes. Corrosion is often thought of as a slow process; however, some forms of corrosion can occur very quickly, in days or even hours. Airplanes exposed to salt air, heavy atmospheric industrial pollution, warm humid environments and/or over water operations will require more stringent corrosion prevention and control programs than airplanes operated in dry environments.
C.
Maintenance of the airplane primary coatings as speciÞed in Chapter 20, Standard Practices Airframe, combined with a constant cycle of cleaning, inspection, preservation and lubrication appropriate to the operational environment, must be incorporated by the operator to prevent corrosion. The basics of a corrosion prevention and control program consists of the following: (1) Personnel trained in the conditions, detection, identiÞcation, cleaning, treatment, and preservation for corrosion. (2) Adequate inspection intervals for detecting corrosion appropriate to the environment. (3) Airplane washing with clean water on regularly scheduled intervals. (4) Keeping drain holes and passages clear and open. (5) Prompt maintenance and repair of the primary coatings as speciÞed in Chapter 20, Standard Practices - Airframe. (6) Prompt corrosion treatment after detection. (7) Inspection and replication of corrosion inhibitive compounds on a scheduled basis. (8) Use of appropriate materials, equipment, and technical publications. NOTE:
2.
For additional general information on corrosion, treatment, repair, damage limits, and corrosion control, refer to FAA Advisory Circular No. 43-4A. For speciÞc information, refer to the 208 Series Structural Repair Manual.
Types of Corrosion A.
Electrochemical Corrosion (1) Refer to Figure 1 for an illustration of electrochemical corrosion. The following conditions must exist for electrochemical corrosion to occur. (a) There must be a metal that corrodes and acts as the anode. (b) There must be a less corrodible metal that acts as the cathode. (c) There must be a continuous liquid path between the two metals which acts as the electrolyte, usually condensation and salt or other contaminations. (d) There must be a conductor to carry the ßow of electrons from the cathode to the anode. This conductor is usually in the form of a metal-to-metal contact (rivets, bolts, welds, etc.). (2) The elimination of any one of the four conditions described above will stop the corrosion reaction process. (3) One of the best ways to eliminate one of the four described conditions is to apply an organic Þlm (such as paint, grease, plastic, etc.) to the surface of the metal affected. This will prevent the electrolyte from connecting the cathode to the anode, and current cannot ßow, therefore, preventing corrosion reaction. (4) At normal atmospheric temperatures, metals do not corrode appreciably without moisture, but the moisture in the air is usually enough to start corrosive action.
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MODEL 208 MAINTENANCE MANUAL (5) (6)
(7)
3.
The initial rate of corrosion is usually much greater than the rate after a short period of time. This slowing down occurs because of the oxide Þlm that forms on the metals surface. This Þlm tends to protect the metal underneath. When components and systems constructed of many different types of metals must perform under various climatic conditions, corrosion becomes a complex problem. Salt on metal surfaces (from sea coast operation) greatly increases the electrical conductivity of any moisture present and accelerates corrosion. Other environmental conditions which contribute to corrosion are: (a) Moisture collecting on dirt particles. (b) Moisture collecting in crevices between lap joints, around rivets, bolts and screws.
B.
Direct Surface Attack - The most common type of general surface corrosion results from direct reaction of a metal surface with oxygen in the atmosphere. Unless properly protected, steel will rust and aluminum and magnesium will form oxides. The attack may be accelerated by salt spray or salt bearing air, industrial pollutants or engine exhaust.
C.
Pitting - While pitting can occur in any metal, it is particularly characteristic of passive materials, such as the alloys of aluminum, nickel and chromium. It is Þrst noticeable as a white or gray powdery deposit similar to dust, which blotches the surface. When the deposits are cleaned away, tiny pits can be seen in the surface.
D.
Dissimilar Metal Corrosion - When two dissimilar metals are in contact and are connected by an electrolyte (continuous liquid or gas path), accelerated corrosion of one of the metals occurs. The most easily oxidized surface becomes the anode and corrodes. The less active member of the couple becomes the cathode of the galvanic cell. The degree of attack depends on the relative activity of the two surfaces; the greater the difference in activity, the more severe the corrosion. Relative activity in descending order is as follows: (1) Magnesium and its alloys. (2) Aluminum alloys 1100, 3003, 5052, 6061, 220, 355, 356, cadmium and zinc. (3) Aluminum alloys 2014, 2017, 2024, 7075 and 195. (4) Iron, lead and their alloys (except stainless steel). (5) Stainless steels, titanium, chromium, nickel, copper, and their alloys. (6) Graphite (including dry Þlm lubricants containing graphite).
E.
Intergranular Corrosion - Selective attack along the grain boundaries in metal alloys is referred to as intergranular corrosion. It results from lack of uniformity in the alloy structure. It is particularly characteristic of precipitation-hardened alloys of aluminum and some stainless steels. Aluminum extrusions and forgings in general can contain nonuniform areas which, in turn can result in galvanic attack along the grain boundaries. When the attack is well advanced, the metal can blister or delaminate and cause exfoliation.
F.
Stress Corrosion - This results from the combined effect of static tensile stresses applied to a surface over a period of time. In general, cracking susceptibility increases with stress, particularly at stresses approaching the yield point; and with increasing temperature, exposure time and concentration of corrosive ingredients in the surrounding environment. Examples of parts which are susceptible to stress corrosion cracking are aluminum alloy bellcranks employing pressed-in taper pins, landing gear shock struts with pipe thread type grease Þttings, clevis joints and shrink Þts.
G.
Corrosion Fatigue - This is a type of stress corrosion resulting from the cyclic stresses on a metal in corrosive surroundings. Corrosion may start at the bottom of a shallow pit in the stressed area. Once attack begins, the continuous ßexing prevents repair of protective surface coating or oxide Þlms and additional corrosion takes place in the area of stress.
Typical Corrosion Areas A.
This section lists typical areas of the airplane which are susceptible to corrosion. These areas should be carefully inspected at periodic intervals to detect corrosion as early as possible. (1) Engine Exhaust Trail Areas. (a) Gaps, seams and fairings on the lower right side of the fuselage, aft of the engine secondary exhaust stack, are typical areas where deposits may be trapped and not reached by normal cleaning methods.
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Corrosion IdentiÞcation Figure 1 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (b)
(2)
(3)
Around rivet heads, skin laps and inspection covers on the airplane lower fuselage, aft of the engine secondary exhaust stack, should be carefully cleaned and inspected. Battery Box and Battery Vent Opening. (a) The battery, battery cover, battery box and adjacent areas, especially areas below the battery box where battery electrolyte may have seeped, are particularly subject to corrosive action. If spilled battery electrolyte is neutralized and cleaned up at the same time of spillage, corrosion can be held to a minimum by using a weak boric acid solution to neutralize the battery electrolyte (ni-cad battery) or baking soda solution to neutralize the lead acid type battery electrolyte. If boric acid or baking soda is not available, ßood the area with water. Steel Control Cables (Including Stainless Steel). (a) Checking for corrosion on control cables is normally accomplished during the preventative maintenance check. During preventative maintenance, broken wire and wear of the control cable is also checked. (b) If the surface of the cable is corroded, carefully force the cable open by reverse twisting and visually inspect the interior. Corrosion on the interior strands of the cable constitutes failure and the cable must be replaced. If no internal corrosion is detected, remove loose external rust and corrosion with a clean, dry, coarse-weave rag or Þber brush. NOTE:
Do not use metallic wools or solvents to clean installed cables. Use of metallic wool will embed dissimilar metal particles in the cables and create further corrosion. Solvents will remove internal cable lubricant, allowing cable strands to abrade and further corrode.
(c)
(4)
After thorough cleaning of the exterior cable surface, apply a light coat of lubricant (VV-L800) to the external cable surface. Piano-Type Hinges. (a) The construction of piano-type hinges forms moisture traps as well as dissimilar metal corrosion between the steel hinge pin and the aluminum hinge. Solid Þlm lubricants are often applied to reduce corrosion problems. (b) Care and replacement of solid Þlm lubricants require special techniques peculiar to the speciÞc solid Þlm being used. Good solid Þlm lubricants conform to SpeciÞcation MIL-L23398D. 1 Solid Þlm lubricants prevent galvanic coupling on close tolerance Þttings and reduce fretting corrosion. Surface preparation is extremely important to the service/wear life of solid Þlm lubricants. 2 Solid Þlm lubricants are usually applied over surfaces precoated with other Þlms such as anodize and phosphate. They have been successfully applied over organic coatings such as epoxy primers.
CAUTION: Solid Þlm lubricants that contain graphite, either alone or in mixture with any other lubricants, may not be used since graphite is cathodic to most metals and will cause dissimilar corrosion in the presence of electrolytes. (5)
Steel Components. (a) The red oxide (rust) will not protect the underlying base metal unlike some other metal oxides. The presence of rust actually promotes additional attack by attracting moisture from the air and acting as a catalyst in causing additional corrosion to take place. Light red rust on bolt heads, hold-down nuts, and other nonstructural hardware is generally not dangerous. However, it is indicative of a general lack of maintenance and possible attack in more critical areas, such as highly stressed landing gear components and ßight surface
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MODEL 208 MAINTENANCE MANUAL actuating components. When paint failures occur or mechanical damage exposes highly stressed steel surfaces to the atmosphere, even small amounts of rusting are potentially dangerous and must be removed.
(6)
4.
If rust is detected on non highly stressed steel surfaces, refer to Chapter 20-3100, Interior and Exterior Finish - Cleaning/Painting, and Chapters 51-10-00 and 51-11-00 of the 208 Series Structural Repair Manual for removal and treatment procedures.
NOTE:
The main landing gear legs, center tube, and nose gear drag link spring are highly stressed components with shot peened surfaces. Refer to Chapter 32, Main Landing Gear - Cleaning and Painting, for instructions on rust removal and treatment of these components.
Internal Fuel Tanks. (a) The internal fuel tanks have the same primary coatings as the other aluminum skins used on the airplane. If fuel contamination is detected or suspected, the internal fuel bays should be inspected for damage to the primary coatings. Repair the coatings in accordance with Chapter 28, Fuel Tanks - Maintenance Practices.
Corrosion Detection A.
5.
NOTE:
The primary means of corrosion detection is visual, but in situations where visual inspection is not feasible, other techniques must be used. The use of liquid dye penetrant, magnetic particle, X-ray and ultrasonic devices can be used, but most of these sophisticated techniques are intended for the detection of physical ßaws within metal objects rather than the detection of corrosion. (1) Visual Inspection. A visual check of the metal surface can reveal the signs of corrosive attack, the most obvious of which is a corrosive deposit. Corrosion deposits of aluminum or magnesium are generally a white or grayish white powder, while the color of ferrous compounds varies from red to dark reddish brown. (a) The indications of corrosive attack are small, localized discolorations of the metal surface. Surfaces protected by paint or plating may only exhibit indications of more advanced corrosive attack by the presence of blisters or bulges in the protective Þlm. Bulges in lap joints are indications of corrosive buildup which is well advanced. (b) In many cases the inspection area is obscured by structural members, equipment installations or, for other reasons, are and is awkward to check visually. In such cases, mirrors, borescope or similar devices can be used to inspect the obscured areas. Any means which allows a thorough inspection can be used. Magnifying glasses are valuable aids for determining whether or not all corrosion products have been removed during cleanup operations. (2) Liquid Dye Penetrant Inspection. Inspection for large stress-corrosion or corrosion fatigue cracks on nonporous or nonferrous metals may be accomplished using dye penetrant processes. The dye applied to a clean metallic surface will enter small openings or cracks by capillary action. After the dye has an opportunity to be absorbed by any surface discontinuity, the excess dye is removed and a developer is applied to the surface. The developer acts like a blotter and draws the dye from cracks or Þssures back to the surface, giving visible indication of any fault that is present on the surface. The magnitude of the fault is indicated by the quantity of dye brought back to the surface by the developer.
Corrosion Repair NOTE:
When corrosion is detected, refer to the 208 Series Structural Repair Manual for damage limits, repair, treatment, and preservation information.
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MODEL 208 MAINTENANCE MANUAL CORROSION SEVERITY MAPS - DESCRIPTION AND OPERATION 1.
General A.
This section contains maps which de fine the severity of potential corrosion on airplane structure.
B.
Corrosion severity zones are affected by atmospheric and other climatic factors. The maps provided in this section are for guidance when determining types and frequency of required inspections and other maintenance. Refer to Figure 1, Figure 2, Figure 3, Figure 4, Figure 5, Figure 6.
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North America Corrosion Severity Map Figure 1 (Sheet 1)
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South America Corrosion Severity Map Figure 2 (Sheet 1)
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Africa Corrosion Severity Map Figure 3 (Sheet 1)
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Asia Corrosion Severity Map Figure 4 (Sheet 1)
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Europe and Asia Minor Corrosion Severity Map Figure 5 (Sheet 1)
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South Pacific Corrosion Severity Map Figure 6 (Sheet 1)
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52 CHAPTER
DOORS
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LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
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52-00-00
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Mar 1/2000
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Dec 1/2006
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Aug 1/1995
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52-Title 52-List of Effective Pages 52-Record of Temporary Revisions 52-Table of Contents 52-List of Tasks
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CONTENTS DOORS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
52-00-00 52-00-00 52-00-00 52-00-00
Page 1 Page 1 Page 1 Page 1
DOORS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Doors and Crew Door Entrance Steps Detailed Inspection. . . . . . . . . . . . . . . . . Passenger/Cargo Doors and Door Frames Detailed Inspection . . . . . . . . . . . . . . . . . .
52-00-00 Page 601 52-00-00 Page 601 52-00-00 Page 601 52-00-00 Page 602
PASSENGER AND CREW DOORS - DESCRIPTION AND OPERATION. . . . . . . . . . . . . . Crew Doors Description and Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Doors Description and Operation (Model 208 and 208B Passenger). . .
52-10-00 Page 1 52-10-00 Page 1 52-10-00 Page 1
CREW DOORS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Doors Latching Mechanism Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . Crew Door Seals Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
52-11-00 Page 201 52-11-00 Page 201 52-11-00 Page 201 52-11-00 Page 201 52-11-00 Page 201 52-11-00 Page 201
PASSENGER DOORS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Passenger Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Passenger Door Latching Mechanism Removal/Installation . . . . . . . . . . . . . . . Lower Passenger Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Door Step Disassembly/Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Entry Door Step Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . Lower Passenger Door Latching Mechanism Removal/Installation . . . . . . . . . . . . . . . Upper and Lower Passenger Doors Adjustment/Test . . . . . . . . . . . . . . . . . . . . . . . . . . . Gas Spring/Upper Door Test Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Door Seals Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Door Seal Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
52-12-00 Page 201 52-12-00 Page 201 52-12-00 Page 201 52-12-00 Page 201 52-12-00 Page 204 52-12-00 Page 204 52-12-00 Page 210 52-12-00 Page 210 52-12-00 Page 211 52-12-00 Page 211 52-12-00 Page 211 52-12-00 Page 213
CARGO DOOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Cargo Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Cargo Door Latching Mechanism Removal/Installation . . . . . . . . . . . . . . . . . . . Lower Cargo Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lower Cargo Door Latching Mechanism Removal/Installation . . . . . . . . . . . . . . . . . . . Upper and Lower Cargo Doors Adjustment/Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Gas Spring Cylinder Disposal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Gas Spring/Upper Door Test Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
52-31-00 Page 201 52-31-00 Page 201 52-31-00 Page 201 52-31-00 Page 201 52-31-00 Page 208 52-31-00 Page 208 52-31-00 Page 208 52-31-00 Page 210 52-31-00 Page 210
IN-FLIGHT MOVABLE (ROLL- UP) DOOR - MAINTENANCE PRACTICES. . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper and Lower Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper and Lower Door Adjustment/Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Side-Mounted Drive Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bottom-Mounted Drive Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . Bottom-Mounted Drive Assembly Motor Removal/Installation. . . . . . . . . . . . . . . . . . . .
52-32-00 Page 201 52-32-00 Page 201 52-32-00 Page 201 52-32-00 Page 201 52-32-00 Page 201 52-32-00 Page 201 52-32-00 Page 206 52-32-00 Page 206
AIR DEFLECTOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Cargo Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Deflector Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL CREW DOOR ENTRANCE STEPS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Door Step Assembly Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Door Step Disassembly/Assembly. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety Walk Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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DOOR WARNING - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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DOOR WARNING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger and Cargo Door Warning Switches Adjustment/Test . . . . . . . . . . . . . . . . .
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52 - CONTENTS © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 52-00-00-220
Crew Doors and Crew Door Entrance Steps Detailed Inspection
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52-00-00-221
Passenger/Cargo Doors and Door Frames Detailed Inspection
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52 - LIST OF TASKS © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL DOORS - GENERAL 1.
Scope A.
2.
Three entry doors, one cargo door, or an optional in-flight movable door are provided for pilot, passengers, and cargo loading and unloading. A single piece entry door for pilot is located on forward left side of fuselage, and an opposite door for front seat passenger is located on forward right side of fuselage. The primary passenger loading and unloading door, located just aft of wing on right side of fuselage, is a two-piece air-stair type door. Cargo loading is accomplished through a large two-piece door on left side of fuselage. Steps that fold and stow just inside left crew door are provided for crew to enter and exit airplane. Optional right crew door steps are available. A door open warning system is also provided as a safety feature. If upper cargo door or upper passenger/air-stair door is not properly latched, a red light, labeled DOOR WARNING, located in annunciator panel, illuminates to alert the pilot.
Tools, Equipment and Materials NOTE:
Equivalent Substitutes may be used for the listed items:
NAME
NUMBER
MANUFACTURER
USE
Scale
5W587
A Grainger V. W., Inc. Grainger Division 2227 Clark Street St. Louis, MO 63107
Test gas spring.
Adhesive
EC-1300L
3M Company St. Paul, MN 55101
To bond door seal corners to door.
Adhesive
RTV-157 Gray
General Electric Waterford, NY
To bond door seals to door.
Commercially Available
Cleaning bonding surfaces and seals.
Methyl n-Propyl Ketone 3.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief description of the sections follows: (1) The section on passenger and crew doors provides a description of components and maintenance practices. (2) The section on cargo door provides a description of components and maintenance practices. (3) The section on roll up door provides a description of components and maintenance practices. (4) The section on air deflector provides a description of components and maintenance practices. (5) The section on door warning provides troubleshooting and maintenance practices.
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MODEL 208 MAINTENANCE MANUAL DOORS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the doors in a serviceable condition.
Task 52-00-00-220 2.
Crew Doors and Crew Door Entrance Steps Detailed Inspection A.
General (1) This task gives the information needed to do a detailed inspection of the crew doors and the crew door entrance steps.
B.
Special Tools (1) Dry Solid Film Lubricant (MIL-L-23398) (2) Isopropyl Alcohol
C.
Access (1) Remove the left and the right crew door upper and lower interior panels. Refer to Crew Door Maintenance Practices. (2) Remove floorboard access panels 231AL, 231CL, 232AR, and 232AC that are adjacent to the crew ladder mounting brackets. Refer to Chapter 6, Access/Inspection Plates - Description and Operation.
D.
Do a Detailed Inspection of the Crew Doors and Crew Door Entrance Steps between FS 128.00 to FS 166.45. (1) Examine the crew door external panel surface for condition, cracks, corrosion, delamination, and security. (2) Examine the areas around the door frame, door hinges hinge pin and attach screws for condition, corrosion, security, and correct attachment. (3) Examine the crew door latch mechanism assembly for condition, corrosion, security, and correct attachment. (a) Make sure that you examine the door handle, roll pin, escutcheon, handle support, and lock pin. (b) Make sure that you examine the bell crank, bell crank bushing, bell crank pin, and door handle spindle. (c) Make sure that you examine the pushrod assembly, clevis, latch bolt, and latch bolt spring. (4) Examine the crew door entrance step assembly for condition, corrosion, and security of assembly to the floor. (5) Examine the upper arm attachments to the floor mounting brackets for condition and security. (a) Examine the bushings for wear. (6) Examine the ladder mounting bracket attach structure for condition, cracks, buckling, bending, and corrosion. (7) Examine the lower arm mount points for condition and security. (a) Examine the bushings for wear. (8) Examine the step anti-skid material for condition and security. (9) Examine the arm rest pad for condition and security. (10) Examine the crew door seals for correct installation, security, cuts, abrasions, and wear. (a) Clean the door seals with a cloth slightly dampened with water or isopropyl alcohol. (11) Examine the door lock assembly for condition, wear, and security. (a) Lubricate the locking mechanism with Moly Sulfide or an equivalent lubricant. (12) Lubricate any pivot point or sliding surface with MIL-L-23398 before you install the crew door interior panels.
E.
Restore Access (1) Install floorboard access panels 231AL, 231CL, 232AR, and 232AC. Refer to Chapter 6, Access/ Inspection Plates - Description and Operation.
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MODEL 208 MAINTENANCE MANUAL (2)
Install the left and the right crew door upper and lower interior panels. Refer to Crew Door Maintenance Practices.
End of task Task 52-00-00-221 3.
Passenger/Cargo Doors and Door Frames Detailed Inspection A.
General (1) This task gives the information needed to do a detailed inspection of the passenger and cargo doors and door frames.
B.
Special Tools (1) None
C.
Access (1) Remove the upper and the lower passenger door interior panels. Refer to Passenger Doors Maintenance Practices. (2) Remove the upper and the lower cargo door interior panels. Refer to Cargo Doors - Maintenance Practices.
D.
Do a Detailed Inspection of the Passenger Door Assembly between FS 234.00 to FS 284.00 for the Model 208 and FS 282.00 to 332.00 for the Model 208B. (1) Examine the upper and lower passenger doors for condition, cracks, corrosion, delamination, and security.
CAUTION: Do not apply too much torque to any of the attaching hardware to the doors. Too much torque can strip the threaded inserts. (2)
Examine the passenger door frames and hinge areas for condition, cracks, and corrosion.
WARNING: If the upper and lower gas cylinders are removed at the same time, do not interchange the upper and lower cylinders. Severe injury and damage to the airplane could occur. (3) (4) (5) (6) (7) (8) (9)
Examine the upper door spring gas cylinders for condition and security. (a) Make sure that the upper door will hold in the open position. Examine the lower door gas spring cylinders for condition and security. (a) Make sure that the cylinders cushion the lower door when the door is released to free-fall from the closed position. Examine all four restraint cables for condition and security. (a) Look closely for broken cable strands in the area where the cable comes out of the clevis end. Examine the entrance step assembly for condition, corrosion, security, and wear. (a) If the step anti-skid material is worn, replace the material. Examine the lower cables for correct adjustment. (a) Both cables must carry all of the load of the lower door when the door is open. (b) The gas cylinders must not be fully extended. Examine the cargo upper door external panel surface, upper door hinge and the fuselage door frame and hinges for condition, corrosion, security, and correct attachment. Examine the upper door latch mechanism assembly to include the following: (a) Door handle, roll pin, escutcheon, handle support, and lock pin for condition, corrosion, security, and correct attachment. (b) Bell crank, bell crank bushing, bell crank pin, and door handle spindle for condition, corrosion, security, and correct attachment. (c) Pushrod assemblies, clevis, latch bolt, and latch bolt spring for condition, corrosion, security, and correct attachment.
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MODEL 208 MAINTENANCE MANUAL (10) Examine the lower door latch mechanism assembly to include the following: (a) Door handle, roll pin, escutcheon, handle support, and lock pin for condition, corrosion, security, and correct attachment. (b) Bell crank, bell crank bushing, bell crank pin, and door handle spindle for condition, corrosion, security, and correct attachment. (c) Pushrod assemblies, clevis, latch bolt, and latch bolt spring for condition, corrosion, security, and correct attachment. (11) Examine the door lock assembly for condition, wear, and security. (a) Lubricate the locking mechanism with Moly Sulfide or an equivalent lubricant. (12) Examine the door seals for correct installation, security, cuts, abrasions, and wear. (a) Clean the door seals with a cloth slightly dampened with water or isopropyl alcohol. E.
Do a Detailed Inspection of the Cargo Door Assembly between FS 234.00 to FS 284.00 for the Model 208 and FS 282.00 to 332.00 for the Model 208B. (1) Examine the upper and the lower cargo doors for condition, cracks, corrosion, delamination, and security.
CAUTION: Do not apply too much torque to any of the attaching hardware to the doors. Too much torque can strip the threaded inserts. (2)
Examine the upper and the lower cargo door hinges for condition, corrosion, security, and correct installation. (3) Examine the cargo door frames for cracks. (4) Examine the upper cargo door spring gas cylinders for condition and security. (a) Make sure that the upper door will hold in the open position. (5) Examine the upper cargo door restraint cables for condition and security. (a) Look closely for broken cable strands in the area where the cable comes out of the clevis end. (6) Examine the areas around the door hinges and the door handle, hinge pin, and attach screws for condition, corrosion, security, and correct attachment. (7) Examine the upper door latch mechanism assembly to include the following: (a) Door handle, roll pin, escutcheon, handle support, and lock pin for condition, corrosion, security, and correct attachment. (b) Bell crank, bell crank bushing, bell crank pin, and door handle spindle for condition, corrosion, security, and correct attachment. (c) Pushrod assemblies, clevis, latch bolt, and latch bolt spring for condition, corrosion, security, and correct attachment. (8) Examine the lower cargo door latch mechanism assembly to include the following: (a) Door handle, roll pin, escutcheon, handle support, and lock pin for condition, corrosion, security, and correct attachment. (b) Bell crank, bell crank bushing, bell crank pin, and door handle spindle for condition, corrosion, security, and correct attachment. (c) Pushrod assemblies, clevis, latch bolt, and latch bolt spring for condition, corrosion, security, and correct attachment. (9) Examine the door lock assembly for condition, wear, and security. (a) Lubricate the locking mechanism with Moly Sulfide or an equivalent lubricant. (10) Examine all door seals for correct installation, security, cuts, abrasions, and wear. (a) Clean all door seals with a cloth slightly dampened with water or isopropyl alcohol. F.
Restore Access (1) Install the upper and the lower cargo door interior panels. Refer to Cargo Doors - Maintenance Practices. (2) Install the upper and the lower passenger door interior panels. Refer to Passenger Doors Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL PASSENGER AND CREW DOORS - DESCRIPTION AND OPERATION 1.
Crew Doors Description and Operation A.
2.
The single piece left crew entry door incorporates a conventional outside door handle, a key operated door lock, a conventional inside door handle, a pull knob to override key lock from inside, and a vent window that opens. The opposite crew entry door on right side is same as left crew door except it does not have vent window or key operated door lock. The right crew door has a manually operated inside door lock. To open either crew door from outside airplane, rotate handle down and forward to OPEN position. To close door from inside airplane, use conventional door handle and door pull. The inside door handle is a three position handle with OPEN, CLOSED, and LATCHED positions. Place handle in CLOSE position and pull door shut, then rotate handle forward to LATCHED position. When handle is rotated to latched position, an over center action will hold it in that position. To lock crew entry doors when leaving airplane, lock right crew door with manually operated door lock, close left crew door and using key, lock the door. To override left crew door from inside airplane, pull and rotate knob located above inside door handle. A folding step attached to floorboard inside left crew door opening, rotates out and unfolds to assist crew entry. The step folds and stows just inside left crew door when not in use.
Passenger Doors Description and Operation (Model 208 and 208B Passenger) A.
The passenger entry door consists of an upper and lower section. When opened, the upper section swings upward and the lower section drops down, exposing steps for entry into the airplane. The upper door incorporates a conventional exterior door handle with a separate key operated lock, a pushbutton type exterior door release, and a conventional interior door handle. Two gas spring cylinders are utilized to lift upper door to full open position. The lower door utilizes a flush handle which is accessible from either inside or outside the airplane. This handle is designed so when the upper door is closed, the handle cannot be rotated to the open position. The lower door also utilizes door support cables and two gas spring cylinder dampeners. A cabin door open warning system is provided so if the upper is not completely latched, a red light, labeled DOOR WARNING, located on the annunciator panel, illuminates to alert the pilot.
WARNING: Outside proximity of lower door must be clear before opening. B.
To enter the airplane through passenger entry door, depress exterior pushbutton door release, rotate exterior door handle on upper door section counterclockwise to open position, and raise upper door section to over center position. Following this action, the gas spring cylinders automatically raises door to full up position. After upper door section is open, release lower door section by pulling up on inside door handle and rotating to open position. Lower the door section until it is supported by door support cables. the door steps deploy automatically from their stowed position.
C.
To close passenger entry door from inside airplane, use support cables to pull lower door section up, then latch lower door section by rotating inside door handle forward to CLOSE position. Using pull strap, close upper door section and latch by rotating inside handle counterclockwise. Then snap handle into its locking receptacle.
WARNING: Outside proximity of lower door must be clear before opening. D.
To exit airplane through passenger door, pull inside door handle on upper door section from its locking receptacle, and then rotate handle clockwise to vertical position. Push door outward to overcenter position, then gas spring cylinders will automatically raise the door to full up position. Next, rotate door handle on lower door section up and aft to OPEN position, and push door outward. The gas spring cylinder dampeners lower door to fully open position, and integral steps will deploy.
E.
To close passenger entry door from outside airplane, close and latch lower door section by rotating inside handle down and forward to CLOSE position. Then close upper door section and latch by rotating outside door handle clockwise to horizontal (LOCKED) position. Use key operated door lock to lock the door.
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MODEL 208 MAINTENANCE MANUAL CREW DOORS - MAINTENANCE PRACTICES 1.
General A.
2.
Tools, Equipment and Materials A.
3.
This section provides maintenance practices, test procedures, and removal/installation for crew doors.
For a list of required tools, equipment and materials, refer to Doors - General.
Crew Door Removal/Installation A.
Remove Crew Doors (Refer to Figure 201). (1) Disconnect oxygen line from outlet above door; disconnect microphone cable from jack on instrument panel; remove line from clips along doorpost; remove oxygen mask from pocket on the door. Refer to Chapter 35. (2) With crew door open and door supported, remove cotter pins (7) and washers (6) from upper and lower hinges (3). (3) Remove hinge pins (2) from upper and lower hinges (3). (4) Remove crew door from airplane.
B.
Install Crew Doors (Refer to Figure 201 ). (1) Align crew door (8) to hinges (3). (2) Install hinge pins (2), washers (6), and cotter pins (7). (3) Connect oxygen line in outlet above door, then install line in clips along doorpost, and put oxygen mask in pocket on door. Connect microphone cable in jack on instrument panel. NOTE:
4.
5.
Normal gap between door rails and jamb is 0.45 inch.
Crew Doors Latching Mechanism Removal/Installation A.
Remove Crew Doors Latching Mechanism (Refer to Figure 201 ). (1) Remove roll pin (43) and inside door handle (42). (2) Remove washer (46). (3) Remove screws (63) and escutcheon (62). (4) Remove knob (61) on left crew door or knob (38) on right crew door. (5) Remove window trim and door upholstery panel by removing screws. (6) Remove four screws (21). (7) Remove two screws (40) and spacers (37). (8) Remove support (22). (9) If further disassembly is required, refer to Figure 201.
B.
Install Crew Doors Latching Mechanism (Refer to Figure 201 ). (1) Position support (22) install two spacers (37) and two screws (40). (2) Install four screws (21) to secure support (22). (3) Install door upholstery panel and window trim moulding. (4) Install knob (61) on left crew door or knob (38) on right crew door. (5) Install escutcheon (62) using two screws (63). (6) Install washer (46). (7) Position inside door handle (42) and secure with roll pin (43).
Crew Door Seals Installation A.
For procedures to install crew door seals, refer to Passenger Doors - Maintenance Practices, Passenger Door Seals Installation.
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Crew Door Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Crew Door Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Crew Door Installation Figure 201 (Sheet 3)
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Crew Door Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL PASSENGER DOORS - MAINTENANCE PRACTICES 1.
General A.
2.
This section provides maintenance practices, test procedures, and removal/installation for passenger doors.
Upper Passenger Door Removal/Installation A.
Remove Upper Passenger Door (Refer to Figure 201). (1) Remove split ring from lower door cable. (2) Remove safety clips (20A) from both end of each gas spring (12) and remove gas springs. (3) Remove screws (8) attaching upper passenger door hinge (4) to fuselage above door jamb. NOTE:
B.
3.
If upper door gas spring is removed, it should be tagged to identify it as being used on upper door. Upper door gas spring and lower door dampener are identical except for pounds of force. Upper gas spring has 45 pounds of force (Airplanes 20800001 Thru 20800058), or 60 pounds of force (Airplanes 20800059 and On). Lower door dampener has 6 pounds of force (Airplanes 20800001 Thru 20800095 Except Airplanes incorporating CAB86-11); or 10 pounds of force (Airplanes 20800096 and On, and 20800001 Thru 20800095 Incorporating CAB 86-11). They both have the same placard installed on cylinder. To determine difference between upper gas spring and lower door snubber, push in on cylinder rod. If rod can be compressed easily, the cylinder is a lower door dampener. If rod cannot be compressed easily, the cylinder is an upper door gas spring. If upper door gas spring or lower door dampener are replaced with a new part, the replaced part must be disposed of in accordance with Cargo Door - Maintenance Practices, Disposal of Gas Spring Cylinder.
Install Upper Passenger Door (Refer to Figure 201). (1) Align holes in upper passenger door hinge (4) with holes in fuselage above door jamb and install screws (8). (2) Place ends of gas springs (12) over ballstuds (20B) and install clips (20A). (3) Install split ring on door pullstrap around lower cable.
Upper Passenger Door Latching Mechanism Removal/Installation A.
Remove Upper Passenger Door Latching Mechanism (Refer to Figure 201 ). (1) Remove roll pin (48), handle (50) and washer (47). (2) Remove door window trim moulding. (3) Remove roll pin (21) outside handle (24) and washer (23). (4) Remove cotter pin (67), washer (66) and lock pin (69). (5) Remove screw (25), nut (27), washer (26), screw (22), and guide (70). (6) Remove escutcheon (72). (7) Remove washers (28), cotter pin (57), pin (30) and spindle (46). (8) Remove two screws (45), nut (43), washer (44), washer (52), screw (51), screw (55) and support (56). (9) Remove bellcrank (41) by removing cotter pins (29) and pins (42). (10) Remove cotter pin (62), pin (60), pushrod (59) and latch pin assembly (61). (11) Remove cotter pin (31), pin (39), pushrod (40) and latch pin assembly (37).
B.
Install Upper Passenger Door Latching Mechanism (Refer to Figure 201). (1) Install latch pin assembly (37) and connect pushrod (40) to latch pin using pin (39) and cotter pin (31). (2) Connect pushrod (40) to bellcrank (41) using pin (42) and cotter pin (29). (3) Install latch pin assembly (61) and connect to pushrod (59) using pin (60) and cotter pin (62). (4) Connect pushrod (59) to bellcrank (41) using pin (42) and cotter pin (29). (5) Install support (56) using two screws (45), screw (55), screw (51), washer (52), washer (44), and nut (43).
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MODEL 208 MAINTENANCE MANUAL
Upper Passenger Door Installation Figure 201 (Sheet 1)
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Upper Passenger Door Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (6)
Insert spindle (46) through bearing (52A), support (56) and bellcrank (41). Secure to bellcrank (41) using pin (30) and cotter pin (57). (7) Install washers (28) on spindle (46). (8) Insert escutcheon (72) through hole in door and onto spindle (46), then install lock pin guide (70) using screw (22), washer (26), nut (27) and screw (25). (9) Insert lock pin (69) into guide (70), and attach to door lock cam using washer (66) and cotter pin (67). (10) Install washer (23) over spindle (46), then install outside handle (24) and secure with roll pin (21). (11) Install door window trim moulding. (12) Install washer (47) and interior handle (50) using roll pin (48). NOTE:
4.
Lower Passenger Door Removal/Installation A.
Remove Lower Passenger Door (Refer to Figure 202). (1) With upper and lower door open, support lower door. (2) Remove nuts (6), lockwashers (7), and washers (8) attaching forward and aft gas spring dampeners (5) to mounting plates (9) on lower door. (3) Disconnect lower ends of forward and aft cable assemblies by removing screws (40) and spacers (43). (4) Remove six bolts (19) and six washers (20), attaching upper hinge (17) to fuselage, then remove step tread (18). (5) Remove four remaining screws (21) attaching upper hinge (17) to fuselage. (6) Slide door and hinge out from between fuselage and door jambs, NOTE:
B.
5.
On airplanes 20800001 Thru 20800100, if the door lock on the passenger door does not operate properly, a new cam (78) may be modified. Refer to Figure 201, Sheet 2, View B-B.
If lower door dampener is removed, it should be tagged to identify it as being used on lower door. Upper gas spring and lower door dampener are identical except for pounds of force. Upper gas spring has 45 pounds of force (Airplanes 20800001 Thru 20800058), or 60 pounds of force (Airplanes 20800059 and On). Lower snubber has 6 pounds of force (Airplanes 20800001 Thru 20800095), except when modified per CAB 86-11. Both have the same placard installed on cylinder. To determine difference between upper gas spring and lower door dampener, push in on cylinder rod. If rod cannot be compressed easily, the cylinder is an upper door gas spring. If upper door gas spring or lower door dampener is replaced with a new part, the replaced part must be disposed of in accordance with Cargo Door - Maintenance Practices, Disposal of Gas Spring Cylinder.
Install Lower Passenger Door (Refer to Figure 202). (1) Slide door hinge in between fuselage and door jambs, align holes in fuselage with holes in hinge and install four screws (21). (2) Align holes in step tread (18) with holes in fuselage and install six washers (20) and six bolts (19). (3) Connect lower ends of forward and aft cable assemblies (42) using spacers (43) and screws (40). (4) Install gas spring dampeners using washers (8), lockwashers (7), and nuts (6).
Passenger Door Step Disassembly/Assembly A.
Disassemble Passenger Entry Door Step Assembly (Refer to Figure 202). (1) Remove step assembly from lower door. Refer to Remove Passenger Entry Door Step Assembly. (2) Remove strap assembly (110) by removing screws (111) and spacers (112). (3) Top step is now free to remove. (4) To remove center step, remove nut (64), washer (63), bushing (62), spacer (61), and screw (59) from both sides of step assembly.
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MODEL 208 MAINTENANCE MANUAL
Lower Passenger Door Installation Figure 202 (Sheet 1)
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Lower Passenger Door Installation Figure 202 (Sheet 2)
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Lower Passenger Door Installation Figure 202 (Sheet 3)
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Lower Passenger Door Installation Figure 202 (Sheet 4)
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Lower Passenger Door Installation Figure 202 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL (5) (6) B.
The lower step is partially disassembled when cable assemblies were removed. To complete the removal of lower step, remove screws (71) and spacer (70) from each side of step assembly.
Assemble Passenger Entry Door Step Assembly (Refer to Figure 202). (1) Partially install lower step by installing screw (71) through step rail with spacer (70) between step rail and step, on each side of step assembly. (2) Install strap (110) on center step assembly using screw (111) and spacer (112). (3) Install center step using screw (59) through step rail, spacer (61) between step rail and step, bushing (62) through mounting hole in step, then install washer (63) and nut (64) on each side of step assembly. (4) To complete assembly of step, refer to Crew Doors - Maintenance Practices. (5) Torque all step assembly screws to 70 inch-pounds, +0 or -5 inch-pounds, plus running torque. NOTE:
6.
Passenger Entry Door Step Assembly Removal/Installation A.
Remove Passenger Entry Door Step Assembly (Refer to Figure 202). (1) With upper and lower door open, support lower door. (2) Disconnect forward and aft cable assemblies (42) from step assembly (69) by removing screws (40) and spacers (43). (3) Remove nut (49), washer (48), bushing (47), spacer (46), spacer (45), washer (30), and screw (31) from each side of step assembly. (4) Remove nut (56), washer (54), spacer (55), washer (32), and screw (33) from each side of step assembly. (5) Remove nut (60), washer (57), spacer (58), washer (34), and screw (35) from each side of step assembly and remove step assembly.
B.
Install Passenger Entry Door Step Assembly (Refer to Figure 202). (1) Align holes in step assembly with holes in door rails. (2) Install screw (31) with washer (30) through door rail, install spacer (45) between door rail and step rail, install spacer (46) between step rail and step, insert bushing (47) into step mounting hole, then install washer (48) and nut (49) on each side of step assembly. (3) Install screw (33) with washer (32) through door rail, install spacer (55) between door rail and step rail, then install washer (54) and nut (56) on each side of step assembly. (4) Install screw (35) with washer (34) through door rail, install spacer (58) between door rail and step rail, then install washer (57) and nut (60) on each side of step assembly. (5) Connect forward and aft cable assemblies using spacers (43) and screws (40). (6) Torque all step assembly mounting screws to 70 inch-pounds, +0 or -5 inch-pounds, plus running torque. NOTE:
7.
Refer to Chapter 20, Torque Data - Maintenance Practices, for information on torquing of bolts.
Refer to Chapter 20, Torque Data - Maintenance Practices, for information on torquing of bolts.
Lower Passenger Door Latching Mechanism Removal/Installation A.
Remove Lower Passenger Door Latching Mechanism (Refer to Figure 202). (1) Remove roll pin (101), remove handle (100), and washer (102). (2) Remove escutcheon (98) by removing two screws (99). (3) Remove door panel. (4) Remove two screws (109), shims (73) and latch pin guide (72). (5) Remove washer (74), spring (75), cup (108), and washer (107) from latch pin (104). (6) Remove cotter pin (77), pin (105), and latch pin (104). (7) Remove cotter pin (79), pin (103), and pushrod (78). (8) Remove two screws (88), shims (73), and latch pin guide (72). (9) Remove cotter pin (86), pin (89), and latch pin (87). (10) Remove cotter pin (79), pin (103), and pushrod (85).
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MODEL 208 MAINTENANCE MANUAL B.
8.
Upper and Lower Passenger Doors Adjustment/Test A.
9.
Adjustment/Test Procedure (Refer to Figure 203). (1) With upper section door handle (50) in locked position, adjust the clevis end of pushrods by loosening jamnut to obtain 0.91 inch, +0.10 or -0.10 inch from outer edge of door rail to the end of latch pins. After adjusting, tighten jamnuts. (2) Adjust plunger in outer end of upper door exterior handle using slot screwdriver. To tighten, turn clockwise; to loosen, turn counterclockwise. Adjust plunger until 10 pounds, + 5 or -0 pounds of force, applied at 0.55 inch from end of inboard handle, is required to disengage plunger from catch plate. (3) To adjust the latch pins in lower door section, refer to step one.
Gas Spring/Upper Door Test Procedure A.
10.
Install Lower Passenger Door Latching Mechanism (Refer to Figure 202). (1) Connect pushrod (85) to bellcrank (83) using pin (103) and cotter pin (79). (2) Insert latch pin (87) into guide (72) and install shims (73) and guide (72) using two screws (88). (3) Connect latch pin (87) to pushrod (85) using pin (89) and cotter pin (79). (4) Connect pushrod (78) to bellcrank (83) using pin (103) and cotter pin (86). (5) Connect latch pin (104) to pushrod (78) using pin (105) and cotter pin (77). (6) Install washer (107), cup (108), spring (75), and washer (74) on latch pin (104). (7) Insert end of latch pin (104) into guide (72), install shims (73) and guide (72) using two screws (72). (8) Install door panel. (9) Install escutcheon using two screws (99). (10) Install washer (102) and handle (100), secure by installing roll pin (101).
When gas spring operation problems are suspected on upper passenger door, the following test procedure should be used. The forces are measured with a spring ambient temperature of 68°F, +2 or -2°F, with piston rod extending downwards. (1) Obtain direct reading sensitive scale. (2) Remove pneumatic extender from airplane. (3) Grasp extender by cylinder in a vertical position with piston rod end down, and place on sensitive scale. (4) Compress device fully four or five times with piston rod end down. (This lubricates seals and piston.) (5) Compress approximately 1.5 inch and relax pressure slightly to allow piston rod to extend slowly until it is approximately 0.20 inch from full extension. Hold steady and read pressure. (6) An acceptable extender will read between 42 and 52 pounds on scale.
Passenger Door Seals Removal/Installation A.
Remove Passenger Door Seals. (1) Remove door seal. A putty knife may be used to aid in the removal of seal. Caution must be taken not to damage door. (2) Clean off old adhesive from bonding surface with a cloth slightly dampened with methyl n-propyl ketone, taking care to apply methyl n-propyl ketone to bonding surface only.
B.
Install Passenger Door Seals. (1) Clean bonding surface with a cloth slightly dampened with methyl n-propyl ketone. Do not allow methyl n-propyl ketone to contact painted surface.
(2)
NOTE:
On Airplanes 20800001 Thru 20800108 and 208B0001 Thru 208B0003, the straight portion of seal is neoprene rubber and corners are silicone. Beginning with Airplanes 20800109 and 208B0004 (and all spares), the entire seal is silicone.
NOTE:
Trim flange of seal as required to clear latch pins and mounting hardware.
Bond neoprene seals to door using EC1300L per manufacturer's instructions.
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Upper and Lower Passenger Doors Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (3) (4) 11.
Bond silicone seals to door using RTV- 157 per manufacturer's instructions. After installation, clean door seals with a cloth slightly dampened with methyl n-propyl ketone. Caution must be taken not to soak seals, as methyl n-propyl ketonewill soften the adhesive.
Door Seal Cleaning A.
Clean Door Seals
CAUTION: Caution must be taken not to oversoak the seals withmethyl npropyl ketone, as methyl n-propyl ketone will soften the adhesive, neoprene rubber seals, and silicone seals. (1)
It is important that all door seals are properly secured and cleaned periodically to ensure an air and water tight seal. Clean door seals with a cloth, slightly dampened with methyl n-propyl ketone.
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MODEL 208 MAINTENANCE MANUAL CARGO DOOR - MAINTENANCE PRACTICES 1.
General A.
2.
The cargo door, located on left side of fuselage just aft of leftwing, consists of an upper and lower section. When opened, upper section swings upward and lower section swings forward. The upper section of cargo door incorporates a conventional outside door handle with a key-operated lock, a pushbutton type exterior door release, and a conventional interior door handle. Two gas spring cylinders are utilized to lift door to the full open position.
Upper Cargo Door Removal/Installation A.
Remove Upper Cargo Door (Refer to Figure 201). (1) Remove cotter pin and pin securing pullstrap to door. (2) Remove screws securing trim plate on upper door jamb and remove trim plate. (3) Remove upholstery trim around upper door jamb. (4) Remove safety clips (18A) from both ends of gas springs (14) and remove gas springs from ballstuds (18B). NOTE: (5)
B.
3.
If gas spring cyclinders are to be replaced with new parts, the replaced parts must be disposed of in accordance with Gas Spring Cylinder Disposal.
Remove screws (3) attaching upper cargo door hinge (4) to fuselage above door jamb.
Install Upper Cargo Door (Refer to Figure 201). (1) Align holes in upper door hinge (4) with holes in fuselage above door jamb and install screws (3). (2) Place ends of gas spring cyclinder (14) over ballstuds (18B) and install clips (18A). (3) Install trimplate on upper door jamb. (4) Install upholstery trim around upper door jamb. (5) Connect pullstrap to door using pin and cotter pin.
Upper Cargo Door Latching Mechanism Removal/Installation A.
Remove Upper Cargo Door Latching Mechanism (Refer to Figure 201). (1) Remove roll pin (42), inside handle (41) or boss (41A), and washer (43). (2) Remove door window trim molding. (3) Remove roll pin (21), handle (19) and washer (20). (4) Remove cotter pin (61), pin (54), and remove latch pin assembly (60). (5) Remove cotter pin (62) and pin (50) to remove pushrod (53). (6) Remove cotter pin (32) and pin (34) and remove latch pin (33). (7) Remove cotter pin (62) and pin (50) to remove pushrod (31). (8) Remove screw (22), nut (65), washer (66), and escutcheon (23). (9) Remove cotter pin (27), washer (28), lockpin (67), and guide (68). (10) Remove cotter pin (51), pin (29), washers (63), and spindle (44). (11) Remove bellcrank (52).
B.
Install Upper Cargo Door Latching Mechanism (Refer to Figure 201). (1) Insert spindle (44) through hole in support assembly (36) and install bellcrank (52) on spindle using pin (29) and cotter pin (51). (2) Install washers (63) on spindle (44), install escutcheon through hole in door and over spindle, then install screw (22) in aft hole. (3) Align holes in lock pin guide with holes in door and escutcheon and secure with screw (64), screw (22), washer (66), and nut (65). (4) Insert locking pin (67) into guide (68) and attach to door lock cam using washer (28) and cotter pin (27). (5) Install washer (20), handle (19) and secure with roll pin (21). (6) Install latch pin (33) and connect to pushrod (31) using pin (34) and cotter pin (32). (7) Connect pushrod (31) to bellcrank (52) using pin (50) and cotter pin (62). (8) Install latch pin assembly (60) and connect to pushrod (53) using pin (54) and cotter pin (61).
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Upper and Lower Cargo Door Installation Figure 201 (Sheet 1)
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Upper and Lower Cargo Door Installation Figure 201 (Sheet 2)
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Upper and Lower Cargo Door Installation Figure 201 (Sheet 3)
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Upper and Lower Cargo Door Installation Figure 201 (Sheet 4)
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Upper and Lower Cargo Door Installation Figure 201 (Sheet 5)
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Upper and Lower Cargo Door Installation Figure 201 (Sheet 6)
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MODEL 208 MAINTENANCE MANUAL (9) Connect pushrod (53) to bellcrank (52) using pin (50) and cotter pin (62). (10) Install door window trim molding. (11) Install washer (43) over spindle (44) and install handle (41) using roll pin (42). NOTE: 4.
5.
6.
On airplanes 20800001 Thru 20800100 if the door lock on the cargo door does not operate properly, cam (74) may be modified.
Lower Cargo Door Removal/Installation A.
Remove Lower Cargo Door (Refer to Figure 201). (1) With upper and lower cargo doors open, support lower door. (2) Remove cotter pins (2) and washers (3) from upper and lower hinge pins (7). (3) Remove hinge pins (7) from upper and lower hinges (6).
B.
Install Lower Cargo Door (Refer to Figure 201). (1) Align hinges on lower cargo door to hinges on fuselage. (2) Install hinge pins (7) through hinges. (3) Secure hinge pins with washers (3) and cotter pins (2).
Lower Cargo Door Latching Mechanism Removal/Installation A.
Remove Lower Cargo Door Latching Mechanism (Refer to Figure 201). (1) Remove roll pin (40), handle (38), and washer (39). (2) Remove two screws (41) and escutcheon (37). (3) Remove lower cargo door upholstery panel. (4) Disconnect pushrod assembly (25) by removing cotter pins (26), (22), and pins (43), (68) attaching pushrod (25) to bellcrank (28) and latch pin assembly (21). (5) Remove latch pin assembly. (6) Disconnect and remove pushrod assembly (65) by removing cotter pins and pins attaching pushrod to bellcrank (50) and bellcrank (67). (7) Remove latch pin assembly (63).
B.
Install Lower Cargo Door Latching Mechanism (Refer to Figure 201). (1) Install latch pin assembly (63) and insert bellcrank (50) into slot in latch pin assembly. (2) Connect pushrod assembly (65) to bellcrank (50) using pin (51) and cotter pin (64). (3) Install latch pin assembly (21) and insert bellcrank (67) into slot in latch pin assembly. (4) Connect pushrod assembly (65) to bellcrank (67) using pin (46) and cotter pin (45). (5) Connect pushrod assembly (25) to latch pin assembly (21) and bellcrank (28) using pins (43), (68) and cotter pins (26), (22). (6) Install lower cargo door upholstery panel. (7) Install escutcheon (37) using two screws (41). (8) Install washer (39), handle (38) and secure with roll pin (40).
Upper and Lower Cargo Doors Adjustment/Test A.
Adjust Upper and Lower Cargo Doors (Refer to Figure 201). (1) With door handles (41 and 19) in their locked position, adjust clevis ends of pushrods by loosening jamb nuts to obtain 0.91 inch, +0.10 or -0.10 inch from outer edge of door rails to ends of latch pins. After adjusting, tighten jam nuts. Refer to Figure 202. (2) Adjust plunger (73) in outer end of handle (19) using slot screwdriver. To tighten, turn clockwise; to loosen, turn counterclockwise. Adjust plunger (73) until 10 pounds, +5 or -0 pounds of force, applied at 0.55 inch from end of handle (19), is required to disengage plunger (73) from catch plate (71).
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Upper and Lower Cargo Doors Adjustment Figure 202 (Sheet 1)
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7.
Gas Spring Cylinder Disposal
WARNING: When removed, depressurize gas spring as described in the following steps before discarding. Protective eye covering must be worn while performing these steps. A.
8.
Dispose of Gas Spring Cylinder (Refer to Figure 203). (1) Place cylinder horizontally in bench vise and tighten vise. (2) Place several layers (4 layers minimum) of shop towels or rags over end of cylinder in vise. (3) Measure 1.50 inches in from fixed end of cylinder, and using a scratch awl or pointed center punch and hammer, drive awl or punch through the towels and into the cylinder until the gas begins to escape. (4) Hold towels and scratch awl in place until all gas has escaped (a few seconds). Then slowly remove scratch awl. Escaping oil will be absorbed by the towels. (5) While still holding towels over hole, push bright shaft completely into cylinder to purge remaining oil. (6) Remove gas spring from vise and discard.
Gas Spring/Upper Door Test Procedure A.
When gas spring operation problems are suspected on upper cargo door, the following test procedure should be used. The forces are measured with a spring ambient temperature of 68°F, +2 or -2°F, with piston rod extending downwards. (1) Obtain direct reading sensitive scale. (2) Remove pneumatic extender from airplane. (3) Grasp extender by cylinder in a vertical position with piston rod end down, and place on sensitive scale. (4) Compress device fully four or five times with piston rod end down. (This lubricates seals and piston.) (5) Compress approximately 1.5 inch and relax pressure slightly to allow piston rod to extend slowly until it is approximately 0.20 inch from full extension. Hold steady and read pressure. (6) An acceptable extender will read between 42 and 52 pounds on scale.
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Gas Spring Cylinder Disposal Figure 203 (Sheet 1)
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Gas Spring Cylinder Disposal Figure 203 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL IN-FLIGHT MOVABLE (ROLL- UP) DOOR - MAINTENANCE PRACTICES 1.
Description A.
2.
Operation A.
3.
4.
The lower door is opened and closed using a chain and sprocket system, driven by an electrical motor; however, if required the door may operated manually. The upper door is opened and closed by movement of the lower door to the open or closed position. Manual operation of doors is accomplished by removing the center fairing and dust cap from the aft fairing. Reach through the hole, pull the pin locking the motor to the torque tube, then rotate torque tube by hand to open or close the doors.
Upper and Lower Door Removal/Installation A.
Remove Upper and Lower Door (Refer to Figure 201 and Figure 202). (1) Disconnect electrical power from the airplane. (2) Remove forward fairing, center fairing, and aft fairing. (3) Remove electrical switch from aft track assembly. (4) Remove chain from sprocket on torque tube. (5) Remove bolts securing torque tube to forward track and aft track. Remove torque tube. (6) Pull lower door and upper door inboard out of the forward track and aft track to remove.
B.
Install Upper and Lower Door (Refer to Figure 201 and Figure 202). (1) Align lower door and upper door in forward track and aft track, pull down on lower door until both the upper and lower doors are completely installed in tracks. (2) Position torque tube on forward track and secure with bolts. (3) Install chain on torque tube sprocket. (a) Rig chain at both forward and aft tracks so that the same number of links are established between door/chain attach point link and the Þrst chain engagement tooth on bottom side of the sprocket which is common to torque tube assembly. (b) Apply highlight yellow paint to common sprocket tooth and chain link to maintain chain timing consistency during future system rigging operations. (4) Install electrical switch on aft track assembly. (5) Install forward, center and aft fairings. (6) Connect electrical power to airplane. (7) Make sure that the doors open and close smoothly. (a) Adjust the doors as necessary until they operate correctly.
Upper and Lower Door Adjustment/Test A.
5.
The inßight movable (roll-up) door is located on the left side of the fuselage, between fuselage station 234.00 and 284.00. The inßight movable (roll-up) door consists of an upper and lower door assembly, forward track, aft track, forward fairing, aft fairing, center fairing, sprockets, chains, chain covers, drive motor, limit switch, torque tube assembly and door jamb.
Adjust Upper and Lower Door. (1) Adjust chain tension to ensure tooth engagement at all sprocket locations throughout chain loop. (2) Optimize tension to minimize wear between chain and track race. Lubricate chain/track with MIL-PRF-81322 grease.
Side-Mounted Drive Assembly Removal/Installation A.
Remove the Side-Mounted Drive Assembly (Refer to Figure 202 ). (1) Remove electrical power from airplane. (2) Disconnect electrical connector from motor. (3) Remove bolts securing drive motor assembly to attach Þtting. (4) Remove drive motor assembly from airplane.
B.
Install the Side-Mounted Drive Assembly (Refer to Figure 202). (1) Position drive motor assembly to attach Þtting and secure with bolts. (2) Connect electrical connector to motor.
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Roll-Up Door Installation Figure 201 (Sheet 1)
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Roll-Up Door Installation Figure 201 (Sheet 2)
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Drive Assembly Installation Figure 202 (Sheet 1)
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Drive Assembly Installation Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (3) 6.
7.
Verify door opens and closes smoothly.
Bottom-Mounted Drive Assembly Removal/Installation A.
Remove the Bottom-Mounted Drive Assembly (Refer to Figure 202 ). (1) Remove the electrical power from the airplane. (2) Disconnect the motor electrical connector. (3) Remove the safety wire from the bolts that attach the drive assembly to the Þtting. (4) Remove the bolts that attach the drive assembly and chain tension plate to the Þtting. (5) Remove drive assembly from airplane.
B.
Install the Bottom-Mounted Drive Assembly (Refer to Figure 202). (1) Put the drive assembly and chain tension plate in position to the Þtting and chain. (a) Loosely install the bolts that attach the drive assembly to the Þtting. (2) Use hand pressure to move the drive assembly forward to tighten the chain. (a) Tighten the bolts that attach the drive assembly to the Þtting. (b) Install safety wire on the bolts. Refer to Chapter 20, Safetying - Maintenance Practices. (3) Connect the motor electrical connector. (4) Apply electrical power to the airplane. (5) Make sure the doors open and close smoothly. (a) Adjust the doors as necessary until they operate correctly.
Bottom-Mounted Drive Assembly Motor Removal/Installation A.
Remove the Bottom-Mounted Drive Assembly Motor (Refer to Figure 202). (1) Remove the bottom-mounted drive assembly. Refer to Bottom-Mounted Drive Assembly Removal/Installation. (2) Remove the nuts that attach the motor to the drive assembly. (3) Remove the motor from the drive assembly. (a) Keep the coupling for installation.
B.
Install the Bottom-Mounted Drive Assembly Motor (Refer to Figure 202). (1) Put the motor and coupling in position to the drive assembly. (2) Install the nuts that attach the motor to the drive assembly. (3) Install the bottom-mounted drive assembly. Refer to Bottom-Mounted Drive Assembly Removal/ Installation.
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MODEL 208 MAINTENANCE MANUAL AIR DEFLECTOR - MAINTENANCE PRACTICES 1.
General A.
2.
A removable cargo door and air deflector may be installed on the airplane. The air de flector is installed to lessen drag on the airplane when airplane is flown with the cargo door removed.
Upper Cargo Door Removal/Installation A.
Remove Upper Cargo Door (Refer to Figure 201). (1) Remove cotter pin and pin securing pull strap to lower door. (2) Remove safety clips from gas springs and remove gas springs from ball studs. (3) Remove cotter pins (1) and pins (3) from hinges (2) and remove door. NOTE:
B.
Install Upper Cargo Door (Refer to Figure 201 ). (1) Position door and install pins (3) in hinges (2), then install cotter pins (1). (2) Install gas springs on ballstuds and install safety clips. NOTE: (3)
3.
Remove lower cargo door.
Install lower cargo door.
Connect pull strap to lower door using pin and cotter pin.
Air Deflector Removal/Installation A.
Remove Air Deflector (Refer to Figure 201). (1) Remove nut (19) and screw (20). (2) Remove screw (18) and washer (17). (3) Remove cotter pin (1) and pin (3). (4) Remove cotter pins (13), washer (12), pins (10), and remove deflector (6).
B.
Install Air Deflector (Refer to Figure 201). (1) Position deflector (6) and install pins (3) and (10). (2) Install washers (12) and cotter pins (13). (3) Install cotter pin (1). (4) Install screw (20) and nut (19). (5) Install washer (17) and screw (18).
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Air Deflector Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL CREW DOOR ENTRANCE STEPS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
4.
The crew door entrance steps consist of metal, ladder-like steps attached to the floorboard just inside each crew door. The steps are hinged at the center, fold in half, then pivot at the floorboard attachments upside down to their stowed position just inside the crew doors.
Crew Door Step Assembly Removal/Installation A.
Remove Crew Door Step Assembly (Refer to Figure 201). (1) Remove nuts (1), Iockwashers (2), washers (3), spacers (4), washers (9), and bolts (10) attaching crew steps to mounting brackets (7) on floorboard.
B.
Install Crew Door Step Assembly (Refer to Figure 201). (1) Align holes in step assembly with holes in mounting brackets (7). (2) Install bolt (10) with washer (9) and spacer (10) through mounting bracket (7), then install spacer (4), washer (3), locknut (2), and nut (1).
Crew Door Step Disassembly/Assembly A.
Disassemble Crew Door Step (Refer to Figure 201). (1) Remove step assembly from airplane. (2) Remove lower arms (14) by removing nut (22), lockwasher (21), washer (20), spacer (19), washer (24), and bolt (25). (3) Remove screws (17) attaching upper step (12) to upper arms (11). (4) Remove screws (17) attaching lower step (15) to lower arms (14).
B.
Assemble Crew Door Step (Refer to Figure 201). (1) Position lower step (15) into slots in lower arms (14) and install screws (17). (2) Position upper step (12) into slots in upper arms (11) and install screws (17). (3) Align holes in lower arms (14) with holes in upper arms (11), install bolts (25) with washers (24), then install spacer (19), washer (20), Iockwasher (21), and nut (22). (4) Install step assembly on airplane.
Safety Walk Replacement A.
Replace Safety Walk. (1) Remove old safety walk material. (2) Clean surface with Acetone or equivalent. (3) Remove backing from new Safety Walk. (4) Apply Safety Walk material to step. (5) Form fillet around edges with 3M Edge Sealing Compound (or equivalent) per instructions on container.
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Crew Door Entrance Step Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL DOOR WARNING - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Door Warning Troubleshooting Chart Figure 101 (Sheet 1)
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Door Warning Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL DOOR WARNING - MAINTENANCE PRACTICES 1.
Description and Operation A.
2.
A cabin door warning system is utilized to provide visual indication on the annunciator panel when the passenger door or cargo door are not securely latched and the battery switch is in the on position. The switches are located just forward of the passenger and cargo door on the forward side of the fuselage frames.
Passenger and Cargo Door Warning Switches Adjustment/Test A.
Adjust Passenger and Cargo Door Warning Switches (Refer to Figure 201). (1) Determine if passenger or cargo door warning switch is out of adjustment. (2) Open passenger or cargo door to be adjusted. (3) Remove plug button (12) in forward passenger or cargo door jamb (13). (4) Position battery switch to ON. (5) Insert a small screwdriver through hole in door jamb and turn adjustment screw (11) counterclockwise to move switch (1) aft. This will cause the light in the annunciator to extinguish sooner when moving the door latch handle from open to the locked position. (6) Turn adjustment screw (11) clockwise to move switch (1) forward. This will cause the light in the annunciator to extinguish later when moving the door latch handle from open to the locked position. (7) Adjust door. (a) Close door and move latch handle towards the locked position. Door warning annunciator shall extinguish when handle lock ball is 1.50 to 2.12 inches between handle plunger and detent in catch plate. Adjust as required. (8) After final adjustment, replace plug button (12) in door jamb. (9) Position battery switch to OFF.
52-71-00 © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Door Warning Switch Installation Figure 201 (Sheet 1)
52-71-00 © Cessna Aircraft Company
Page 202 Mar 1/1999
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Door Warning Switch Installation Figure 201 (Sheet 2)
52-71-00 © Cessna Aircraft Company
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53 CHAPTER
FUSELAGE
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
53-00-00
Page 1
Aug 1/1995
53-10-00
Pages 1-5
Aug 1/1995
53-10-00
Pages 601-617
Jun 1/2011
53-20-06
Pages 201-203
Dec 1/2006
53-20-07
Pages 601-603
Jun 1/2011
53-21-00
Pages 201-204
Aug 1/1995
53-22-00
Pages 1-2
Aug 1/1995
53-23-00
Page 1
Aug 1/1995
53-25-00
Pages 201-202
Aug 1/1995
53-25-00
Pages 601-602
Jun 1/2011
53-50-00
Pages 201-202
Aug 1/1995
53-Title 53-List of Effective Pages 53-Record of Temporary Revisions 53-Table of Contents 53-List of Tasks
53 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS FUSELAGE - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-00-00 53-00-00 53-00-00 53-00-00
Page 1 Page 1 Page 1 Page 1
FUSELAGE - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-10-00 Page 1 53-10-00 Page 1 53-10-00 Page 1
FUSELAGE - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . External Fuselage Zonal Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Internal Cockpit Zonal Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Internal Cabin Zonal Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Internal Tail Cone Zonal Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Empennage and Horizontal Stabilizer Zonal Inspection . . . . . . . . . . . . . . . . . . . . . . . . . Carry-Through Root Rib Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Door Frames Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger and Cargo Door Frames Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . Firewall Brace and Doubler Assemblies Detailed Inspection. . . . . . . . . . . . . . . . . . . . . Stringers at Intersections with Forward and Aft Carry - Thru Bulkheads Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage Skin Doubler at Main Landing Gear Cutout Detailed Inspection. . . . . . . . . Fuselage Engine Mount Fittings Special Detailed Inspection . . . . . . . . . . . . . . . . . . . . Cargo and Passenger Door Doublers Special Detailed Inspection . . . . . . . . . . . . . . . Fuselage to Wing Attach Fitting Lugs Special Detailed Inspection . . . . . . . . . . . . . . . Lower Forward Carry-Thru Bulkhead Special Detailed Inspection . . . . . . . . . . . . . . . . Main Landing Gear Fitting Special Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Attach Fittings and Aft Carry-Thru Bulkhead Special Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage to Wing Carry-Thru Attach Fitting and Bulkhead Special Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage to Horizontal Stabilizer Attach Fittings Special Detailed Inspection . . . . . . Vertical Stabilizer Attach Points Special Detailed Inspection (Typical Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vertical Stabilizer Attach Points Special Detailed Inspection (Severe Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-10-00 Page 601 53-10-00 Page 601 53-10-00 Page 601 53-10-00 Page 602 53-10-00 Page 603 53-10-00 Page 606 53-10-00 Page 607 53-10-00 Page 608 53-10-00 Page 609 53-10-00 Page 609 53-10-00 Page 610 53-10-00 Page 611 53-10-00 Page 611 53-10-00 Page 612 53-10-00 Page 612 53-10-00 Page 613 53-10-00 Page 613 53-10-00 Page 614 53-10-00 Page 615 53-10-00 Page 615 53-10-00 Page 616 53-10-00 Page 616 53-10-00 Page 617
CARRY-THRU BULKHEAD FITTINGS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bolt Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-20-06 Page 201 53-20-06 Page 201 53-20-06 Page 201
FUSELAGE TO STRUT ATTACH FITTING LUGS - INSPECTION/CHECK . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Nominal Standard Bolt Size) (Typical Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Nominal Standard Bolt Size) (Severe Inspection Compliance). . . . . . . . . . . . . . . . . . . . . . . . . Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Oversize 1/64 Inch Bolt Size) (Severe Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Oversize 1/32Inch Bolt Size) (Severe Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-20-07 Page 601 53-20-07 Page 601 53-20-07 Page 601 53-20-07 Page 601 53-20-07 Page 602 53-20-07 Page 603
53 - CONTENTS © Cessna Aircraft Company
Page 1 of 2 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FLOORBOARDS AND ACCESS PLATES - MAINTENANCE PRACTICES . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Floorboard Removal/Installation (Passenger Airplanes). . . . . . . . . . . . . . . . . . . . . . . . . Floorboard Removal/Installation (Cargo Airplanes) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Access Plate Removal/Installation (Passenger Airplanes) . . . . . . . . . . . . . . . . . . . . . . . Access Plate Removal/Installation (Cargo Airplanes and Optional on 208B Passenger) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-21-00 Page 201 53-21-00 Page 201 53-21-00 Page 201 53-21-00 Page 201 53-21-00 Page 201 53-21-00 Page 201
PEDESTAL - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-22-00 Page 1 53-22-00 Page 1 53-22-00 Page 1
PLATES/SKIN - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-23-00 Page 1 53-23-00 Page 1
SEAT RAILS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Compartment Seat Rail Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Seat Rail Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-25-00 Page 201 53-25-00 Page 201 53-25-00 Page 201 53-25-00 Page 201
SEAT RAILS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Seat Rails and Attachment Structure Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . Bulkheads and Stiffeners Below the Seat Rail Attachments at FS 143.00 and FS 158.00 Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-25-00 Page 601 53-25-00 Page 601 53-25-00 Page 601 53-25-00 Page 602
TAIL STINGER - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stinger Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53-50-00 Page 201 53-50-00 Page 201 53-50-00 Page 201
53-21-00 Page 204
53 - CONTENTS © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 53-10-00-210
External Fuselage Zonal Inspection
53-10-00 Page 601
53-10-00-211
Internal Cockpit Zonal Inspection
53-10-00 Page 602
53-10-00-212
Internal Cabin Zonal Inspection
53-10-00 Page 603
53-10-00-213
Internal Tail Cone Zonal Inspection
53-10-00 Page 606
53-10-00-214
Empennage and Horizontal Stabilizer Zonal Inspection
53-10-00 Page 607
53-10-00-220
Carry-Through Root Rib Detailed Inspection
53-10-00 Page 608
53-10-00-221
Crew Door Frames Detailed Inspection
53-10-00 Page 609
53-10-00-222
Passenger and Cargo Door Frames Detailed Inspection
53-10-00 Page 609
53-10-00-223
Firewall Brace and Doubler Assemblies Detailed Inspection
53-10-00 Page 610
53-10-00-224
Stringers at Intersections with Forward and Aft Carry - Thru Bulkheads Detailed Inspection
53-10-00 Page 611
53-10-00-225
Fuselage Skin Doubler at Main Landing Gear Cutout Detailed Inspection
53-10-00 Page 611
53-10-00-250
Fuselage Engine Mount Fittings Special Detailed Inspection
53-10-00 Page 612
53-10-00-251
Cargo and Passenger Door Doublers Special Detailed Inspection
53-10-00 Page 612
53-10-00-252
Fuselage to Wing Attach Fitting Lugs Special Detailed Inspection
53-10-00 Page 613
53-10-00-253
Lower Forward Carry-Thru Bulkhead Special Detailed Inspection
53-10-00 Page 613
53-10-00-254
Main Landing Gear Fitting Special Detailed Inspection
53-10-00 Page 614
53-10-00-255
Main Landing Gear Attach Fittings and Aft Carry-Thru Bulkhead Special Detailed Inspection
53-10-00 Page 615
53-10-00-256
Fuselage to Wing Carry-Thru Attach Fitting and Bulkhead Special Detailed Inspection
53-10-00 Page 615
53-10-00-257
Fuselage to Horizontal Stabilizer Attach Fittings Special Detailed Inspection
53-10-00 Page 616
53-10-00-258
Vertical Stabilizer Attach Points Special Detailed Inspection (Typical Inspection Compliance)
53-10-00 Page 616
53-10-00-259
Vertical Stabilizer Attach Points Special Detailed Inspection (Severe Inspection Compliance)
53-10-00 Page 617
53-20-07-250
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Nominal Standard Bolt Size) (Typical Inspection Compliance)
53-20-07 Page 601
53-20-07-251
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Nominal Standard Bolt Size) (Severe Inspection Compliance)
53-20-07 Page 601
53-20-07-252
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Oversize 1/64 - Inch Bolt Size) (Severe Inspection Compliance)
53-20-07 Page 602
53-20-07-253
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Oversize 1/32- Inch Bolt Size) (Severe Inspection Compliance)
53-20-07 Page 603
53-25-00-220
Seat Rails and Attachment Structure Detailed Inspection
53-25-00 Page 601
53-25-00-221
Bulkheads and Stiffeners Below the Seat Rail Attachments at FS 143.00 and FS 158.00 Detailed Inspection
53-25-00 Page 602
53 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FUSELAGE - GENERAL 1.
Scope A.
2.
This chapter describes structural units and associated components which make up the compartments for equipment, passengers, crew and cargo.
Tools, Equipment and Materials
NAME
NUMBER
MANUFACTURER
USE
Scotch Speed Tape
Y435-3M
3M Co. Industrial Tape Division 3M Center St.Paul, MN 55101
To tape across openings in cargo compartment area.
3.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief de finition of the sections incorporated in this chapter is as follows: (1) The section on interior floorboards and attach fittings provides description, removal and installation procedures on cockpit, cabin and cargo floorboards, as well as fixtures attached to the floorboards. (2) The section on aerodynamic fairings provides removal and installation procedures for external fairings.
53-00-00 © Cessna Aircraft Company
Page 1 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FUSELAGE - DESCRIPTION AND OPERATION 1.
General A.
2.
This section describes the fuselage structure of the 208 and 208B airplanes in both the passenger and cargo configurations.
Description NOTE:
Refer to Figure 1 for a description of the major sections and fuselage station (FS) locations.
A.
The fuselage is of all- metal, semi-monocoque construction, with the skin carrying a portion of the structural load. The fuselage consists of the forward section, center section, tailcone section and stinger. Construction consists of formed bulkheads, longitudinal stringers, reinforcing channels and skins. (1) The fuselage forward section consists of the instrument panel, pedestal, left sidewall circuit breaker panel, seat rails, floorboards, access plates and two avionic equipment racks. One is located behind the right side instrument panel and the other is located under the copilot's floorboard. (2) The center section contains cargo/passenger compartment structures, floorboards, seat rails and access plates. (3) The tail cone section contains the oxygen cylinder, oxygen filler valve access plate, flight control cables, emergency locator transmitter, baggage partition, access plates and stinger.
B.
On the 208 and 208 Cargomaster, the fuselage forward section is all the fuselage structure from FS 100.00 to FS 166.45; the center section is from FS 166.45 to FS 284.00; and the tail cone section is from FS 284.00 to FS 427.88.
C.
On the 208B, 208B Super Cargomaster and 208B Passenger airplanes, the fuselage forward section is all the fuselage structure from FS 100.00 to FS 166.45; the center section is from FS 166.45 to FS 332.00; and the tail cone section is from FS 332.00 to FS 475.88. NOTE:
Refer to Figure 2 for an illustration of fuselage main frames.
D.
The main frame of the airplane fuselage includes transverse frames (bulkheads), formers, longerons, stringers, carry-thru spars and frames around openings. (1) The auxiliary structure consists of avionics equipment racks, floorboards, access plates and the pedestal. (2) The avionic equipment rack is located in front of the copilot’s seat, just forward of the door post.
E.
Attach fittings are provided on the fuselage for the horizontal stabilizer, wings, landing gear and seats. Carry-thru spars are provided through the fuselage for attachment of wings. Refer to appropriate chapters for a more complete description of how these attach fittings interface with the various components.
53-10-00 © Cessna Aircraft Company
Page 1 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Fuselage Sections Figure 1 (Sheet 1)
53-10-00 © Cessna Aircraft Company
Page 2 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Fuselage Sections Figure 1 (Sheet 2)
53-10-00 © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL
Fuselage Main Frame Figure 2 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Fuselage Main Frame Figure 2 (Sheet 2)
53-10-00 © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL FUSELAGE - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fuselage in a serviceable condition.
Task 53-10-00-210 2.
External Fuselage Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for an external zonal inspection of the fuselage. NOTE:
An external zonal GVI is a general visual examination of an exterior area, and/or an open installation or assembly to find damage, failure or defects. This level of inspection is made during typical lighting conditions such as daylight, hangar light or flashlight by approximately an arm-length distance to the inspection object. Unless it is specified, it is not necessary to remove or open access panels or doors to do an external GVI. You can use an inspection mirror to help with visual access to all opened surfaces in the inspection area. You can use maintenance stands, ladders, or platforms to get near the inspection area.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do the External Fuselage Zonal Inspection. NOTE: (1) (2) (3) (4) (5)
This inspection is from the forward tip of the nose spinner to the aft tip of the tailcone.
Examine the external fuselage for damage, failure, and signs of overheating. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, External Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. Examine all tubing, hose and fluid fittings for evidence of leaks, damage and chafing, and correct clamp installation. Examine all placards and markings for security of installation, legibility, and correct location. For the correct placards and placard locations, refer to the Model 208 Illustrated Parts Catalog or the Pilot's Operating Handbook. Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
E.
Restore Access (1) None End of task
53-10-00 © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL Task 53-10-00-211 3.
Internal Cockpit Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for an internal zonal inspection of the complete cockpit above and below the floorline. NOTE:
B.
Special Tools (1) None
C.
Access NOTE:
(1) (2) (3) (4) D.
An internal zonal GVI is a general visual examination that includes all of the systems and the structural components of an interior area, installation, or assembly. This includes a check for signs of corrosion, cracks, chafing of tubing, loose duct support, wiring damage, cable and pulley wear, fluid leaks, drainage that is not sufficient, and other conditions that can cause corrosion or damage. This level of inspection is made during typical lighting conditions such as daylight, hangar light, flood light, or flashlight by approximately an arm-length distance to the inspection object. It can be necessary to remove and/or open access panels or doors to complete an internal GVI. You can use an inspection mirror to help with visual access to all opened surfaces in the inspection area. You can use maintenance stands, ladders, or platforms to get near the inspection area.
The removal of the Primary Flight Displays (PFDs),and the Muti-Function Flight Display (MFD) from the instrument panels is not necessary, but it will help get access to the areas of this inspection.
Remove the flight crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. Remove the carpet in the cockpit to get access to the necessary floorboard panels. Remove center pedestal panels 226A, 226B, 226C, and 226D. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. Remove cockpit floorboard panels 211EL, 212FR, 231BL, 231DL, 232AC, 232BC, 232BR, and 232DR. Refer to Chapter 6, Access/Inspection Plates - Description and Operation.
Do the Internal Cockpit Zonal Inspection. NOTE:
(1)
(2)
This inspection is for the cockpit, and starts at and includes the aft side of the forward bulkhead (FS 100.00) to the aft end of the seat tracks (FS 166.45) above and below the floorline.
Examine all of the wire bundle assemblies and the electrical components for signs of overheating, correct installation, frayed or chafed wiring insulation, electrical bonding, damage, and corrosion. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. (a) Make sure that you examine the areas that follow between FS 100.00 to FS 118.00. The left and right fuselage skin areas, the left and right side longerons at BL 8.00 and BL 19.00 and outboard longerons. the channels and stiffeners common to the BL 8.00 longeron, the bottom bulkhead segment at FS 118.00, and the firewall support structures, brackets stiffeners, and doublers. (b) Make sure that you examine the areas that follow between FS 118.00 to FS 128.00. The left and right fuselage skin areas, the left and right side longerons at BL 8.00 and BL 19.00 and outboard longerons. the bulkhead segment at FS 128.00, the left and right side inboard control column support and pulley support structures.
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(3) (4) (5)
Make sure that you examine the areas that follow between FS 128.00 to FS 166.45. The top right and left fuselage side skin surface, the bottom forward skin surface, the center right side and center left side skin surface, the left and right side longerons at BL 8.00 and BL 19.00 and outboard longerons, the bulkhead segments at FS 143.00 and FS 158.00, the elevator bellcrank support assembly, the left and right side crew door sill assembly, the elevator trim and aileron and rudder pulley brackets. Examine all tubing, hose and fluid fittings for evidence of leaks, damage and chafing, and correct clamp installation. Examine all placards and markings for security of installation, legibility, and correct location. For the correct placards and placard locations, refer to the Model 208 Illustrated Parts Catalog or the Pilot's Operating Handbook. Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
E.
Restore Access (1) Install cockpit floorboard panels 211EL, 212FR, 231BL, 231DL, 232AC, 232BC, 232BR, and 232DR. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (2) Install the carpet. (3) Install center pedestal panels 226A, 226B, 226C, and 226D. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (4) Install the flight crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. End of task Task 53-10-00-212 4.
Internal Cabin Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for an internal zonal inspection of the cabin. NOTE:
An internal zonal GVI is a general visual examination that includes all of the systems and the structural components of an interior area, installation, or assembly. This includes a check for signs of corrosion, cracks, chafing of tubing, loose duct support, wiring damage, cable and pulley wear, fluid leaks, drainage that is not sufficient, and other conditions that can cause corrosion or damage. This level of inspection is made during typical lighting conditions such as daylight, hangar light, flood light, or flashlight by approximately an arm-length distance to the inspection object. It can be necessary to remove and/or open access panels or doors to complete an internal GVI. You can use an inspection mirror to help with visual access to all opened surfaces in the inspection area. You can use maintenance stands, ladders, or platforms to get near the inspection area.
B.
Special Tools (1) None
C.
Access (1) Remove the cabin seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. (2) Remove the aft bulkhead cabin partition. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices.
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D.
Open (unzip) the fabric headliner (passenger) or remove the hard shelled headliner (cargo) to get access to the areas of this inspection. Refer to Chapter 25, Cabin Upholstery - Maintenance Practices. Remove the carpet in the cabin to get access to the necessary floorboard panels. Remove the cabin floorboard panels. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (a) For the Models 208, and 208 Cargomaster, remove panels 251FL, 252AR, 252CR, 252FR, 255AL, 255DL, 255GL, 255JL, 255LL, 255NL, 255QL, 255SL, 256BR, 256ER, 256HR, 256KR, 256MR, 256PR, 256RR, 256TR. (b) For the models 208B, 208B Super Cargomaster, and 208B Passenger, remove panels 251BL, 251EL, 251ML, 255AL, 255DL, 255GL, 255KL, 255NL, 255RL, 255TL, 255VL, 255XL, 255ZL, 255ACL, 252BR, 252ER, 252GR, 252JR, 252MR, 256BR, 256ER, 256HR, 256LR, 256PR, 256SR, 256UR, 256WR, 256YR, 256AAR, 256ACR.
Do the Internal Cabin Zonal Inspection. NOTE:
(1)
(2)
This inspection is for the cabin, and starts at the forward side of (F.S. 166.45) to and including the forward side of the aft bulkhead (F.S.308.00 for the 208 or F.S. 356.00 for the 208B) above and below the floorline.
Examine all of the wire bundle assemblies and the electrical components for signs of overheating, correct installation, frayed or chafed wiring insulation, electrical bonding, damage, and corrosion. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. (a) Make sure that you examine the areas that follow between FS 166.45 to FS 208.00 for the Model 208 and FS 166.45 to FS 228.00 for the Model 208B. The fuselage left and right side skin surface, the upper forward left and right skin surface. (b) Make sure that you examine the areas that follow between FS 168.70 to FS 194.40 for the Model 208 and FS 188.70 to FS 214.40 for the Model 208B. The lower fuselage internal structure, the left and right side longerons at BL 9.00, BL 14.00 and BL 23.47. The lower carry-thru bulkhead segment and lower main landing gear bulkhead segment. The lower left and right side attach angle and stiffener. The center stiffener assembly and fuel reservoir support assembly. The belly skin internal surface between the forward carry-thru bulkhead structure and main landing gear bulkhead structure. (c) Make sure that you examine the areas that follow between FS 194.40 to FS 208.00 for the Model 208 and FS 214.40 to FS 228.00 for the Model 208B. The lower fuselage internal structure, the longitudinal bulkheads at BL 0.00, BL 13.97 and BL 23.50. The lower forward and aft main landing gear bulkhead segments. The sealing skin internal surface between the forward and aft main landing gear bulkhead. (d) Make sure that you examine the areas that follow between FS 166.45 to FS 186.45 for the Model 208 Only. The lower fuselage internal structure, the longitudinal bulkheads at BL 0.00, BL 13.97 and BL 23.50. The lower skin internal surface between the bulkhead at FS 166.45 and the forward carry-thru bulkhead at FS 186.45 and the seat tracks. (e) Make sure that you examine the areas that follow between FS 166.45 to FS 208.00 for the Model 208 and FS 166.45 to FS 228.00 for the Model 208B. The lower fuselage left and right side forward skin surface. The fuselage left and right lower center skin surface. The fuselage lower forward skin and lower center skin surface. The fuselage left and right side doublers. The left and right main landing gear bay stiffener. (f) Make sure that you examine the areas that follow between FS 208.00 to FS 322.80 for the Model 208 and FS 228.00 to FS 365.00 for the Model 208B. The fuselage upper aft skin surface. The fuselage left and right side aft skin surface. (g) Make sure that you examine the areas that follow between FS 208.00 to FS 284.00 for the Model 208 and FS 228.00 to FS 332.00 for the Model 208B. The lower fuselage internal structure, the longerons at BL 0.00, BL 13.97 and BL 23.47. The lower left and right side stiffener. The lower bulkhead segments between the aft main landing gear bulkhead and aft
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cargo/passenger door jamb bulkhead. The doublers at the longeron to door jamb bulkhead attachment. The lower skin internal surface between the aft main landing gear bulkhead and aft cargo/passenger door jamb bulkhead and the seat tracks. (h) Make sure that you examine the areas that follow between FS 284.00 to FS 308.00 for the Model 208 and FS 332.00 to FS 356.00 for the Model 208B. The lower fuselage internal structure, the outboard, inboard and center longerons. The lower bulkhead segments between the aft cargo/passenger door jamb bulkhead and curtain attach bulkhead. The lower skin internal surface between the aft cargo/passenger door jamb bulkhead and the curtain attach bulkhead. (i) Make sure that you examine the areas that follow between FS 284.00 to FS 308.00 for the Model 208 and FS 332.00 to FS 365.00 for the Model 208B. The fuselage lower-center aft-skin surface. The fuselage lower aft-skin surface. Examine all tubing, hose and fluid fittings for evidence of leaks, damage and chafing, and correct clamp installation. Examine all placards and markings for security of installation, legibility, and correct location. For the correct placards and placard locations, refer to the Model 208 Illustrated Parts Catalog or the Pilot's Operating Handbook. Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
E.
Restore Access (1) Install the cabin floorboard panels. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (a) For the Models 208, and 208 Cargomaster, install panels 251FL, 252AR, 252CR, 252FR, 255AL, 255DL, 255GL, 255JL, 255LL, 255NL, 255QL, 255SL, 256BR, 256ER, 256HR, 256KR, 256MR, 256PR, 256RR, 256TR. (b) For the models 208B, 208B Super Cargomaster, and 208B Passenger, install panels 251BL, 251EL, 251ML, 255AL, 255DL, 255GL, 255KL, 255NL, 255RL, 255TL, 255VL, 255XL, 255ZL, 255ACL, 252BR, 252ER, 252GR, 252JR, 252MR, 256BR, 256ER, 256HR, 256LR, 256PR, 256SR, 256UR, 256WR, 256YR, 256AAR, 256ACR. (2) Install the carpet in the cabin. (3) Close (zip) the fabric headliner (passenger) or install the hard shelled headliner (cargo). Refer to Chapter 25, Cabin Upholstery - Maintenance Practices. (4) Install the aft bulkhead cabin partition. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices. (5) Install the cabin seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. End of task
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Internal Tail Cone Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for an internal zonal inspection of the tail section. NOTE:
An internal zonal GVI is a general visual examination that includes all of the systems and the structural components of an interior area, installation, or assembly. This includes a check for signs of corrosion, cracks, chafing of tubing, loose duct support, wiring damage, cable and pulley wear, fluid leaks, drainage that is not sufficient, and other conditions that can cause corrosion or damage. This level of inspection is made during typical lighting conditions such as daylight, hangar light, flood light, or flashlight by approximately an arm-length distance to the inspection object. It can be necessary to remove and/or open access panels or doors to complete an internal GVI. You can use an inspection mirror to help with visual access to all opened surfaces in the inspection area. You can use maintenance stands, ladders, or platforms to get near the inspection area.
B.
Special Tools (1) None
C.
Access (1) Remove the aft bulkhead cabin partition. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices. (2) Remove tailcone panel 320A. Refer to Chapter 6, Access Plates And Panels Identification Description and Operation.
D.
Do the Internal Tail Cone Zonal Inspection. NOTE: (1)
(2)
(3) (4)
This inspection is starts and includes the aft side of the aft cabin bulkhead (F.S. 308.00 for the 208 and F.S. 356.00 for the 208B) and goes to the tip of the tailcone.
Examine all of the wire bundle assemblies and the electrical components for signs of overheating, correct installation, frayed or chafed wiring insulation, electrical bonding, damage, and corrosion. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. (a) Make sure that you examine the areas that follow between FS 322.00 to FS 388.68 for the Model 208 and FS 365.00 to FS 436.68 for the Model 208B. The tailcone upper and lower skin surface. The tailcone left and right side skin surface. The tailcone dorsal skin surface. (b) Make sure that you examine the areas that follow, FS 388.68 for the Model 208 and FS 436.68 for the Model 208B. The upper and lower tailcone canted bulkhead, The elevator bellcrank bracket assembly including the bracket, angles, doublers and stiffeners. The left and right stabilizer attach fittings. (c) Make sure that you examine the areas that follow between FS 388.68 to FS 427.88 for the Model 208 and FS 436.68 to FS 475.88 for the Model 208B. The left and right side aft tailcone skin surface. The lower aft tailcone skin surface. Examine all tubing, hose and fluid fittings for evidence of leaks, damage and chafing, and correct clamp installation. Examine all placards and markings for security of installation, legibility, and correct location. For the correct placards and placard locations, refer to the Model 208 Illustrated Parts Catalog or the Pilot's Operating Handbook.
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Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
E.
Restore Access (1) Install tailcone panel 320A. Refer to Chapter 6, Access Plates And Panels Identification Description and Operation. (2) Install the aft bulkhead cabin partition. Refer to Chapter 25, Rear Compartment Wall Maintenance Practices. End of task Task 53-10-00-214 6.
Empennage and Horizontal Stabilizer Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for an external zonal inspection of the empennage and horizontal stabilizer. NOTE:
An external zonal GVI is a general visual examination of an exterior area, and/or, an open installation or assembly to find damage, failure or defects. This level of inspection is made during typical lighting conditions such as daylight, hangar light or flashlight by approximately an arm-length distance to the inspection object. Unless it is specified, it is not necessary to remove or open access panels or doors to do an external GVI. You can use an inspection mirror to help with visual access to all opened surfaces in the inspection area. You can use maintenance stands, ladders, or platforms to get near the inspection area.
B.
Special Tools (1) None
C.
Access (1) Remove the tail stinger Refer to Tail Stinger - Maintenance Practices. (2) Remove the horizontal stabilizer fairings and inspection access panels. Refer to Chapter 55, Horizontal Stabilizer - Removal/Installation. (3) Remove vertical access panels 340A, 341A, 341B, 341C, and rudder access panel 343A. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (4) Remove Horizontal panels 373AL, 373BL, 374AR, 374BR, and tail cone access panel 320A. Refer to Chapter 6, Access/Inspection Plates - Description and Operation.
D.
Do the External Zonal Inspection of the Empennage and Horizontal Stabilizer. NOTE: (1) (2) (3)
This inspection is external from the forward tip of the vertical stabilizer forward fin to the aft tip of the tailcone, and from the upper tip to the bottom surface of the horizontal stabilizers.
Examine the external horizontal stabilizer, vertical stabilizer, and empennage for damage and signs of overheating. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/ Check, External Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. Examine all placards and markings for security of installation, legibility, and correct location. For the correct placards and placard locations, refer to the Model 208 Illustrated Parts Catalog or the Pilot's Operating Handbook.
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E.
Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
Do the Internal Zonal Inspection of the Empennage and Horizontal Stabilizer. NOTE:
(1) (2)
(3)
(4) (5) (6)
This inspection is internal from the forward tip of the vertical stabilizer forward fin to the aft tip of the tailcone, and from the upper tip to the bottom surface of the vertical stabilizer to include attach points accessible through the tail cone.
Examine all horizontal and vertical stabilizer attach points, attach fasteners, bolts, hardware, and related attach fitting structures for damage, corrosion, cracks, loose fasteners, loose or unsafetied hardware, and correct installation. Examine all of the wire bundle assemblies and the electrical components for signs of overheating, correct installation, frayed, or chafed wiring insulation, electrical bonding, damage, and corrosion. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and related structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. (a) Make sure that you examine the areas that follow, FS 427.88 for the Model 208 and FS 475.88 for the Model 208B. The tailcone aft canted bulkhead. The left and right side doublers. The support bracket and stiffeners. The forward and aft stabilizer attach fittings. Examine all tubing, hose, and fluid fittings for signs of leaks, damage, chafing, and correct clamp installation. Examine all placards and markings for security of installation, legibility, and correct location. For the correct placards and placard locations, refer to the Model 208 Illustrated Parts Catalog or the Pilot's Operating Handbook. Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
F.
Restore Access (1) Install the tail stinger Refer to Tail Stinger - Maintenance Practices. (2) Install the horizontal stabilizer fairings and inspection access panels. Refer to Chapter 55, Horizontal Stabilizer - Removal/Installation. (3) Install vertical access panels 340A, 341A, 341B, 341C, and rudder access panel 343A. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. (4) Install Horizontal panels 373AL, 373BL, 374AR, 374BR, and tail cone access panel 320A. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. End of task Task 53-10-00-220 7.
Carry-Through Root Rib Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the carry-through root rib in a serviceable condition.
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Special Tools (1) None
C.
Access (1) Remove the wing from the airplane. Refer to Chapter 57, Wings - Removal/Installation.
D.
Do a Carry-Through Root Rib Detailed Inspection. (1) Do a visual inspection of the root rib for cracks. (2) If no cracks are found, install the wing on the airplane. Refer to Chapter 57, Wings - Removal/ Installation. (3) If cracks are found, repair or replace the root rib. Refer to Chapter 57, Wings - Removal/ Installation or the Model 208 Structural Repair Manual.
E.
Restore Access (1) Install the wing on the airplane. Refer to Chapter 57, Wings - Removal/Installation. End of task Task 53-10-00-221 8.
Crew Door Frames Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the crew door frames in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the left and the right crew door upper and lower interior panels. Refer to Crew Door Maintenance Practices.
D.
Do a visual inspection of the crew door frames for cracks. Refer to Chapter 52, Crew Doors Maintenance Practices. (1) With the crew doors open, examine the corners and around the jamb assembly and the area around the hinges for cracks, corrosion or damage. (a) Replace the jamb assembly if cracks or damage are found. Refer to Chapter 52, Crew Doors - Maintenance Practices. (b) If corrosion is found, refer to Chapter 51, Corrosion Prevention and Control Program Description and Operation. (2) Examine all exposed frame areas for cracks. (3) If cracks are found, repair or replace the damaged part(s). Refer to Chapter 52, Crew Doors Maintenance Practices.
E.
Restore Access (1) Install the left and the right crew door upper and lower interior panels. Refer to Crew Door Maintenance Practices. End of task Task 53-10-00-222 9.
Passenger and Cargo Door Frames Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the passenger and cargo door frames in a serviceable condition.
B.
Special Tools (1) None
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Access (1) Open the passenger and cargo doors.
D.
Examine the Passenger and Cargo Door Frames. (1) Do a visual inspection of the passenger door frames for cracks. (a) If cracks are found, repair or replace the part(s). Refer to Chapter 52, Passenger Doors Maintenance Practices. (b) Examine the corners and around the jamb assembly for cracks or damage. (c) Replace the jamb assembly if it is cracked or damaged. Refer to Chapter 52, Passenger Doors - Maintenance Practices. (2) Do a visual inspection of the cargo door frames for cracks. (a) If cracks are found, repair or replace the part(s). Refer to Chapter 52, Cargo Doors Maintenance Practices. (b) Examine the corners and hinge areas around the jamb assembly for cracks, corrosion or damage. (c) Replace the jamb assembly if it is cracked or damaged. Refer to Chapter 52, Cargo Doors - Maintenance Practices. (d) If corrosion is found, refer to Chapter 51, Corrosion Prevention and Control Program Description and Operation.
E.
Restore Access (1) Close the passenger and cargo doors. End of task Task 53-10-00-223 10.
Firewall Brace and Doubler Assemblies Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the firewall brace and doubler assemblies in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the engine cowling to get access to the firewall engine mount assemblies. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Examine the Firewall Brace and Doubler Assemblies. (1) Do a visual inspection of the doubler and supports for cracks on the forward and the aft sides of the firewall. (a) If cracks are found, repair or replace the part(s). Refer to Chapter 71, Engine Mount Maintenance Practices and the Model 208 Structural Repair Manual. (2) Do a visual inspection of the firewall brace and the adjacent web for cracks that come from the fastener holes on the forward and the aft sides of the firewall. (a) If cracks are found, repair or replace the part(s). Refer to Chapter 71, Engine Mount Maintenance Practices and the Model 208 Structural Repair Manual.
E.
Restore Access (1) Install the engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task
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Stringers at Intersections with Forward and Aft Carry - Thru Bulkheads Detailed Inspection NOTE:
Models 208/208A have the stringers installed between FS 168.00 and FS 195.00.
NOTE:
The Model 208B has the stringers installed between FS 188.00 and FS 215.00.
A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the stringers at intersections with forward and aft carry - thru bulkheads in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the passenger seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. (2) Remove the upholstery and floorboard panels below the passenger seat locations. Refer to Chapter 53, Floorboards and Access Plates - Maintenance Practices.
D.
Examine the stringers at Intersections with Forward and Aft Carry - Thru Bulkheads. (1) Do a visual inspection of the stringers for cracks. (a) Do a visual inspection of the corners around the attach fittings for cracks. 1 Do a visual inspection of the intersections with the forward and aft carry-thru bulkheads. (2) If cracks are found in the stringer(s), repair or replace the stringer(s). Refer to the Model 208 Structural Repair Manual. (3) If no cracks are found, restore access.
E.
Restore Access (1) Install the upholstery and floorboard panels below the passenger seat locations. Refer to Chapter 53, Floorboards and Access Plates - Maintenance Practices. (2) Install the passenger seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. End of task Task 53-10-00-225 12.
Fuselage Skin Doubler at Main Landing Gear Cutout Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the fuselage skin doubler at the main landing gear cutout in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the main landing gear fairing. Refer to Chapter 32, Main Landing Gear - Maintenance Practices.
D.
Examine the Fuselage Skin Doubler at the Main Landing Gear Cutout. (1) Do a visual inspection for cracks and gouges in the doubler. (a) Do a visual inspection of the area where the main landing gear fairing attaches. (2) If cracks or gouges are found, repair the cracks or gouges with the instructions given in Service Bulletin CAB03-6. (3) If no cracks or gouges are found, restore access.
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Restore Access (1) Install the main landing gear fairing. Refer to Chapter 32, Main Landing Gear - Maintenance Practices. End of task Task 53-10-00-250 13.
Fuselage Engine Mount Fittings Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the fuselage engine mount fittings in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the engine cowling. Refer to Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a Special Detailed Inspection of the Fuselage Engine Mount Fittings. (1) Do a nondestructive testing (NDT) inspection of the four truss assembly attachment points. Refer to the Model 208 Nondestructive Testing Manual, Part 6, Eddy Current, Fuselage Engine Mount Fittings - Description And Operation. (2) Do a visual inspection of the gusset for cracks around the engine truss assembly attachment points. (3) Do a visual inspection of the flange rings for cracks. (4) Do an NDT inspection of the upper engine mount attachment. Refer to the Model 208 Nondestructive Testing Manual, Part 6, Eddy Current, Fuselage Engine Mount Fittings Description And Operation. (5) If no cracks are found, restore access. (6) If cracks are found, repair or replace the damaged part(s). Refer to Chapter 71, Engine Mount - Maintenance Practices or the Model 208 Structural Repair Manual.
E.
Restore Access (1) Install the engine cowling. Refer to Engine Cowling and Nose Cap - Maintenance Practices. End of task Task 53-10-00-251 14.
Cargo and Passenger Door Doublers Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the cargo and passenger door doublers in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the upper and the lower passenger door interior panels. Refer to Chapter 52, Passenger Doors - Maintenance Practices. (2) Remove the upper and the lower cargo door interior panels. Refer to Chapter 52, Cargo Doors - Maintenance Practices.
D.
Do a Special Detailed Inspection of the Cargo and Passenger Door Doublers. (1) Do a nondestructive testing (NDT) inspection of the upper passenger frame doublers and corners for cracks. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Cargo and Passenger Door Doublers - Description And Operation.
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Do an NDT inspection of the upper cargo frame doublers and corners for cracks. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Cargo and Passenger Door Doublers - Description And Operation. If no cracks are found, restore access. If cracks are found, repair or replace the damaged part(s). (a) For the passenger doors, refer to Chapter 52, Passenger Doors - Maintenance Practices or the Model 208 Structural Repair Manual. (b) For the cargo doors, refer to Chapter 52, Cargo Doors - Maintenance Practices or the Model 208 Structural Repair Manual.
E.
Restore Access (1) Install the upper and the lower cargo door interior panels. Refer to Chapter 52, Cargo Doors Maintenance Practices. (2) Install the upper and the lower passenger door interior panels. Refer to Chapter 52, Passenger Doors - Maintenance Practices. End of task Task 53-10-00-252 15.
Fuselage to Wing Attach Fitting Lugs Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the fuselage to wing attach fitting lugs in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing. Refer to Chapter 57, Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Fuselage to Wing Attach Fitting Lugs. (1) Do a nondestructive testing (NDT) inspection of the bolt holes in the attach fittings lugs. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Fuselage to Wing Attach Fitting Lugs - Description And Operation. (2) If no cracks are found, restore access. (3) If cracks are found, replace the fittings.
E.
Restore Access (1) Install the wing. Refer to Chapter 57, Wings - Removal/Installation. End of task Task 53-10-00-253 16.
Lower Forward Carry-Thru Bulkhead Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the lower forward carry-thru bulkhead in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the passenger seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. (2) Remove the upholstery and floorboard panels below the passenger seat locations. Refer to Floorboards and Access Plates - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL D.
Do a Special Detailed Inspection of the Lower Forward Carry-Thru Bulkhead. NOTE:
The lower forward carry-thru bulkhead on Models 208/208A are installed at FS 166.00 and FS 168.00.
NOTE:
The lower forward carry-thru bulkhead on Model 208B are installed at FS 186.00 and FS 188.00.
(1) (2) (3) (4) (5)
Do a visual inspection of the bulkhead and the frames for cracks. Do a visual inspection of the corners of the frames around the strut attach fittings. Do a nondestructive testing (NDT) inspection of the lower forward carry-thru bulkhead. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Lower Forward Carry-Thru Bulkhead - Description And Operation. If no cracks are found, restore access. If cracks are found in the attach fitting(s), replace the attach fitting(s).
E.
Restore Access (1) Install the upholstery and floorboard panels below the passenger seat locations. Refer to Floorboards and Access Plates - Maintenance Practices. (2) Install the passenger seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. End of task Task 53-10-00-254 17.
Main Landing Gear Fitting Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the main landing gear fitting in a serviceable condition.
B.
Special Tools (1) Airplane Jacks (2) Tail Stand
C.
Access (1) Remove the main landing gear fairing. Refer to Chapter 32, Main Landing Gear - Maintenance Practices, Main Gear Fairing Removal/Installation.
D.
Do a Special Detailed Inspection of the Main Landing Gear Fitting. (1) Use jacks to lift the airplane. Refer to Chapter 7, Jacking - Maintenance Practices. (2) Do a nondestructive testing (NDT) inspection of the main landing gear fitting attachment holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Main Landing Gear Fitting - Description And Operation. (3) If no cracks are found, do the steps that follow: (a) Lower the airplane and remove the jacks. Refer to Chapter 7, Jacking - Maintenance Practices. (b) Restore access. (4) If cracks are found in the attach fitting(s), do the steps that follow: (a) Replace the main landing gear attach fitting. Refer to Chapter 32, Main Landing Gear Maintenance Practices. (b) Lower the airplane and remove the jacks. Refer to Chapter 7, Jacking - Maintenance Practices. (c) Restore access.
E.
Restore Access (1) Install the main landing gear fairing. Refer to Chapter 32, Main Landing Gear - Maintenance Practices, Main Gear Fairing Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL Task 53-10-00-255 18.
Main Landing Gear Attach Fittings and Aft Carry-Thru Bulkhead Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the main landing gear attach fittings and aft carry-thru bulkhead in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the passenger seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. (2) Remove the upholstery and floorboard panels below the passenger seat locations. Refer to Floorboards and Access Plates - Maintenance Practices.
D.
Do a Special Detailed Inspection of the Main Landing Gear Attach Fittings and Aft Carry-Thru Bulkhead. NOTE:
The main landing gear attach fittings and aft carry-thru bulkhead on Models 208/208A are installed at FS 194.40 and FS 208.00.
NOTE:
The main landing gear attach fittings and aft carry-thru bulkheads on Airplanes Model 208B are installed at FS 214.40 and FS 228.00.
(1) (2) (3)
(4) (5) (6)
Do a visual inspection of the corners around the attach fittings for cracks. Do a visual inspection of the corners of the frames around the strut attach fittings. Do a nondestructive testing (NDT) inspection of the main landing gear attach fittings and lower aft carry-thru bulkhead. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Main Landing Gear Attach Fittings and Aft Carry-Thru Bulkhead - Description And Operation. If no cracks are found, restore access. If cracks are found in the lower aft carry-thru bulkhead, repair the bulkhead. Refer to the Model 208 Structural Repair Manual. If cracks are found in the attach fitting(s), replace the attach fitting(s).
E.
Restore Access (1) Install the upholstery and floorboard panels below the passenger seat locations. Refer to Floorboards and Access Plates - Maintenance Practices. (2) Install the passenger seats. Refer to Chapter 25, Passenger Seats - Maintenance Practices. End of task Task 53-10-00-256 19.
Fuselage to Wing Carry-Thru Attach Fitting and Bulkhead Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the fuselage to wing carry-thru attach fitting and bulkhead in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the upholstery and floorboard panels to get access to the fuselage to wing carry-thru bulkhead. Refer to Floorboards and Access Plates - Maintenance Practices.
D.
Do a Special Detailed Inspection of the Fuselage to Wing Carry-Thru Attach Fitting and Bulkhead. (1) Do a visual inspection of the corners around the attach fittings for cracks.
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5)
Do a nondestructive testing (NDT) inspection of the fuselage to wing carry-thru attach fitting. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Fuselage to Wing Carry-Thru Attach Fitting - Description And Operation. If no cracks are found, restore access. If cracks are found in the fuselage to wing carry-thru bulkhead, repair the bulkhead. Refer to the Model 208 Structural Repair Manual. If cracks are found in the wing carry-thru attach fitting, replace the attach fitting.
E.
Restore Access (1) Install the upholstery and floorboard panels. Maintenance Practices. End of task
Refer to Floorboards and Access Plates -
Task 53-10-00-257 20.
Fuselage to Horizontal Stabilizer Attach Fittings Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the fuselage to horizontal stabilizer attach fittings in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the horizontal stabilizer. Installation.
D.
Refer to Chapter 55, Horizontal Stabilizer - Removal/
Do a Special Detailed Inspection of the Fuselage to Horizontal Stabilizer Attach Fittings. (1) Do a visual inspection for cracks in the forward side of the horizontal stabilizer forward attach fitting. (2) Do a nondestructive testing (NDT) inspection for cracks at the horizontal stabilizer forward attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Fuselage to Horizontal Stabilizer Attach Fittings - Description And Operation. (3) Do a visual inspection for cracks in the fuselage side of the horizontal stabilizer aft attach fitting. (4) Do a NDT inspection for cracks at the horizontal stabilizer aft attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Fuselage to Horizontal Stabilizer Attach Fittings - Description And Operation. (5) If no cracks are found, restore access. (6) If cracks are found, replace the damaged parts. Refer to Chapter 55, Horizontal Stabilizer Removal/Installation.
E.
Restore Access (1) Install the horizontal stabilizer. Refer to Chapter 55, Horizontal Stabilizer - Removal/Installation. End of task Task 53-10-00-258 21.
Vertical Stabilizer Attach Points Special Detailed Inspection (Typical Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the vertical stabilizer attach points in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation.
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MODEL 208 MAINTENANCE MANUAL D.
Do a Special Detailed Inspection of the Vertical Stabilizer Attach Points. (1) Do a visual inspection for cracks in the fuselage side of the vertical stabilizer forward attach point. (2) Do a nondestructive testing (NDT) inspection for cracks at the vertical stabilizer forward attach point holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Attach Points - Description And Operation. (3) Do a visual inspection for cracks in the fuselage side of the vertical stabilizer aft attach point. (4) Do a NDT inspection for cracks at the vertical stabilizer aft attach point holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Attach Points - Description And Operation. (5) If no cracks are found, restore access. (6) If cracks are found, replace the damaged parts or contact Cessna Propeller Aircraft Product Support for repair procedures. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation.
E.
Restore Access (1) Install the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation. End of task Task 53-10-00-259 22.
Vertical Stabilizer Attach Points Special Detailed Inspection (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the vertical stabilizer attach points in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation.
D.
Do a Special Detailed Inspection of the Vertical Stabilizer Attach Points. (1) Do a visual inspection for cracks in the fuselage side of the vertical stabilizer forward attach point. (2) Do a nondestructive testing (NDT) inspection for cracks at the vertical stabilizer forward attach point holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Attach Points - Description And Operation. (3) Do a visual inspection for cracks in the fuselage side of the vertical stabilizer aft attach point. (4) Do a NDT inspection for cracks at the vertical stabilizer aft attach point holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Attach Points - Description And Operation. (5) If no cracks are found, restore access. (6) If cracks are found, replace the damaged parts or contact Cessna Propeller Aircraft Product Support for repair procedures. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation.
E.
Restore Access (1) Install the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL CARRY-THRU BULKHEAD FITTINGS - MAINTENANCE PRACTICES 1.
2.
General A.
Supplemental inspections 53-20-03 and 53-20-06 require inspection of certain bolt holes for cracks by the eddy current method. Because access to the nuts for torquing is limited, the production conÞguration of the bolts and nuts may not be practical for reinstallation of the bolts.
B.
This procedure allows and provides a method of installation of bolts that will make installation easier.
Bolt Replacement A.
This procedure does not identify the correct part number bolt for each installation. It is the responsibility of the technician to correctly identify each bolt removed and reinstall the correct part number.
CAUTION: During the inspection of the bolt holes it is critical that only one bolt be removed at a time. Complete the inspection for one hole and then install the bolt before you remove the next bolt. This will help prevent damage to the Þttings. B.
Replace bolt (Refer to Figure 201). (1) Figure 201 illustrates the bolts to be inspected, and shows the production conÞguration of the bolts. (2) You can install each bolt shown with the head on the opposite side. (3) Because the clearance between the bulkhead frames is limited for inserting the short bolts that were removed, the following is suggested: (a) From the outside of the bulkhead, insert a segment of safety wire until the end is visible. (b) Attach the end of the wire to the tip of the bolt with 5-minute epoxy or similar material. (c) Allow the epoxy to cure, and use the wire to guide the bolt through the bolt hole. (d) Install nuts and washers and torque as required to standard torque.
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Bolt IdentiÞcation Figure 201 (Sheet 1)
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Bolt IdentiÞcation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL FUSELAGE TO STRUT ATTACH FITTING LUGS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the fuselage to strut attach fitting lugs in a serviceable condition. General
Task 53-20-07-250 2.
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Nominal Standard Bolt Size) (Typical Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to do a special detailed inspection of the fuselage to strut attach fitting lugs.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. (2) Remove the struts. Refer to Chapter 57, Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Fuselage to Strut Attach Fitting Lugs. NOTE:
The Model 208/208A airplanes have the fuselage to strut attach fitting lugs installed between FS 166.00 and FS 168.00.
NOTE:
The Model 208B airplane has the fuselage to strut attach fitting lugs installed between FS 186.00 and FS 188.00.
(1) (2) (3)
Do a nondestructive testing (NDT) inspection for cracks in the fuselage attach fittings. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Fuselage to Strut Attach Fitting Lugs - Description And Operation. If no cracks are found, restore access. If cracks are found, replace the strut attach fitting lug. Refer to Chapter 57, Wings - Removal/ Installation.
E.
Restore Access (1) Install the struts. Refer to Chapter 57, Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. End of task Task 53-20-07-251 3.
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Nominal Standard Bolt Size) (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to do a special detailed inspection of the fuselage to strut attach fitting lugs.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. (2) Remove the struts. Refer to Chapter 57, Wings - Removal/Installation.
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MODEL 208 MAINTENANCE MANUAL D.
Do a Special Detailed Inspection of the Fuselage to Strut Attach Fitting Lugs. NOTE:
The Model 208/208A airplanes have the fuselage to strut attach fitting lugs installed between FS 166.00 and FS 168.00.
NOTE:
The Model 208B airplane has the fuselage to strut attach fitting lugs installed between FS 186.00 and FS 188.00.
(1) (2) (3)
Do a nondestructive testing (NDT) inspection for cracks in the fuselage attach fittings. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Fuselage to Strut Attach Fitting Lugs - Description And Operation. If no cracks are found, restore access. If cracks are found, replace the strut attach fitting lug. Refer to Chapter 57, Wings - Removal/ Installation.
E.
Restore Access (1) Install the struts. Refer to Chapter 57, Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. End of task Task 53-20-07-252 4.
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Oversize 1/64 - Inch Bolt Size) (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to do a special detailed inspection of the fuselage to strut attach fitting lugs.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. (2) Remove the struts. Refer to Chapter 57, Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Fuselage to Strut Attach Fitting Lugs. NOTE:
The Model 208/208A airplanes have the fuselage to strut attach fitting lugs installed between FS 166.00 and FS 168.00.
NOTE:
The Model 208B airplane has the fuselage to strut attach fitting lugs installed between FS 186.00 and FS 188.00.
(1) (2) (3)
Do a nondestructive testing (NDT) inspection for cracks in the fuselage attach fittings. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Fuselage to Strut Attach Fitting Lugs - Description And Operation. If no cracks are found, restore access. If cracks are found, replace the strut attach fitting lug. Refer to Chapter 57, Wings - Removal/ Installation.
E.
Restore Access (1) Install the struts. Refer to Chapter 57, Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL Task 53-20-07-253 5.
Fuselage to Strut Attach Fitting Lugs Special Detailed Inspection (Oversize 1/32- Inch Bolt Size) (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to do a special detailed inspection of the fuselage to strut attach fitting lugs.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. (2) Remove the struts. Refer to Chapter 57, Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Fuselage to Strut Attach Fitting Lugs. NOTE:
The Model 208/208A airplanes have the fuselage to strut attach fitting lugs installed between FS 166.00 and FS 168.00.
NOTE:
The Model 208B airplane has the fuselage to strut attach fitting lugs installed between FS 186.00 and FS 188.00.
(1) (2) (3)
Do a nondestructive testing (NDT) inspection for cracks in the fuselage attach fittings. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Fuselage to Strut Attach Fitting Lugs - Description And Operation. If no cracks are found, restore access. If cracks are found, replace the strut attach fitting lug. Refer to Chapter 57, Wings - Removal/ Installation.
E.
Restore Access (1) Install the struts. Refer to Chapter 57, Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Chapter 57, Wings - Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL FLOORBOARDS AND ACCESS PLATES - MAINTENANCE PRACTICES 1.
General A.
2.
Tools, Equipment and Materials A.
3.
4.
This section provides removal and installation instructions for floorboards and access panels used in the cargo and passenger airplanes.
For a list of required tools, equipment and materials, refer to Fuselage - General.
Floorboard Removal/Installation (Passenger Airplanes) A.
Remove Floorboards (Refer to Figure 201). (1) Remove crew and commuter/utility seats. Refer to Chapter 25, Flight Compartment Maintenance Practices and Chapter 25, Passenger Seats - Maintenance Practices. (2) Remove vinyl floor covering. (3) Remove attaching screws and remove access plates. (4) Remove rivets, as required, securing floorboards to floor supports, and remove floorboards.
B.
Install Floorboards (Refer to Figure 201). (1) Position floorboards to airplane, and install rivets securing floorboards to floor supports. (2) Position access plates and tighten screws. (3) Install vinyl floor covering. (4) Install crew and commuter/utility seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices and Chapter 25, Passenger Seats - Maintenance Practices.
Floorboard Removal/Installation (Cargo Airplanes) A.
Remove Floorboards (Refer to Figure 201). (1) Remove crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (2) Remove screws securing plywood floorboard coverings in cargo area, and remove coverings. NOTE: (3) (4) (5)
B.
5.
Note dimensions of tape over access openings under plywood floorboard coverings and install tape of the same size on reinstallation, with a minimum of 0.60 inch overlap.
Remove tape from access openings. Remove screws securing access plates in flight crew area, and remove plates. Remove rivets, as required, securing floorboards to floor supports and remove floorboards.
Install Floorboards (Refer to Figure 201). (1) Position floorboards to airplane and install rivets securing floorboards to floor supports. (2) Position access plates and install screws. (3) Tape access openings in cargo compartment with Scotch Speed Tape, using a minimum of 0.60 inch overlap. (4) Position plywood floorboard coverings and secure with screws. (5) Install crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices.
Access Plate Removal/Installation (Passenger Airplanes) A.
Remove Access Plates (Refer to Figure 201). (1) Remove crew and commuter/utility seats. Refer to Chapter 25, Flight Compartment Maintenance Practices and Chapter 25, Passenger Seats - Maintenance Practices. (2) Remove vinyl floor covering. (3) Remove attaching screws and remove access plates.
B.
Install Access Plates (Refer to Figure 201). (1) Position access plates in fuselage and tighten using screws. (2) Install vinyl floor covering. (3) Install crew and commuter/utility seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices and Chapter 25, Passenger Seats - Maintenance Practices.
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Floorboard Access Plates and Seat Rails Figure 201 (Sheet 1)
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Floorboard Access Plates and Seat Rails Figure 201 (Sheet 2)
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6.
Access Plate Removal/Installation (Cargo Airplanes and Optional on 208B Passenger) A.
Remove Access Plates (Refer to Figure 201). (1) Remove crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (2) Remove screws securing plywood floorboard coverings in cargo area, and remove coverings. NOTE: (3) (4)
B.
Note dimensions of tape over access openings under plywood floorboard coverings and install tape of the same size on reinstallation, with a minimum of 0.60 inch overlap.
Remove tape from access openings. Remove screws securing access plates in flight crew area, and remove plates.
Install Access Plates (Refer to Figure 201). (1) Position access plates in fuselage and tighten using screws. (2) Tape access openings in cargo compartment with Scotch Speed Tape, using a minimum of 0.60 inch overlap. (3) Position plywood floorboard coverings and secure with screws. (4) Install crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL PEDESTAL - DESCRIPTION AND OPERATION 1.
General A.
2.
This section provides a description of the cockpit mounted pedestal.
Description A.
The pedestal is located in the flight compartment between crew seats. The pedestal contains flight controls, power levers, flap control, propeller feather, fuel lever, emergency power lever and flood lights which illuminate the lower center portion of the instrument panel. The pedestal incorporates removable access plates which allow access to various controls. (1) For removal and installation of flight controls, refer to Chapter 27, Flight Controls - General. (2) For removal and installation of fuel shutoff, refer to Chapter 28, Fuel - General. (3) For removal and installation of engine controls, refer to Chapter 76, Engine Controls - General.
B.
Refer to Figure 1 for an illustration of the pedestal.
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MODEL 208 MAINTENANCE MANUAL
Pedestal Figure 1 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PLATES/SKIN - DESCRIPTION AND OPERATION 1.
General A.
The fuselage exterior covering consists of skins of aluminum alloy. The skins are attached to bulkheads, stringers, and doublers with permanent fasteners. A removable plate located in right aft exterior area of fuselage is provided for servicing oxygen system. An additional plate is located on the bottom side of the empennage for access to the tailcone. For exact locations and callouts of all plates used on the airplane, refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
B.
The airplane receives an application of LPS-3 Metal Protector in the bilge area beneath the cargo floorboards. This application provides corrosion protection to the plates and skins in the bilge area. For additional information about corrosion and corrosion protection, refer to Chapter 51, Corrosion and Corrosion Control.
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MODEL 208 MAINTENANCE MANUAL SEAT RAILS - MAINTENANCE PRACTICES 1.
General A.
2.
Seat rails are located in crew and passenger compartment. They provide support for crew and passenger seats. The passenger seat rails are installed as an integral part of airplane structure and are attached to floor boards with rivets. The flight compartment seat rails are attached with both screws and rivets.
Flight Compartment Seat Rail Removal/Installation A.
Remove Flight Compartment Seat Rails (Refer to Figure 201).
WARNING: When it becomes necessary to replace seat rail rivets, great care should be taken in their removal so that rivet holes will retain their original size and not require larger size rivets. (1) (2) (3) (4) B.
3.
Remove flight compartment crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices. Remove vinyl floor covering. Remove screws (1) securing seat rails to supports. Remove rivets (2) securing seat rails to supports and remove seat rails.
Install Flight Compartment Seat Rails (Refer to Figure 201). (1) Position seat rails and install screws (1) securing seat rails to supports. (2) Install rivets (2) securing seat rails to supports. (3) Install vinyl floor covering. (4) Install flight compartment crew seats. Refer to Chapter 25, Flight Compartment - Maintenance Practices.
Passenger Seat Rail Removal/Installation A.
Remove Passenger Seat Rail (Refer to Figure 201).
WARNING: When it becomes necessary to replace seat rail rivets, great care should be taken in their removal so that rivet holes will retain their original size and not require larger size rivets. (1) (2) (3) B.
Remove passenger compartment seats. Refer to Chapter 25, Passenger Seats/Seat Belts Maintenance Practices. Remove vinyl floor covering. Remove rivets and floorboards as required to gain access to seat rails, and remove seat rails.
Install Passenger Seat Rail (Refer to Figure 201). (1) Position floorboards and install rivets securing seat rails to floorboards. (2) Install vinyl floor covering. (3) Install passenger compartment seats. Refer to Chapter 25, Passenger Seats/Seat Belts Maintenance Practices.
53-25-00 © Cessna Aircraft Company
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Seat Rails Installation Figure 201 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SEAT RAILS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the seat rails in a serviceable condition.
Task 53-25-00-220 2.
Seat Rails and Attachment Structure Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to do a detailed inspection of the seat rails.
B.
Special Tools (1) None
C.
Access (1) Remove the pilot and the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (2) Remove floorboard access panels 231AL, 231BL, 231CL, 231DL, 232BC, 212FR, 216BC, 232AR, 232BR, 232AC, 232CR, and 232DR. Refer to Chapter 6, Access Plates and Panels Description and Operation. (3) Remove the passenger seats and floor covering. Refer to Chapter 25, Passenger Seats Maintenance Practices. (4) Remove necessary floor panels adjacent to passenger seat rails to examine structure beneath the passenger seat rails for cracks. Refer to Chapter 6, Access Plates and Panels - Description and Operation.
D.
Examine the Pilot and Copilot's Seat Rails and Attachment Structure. (1) Do a visual inspection of the pilot and copilot's seat rails for cracks or holes that are stretched more than 0.21 inch (5.33 mm). (a) If any of the seat rails have cracks or holes that are stretched more than 0.21 inch (5.33 mm), replace the seat rails. Refer to Service Kit SK208-138 and Seat Rails - Maintenance Practices. (b) If there are no cracks or holes that are stretched more than 0.21 inch (5.33 mm), continue with the inspection. (2) Do a visual inspection of the structure beneath the pilot and copilot's seat rails for cracks. (a) If there are no cracks in the structure beneath the pilot or copilot's seat rails, restore access. (b) If the structure beneath the pilot and copilot's seats has cracks, repair the structure. Refer to Service Kit SK208-138 and the Model 208 Structural Repair Manual.
E.
Examine the Passenger Seat Rails and Attachment Structure. (1) Do a visual inspection of the passenger seat rails for cracks. (a) If any of the seat rails have cracks, replace the seat rails. Refer to Seat Rails - Maintenance Practices. (b) If there are no cracks in the passenger seat rails, continue with the inspection. (2) Do a visual inspection of the structure under the passenger seat rails for cracks. (a) If the structure under the passenger seats has cracks, repair the structure. Refer to the Model 208 Structural Repair Manual. (b) If there are no cracks in the structure under the passenger seat rail restore access.
F.
Restore Access (1) Install access panels 231AL, 231BL, 231CL, 231DL, 232BC, 212FR, 216BC, 232AR, 232BR, 232AC, 232CR, and 232DR. Refer to Chapter 6, Access Plates and Panels - Description and Operation. (2) Install the pilot and the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (3) Install floor panels adjacent to passenger seat rails that were removed. Refer to Chapter 6, Access Plates and Panels - Description and Operation.
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MODEL 208 MAINTENANCE MANUAL (4)
Install the passenger seats and floor covering. Maintenance Practices.
Refer to Chapter 25, Passenger Seats -
End of task Task 53-25-00-221 3.
Bulkheads and Stiffeners Below the Seat Rail Attachments at FS 143.00 and FS 158.00 Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to do a detailed inspection of the bulkheads and stiffeners below the seat rail attachments at FS 143.00 and FS 158.00.
B.
Special Tools (1) None
C.
Access (1) Remove the pilot and the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. (2) Remove floorboard access panels 231CL, 231DL, 232BC, 251PL, 251AL, 252PR, 232CR, and 232DR. Refer to Chapter 6, Access Plates and Panels - Description and Operation.
D.
Examine the Bulkheads and Stiffeners Below the Seat Rail Attachments. (1) Examine the bulkheads and stiffeners for cracks at FS 143.00 and FS 158.00. (a) Do a visual inspection of the bulkheads and stiffeners for cracks. (b) Do a visual inspection of the bulkhead structure near the lightening holes for cracks. (c) Do a visual inspection of the bulkhead structure near the intersections with the longitudinal bulkheads for cracks. (2) Examine the longitudinal bulkheads and stiffeners for cracks at FS 143.00 and FS 158.00. (a) Do a visual inspection of the longitudinal bulkheads and stiffeners for cracks. (b) Do a visual inspection of the bulkhead structure near the lightening holes for cracks. (c) Do a visual inspection of the bulkhead structure near the intersections with the longitudinal bulkheads for cracks. (3) If there are cracks in the bulkheads, replace the bulkheads. Refer to the Model 208 Structural Repair Manual. (4) If there are no cracks, restore access.
E.
Restore Access (1) Install floorboard access panels 231CL, 231DL, 232BC, 251PL, 251AL, 252PR, 232CR, and 232DR. Refer to Chapter 6, Access Plates and Panels - Description and Operation. (2) Install the pilot and the copilot's seat. Refer to Chapter 25, Flight Compartment - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL TAIL STINGER - MAINTENANCE PRACTICES 1.
General A.
2.
Aerodynamic smoothness for empennage is provided by stinger. This section provides removal and installation instructions for the stinger.
Stinger Removal/Installation A.
Remove Stinger (Refer to Figure 201). (1) Remove screws (1) securing stinger to fuselage. (2) Disconnect electrical leads and remove stinger.
B.
Install Stinger (Refer to Figure 201). (1) Position stinger and connect electrical leads. (2) Install screws (1) securing stinger to fuselage.
53-50-00 © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL
Stinger Installation Figure 201 (Sheet 1)
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55 CHAPTER
STABILIZERS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
55-00-00
Page 1
Aug 1/1995
55-10-00
Pages 401-402
Mar 3/1997
55-10-00
Pages 601-603
Jun 1/2011
55-10-01
Pages 401-402
Sep 4/2001
55-20-00
Pages 401-405
Aug 1/1995
55-30-00
Pages 401-402
Aug 1/1995
55-30-00
Pages 601-602
Jun 1/2011
55-40-00
Pages 401-404
Aug 1/1995
55-Title 55-List of Effective Pages 55-Record of Temporary Revisions 55-Table of Contents 55-List of Tasks
55 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS STABILIZERS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-00-00 55-00-00 55-00-00 55-00-00
Page 1 Page 1 Page 1 Page 1
HORIZONTAL STABILIZER - REMOVAL/INSTALLATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Horizontal Stabilizer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-10-00 Page 401 55-10-00 Page 401 55-10-00 Page 401
HORIZONTAL STABILIZER - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Horizontal Stabilizer Forward and Aft Attach Points Special Detailed Inspection . . . Horizontal Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Horizontal Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-10-00 Page 601 55-10-00 Page 601 55-10-00 Page 601
STABILIZER ABRASION BOOTS - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation and Application of Bonding Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stabilizer Abrasion Boots Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-10-01 Page 401 55-10-01 Page 401 55-10-01 Page 401 55-10-01 Page 401
ELEVATOR - REMOVAL/INSTALLATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Trim Tab Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-20-00 Page 401 55-20-00 Page 401 55-20-00 Page 401 55-20-00 Page 401
VERTICAL STABILIZER - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vertical Stabilizer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-30-00 Page 401 55-30-00 Page 401 55-30-00 Page 401
VERTICAL STABILIZER - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vertical Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vertical Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-30-00 Page 601 55-30-00 Page 601
RUDDER - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55-40-00 Page 401 55-40-00 Page 401 55-40-00 Page 401
55-10-00 Page 601 55-10-00 Page 602
55-30-00 Page 601 55-30-00 Page 601
55 - CONTENTS © Cessna Aircraft Company
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LIST OF TASKS 55-10-00-250
Horizontal Stabilizer Forward and Aft Attach Points Special Detailed Inspection
55-10-00 Page 601
55-10-00-251
Horizontal Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance)
55-10-00 Page 601
55-10-00-252
Horizontal Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance)
55-10-00 Page 602
55-30-00-250
Vertical Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance)
55-30-00 Page 601
55-30-00-251
Vertical Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance)
55-30-00 Page 601
55 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL STABILIZERS - GENERAL 1.
Scope A.
2.
This chapter describes the horizontal and vertical stabilizers.
Tools and Equipment NOTE:
A suitable substitute may be used for the listed items:
NAME
NUMBER
MANUFACTURER
USE
Methyl Ethyl Ketone
Commercially available
Clean metal surfaces.
Toluene
Commercially available
Clean metal surfaces.
Strypeese
Commercially available
Remove topcoat finishes.
Easy-Strip
19A
Pennwalt 151 Old New Brunswick Rd. Piscataway, NJ 88854-3788
Remove topcoat finishes.
Cement
EC-1300L
3M Co. 3M Center St. Paul, MN 55144-1000
Adhesive.
3.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. A brief description of the sections follows: (1) The section on the horizontal stabilizer provides a description of components and removal/ installation procedures. (2) The section on the stabilizer abrasion boots provides a description of maintenance practices. (3) The section on the elevator provides a description of components and removal/installation procedures. (4) The section on the vertical stabilizer provides a description of components and removal/ installation procedures. (5) The section on the rudder provides a description of components and removal/installation procedures.
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MODEL 208 MAINTENANCE MANUAL HORIZONTAL STABILIZER - REMOVAL/INSTALLATION 1.
General A.
2.
The horizontal stabilizers are of all metal, fully cantilever, semimonocoque design consisting of spars, stringers, ribs, and skins and attach fittings. Skins are riveted to supporting structure with conventional MS20470AD rivets. Spar caps and attach fittings are of extruded and forged 7075 aluminum alloy material, respectively, while the remainder of the stabilizer structure is of formed 2024 sheet Alclad material, heat-treated after forming. The entire horizontal stabilizer structure is riveted together using standard MS20470AD universal head rivets. An aerodynamically balanced elevator and rudder are hinged to the trailing edge of the horizontal and vertical stabilizer, respectively. Left and right elevator trim tabs are attached to the trailing edges of the elevator via piano-type hinges.
Horizontal Stabilizer Removal/Installation A.
Remove Horizontal Stabilizer (Refer to Figure 401). (1) Remove stinger by removing attach screws. (2) Disconnect tail navigation light wire at quick-disconnect. (3) Remove rudder in accordance with Rudder - Removal/Installation. (4) Remove vertical stabilizer in accordance with Vertical Stabilizer - Removal/Installation. (5) Remove elevator pushrod at pushrod arm assembly. Refer to Elevator - Removal/Installation. (6) Remove four inspection covers (9) on top of horizontal stabilizer near center. (7) Remove access panel from aft lower surface of tailcone. NOTE:
Removal of access panels on the lower left and right sides of the horizontal stabilizer at Stabilizer Stations 80.60 provides access to the left and right elevator trim tab actuators.
(8)
Cut safety wire from elevator trim tab cable turnbuckles inside tailcone, tag cables for identification, and disconnect turnbuckles. (9) Remove attach bolts (4) and (6). (10) Remove horizontal stabilizer (1) from airplane. Retain shims (7).
B.
Install Horizontal Stabilizer (Refer to Figure 401). (1) With shims (7) positioned in same relative position as removed, line up attach holes in fittings (10) and supports (11) with attach holes in tail cone structure. (2) Install bolts (4) and (6). (3) Torque nuts (5) to 480 to 690 inch-pounds. (4) Torque nuts (8) to exactly 70 inch-pounds.
WARNING: Upon completion of all systems installations and/or rigging, ensure that all bolts, nuts, fittings, connections, etc., are tightened and secured properly. Check installations for freedom of movement. (5)
Connect elevator trim tab cable turnbuckles inside tail cone and rig trim tab in accordance with Chapter 27. Safety turnbuckles. (6) Connect elevator pushrod at elevator horn and safety. (7) Install four inspection covers (9). (8) Install vertical stabilizer in accordance with Vertical Stabilizer - Removal/Installation. (9) Install rudder in accordance with Rudder - Removal/Installation. (10) Connect tail navigation light wire at quick-disconnect. (11) Install stinger.
55-10-00 © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL
Horizontal Stabilizer Installation Figure 401 (Sheet 1)
55-10-00 © Cessna Aircraft Company
Page 402 Mar 3/1997
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL HORIZONTAL STABILIZER - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the horizontal stabilizer in a serviceable condition.
Task 55-10-00-250 2.
Horizontal Stabilizer Forward and Aft Attach Points Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the horizontal stabilizer forward and aft attach points in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the horizontal stabilizer. Installation.
D.
Refer to Chapter 55, Horizontal Stabilizer - Removal/
Do a Special Detailed Inspection of the Horizontal Stabilizer Forward and Aft Attach Points. (1) Do a visual inspection for cracks in the forward side of the horizontal stabilizer forward spar attach fitting. (2) Do a nondestructive testing (NDT) inspection for cracks in the horizontal stabilizer forward spar attach fitting. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Forward and Aft Attach Points - Description And Operation. (3) Do a visual inspection for cracks in the forward side of the horizontal stabilizer aft spar attach fitting. (4) Do a NDT inspection for cracks in the horizontal stabilizer forward spar attach fitting. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Forward and Aft Attach Points - Description And Operation. (5) If no cracks are found, restore access. (6) If cracks are found, replace the horizontal stabilizer attach fitting. Refer to Chapter 55, Horizontal Stabilizer - Removal/Installation.
E.
Restore Access (1) Install the horizontal stabilizer. Refer to Chapter 55, Horizontal Stabilizer - Removal/Installation. End of task Task 55-10-00-251 3.
Horizontal Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the horizontal stabilizer spars in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the horizontal stabilizer. Installation.
D.
Refer to Chapter 55, Horizontal Stabilizer - Removal/
Do a Special Detailed Inspection of the Horizontal Stabilizer Forward Spar. (1) Do a visual inspection for cracks in the horizontal stabilizer forward spar upper cap.
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4)
(5) E.
Do a nondestructive testing (NDT) inspection for cracks in the horizontal stabilizer forward spar upper cap between SS 0.00 and SS 9.90. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. Do a visual inspection for cracks in the horizontal stabilizer forward spar lower cap. Do a NDT inspection for cracks in the horizontal stabilizer forward spar lower cap between SS 0.00 and SS 9.90. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. (a) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures. If no cracks are found, continue with the inspection.
Do a Special Detailed Inspection of the Horizontal Stabilizer Aft Spar. (1) Do a visual inspection for cracks in the horizontal stabilizer aft spar upper cap. (2) Do a NDT inspection for cracks in the horizontal stabilizer aft spar upper cap between SS 0.00 and SS 10.60. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. (3) Do a visual inspection for cracks in the horizontal stabilizer aft spar lower cap. (4) Do a NDT inspection for cracks in the horizontal stabilizer aft spar lower cap between SS 0.00 and SS 10.60. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. (a) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures. (5) If no cracks are found, restore access.
F.
Restore Access (1) Install the horizontal stabilizer. Refer to Chapter 55, Horizontal Stabilizer - Removal/Installation. End of task Task 55-10-00-252 4.
Horizontal Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the horizontal stabilizer spars in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the horizontal stabilizer. Installation.
Refer to Chapter 55, Horizontal Stabilizer - Removal/
D.
Do a Special Detailed Inspection of the Horizontal Stabilizer Forward Spar. (1) Do a visual inspection for cracks in the horizontal stabilizer forward spar upper cap. (2) Do a nondestructive testing (NDT) inspection for cracks in the horizontal stabilizer forward spar upper cap between SS 0.00 and SS 9.90. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. (3) Do a visual inspection for cracks in the horizontal stabilizer forward spar lower cap. (4) Do a NDT inspection for cracks in the horizontal stabilizer forward spar lower cap between SS 0.00 and SS 9.90. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. (a) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures. (5) If no cracks are found, continue with the inspection.
E.
Do a Special Detailed Inspection of the Horizontal Stabilizer Aft Spar. (1) Do a visual inspection for cracks in the horizontal stabilizer aft spar upper cap.
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4)
(5)
Do a NDT inspection for cracks in the horizontal stabilizer aft spar upper cap between SS 0.00 and SS 10.60. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. Do a visual inspection for cracks in the horizontal stabilizer aft spar lower cap. Do a NDT inspection for cracks in the horizontal stabilizer aft spar lower cap between SS 0.00 and SS 10.60. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Horizontal Stabilizer Spars - Description And Operation. (a) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures. If no cracks are found, restore access.
F.
Restore Access (1) Install the horizontal stabilizer. Refer to Chapter 55, Horizontal Stabilizer - Removal/Installation. End of task
55-10-00 © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL STABILIZER ABRASION BOOTS - REMOVAL/INSTALLATION 1.
General A.
2.
Airplanes may be equipped with two extruded rubber abrasion boots, one on the leading edge of each horizontal stabilizer. These boots are installed to protect the stabilizer leading edge from damage by rocks thrown back by propeller.
Preparation and Application of Bonding Material A.
Procedure. (1) Adhesive EC1300L must be thoroughly stirred prior to application. A uniform coat of adhesive shall be brushed onto the masked off metal surface and onto the faying surface of the deice boot. When brushing on adhesives, use good quality, clean nylon brushes. Avoid hot air drafts from heaters or fans which can cause dragging and produce a very rough surface. The adhesive shall be allowed to dry thoroughly (at least one hour at 77°F and 50 percent relative humidity lower temperatures and/or higher humidities require longer drying times to completely dry) and should not have any tack. A second uniform coat of adhesive shall be brushed onto each of the faying surfaces and allowed to dry like the first coat. NOTE: (2)
3.
Adhesive EC1300L may be thinned by adding 1.5 fluid ounces of Toluene to 16 ounces (1 fluid pint) of adhesive to achieve a more applicable consistency.
The dry adhesive shall be covered and kept clean until it is reactivated. The adhesive shall be reactivated within 48 hours by wiping lightly with clean cheesecloth slightly moistened with Toluene. Only a small area, approximately 3 inches by 18 inches or less, shall be reactivated at one time. Do not allow the adhesives to become too dry before placing the deice boot in contact with the metal. Excessive rubbing or excessive solvent usage should be avoided so that adhesive will not be removed.
Stabilizer Abrasion Boots Removal/Installation A.
Remove Stabilizer Abrasion Boots (Refer to Horizontal Stabilizer - Removal/Installation, Figure 401). (1) Toluene shall be used to soften cement line. A minimum amount of this solvent should be applied to cement line as tension is applied to pull back boot. The removal should be slow enough to allow solvent to undercut boot so that parts will not be damaged. Excessive quantities of solvent on airplane must be avoided.
B.
Install Stabilizer Abrasion Boots (Refer to Horizontal Stabilizer - Removal/Installation, Figure 401). (1) Mask off boot area on leading edge of stabilizer with one inch masking tape, allowing a half-inch margin from boot edge.
CAUTION: Ensure that corrosion protection is not removed from metal surfaces during cleaning process. (2)
Clean metal surfaces of stabilizer where boot is to be installed and clean inside surface of abrasion boot thoroughly with Methyl n-Propyl Ketone or Toluene. Wipe surface with clean cloth saturated in solvent. Cloth should be folded each time surface is wiped to present a clean area and avoid redepositing of grease. Wipe surface immediately with clean dry cloth. Do not allow solvent to dry on surface. NOTE:
(3)
Boots may be applied over properly cured epoxy primer. Boots shall not be applied over topcoat finishes. Topcoat finishes shall be removed with Strypeese or EasyStrip.
Apply EC1300 adhesive. Refer to Preparation and Application of Bonded Materials.
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MODEL 208 MAINTENANCE MANUAL (4)
Place a straight line along leading edge line of stabilizer and a corresponding line on inside of boot. Position center line of boot with leading edge line of stabilizer and, using a clean, lint free cloth heavily moistened with toluene, reactivate surface of cement on stabilizer and boot, beginning at inboard end. NOTE:
Avoid excessive rubbing of cement, which would remove it from surface. Have enough help to hold boot steady during installation and avoid handling cemented surfaces.
Roll boot firmly against leading edge, beginning at inboard end, being careful not to trap any air between it and stabilizer. (6) If the boot should attach "off-course" (reference centerline on leading edge not coinciding with reference centerline on boot), apply Toluene with a small brush or squirt bottle to soften the bond line. (a) Apply only a small amount of Toluene while applying sufficient tension to peel back the softened adhesive. (b) To prevent damage to the boot, avoid twisting, sharply bending, or jerking the boot loose from the bonded area. Allow solvent wetted area to dry thoroughly before continuing with applications. Reapply EC1300L adhesive as needed. (7) Roll entire surface of boot applying pressure. Should an air pocket be encountered, carefully insert a hypodermic needle and allow air to escape. (8) Apply a coat of GACO N700A Neoprene Coating or equivalent along trailing edge of boot to surface of skin, forming a neat, straight fillet. (9) Remove masking tape and clean surfaces with toluene. (10) Mask edge of boot for painting stabilizer.
(5)
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MODEL 208 MAINTENANCE MANUAL ELEVATOR - REMOVAL/INSTALLATION 1.
General A.
2.
A mechanically actuated elevator is attached to the rear spar of the horizontal stabilizer via hinges and hinge bolts. The elevator provides pitch control for the airplane. Structurally, the elevator consists of two sections bolted together near the airplane centerline via torque tubes. The individual sections are made up of spars, ribs, and skins riveted together. A trim tab is attached to each elevator section trailing edge by piano-type hinges. A balance weight is located in the outboard end of each elevator section, forward of the hinge line, and is accessible through a removable cover.
Elevator Removal/Installation A.
Remove Elevator (Refer to Figure 401). (1) Remove stinger by removing attaching screws. (2) Disconnect tail navigation light wire at quick-disconnect. (3) Remove three bolts (20) attaching left torque tube (11), right torque tube (6), and adaptors (15) to pushrod arm assembly (25). (4) Disconnect elevator trim tab pushrods at trim tab horns (12), both sides. Retain pushrod clevis bushings, two each per side. NOTE: (5) (6) (7) (8)
B.
Roll elevator trim tab control wheel in cockpit to full nose-up position in order to withdraw elevator trim tab pushrods as far as possible into elevator. Remove cotter pins (21) from nuts (22) at each outboard and inboard hinge location of both elevator section and remove nuts (22). Support each elevator section and remove inboard hinge bolt (27) and outboard hinge bolt (29) from both sides. Withdraw each elevator section from horizontal stabilizer, being careful not to damage elevator trim tab pushrods and torque tube flanges.
Install Elevator (Refer to Figure 401). (1) Support each elevator section in turn and carefully guide torque tubes (6) and (12) into mating position with adaptors (15), while guiding elevator trim tab pushrods into elevator recesses forward of horns (12). NOTE:
(2) (3) (4) (5) (6) 3.
To prevent loss of trim tab control system rigging, safety-wire pushrods together on each side.
Instructions for fabricating an elevator attach bolt installation tool are shown on Figure 402. The tool may be used to securely hold elevator attach bolt when inserting it into elevator hinge. To use tool, insert head of bolt between clip and handle of tool.
Align holes in hinge bearings (17) with holes in hinge brackets (28) and install outboard hinge bolts (29) and then inboard hinge bolts(27). Install three bolts (20) through aligned holes of torque tube flanges (11) and (6), adaptors (15), and arm assembly (25). Install nuts (22) and washers on inboard and outboard hinge bolts (27) and (29), respectively. Safety nuts (22) with cotter pins (21). Install elevator trim tab pushrod bushings through aligned holes in pushrod clevises and pushrod horns (12). Install pushrod bolts, washers, and nuts, and safety four nuts with cotter pins. Connect tail navigation light wire at quick-disconnect and install stinger using attach screws.
Trim Tab Removal/Installation A.
Remove Trim Tab (Refer to Figure 401) (1) Disconnect elevator trim tab pushrods at trim tab horns (12). Retain pushrod clevis bushings, two each per side. (2) Remove retaining screw (3) at elevator tip securing hinge pin (4) and remove hinge pin. (3) Remove trim tab (5).
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Elevator Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Elevator Installation Figure 401 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Elevator Bolt Installation Tool Fabrication Figure 402 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL B.
Install Trim Tab (Refer to Figure 401) (1) Align piano-type hinge halves of trim tab (5) with those of elevator trailing edge and install hinge pin (4). (2) Ensure hinge pin (4) has bottomed out in piano-type hinge and install retaining screw (3). (3) Align holes in trim tab pushrod clevises with holes in pushrod horn (12) and install push-rod clevis bushings, pushrod bolts, washers, and nuts. Safety nuts with cotter pins.
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MODEL 208 MAINTENANCE MANUAL VERTICAL STABILIZER - REMOVAL/INSTALLATION 1.
General A.
2.
The vertical stabilizer is of conventional sweptback design of semimonocoque construction utilizing spars, semi-spars, ribs, and skins. The assembly is riveted together with conventional universal head rivets. A dorsal fin is attached with screws to the forward section of the vertical fin and to the top of the fuselage. An aerodynamically balanced rudder is hinged, using bolts, to the trailing edge of the vertical stabilizer at three hinge points.
Vertical Stabilizer Removal/Installation A.
Remove Vertical Stabilizer (Refer to Figure 401). NOTE: (1) (2) (3) (4) (5) (6) (7)
B.
An access cover located on the bottom aft side of the tailcone provides access to the vertical stabilizer mounting points.
Remove rudder in accordance with Rudder - Removal/Installation. Remove dorsal fin (3) by removing screws (4). Disconnect balanced loop VOR NAV antenna lead (if installed) in tailcone. Disconnect tail surface de-ice system plumbing (if installed) in tailcone. Working through tailcone access opening, remove both front spar attach bolts (5) and (6). While supporting vertical stabilizer, remove both rear spar attach bolts (9) and (10). Remove vertical stabilizer (2).
Install Vertical Stabilizer (Refer to Figure 401). (1) Position vertical stabilizer (2) on airplane, lining up front and rear spar attachment holes with corresponding holes in tailcone bulkheads. (2) Support vertical stabilizer (2) and install spar attachment bolts (5), (6), (9), and (10). (3) Connect tail surface de-ice system plumbing (if removed) in tailcone. (4) Connect balanced loop VOR NAV antenna lead (if removed) in tailcone. (5) Install dorsal fin (3) with screws (4). (6) Install rudder in accordance with Rudder - Removal/Installation. (7) Check rudder system for proper travels and cable tension. Rerig as required. Refer to Chapter 27.
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Vertical Stabilizer and Dorsal Fin Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL VERTICAL STABILIZER - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the vertical stabilizer in a serviceable condition.
Task 55-30-00-250 2.
Vertical Stabilizer Spars Special Detailed Inspection (Typical Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the vertical stabilizer spars in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation.
D.
Do a Special Detailed Inspection of the Vertical Stabilizer Spars. (1) Do a visual inspection for cracks in the vertical stabilizer forward spar cap. (2) Do a nondestructive testing (NDT) inspection for cracks at the vertical stabilizer forward spar attach holes at WL 126.03. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (3) Do a NDT inspection for cracks at the vertical stabilizer forward spar attach holes at WL 134.28. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (4) Do a NDT inspection for cracks in the vertical stabilizer forward spar from WL 134.38 to WL 138.00. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (5) Do a visual inspection for cracks in the vertical stabilizer aft spar cap. (6) Do a NDT inspection for cracks at the vertical stabilizer aft spar attach holes at WL 116.40. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (7) Do a NDT inspection for cracks at the vertical stabilizer aft spar attach holes at WL 126.12. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (8) Do a NDT inspection for cracks in the vertical stabilizer aft spar from WL 126.12 to WL 129.00. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (9) If no cracks are found, restore access. (10) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures.
E.
Restore Access (1) Install the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation. End of task Task 55-30-00-251 3.
Vertical Stabilizer Spars Special Detailed Inspection (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the vertical stabilizer spars in a serviceable condition.
B.
Special Tools (1) None
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MODEL 208 MAINTENANCE MANUAL C.
Access (1) Remove the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation.
D.
Do a Special Detailed Inspection of the Vertical Stabilizer Spars. (1) Do a visual inspection for cracks in the vertical stabilizer forward spar cap. (2) Do a nondestructive testing (NDT) inspection for cracks at the vertical stabilizer forward spar attach holes at WL 126.03. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (3) Do a NDT inspection for cracks at the vertical stabilizer forward spar attach holes at WL 134.28. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (4) Do a NDT inspection for cracks in the vertical stabilizer forward spar from WL 134.38 to WL 138.00. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (5) Do a visual inspection for cracks in the vertical stabilizer aft spar cap. (6) Do a NDT inspection for cracks at the vertical stabilizer aft spar attach holes at WL 116.40. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (7) Do a NDT inspection for cracks at the vertical stabilizer aft spar attach holes at WL 126.12. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (8) Do a NDT inspection for cracks in the vertical stabilizer aft spar from WL 126.12 to WL 129.00. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current Inspection, Vertical Stabilizer Spars - Description And Operation. (9) If no cracks are found, restore access. (10) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures.
E.
Restore Access (1) Install the vertical stabilizer. Refer to Chapter 55, Vertical Stabilizer - Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL RUDDER - REMOVAL/INSTALLATION 1.
General A.
2.
The rudder consists of a spar, ribs, leading and trailing edge skins, hinge brackets, and a torque tube. These assemblies are riveted together using conventional universal head rivets. A balance weight is attached to upper forward leading edge.
Rudder Removal/Installation A.
Remove Rudder (Refer to Figure 401 ). (1) Remove screws attaching stinger to tailcone and remove stinger. (2) Disconnect tail navigation light wire at quick-disconnect. (3) Remove cotter pins and clevis bolt nuts from rudder cable attach bolts at torque tube (12). (4) Remove cotter pins from rudder hinge bolt nuts (7), (14), and (15), and remove nuts. (5) With rudder supported, remove hinge bolts (6), (11), and (17). (6) Remove rudder (1).
B.
Install Rudder (Refer to Figure 401). (1) Position rudder so that hinge bolt holes in rudder align with hinge bearings (5), (9), and (16). NOTE:
(2) (3) (4) (5) (6)
Instructions for fabricating an elevator attach bolt installation tool are shown on Figure 402. The tool may be used to securely hold rudder attach bolt when inserting it into rudder hinge. To use tool, insert head of bolt between clip and handle of tool.
Install hinge bolts (6), (11), and (17) and nuts (7), (14), and (15), respectively. Safety with cotter pins. Install clevis bolts attaching rudder cables to torque tube (12). Safety nuts with cotter pins (8). Connect tail navigation light wire at quick-disconnect. Install stinger to tailcone with attach screws. Check rudder travel in accordance with Chapter 27. Rerig as required.
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Rudder Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Rudder Installation Figure 401 (Sheet 2)
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Rudder Bolt Installation Tool Fabrication Figure 402 (Sheet 1)
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56 CHAPTER
WINDOWS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
56-00-00
Pages 1-3
Jan 2/2006
56-00-01
Pages 1-3
Jan 2/2006
56-00-01
Pages 201-203
Apr 1/2010
56-00-01
Pages 601-608
Jun 1/2011
56-10-00
Pages 401-404
Mar 1/1999
56-20-00
Pages 401-403
Aug 1/1995
56-21-00
Pages 401-403
Aug 1/1995
56-30-00
Pages 401-405
Aug 1/1995
56-Title 56-List of Effective Pages 56-Record of Temporary Revisions 56-Table of Contents 56-List of Tasks
56 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS WINDOWS - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56-00-00 56-00-00 56-00-00 56-00-00
WINDSHIELDS AND WINDOWS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56-00-01 Page 1 56-00-01 Page 1 56-00-01 Page 1
WINDSHIELDS AND WINDOWS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning Instructions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield and Window Preventive Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield and Window Installation Techniques . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Window Plug Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Rain Repellent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56-00-01 Page 201 56-00-01 Page 201 56-00-01 Page 201 56-00-01 Page 201 56-00-01 Page 201 56-00-01 Page 202 56-00-01 Page 202 56-00-01 Page 203
WINDSHIELDS AND WINDOWS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield and Attachment Structure Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . Windshield Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56-00-01 Page 601 56-00-01 Page 601 56-00-01 Page 601 56-00-01 Page 601
FLIGHT COMPARTMENT WINDOWS - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshield Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56-10-00 Page 401 56-10-00 Page 401 56-10-00 Page 401 56-10-00 Page 401
CABIN WINDOWS - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cabin Window Removal/installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56-20-00 Page 401 56-20-00 Page 401 56-20-00 Page 401
WINDOW PLUGS - REMOVAL/INSTALLATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Window Plug Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spring Clip Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
56-21-00 Page 401 56-21-00 Page 401 56-21-00 Page 401 56-21-00 Page 401
DOOR WINDOWS - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Crew Door Window Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pilot’s Vent Window Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cargo and Passenger Door Window Removal/Installation. . . . . . . . . . . . . . . . . . . . . . .
56-30-00 Page 401 56-30-00 Page 401 56-30-00 Page 401 56-30-00 Page 401 56-30-00 Page 401 56-30-00 Page 403
56 - CONTENTS © Cessna Aircraft Company
Page 1 Page 1 Page 1 Page 3
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 56-00-01-220
Windshield and Attachment Structure Detailed Inspection
56-00-01 Page 601
56-00-01-720
Windshield Functional Check
56-00-01 Page 601
56 - LIST OF TASKS © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL WINDOWS - GENERAL 1.
Scope A.
2.
This chapter covers windows used in the flight compartment, cargo compartment and passenger compartment.
Tools, Equipment and Materials NOTE:
Equivalents are permitted for the items that follow:
NAME
NUMBER
MANUFACTURER
USE
Aliphatic Naphtha
Type II Federal Specification TT-N-95
Commercially Available
To remove deposits from windshields and windows.
Chogel-20
Echo Laboratories RD#1, Box 297 Titusville, PA 16354
To couple the prism to the window.
Great Reflections Paste Wax
E.I. DuPont DeNemours and Co. (Inc.) Wilmington, DE 19898
To wax acrylic windshields and windows.
Inspection Prism Kit
AWR P-17
Aircraft Windows Repairs Company 2207 Border Ave. Torrance, CA 90501-3612
To do the optical inspection of the windows. Note 2
Masking Paper
WPL-3
St. Regis Paper Co. 156 Oak St. Newton, MA 02164-1440
To protect surfaces from solvent attack.
Commercially Available
To clean windshields and windows.
Meguiars Mirror Bright Polish 210 N First Ave. Arcadia, CA 91006
To clean and polish acrylic windshields and windows.
Fabricate
To do the optical prism inspection of the windows.
Mild Soap or Detergent (hand dishwashing type without abrasives) Mirror Glaze
MGH-7
Optical Prism Optical Prism
6580000-1 NOTE: The 6580000-1 Optical Prism will not look exactly like the prism illustrated in this manual
Cessna Aircraft Company Cessna Parts Distribution 5800 E. Pawnee 1521 Wichita, KS 67218
To do the optical prism inspection of the windows.
Optical Prism
AWR P-17
Aircraft Window Repairs Company 2207 Border Ave. Torrance, CA 90501
Alternate prism to do the optical prism inspection of the windows.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Permatex
403D
Loctite Corp. North America Group 1001 Trout Brook Crossing Rocky Hill, CT 06067
To clean and polish acrylic windshields and windows.
Poly-Spotstick
5XN
St. Regis Paper Co.
To protect surfaces from solvent attack.
Propylene Glycol
PR130USP
Del Amo Chemical Co. 535 W 152 St. Gardena, CA 90248 Phone: (310) 532-9214
A couplant used between the prism and the window.
Protex 10VS
Mask Off Company
To protect surfaces from solvent attack.
Protex 40
Mask Off Company 345 W. Maple Avenue Monrovia, CA 91016-3331
To protect surfaces from solvent attack.
Scotch Black Tape
344
Commercially Available
To protect surfaces from solvent attack.
Sealant
FS-4291
H.B. Fuller Company 3530 Lexington Ave. North St. Paul, MN 55126-8076
To seal windshield and windows.
Sealant
890/890A
PRC-DeSoto International 5454 SanFernando Rd. Glendale, CA 91203
To seal windshield and windows.
Slip-Stream Wax
Classic Chemical 3131 Turtle Creek Suite 1010 Dallas, TX 75219-5415
To wax acrylic windshields and windows.
Soft cloth, such as cotton flannel or cotton terry cloth
Commercially Available
To apply and remove wax and polish.
Turtle wax (paste)
Turtle Wax, Inc. 5655W 73rd St Chicago, IL 60638-6211
To wax acrylic windshields and windows.
Ultragel II
Sonotech, Inc. 774 Marine Drive Bellingham, WA 98225 Email: www.sonotech-inc.com
To couple the prism to the window.
White Light Source (Flash or Penlight)
Commercially Available
To supply light to the area that you inspect.
Commercially Available
To protect surfaces from solvent attack.
White Spary Lab
MIL-C-6799 Type I, Class II
NOTE:
The prism will let you examine the windshield fastener holes without the removal of the shroud.
NOTE:
You can use equivalent couplants. However the operator or the inspector will make sure that the material will not cause damage to the window surface, painted surface, or airplane structure.
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3.
Definition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief de finition of the sections incorporated in this chapter is as follows: (1) The section on windshields and windows provides general information on the care and cleaning of all acrylic sheets used on the airplane. (2) The section on flight compartment windows provides information on the windshields and side windows in the flight compartment. (3) The section on cabin windows provides information on windows used in the cabin area. (4) The section on door windows provides information on windows used in various doors.
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MODEL 208 MAINTENANCE MANUAL WINDSHIELDS AND WINDOWS - DESCRIPTION AND OPERATION 1.
2.
General A.
This section covers windshields and windows used on various Model 208 airplanes. For an illustration of window locations, refer to Figure 1.
B.
If it is necessary to inspect the windshield, refer to Windshields and Windows - Inspection/Check.
Description A.
Flight compartment windows consist of left and right windshields, right forward crew door window and left forward crew door window including a hinged vent window. (1) Flight compartment windshields are comprised of two-piece, 0.312 inch thick, green-tinted acrylic.
B.
Cabin windows are made of 0.187 inch thick, green-tinted acrylic. Locations and numbers vary with different Model 208 configurations. (1) On the Model 208 only, windows in the cabin consist of four observation windows on the right side of the airplane and three observation windows on the left side of the airplane. (2) On the Model 208B Passenger only, windows in the cabin consist of six observation windows on the right side and five observation windows on the left side. (3) The airplane door windows in the cabin section consist of one window in the right-hand upper passenger door and two windows in the left-hand upper cargo door.
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Window Locations Figure 1 (Sheet 1)
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Window Locations Figure 1 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL WINDSHIELDS AND WINDOWS - MAINTENANCE PRACTICES 1.
General A.
2.
Tools, Equipment and Materials A.
3.
This section provides instructions and tips for cleaning and installing windshields and windows installed in the crew, passenger and/or cargo compartments of the airplane. Also included in this section are cleaning instructions for the window plugs.
For a list of required tools, equipment and materials, refer to Windows - General.
Cleaning Instructions
CAUTION: Windshields and windows (acrylic-faced) are easily damaged by improper handling and cleaning techniques. CAUTION: Do not use methanol, denatured alcohol, gasoline, benzene, xylene, methyl n-propyl ketone, acetone, carbon tetrachloride, lacquer thinners, commercial or household window cleaning sprays on windshields or windows. A.
Instructions For Cleaning. (1) Place airplane inside hangar or in shaded area and allow to cool from heat of suns direct rays. (2) Using clean (preferably running) water, flood the surface. Use bare hands with no jewelry to feel and dislodge any dirt or abrasive materials. (3) Using a mild soap or detergent (such as a dishwashing liquid) in water, wash the surface. Again use only the bare hand to provide rubbing force. (A clean cloth may be used to transfer the soap solution to the surface, but extreme care must be exercised to prevent scratching the surface.) (4) When contaminants on acrylic windshields and windows cannot be removed by a mild detergent, Type Il aliphatic naphtha, applied with a soft clean cloth, may be used as a cleaning solvent. Be sure to frequently refold cloth to avoid redepositing contaminants and/or scratching windshield with any abrasive particles. (5) Rinse surface thoroughly with clean fresh water and dry with a clean cloth. (6) Hard polishing wax should be applied to acrylic surfaces. (The wax has an index of refraction nearly the same as transparent acrylic and will tend to mask any shallow scratches on the windshield surface). (7) Acrylic surfaces may be polished using a polish meeting Federal Specification P-P-560 applied per the manufacturers instructions. NOTE:
4.
When applying and removing wax and polish, use a clean, soft cloth, such as cotton or cotton flannel.
Windshield and Window Preventive Maintenance NOTE: A.
Utilization of the following techniques will help minimize windshield and window crazing.
General Notes and Techniques For Acrylic Windshields. (1) Keep all surfaces of windshields and windows clean. (2) If desired, wax acrylic surfaces. (3) Carefully cover all surfaces during any painting, powerplant cleaning or other procedure that calls for use of any type of solvents or chemicals. Refer to Windows - General for approved covering materials. (4) Do not park or store airplane where it might be subjected to direct contact with or vapors from: methanol, denatured alcohol, gasoline, benzene, xylene, methyl n-propyl ketone, acetone, carbon tetrachloride, lacquer thinners, commercial or household window cleaning sprays, paint strippers, or other types of solvents.
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MODEL 208 MAINTENANCE MANUAL (5) (6) 5.
Do not leave sun visors up against windshield when not in use. The reflected heat from these items causes elevated temperatures on the windshield. If solar screens are installed on the inside of the airplane, make sure they are the silver appearing, reflective type. Do not use a power drill motor or other powered device to clean, polish, or wax surfaces.
Windshield and Window Installation Techniques A.
Installation Techniques. (1) Special drills must be used when drilling holes in acrylic. Standard drills will cause the hole to be oversized, distorted, or excessively chipped. (2) Whenever possible, a coolant such as a plastic drilling wax should be used to lubricate the drill bit. Cessna recommends "Reliance" drill wax or Johnson No. 140 Stick Wax. (3) Drilled holes should be smooth with a finish of 125 rms (root mean square). (4) The feed and speed of the drill is critical. Refer to Table 201 for thickness verses drill speed information.
Table 201. Material Thickness vs. Drill Speed. Thickness (in inches)
Drill Speed (RPM)
0.062 to 0.1875
1500 to 4500
0.250 to 0.375
1500 to 2000
0.4375
1000 to 1500
0.500
500 to 1000
0.750
500 to 800
1.00
500 (5)
(6)
B. 6.
In addition to feed and speed of the drill bit, the tip configuration is of special importance when drilling through acrylic windows and windshields. Tip configuration varies with hole depth, and the following information applies when drilling through acrylic: (a) Shallow Holes - When hole depth to hole diameter ratio is less than 1.5 to 1, the drill shall have an included tip angle of 55 degrees to 60 degrees and a lip clearance angle of 15 degrees to 20 degrees. (b) Medium Deep Holes - When hole depth to hole diameter ratio is from 1.5 to 1 up to 3 to 1, the drill shall have an included tip angle of 60 degrees to 140 degrees and a lip clearance angle of 15 degrees to 20 degrees. (c) Deep Holes - when hole depth of hole diameter ratio is greater than 3.0 to 1, the drill shall have an included tip angle of 140 degrees and a lip clearance of 12 degrees to 15 degrees. Parts which must have holes drilled shall be backed up with a drill fixture. Holes may be drilled through the part from one side. However, less chipping around holes will occur if holes are drilled by drilling the holes from both sides. This is accomplished by using a drill with an acrylic backup piece on the opposite side. Remove the drill from the hole and switch the backup plate and finish drilling from the opposite side.
If it is necessary to inspect the windshield, refer to Windshields and Windows - Inspection/Check.
Window Plug Cleaning A.
Window Plug Cleaning Instructions. (1) Remove window plugs. (2) Dust or vacuum both sides of plug with a clean, dry cotton cloth. (3) Clean both sides of plug with a clean, damp cotton cloth. (4) Oily spots and stains may be cleaned with household spot removers, used sparingly. Before using any solvent, read the instructions on the container and test it on an obscure place on the plug to be cleaned. Never saturate the felt trim around the plug with a violent solvent; it may damage the felt material. (5) Scrape off stuck materials with a dull knife, then clean the area.
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7.
Windshield Rain Repellent A.
A Cessna approved rain repellent and surface conditioner may be used to increase the natural cleaning of the windshield during rain. Apply in accordance with manufacturers instructions. NOTE:
REPCON is the only rain repellent conforming to Federal Specification MIL-W-6862 that is approved to use on Cessna Model 208 series airplanes.
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MODEL 208 MAINTENANCE MANUAL WINDSHIELDS AND WINDOWS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the windshield and the windows in a serviceable condition.
Task 56-00-01-220 2.
Windshield and Attachment Structure Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the windshield and attachment structure in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do a Detailed Inspection of the Windshield and Attachment Structure. (1) Use the optical prism procedure to examine for cracks in the windshield that start at the fastener hole(s) and extend to the adjacent fastener holes into the window area that the operator views out of or the edge of the window. Refer to Windshield Functional Check in this section. (2) Do a visual inspection of the windshield retainer for cracks. (3) Do a visual inspection for cracks in the skin around the windshield. (4) Do a visual inspection of the door post for cracks out of the fastener holes. (5) If cracks are found, replace or repair the damaged part(s). Refer to Flight Compartment Windows - Removal/Installation or the Model 208 Structural Repair Manual.
E.
Do a Detailed Inspection of the Cabin and Door Windows. (1) Do a visual inspection of the cabin and door window retainers and fasteners for cracks and corrosion. (2) If cracks are found, replace or repair the damaged part(s). Refer to Flight Compartment Windows - Removal/Installation or the Model 208 Structural Repair Manual. (3) If corrosion is found, repair or replace the damaged part(s). Refer to Chapter 51, Corrosion Prevention and Control Program - Description and Operation for more information.
F.
Restore Access (1) None End of task Task 56-00-01-720 3.
Windshield Functional Check NOTE:
The optical inspection procedure included in this section will find voids and cracks in the area of the fastener holes of the acrylic windows without the removal of the edge retainers or their related fasteners. The inspection will look for cracks that start at the fastener hole(s) and go to adjacent fastener holes, into the viewable area, or to the edge of the window.
NOTE:
An optical prism can be purchased or locally fabricated. Refer to Figure 601 for information on how to make the optical prism.
A.
General (1) This task gives the procedures to do a functional check of the windshield.
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MODEL 208 MAINTENANCE MANUAL B.
Special Tools (1) Optical Prism NOTE: (2) (3) (4)
The 70-degree or the 6580000-1 optical prisms are permitted to use for this functional check.
Aliphatic Naphtha Mild Soap or Detergent (Hand Dishwashing Type Without Abrasives) Couplant
C.
Access (1) Remove the windshield deicing ducts.
D.
Do the Functional Check of the Windshield (Procedure with a 70-Degree Prism) (Refer to Figure 602).
CAUTION: The use of cleaning materials other than aliphatic naphtha followed by a solution of liquid soap and water solution can cause crazing of the acrylic windows. (1) (2)
Use aliphatic naphtha followed by a solution of liquid soap and water solution to thoroughly clean dust and unwanted material from the window. Clean the acrylic window area a minimum of six to eight inches from the fastener holes. Apply the couplant to the 70-degree face of the prism and the inspection area of the window. NOTE:
(3) (4)
The inspections are done from the outside surface of the windows.
Put the prism on the window, refer to Figure 602. Use the light source to add light at an angle of 30 to 60 degrees from the vertical of the prism and examine the fastener holes. NOTE:
To get a clear view of both the top and the bottom surfaces of the fastener hole, move the prism toward, or away from the fastener.
(a) (b)
(5) (1)
E.
The image of an undamaged hole will show as a cylinder that is not transparent. The image of a fastener hole with a crack that extends from one surface of the material under inspection into the hole will show as a reflection. The reflection is not transparent and it extends from the fastener hole as in View A-A of Figure 602. (c) The image of a crack from one fastener hole to an adjacent fastener hole will show as an irregular surface that is not transparent. View B-B of Figure 602 shows a crack from hole to hole. After the inspection is completed, remove the couplant from the window with aliphatic naphtha followed by a weak soap and water solution. If you find a crack, contact Cessna Propeller Aircraft Product Support, P.O. Box 7706, Wichita, KS 67277 USA. Telephone 316-517-5800. Provide the following information: (a) Crack location (b) Crack length (c) Crack orientation
Do the Functional Check of the Windshield (Procedure with a 6580000-1 Prism) (refer to Figure 603.
CAUTION: The use of cleaning materials other than aliphatic naphtha followed by a solution of liquid soap and water solution can cause crazing of the acrylic windows. (1)
Use aliphatic naphtha followed by a solution of liquid soap and water solution to thoroughly clean dust and unwanted material from the window. Clean the acrylic window area a minimum of six to eight inches from the fastener holes.
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Fabrication of Optical Prism Figure 601 (Sheet 1)
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Optical Inspection Using 70-Degree Prism Figure 602 (Sheet 1)
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Optical Inspection Using 70-Degree Prism Figure 602 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (2)
Apply the couplant to the face of the prism and the inspection area of the window. NOTE:
(3) (4) (1)
The inspections are done from the outside surface of the windows.
Put the prism to the window as shown in Figure 603, and with the light source to add light at an angle of 30 to 60 degrees from the vertical of the prism, examine the fastener holes. After the inspection is completed, remove the couplant from the window with aliphatic naphtha followed by a weak soap and water solution. If you find a crack, contact Cessna Propeller Aircraft Product Support, P.O. Box 7706, Wichita, KS 67277 USA. Telephone 316-517-5800. Provide the following information: (a) Crack location (b) Crack length (c) Crack orientation
F.
Restore Access (1) Install the windshield deicing ducts. End of task
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Optical Inspection Using 6580000-1 Prism Figure 603 (Sheet 1)
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Optical Inspection Using 6580000-1 Prism Figure 603 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL FLIGHT COMPARTMENT WINDOWS - REMOVAL/INSTALLATION 1.
General A.
2.
Tools, Equipment and Materials A.
3.
This section covers removal and installation procedures for the left and right flight compartment windshields.
For a list of required tools, equipment and materials, refer to Windows - General.
Windshield Removal/Installation A.
Remove Windshield (Refer to Figure 401). (1) Remove OAT gage. Refer to Chapter 34, Outside Air Temperature Gage - Maintenance Practices. (2) Remove windshield anti-ice panel, if installed. Refer to Chapter 30, Windshield Anti-Ice Maintenance Practices.
CAUTION: Do not use any tool, abrasive or cleaner which might damage windshield. (3) (4) (5) (6)
Cover the left (9) and right (2) windshield halves with a protective covering. Remove cover strip (15) from windshield post (5). Remove and retain all screws (3), tubing spacers/rubber grommets (12), and nuts securing left (8), and right (1) outboard retainers and lower left (10) and right (11) retainers. Remove and retain screws (3), washers (6) and nuts (7) securing center retainer (4) to windshield post (5), and left (9) and right (2) windshield halves and. Retain center retainer (4). NOTE:
(7) (8)
B.
To remove only left (9) or right (2) windshield half, remove attaching screws (3), tubing spacers/grommets (12), washers (6), and nuts (7), securing applicable windshield half. Tubing spacers/rubber grommets (12) are not installed in the windshield center retainer (4) or windshield post (5), but rather, around outer periphery of windshield halves.
Remove windshield halves as desired. Carefully remove existing sealant, solution, tape, and foreign material from left (9) and right (2) windshield halves, center retainer (4), windshield post (5), left (1) and right (8) outboard retainers, cowling deck (16), and upper cabin skin (17).
Install Windshield (Refer to Figure 401). NOTE: (1) (2) (3) (4) (5) (6)
Pro Seal 890 can be combined with either a B2 or B4 accelerator.
Ensure left (9) and right (2) windshield halves, center retainer (4), windshield post (5), left (1) and right (8) outboard retainers, cowling deck (16), and upper cabin skin (17) are free of foreign material, old sealant and solutions. Scuff periphery of left (9) and right (2) windshield halves where center retainer (4), left (1) and right (8) outboard retainers, lower left (10) and lower right (11) retainers overlap with Scotch Brite abrasive pad or equivalent. Spread a thin layer of Pro Seal 890 to inside surface of upper cabin skin (17) where skin laps periphery of left (9) and right (2) windshield halves. Apply FS4291 Sealant Tape, as required, to airframe where periphery of left (9) and right (2) windshield halves lap the airframe. Do not remove sealant tape backing common to forward lower region of windshield. Peel ends of FS4291 Sealant Tape backing to allow a portion of backing to protrude from under windshield halves. Install left (9) and right (2) windshield halves and secure upper portion of windshield to upper cabin skin (17) using screws (3), washers (6), and nuts (7). Finger tighten nuts.
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9) (10) (11) (12) (13) (14)
Once windshield halves have been properly positioned, gently grasp sealant tape backing and pull from between windshield halves and airframe. Firmly press left (9) and right (2) windshield halves against airframe to provide an adequate seal. Fill gaps resulting from contour difference using additional FS4291 Sealant Tape. Carefully trim ends of FS4291 Sealant Tape left protruding in step 3.B.(5). Fill void between left (9) and right (2) windshield halves and airframe with Pro Seal 890. Smooth sealant to provide flush transition between edges of left (9) and right (2) windshield halves and airframe. Spread a thin layer of Pro Seal 890 over exterior surface of left (9) and right (2) windshield halves where lower left (10) and right (11) lower retainers lap left (9) and right (2) windshield halves. Apply a bead of Pro Seal 890 to airframe where left (10) and right (11) retainers lap airframe. Spread a thin layer of Pro Seal 890 over interior surface of lower left (10) and right (11) retainers.
(15) Hand form and install lower left (10) and right (11) lower retainers using previously retained screws (3), tubing spacers/rubber grommets (12), and nuts (7). (16) Tighten loosely installed fasteners from step 3.B.(6). (17) Apply two beads of Pro Seal 890 to exterior surfaces of left (9) and right (2) windshield halves in the region lapped by the center retainer (4). (18) Fill void between left (9) and right (2) windshield halves using Pro Seal 890. Avoid obstructing holes common to windshield post. (19) Fill any gaps between lower left (10) and right (11) lower retainers with Pro Seal 890 if necessary. (20) Install and secure center retainer (4) to windshield post (5, to left (9) and right (2) windshield halves using screws (3), washers (6) and nuts (7). (21) Spread a thin layer of Pro Seal 890 over interior surface of left (8) and right (1) outboard retainers. (22) Install left (8) and right (1) outboard retainers. Secure using screws (3), tubing spacers/rubber grommets (12), and nuts (7). (23) Remove excess sealant from windshield and airframe.
CAUTION: Do not use any tool, abrasive or cleaner which might damage windshield. (24) Remove protective covering from left (9) and right (2) windshield halves. (25) Install cover strip (15) on windshield post (5). (26) Install windshield anti-ice panel if removed. Refer to Chapter 30, Windshield Anti-Ice Maintenance Practices. (27) Install OAT gage. Refer to Chapter 34, Outside Air Temperature Gage - Maintenance Practices.
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Windshield Installation Figure 401 (Sheet 1)
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Windshield Installation Figure 401 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL CABIN WINDOWS - REMOVAL/INSTALLATION 1.
General A.
2.
This section provides removal and installation procedures for the cabin windows.
Cabin Window Removal/installation A.
Remove Cabin Windows (Refer to Figure 401). (1) Cover window to protect from scratches. (2) Remove upholstery trim around window by removing the upholstery attaching screws. (3) Drill out attaching rivets (3) and remove retainer rings (4) to remove window. Retain retainer rings (4). (4) Remove window from inside cabin opening.
B.
Install Cabin Windows (Refer to Figure 401). (1) Cover window to protect from scratches. (2) Ensure window and cabin skin are free from foreign material, old sealant and solutions. (3) Apply Pro-Seal 890 with 890A Accelerator around outer periphery of window in a bead 0.85 inches wide. (4) Position window in cabin opening. (5) Install rivets through cabin skin and window. (6) Install retainer rings (4) on rivet shanks. (7) Buck rivet shanks just enough to swell the rivet slightly. (8) Install upholstery trim around cabin window with attaching screws, and remove protective cover from window.
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Cabin Window Installations Figure 401 (Sheet 1)
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Cabin Window Installations Figure 401 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL WINDOW PLUGS - REMOVAL/INSTALLATION 1.
General A.
2.
3.
Optional window plugs are available on the Model 208 cargo version beginning with Airplanes 20800114 and On, to block-out and protect the seven passenger windows and three windows installed in the cargo door and passenger air-stair door. The window plugs are equipped with a handle (strap assembly) to facilitate clipping the plugs in place. The textured surface of the plug (opposite side of the handle) faces outboard and the plugs are retained in place by three spring clips on each window.
Window Plug Removal/Installation A.
Remove Window Plugs (Refer to Figure 401). (1) Grasp the window plug (7) by the strap assembly handle (9) and slip the plug from under the three retaining spring clips (6) which are attached to the window trim molding.
B.
Install Window Plugs (Refer to Figure 401). (1) Grasp the window plug (7) by the strap assembly handle (9) and slip the plug under the three spring clips (6) one at a time while holding the spring clips open one at a time. (2) Ensure the window plug (7) is properly positioned on the window and the felt trim line is symmetrically aligned with the window trim molding (5). Spring clips (6) should fasten over the reinforcement strips (8) on the window plug assembly (7).
Spring Clip Removal/Installation A.
Remove Spring Clip (Refer to Figure 401). (1) Remove and retain nut (15), washer (14) and screw (10). The cover (11), spring clip (12) and stiffener (13) will be free fro removal/ Retain all good parts for reinstallation.
B.
Install Spring Clip (Refer to Figure 401). (1) Align cover (11), spring clip (12) and stiffener (13) with two mounting holes in the window trim molding (5) and attach securely with screw (10), washer (14) and nut (15).
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Passenger Window Plug Installation Figure 401 (Sheet 1)
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Passenger Window Plug Installation Figure 401 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL DOOR WINDOWS - REMOVAL/INSTALLATION 1.
2.
General A.
The left and right crew doors incorporate windows made of 0.250 inch thick, green-tinted acrylic. The left crew door incorporates a vent window hinged at the bottom which opens inward. The vent window is made of 0.375 inch thick, green-tinted acrylic.
B.
Two cargo door windows are mounted in the top half of the cargo door. The cargo door windows are made of 0.187 inch thick, green-tinted acrylic.
C.
One window is mounted in the top half of the right side passenger door. The passenger door window is made of 0.187 inch thick, green-tinted acrylic.
D.
This section provides instructions for removal and installation of the windows located within both passenger and crew doors.
Tools, Equipment and Materials A.
3.
4.
For a list of required tools, equipment and materials, refer to Windows - General.
Crew Door Window Removal/Installation A.
Remove Crew Door Window (Refer to Figure 401). (1) Cover crew door window to protect from scratches. (2) Remove upholstery trim around window and door assist strap by removing the upholstery and door assist strap attaching screws. (3) Drill out attaching rivets (21) around the periphery of the window (22) securing retainer rings (20). (4) Remove crew window (22) and sealant (19).
B.
Install Crew Door Window (Refer to Figure 401). (1) Cover window to protect from scratches. (2) Ensure opening around window is free from foreign material, old sealant and solutions. (3) Apply FS-4291 sealant around outer periphery crew door window (22). (4) Position crew door window (22) in cabin door skin (18) opening. (5) Install rivets (21) through cabin door skin (18) and crew window (22). (6) Install retainer ring (20) on rivet shanks (21). (7) Buck rivet shanks just enough to swell the rivet slightly. (8) Install upholstery trim around cabin window (22) with attaching screws and remove protective cover from window.
Pilot’s Vent Window Removal/Installation A.
Remove Pilot’s Vent Window (Refer to Figure 401). (1) Cover window to protect from scratches. (2) Move handle (8) to disengage vent window (6) from window stop. (3) Remove nuts (23), screws (25) and spacers (24), and remove vent window. (4) Retain seal (4) for reinstallation. (5) If hinge straps (13) are to be removed, remove attaching hardware. (6) If required, remove handle (8) by removing nut (10), handle (8), spacer (7) and screw.
B.
Install Pilot’s Vent Window (Refer to Figure 401). (1) Cover window to protect from scratches. (2) Install seal (5) on vent window (6). (3) Align holes in vent window (6) with holes in hinge strap assembly (13). (4) Install hardware securing hinge strap assembly to vent window (6). (5) If handle (8) was removed, install screw, spacer (7), handle (8), washer (9) and nut (10) on vent window (6). (6) Remove protective cover from vent window. (7) Install spacers (24), screws (25) and nuts (23). (8) Engage handle (8) to secure vent window.
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Crew Door Windows and Pilot’s Vent Window Installation Figure 401 (Sheet 1)
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5.
Cargo and Passenger Door Window Removal/Installation A.
Remove Cargo and Passenger Door Windows (Refer to Figure 402). (1) Cover window (5) to protect from scratches. (2) Remove upholstery around window by removing the upholstery attaching screws. (3) Drill out attaching rivets (3) around periphery of window (1) or (2) securing retainer rings (4). (4) Remove window from cabin opening and remove sealant (6).
B.
Install Cargo and Passenger Door Windows (Refer to Figure 402). (1) Cover window (5) to protect from scratches. (2) Ensure opening around window is free from foreign material, old sealant and solutions. (3) Apply FS-4291 sealant (6) around outer periphery of window (5). (4) Position window in door opening and install rivets (3) through door frame and window (5). (5) Install retainer ring (4) on rivet shafts (3). (6) Buck rivet shanks (3) just enough o swell the rivet slightly. (7) Install upholstery trim around window (5) with attaching screws and remove protective cover from window.
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MODEL 208 MAINTENANCE MANUAL
Cargo and Passenger Door Window Installation Figure 402 (Sheet 1)
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Cargo and Passenger Door Window Installation Figure 402 (Sheet 2)
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57 CHAPTER
WINGS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
57-00-00
Page 1
Apr 1/2010
57-10-00
Pages 401-409
Jun 1/2011
57-10-00
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Jun 1/2011
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Nov 3/2003
57-10-01
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Jun 1/2011
57-11-00
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Aug 1/1995
57-20-00
Pages 201-203
Apr 1/2010
57-40-00
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Aug 1/1995
57-50-00
Pages 201-206
Apr 1/2010
57-60-00
Pages 201-203
Apr 1/2010
57-70-00
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Jan 3/2005
57-Title 57-List of Effective Pages 57-Record of Temporary Revisions 57-Table of Contents 57-List of Tasks
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS WINGS - GENERAL. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57-00-00 57-00-00 57-00-00 57-00-00
Page 1 Page 1 Page 1 Page 1
WINGS - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Incidence Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Tip Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bullet Fabrication and Use . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57-10-00 Page 401 57-10-00 Page 401 57-10-00 Page 401 57-10-00 Page 408 57-10-00 Page 408 57-10-00 Page 408
WINGS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Zonal Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing to Carry - Thru Front Spar Attachment Fittings Special Detailed Inspection . . Wing to Carry - Thru Rear Spar Attachment Fittings Special Detailed Inspection . . Front Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection . . . . . . . . Rear Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection. . . . . . . . . Center Flap Track and Inboard Flap Track Special Detailed Inspection . . . . . . . . . . . Outboard Flap Track Special Detailed Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57-10-00 Page 601 57-10-00 Page 601 57-10-00 Page 601 57-10-00 Page 604 57-10-00 Page 604 57-10-00 Page 605 57-10-00 Page 605 57-10-00 Page 606 57-10-00 Page 606
WING STRUT ATTACH FITTING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Strut Channel Modification. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57-10-01 Page 201 57-10-01 Page 201 57-10-01 Page 201 57-10-01 Page 201
WING STRUT ATTACH FITTING - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Strut Fittings Special Detailed Inspection (Typical Inspection Compliance) . . Wing Strut Fittings Special Detailed Inspection (Severe Inspection Compliance) . . Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/ Standard Bolt Size) (Typical Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/ Standard Bolt Size) (Severe Inspection Compliance). . . . . . . . . . . . . . . . . . . . . . . . . Wing Strut Attachment to Front Spar Special Detailed Inspection (1/64 Inch Oversize Bolt Size) (Severe Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . . Wing Strut Attachment to Front Spar Special Detailed Inspection (1/32 Inch Oversize Bolt Size) (Severe Inspection Compliance) . . . . . . . . . . . . . . . . . . . . . . . . .
57-10-01 Page 601 57-10-01 Page 601 57-10-01 Page 601 57-10-01 Page 601
WING PLATES/SKINS - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Plates/Skins Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57-11-00 Page 401 57-11-00 Page 401 57-11-00 Page 401
VORTEX GENERATORS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vortex Generator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57-20-00 Page 201 57-20-00 Page 201 57-20-00 Page 201 57-20-00 Page 201
LEADING EDGE - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Leading Edge Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57-40-00 Page 401 57-40-00 Page 401 57-40-00 Page 401
FLAPS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment of Fuselage-Mounted Flap Roller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Vortex Generator Boots Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL AILERONS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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SPOILERS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spoiler Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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LIST OF TASKS 57-10-00-210
Wing Zonal Inspection
57-10-00 Page 601
57-10-00-250
Wing to Carry - Thru Front Spar Attachment Fittings Special Detailed Inspection
57-10-00 Page 604
57-10-00-251
Wing to Carry - Thru Rear Spar Attachment Fittings Special Detailed Inspection
57-10-00 Page 604
57-10-00-252
Front Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection
57-10-00 Page 605
57-10-00-253
Rear Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection
57-10-00 Page 605
57-10-00-254
Center Flap Track and Inboard Flap Track Special Detailed Inspection
57-10-00 Page 606
57-10-00-255
Outboard Flap Track Special Detailed Inspection
57-10-00 Page 606
57-10-01-250
Wing Strut Fittings Special Detailed Inspection (Typical Inspection Compliance)
57-10-01 Page 601
57-10-01-251
Wing Strut Fittings Special Detailed Inspection (Severe Inspection Compliance)
57-10-01 Page 601
57-10-01-252
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Typical Inspection Compliance)
57-10-01 Page 602
57-10-01-253
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Severe Inspection Compliance)
57-10-01 Page 603
57-10-01-254
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/64 Inch Oversize Bolt Size) (Severe Inspection Compliance)
57-10-01 Page 603
57-10-01-255
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/32 Inch Oversize Bolt Size) (Severe Inspection Compliance)
57-10-01 Page 604
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MODEL 208 MAINTENANCE MANUAL WINGS - GENERAL 1.
Scope A.
2.
This chapter gives a description for those wing structures and associated components which support the airplane in flight. Included are the attach fittings, ailerons, and flaps.
Tools, Equipment and Materials A.
Equivalent alternatives may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Installation Cement
Scotch Grip EC1300L cement (FSN 8040-628-4199)
3M Co. St. Paul, MN 55144-1000
To bond the de-ice boot to the flap.
Commercially Available
To clean the boot and mating surface.
Methyl n-Propyl Ketone Cleaning Solvent
Hexane
Commercially Available
To clean the mating surface.
Cleaning Solvent
Toluol FSN TT-T548
Commercially Available
To remove the boot, to remove vortex generators and clean wing surface where they are installed.
Neoprene Coating
GACO N-700-A
Gates Engineering Co. Wilmington, DE
To apply a finish to the edges of the boot.
Rubber Roller
Two inch (5 cm) Rubber
Gates Engineering Co. Wilmington, DE
To install the boots.
Metal Stitcher Roller
1/4 inch (0.63 cm)
Commercially Available
To install the boots.
Masking Tape
1 inch
Commercially Available
To isolate the boot area.
Carpenter’s Chalk Line
Commercially Available
To put a mark on the centerline of the flap and boot.
Measuring Tape
Commercially Available
Sharp Knives
Commercially Available
Cleaning Cloth (Lint Free)
Commercially Available
3.
Definition A.
This chapter is divided into sections to help maintenance personnel find data. The Table of Contents will also help find specified data. A brief definition of the sections included in this chapter is as follows: (1) The section on Wings gives a description of the components and the removal and installation procedures. (2) The section on Wing Plates/Skins gives a description of the components and the removal and installation procedures. (3) The section on Leading Edge gives a description of the components and the removal and installation procedures. (4) The section on Flaps gives a description of the components and the maintenance practices. (5) The section on Ailerons gives a description of the components and the maintenance practices. (6) The section on Spoilers gives a description of the components and the maintenance practices.
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MODEL 208 MAINTENANCE MANUAL WINGS - REMOVAL/INSTALLATION 1.
2.
General A.
The wing assembly consists of all metal left and right removable panels, each braced by a lift strut and attached to fuselage on each side with bolts through holes in attach fittings in both fuselage and inboard ends of wing spars. The structure of each wing panel is of conventional, semi-monocoque design employing a front and rear spar, ribs, stringers, and skins. Each wing panel also incorporates a sealed, integral fuel bay, a flap, a balanced aileron, and a slot-lip spoiler. Access holes in lower wing skins between ribs provide access to fuel bay, flight controls, electrical components, deice system plumbing, and ventilation system components. Wing tips of sheet metal construction attach to each wing panel with screws. The wing tips contain navigation lights and provisions for strobe lights. The airplane landing and taxi lights are located in leading edge of each wing panel, between Wing Stations 185.30 and 201.75. A stall warning detector and a pitot static probe are incorporated into leading edge of the left wing panel.
B.
The wing panel main frame structure consists of a front and rear spar assembly, center ribs, and upper and lower skins. The spars are of bonded and riveted construction and stringers between spars are bonded to interior wing skins. Spar caps are extruded angles riveted and bonded to sheet metal webs. The front spar incorporates a special forged fitting and formed channel assembly for lift strut attachment. Access openings with covers are provided between ribs to allow access to fuel bay and flight control system.
C.
The wings are attached to the fuselage with attach fittings on the forward and aft spar on each side of the fuselage. The forward spar also has fittings for the attachment of the lift strut. The wing trailing edge structure contains fittings for flap and aileron attachments. Wing spar fittings and fuselage fittings are shown in Figure 401.
D.
Sheet metal wing tips attach to the wing structure with screws, and contain navigation lights plus provisions for strobe lights.
Wing Removal/Installation A.
Preparation for Removal of Wing (Refer to Figure 401). (1) Turn off all electrical power, and ground the airplane structure. (2) Defuel the airplane in accordance with Chapter 28, Fuel System - Maintenance Practices. (3) Remove the attach screws from the lower wing root access covers (1) and (2) and remove the covers. (4) On the 208, Remove the attach screws from the strap fairing assembly (17) and the fairing assembly (16) and remove the fairings. (5) On 208B, remove the attach screws from the fairing assembly (17A) and strap fairing assembly (17B) and remove the fairings. (6) Remove the attach screws from the air inlet cover (5) and the upper lift strut fairing (6) and remove the cover and the fairing. (7) Remove the attach screws from the lower lift strut fairing (3) and remove the fairing. (8) Disconnect the lift strut deice system plumbing (if installed) at connections inside the wing at the upper lift strut attachment location. (9) Disconnect the wing deice system plumbing (if installed) in the wing root area. (10) Loosen and remove the forward and aft hose connections of fuel bay supply lines in the wing root area. Drain the residual fuel. (11) Remove the hose connection from fuel vent system crossover (right wing) or disconnect the three hoses of the fuel vent system (left wing) at the vent system cross inside the inboard wing bay. (12) Loosen and separate the electrical wiring connector in the wing root leading edge. (13) Disconnect the flap motor and crossover pushrods at the inboard flap bell cranks. (14) Disconnect the pitot/static lines at the connections in the leading edge root area (left wing only). Cap the lines to prevent contamination. (15) Remove the headliner and cut the safety wire and disconnect turnbuckles to relieve tension on the aileron control system carry-thru cables.
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Wing Installation Figure 401 (Sheet 1)
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Wing Installation Figure 401 (Sheet 2)
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Wing Installation Figure 401 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (16) If the right wing is to be removed, disconnect the aileron trim cables at the quick-disconnects located in the inboard leading edge wing bay. Refer to Chapter 27, Aileron Trim System Maintenance Practices. (17) Disconnect the wiring, the shutoff valve linkage, and the inlet duct from the cabin air blower system in wing root area. Refer to Chapter 21, Fresh Air Distribution - Maintenance Practices. (18) Disconnect the fuel tank shutoff valve control at the clevis on the aft valve lever arm that is located in the wing root area. Remove the clevis attaches the interconnect link to the forward valve lever arm and remove the safety wire that attaches the control to mounting bracket. B.
Remove the Wing
CAUTION: The wing must be separated from the fuselage at as near normal a dihedral angle as possible. Any motion up or down at the wing tip greater than +1.00 or -1.00 inch as the fittings are separated will damage attach fittings. CAUTION: Support of the wing during the spar attach bolt and the strut support bolt removal is critical. If the bolts are loaded, it is impossible to remove them without damaging the fitting. (1)
Remove the wing struts (4). NOTE: (a) (b) (c) (d)
Use a wing jack or a hoist to raise or lower the wing tip.
Make sure that all the load is removed from the strut-to-wing attach bolt. Remove the nuts from the upper and lower strut fittings. Apply a wrench to the head of the upper and lower strut attach bolts (30, 40) and attempt to rotate the bolts in the fittings. Raise or lower the wing tip in very small increments until the force required to rotate the bolts is at a minimum. NOTE:
(e) (f) (g) (h)
A torque wrench can be useful to determine the minimum rotational force.
If the bolts will not rotate, there may be corrosion between the bolts and the fittings. 1 Apply penetrating oil such as Kroil or Mouse Milk to the area and allow it to penetrate and dissolve the corrosion. When the bolts rotate with minimum force, the wing is in the proper position for removal of the attach bolts. Make sure the inboard and outboard areas of the wings are supported. Remove the upper strut-to-wing attach bolt (40). NOTE:
If you cannot remove the bolts with your fingers, a "bullet" can be fabricated to help drive the bolt out. Refer to Bullet Fabrication and Use.
Support the wing struts and remove the lower fuselage-to-strut bolts and then remove the struts. Remove the wings from the fuselage attach fittings. (a) Mark the location of the incidence setting of the index marker (arrow) on the head of the rear attach bolt (24) to the face of fitting (29) so the incidence setting is kept when wing is reinstalled. (b) Remove the forward and then the aft wing spar attach bolts (18) and (24). (i)
(2)
NOTE:
If you cannot remove the bolts with your fingers, a "bullet" can be fabricated to help drive the bolt out. Refer to Bullet Fabrication and Use.
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CAUTION: The wing must be separated from the fuselage at as near normal a dihedral angle as possible. Any motion up or down at the wing tip greater than +1.00 or -1.00 inch as the fittings are separated will damage attach fittings. (c) (d) (e) C.
Carefully move the wing outboard until the lugs on the wings and fuselage are separated. Place the removed wing on padded support. Remove the eccentrics (26, 27) from the lugs if necessary.
Install the Wing (Refer to Figure 401). (1) Install the eccentrics in the aft spar wing/fuselage attachment fittings. NOTE:
The aft spar/fuselage attachment incorporates eccentrics to allow the wing incidence adjustments to trim out the wing heaviness.
NOTE:
Correct positioning of eccentrics in the fittings is critical.
(a)
(2)
Assemble the aft attach fittings incidence eccentrics (26) into the fuselage attach fittings (29), with the keyway slot of the eccentrics approximately 90 degrees clockwise from the marked location of the bolt arrow and the thick side of the eccentric inboard. Refer to Figure 401. (b) Assemble the incidence eccentrics (27) into the aft wing attach fittings (28), with the keyway slot of the eccentrics aligned with those in the fuselage fittings and the thick side of the eccentrics outboard. Refer to Figure 401. Install the wings to the fuselage . (a) Apply MIL-G-21164 grease to each fitting lug face and bore.
CAUTION: Wing must be mated to fuselage at as near normal a dihedral angle as possible. Motion up or down at wing tip as or after fittings are mated must be limited to 1.00 inch up or down to avoid damage to attach fittings. CAUTION: Support of the wing during the spar attach bolt and the strut support bolt installation is critical. If the bolts are loaded, it is impossible to install them without damage to the fitting. (b)
(3)
Carefully position the wings to mate the spar attach fittings (23) and (28) to the fuselage attach fittings (20) and (29). Install the forward spar wing/fuselage attach fittings. (a) Install forward attach fitting bolt (18). NOTE:
It is important that the wing fittings are not spread out or bent in by the process of bolt installation.
Install the nut (21A). Tighten the nut until it is snug, then loosen until the washer under the nut is free to turn, and the cotter pin can be installed. Install the aft spar wing/fuselage attach fittings. (a) Apply MIL-G-21164 grease to the bolt shank before you insert it. 1 Put a keyway washer (29B) on the bolt (24). 2 Install the rear spar attach bolt (24) so that the index arrow is pointed as it was marked in the removal procedure. This will make sure wing incidence rigging is not disturbed. This bolt is installed with the head aft. (b)
(4)
NOTE:
The bullet used for disassembly should be inserted first to precisely align the fittings so the bolt threads do not damage the fittings.
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MODEL 208 MAINTENANCE MANUAL 3
Install five washers (25) and nut (21B), or the lifting hinge (29C) with one washer (25), which ever configuration applies.
WARNING: This nut tightening procedure is for the aft wing attach fittings only. Tighten the nut until it is snug and the washer (25) is no longer free to rotate. Then tighten the nut not more than one more full turn and line up the cotter pin hole. Install the cotter pin (22). Install the wing struts. (a) Put the lift strut (4) lower fitting into the fuselage strut fitting. (b) Install the bolt (30) into the lower end of the strut and fuselage fitting. (c) With the outboard section of the wing supported at a normal dihedral angle, raise the lift strut (4) to align the upper attach fitting holes (38) with the strut wing fitting (39). (d) With the strut held against one face of the wing fitting, measure the gap between the strut fitting and the wing fitting. 1 If the gap is 0.025 to 0.070 inch, it will be necessary to install a single 2622246-X shim. Refer to the Illustrated Parts Catalog for dash number and thickness of shim. a Choose the shim that will provide a minimum gap of 0.005 inch. b Apply a fay seal of Type 1 Class C sealant to one face only of the shim. Refer to Chapter 20, Fuel, Weather, and High Temperature Sealing. c Adhere the sealant to the wing attach fitting. 2 If the gap is less than 0.025 inch, proceed to the next step. (e) Install the spacer (37) and carefully tap lift strut-to-wing attach bolt (40) into place. 4
(5)
NOTE: (f)
Install the nut (33). Tighten the nut until it is snug, then loosen until the washer under the nut is free to turn, and the cotter pin can be installed. NOTE:
(6) (7)
(8) (9) (10) (11) (12)
(13) (14) (15) (16) (17) (18) (19)
The bullet used for disassembly should be inserted first to precisely align the fittings so the bolt threads do not damage the fittings.
It is important that the wing fittings are not spread out or bent in by the process of bolt installation.
Reconnect wiring, shutoff valve linkage, and inlet duct from cabin air blower system. Refer to Chapter 21, Fresh Air Distribution System - Maintenance Practices. Connect aileron control system cables at turnbuckles in overhead cabin area. If right wing is being installed, connect aileron trim tab cables at quick-disconnects in right inboard leading edge bay. Adjust cable tensions and rig aileron and aileron trim systems in accordance with Chapter 27, Aileron Trim System - Maintenance Practices. Connect pitot/static system lines at connections in wing leading edge root area (left wing only). Connect flap motor and crossover pushrods at inboard flap bell cranks. Rig flaps in accordance with Chapter 27. Connect electrical wiring connector in wing root leading edge. In left wing, connect fuel vent system crossover hose, or in right wing, connect three vent hoses to vent system cross inside inboard wing bay. Connect fuel system forward and aft supply lines in wing root area. Wire fuel tank shutoff valve control to mounting bracket. Attach interconnect link to forward valve lever arm and connect control to aft valve lever arm. Insert cotter pin to clevis pins of both lever arm. Fuel airplane and check for leaks. Connect wing deice system plumbing (if installed) in wing root area. Connect lift strut deice boot plumbing at connections inside lift strut upper attachment location. Install lift strut-to-fuselage fairing (3). On 208, install strap fairing assembly (17) and fairing assembly (16). On 208B, install fairing assembly (17A) and strap fairing assembly (17B). Install lower wing root access covers (1) and (2). Install lift strut-to-wing fairing (6) and air inlet cover (5).
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MODEL 208 MAINTENANCE MANUAL (20) Remove the electrical ground and apply power to airplane electrical system. (a) Examine the landing/taxi lights, stall warning system, pitot/static system heaters, navigation lights, fuel quantity indicating system, air conditioning or vent blowers, and deice system (if installed) for correct operation. (21) Complete the Pitot System Inspection and Leak Test and the Pitot/Static System Functional Check. Refer to Chapter 34, Pitot/Static System - Inspection/Check. 3.
Incidence Adjustment A.
General (1) The Incidence of the wing can be adjusted to reduce a "wing heavy" condition by adjustment of the aft fuselage-to-wing attach bolt (a) This procedure is normally done after a flight test. It should not be required unless major wing damage to a wing has occurred. (b) If adjustment is required, make the adjustment to both wings in opposite directions. (2) Incidence Adjustment (a) Increase incidence is by rotating bolt head so that arrow points down. To decrease incidence, rotate bolt head so that arrow points up. NOTE: (b)
4.
5.
Lettering on bolt head of the left wing will be upside down when arrow points outboard.
Do not rotate the arrow past vertical in either direction.
Wing Tip Removal/Installation A.
Remove Wing Tip (Refer to Figure 401). (1) Turn off all electrical power. (2) Remove attach screws and slide wing tip outboard slightly to gain access to electrical plug of tip navigation light. (3) Disconnect electrical plug(s) and remove wing tip. (4) Remove navigation light from wing tip.
B.
Install Wing Tip (Refer to Figure 401). (1) Install navigation light in wing tip assembly. (2) Connect navigation light electrical plug and install wing tip assembly with attach screws. (3) Restore electrical power and check navigation light operation.
Bullet Fabrication and Use A.
Fabricate a Bullet (Refer to Figure 402). (1) Get a NAS464P14-76 and NAS464P12-37 bolt. (2) Grind to remove the threads and head of the bolts as shown. (3) Make sure there are no burrs or sharp edges at the transition to the full bolt diameter.
B.
Use the Bullet. (1) When you drive an existing bolt out of the fittings, the blunt end must be against the threaded end of the bolt so the threaded part of the bolt does not cause gouges in the fittings. NOTE:
(2)
It is possible that shifting of the fittings can occur as the bolt is removed if the bullet is not used.
(a) Always lubricate the shank of the bullet before you use it. (b) After its use it is necessary to drive the bullet out of the fitting with a drift punch. When you use a bullet to align the fittings for the installation of a bolt, put the more tapered end into the lug holes from the direction the bolt is to be installed. The bolt will then drive the bullet out of the lugs as it is forced into place.
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MODEL 208 MAINTENANCE MANUAL
Bullet Fabrication Figure 402 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL WINGS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the wings in a serviceable condition.
Task 57-10-00-210 2.
Wing Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for an zonal inspection of the wings. NOTE:
B.
Special Tools (1) None
C.
Access NOTE: (1) (2) (3)
D.
An external zonal GVI is a general visual examination of an exterior area, and/or an open installation or assembly to find damage, failure or defects. This level of inspection is made during typical lighting conditions such as daylight, hangar light or flashlight by approximately an arm-length distance to the inspection object. Unless it is specified, it is not necessary to remove or open access panels or doors to do an external GVI. You can use an inspection mirror to help with visual access to all opened surfaces in the inspection area. You can use maintenance stands, ladders, or platforms to get near the inspection area.
The lower wing fuel access panels and lower wing dry bay panels are removed after the inspection steps for removing the fuel.
Lower wing fuel access panels. Lower wing dry bay panels. Wing and strut fairing panels.
Do the External Zonal Inspection of the Wings. (1) Examine the external wings for loose fasteners, corrosion, cracks, wrinkles, and dents. NOTE:
If you suspect corrosion under the deice boots (if installed), remove the deice boots for inspection. Refer to Chapter 30, Pneumatic Deice Boots Removal/Installation.
(a)
(2)
Make sure that you examine the areas that follow between WS 35.00 to WS 155.90. The wing leading edge skin surface. The access covers around the screw attachments. (b) Make sure that you examine the areas that follow between WS 155.90 to WS 308.00. The wing leading edge skin surface. The access covers around the screw attachments. (c) Make sure that you examine the areas that follow between WS 53.00 to WS 214.30. The wing upper and lower forward skin surface. The access covers around the screw attachments. (d) Make sure that you examine the areas that follow between WS 214.30 to WS 308.00. The wing upper and lower forward skin surface. The access covers around the screw attachments. (e) Make sure that you examine the areas that follow between WS 35.00 to WS 229.00. The wing upper and lower aft skin surface. The access covers around the screw attachments. (f) Make sure that you examine the areas that follow between WS 35.00 to WS 228.00. The flap skin surface and the flap leading edge skin surface. (g) Make sure that you examine the areas that follow at WS 53.00, WS 126.50, and WS 214.30. The flap track structure including the inboard, center, and outboard flap tracks. The inboard, center, and outboard flap support attach bolt, bracket and attach bolt. Examine the attach points for condition and security of installation.
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MODEL 208 MAINTENANCE MANUAL (3) (4)
Examine the full wing surface for fuel leaks. Examine the wing spar fittings and bolts for corrosion, condition and security of installation. (a) Make sure that you examine the areas that follow at WS 35.00. The forward spar fitting and lug surface. The rear spar fitting and lug surface. NOTE:
(5) (6)
Examine the wing struts for signs of damage, condition, and security of installation. Examine the upper and lower wing strut fittings, fairings and bolts for corrosion, condition and security of installation. (a) Make sure that you examine the areas that follow at FS 168.70 for the Model 208 and FS 188.70 for the Model 208B. The lower wing strut to fuselage attach fitting and lug surface. The wing strut to wing attach fitting and lug surface. NOTE:
(7) (8) (9) (10)
E.
If corrosion is found on the lug surface or the attaching hardware (bolt, nut, or cotter pin), remove the attach bolt and inspect the lug bore.
If corrosion is found on the lug surface or the attaching hardware (bolt, nut, or cotter pin), remove the attach bolt and inspect the lug bore.
Examine the drain openings and vent holes in the bottom of the wing for obstructions. Examine all wing access panels for security of installation and signs of damage. Examine the fuel access panels for signs of leaks. Examine the external wing surface for damage and signs of overheating. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, External Zonal Visual Inspection of Lightning and High Intensity Radiated Fields.
Do the Internal Zonal Inspection of the Wing Wet Bays. NOTE:
An internal zonal GVI is a general visual examination that includes all of the systems and the structural components of an interior area, installation, or assembly. This includes a check for signs of corrosion, cracks, chafing of tubing, loose duct support, wiring damage, cable and pulley wear, fluid leaks, drainage that is not sufficient, and other conditions that can cause corrosion or damage. This level of inspection is made during typical lighting conditions such as daylight, hangar light, flood light, or flashlight by approximately an armlength distance to the inspection object. It can be necessary to remove and/or open access panels or doors to complete an internal GVI. You can use an inspection mirror to help with visual access to all opened surfaces in the inspection area. You can use maintenance stands, ladders, or platforms to get near the inspection area.
WARNING: Before you do maintenance on the fuel system, you must read and understand all of the fuel system maintenance, fire precautions, and safety practices. Refer to Fuel System - Maintenance Practices and Chapter 12, Fuel – Servicing. (1) (2)
Defuel the airplane. Refer to Chapter 12, Fuel – Servicing. (a) Remove the remaining fuel from the fuel storage areas with the fuel drain valves. Refer to Chapter 12, Fuel – Servicing. Remove lower wing fuel access panels 521AB, 521BB, 521DB, 521EB left, and 621AB, 621BB, 621DB and 621EB right. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation.
CAUTION: Be careful to not separate the wing skin doubler from the wing skin. (3) (4) (5)
(a) Purge the fuel tanks. Refer to Chapter 12, Fuel – Servicing. Purge fuel tanks. Refer to Chapter 12, Fuel - Servicing. Examine the eight (8) quantity transmitter mounting plates for condition, leaks, and security. Examine the transmitters wire harnesses and terminals at the transmitters for condition and security.
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MODEL 208 MAINTENANCE MANUAL (6) (7)
(8) (9) F.
Do the Internal Zonal Inspection of the Wing Dry Bays. (1) Remove lower wing dry bay panels 501AB, 501BB, 501CB, 501DB, 501EB, 503AB,503BB, 503CB, 503DB, 503EB, 503FB, 503GB, 503HB, 503JB, 511AB, 525AB, 525BB, 525CB, 525DB, 525EB, 525FB, 525GB, 551AB, and 575AB left, and 601AB, 601BB, 601CB, 601DB, 601EB, 603AB, 603BB, 603CB, 603DB, 603EB, 603FB, 603GB, 603HB, 603JB, 611AB, 621CB, 623AB, 625AB, 625BB, 625CB, 625DB, 625EB, 625FB, 625GB, 651AB, 675AB right, for the internal zonal inspection. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. (2) Examine all of the wire bundle assemblies and the electrical components for signs of overheating, correct installation, frayed or chafed wiring insulation, electrical bonding, damage, and corrosion. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. (3) Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. (4) Examine all tubing, hose, and fluid fittings for signs of leaks, damage, chafing, and correct clamp installation. (5) Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected.
(6) (7)
(8) (9) G.
Examine the tank drains for condition, leaks, and security. Examine all of the wire bundle assemblies and the electrical components for signs of overheating, correct installation, frayed or chafed wiring insulation, electrical bonding, damage, and corrosion. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, Internal Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, and correct installation. Examine all tubing, fuel shut-off-valves, hose, and fluid fittings for signs of leaks, damage, chafing, correct clamp installation, condition, and security.
NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
Install lower wing fuel access panels 521AB, 521BB, 521DB, 521EB left, and 621AB, 621BB, 621DB and 621EB right. Refer to Chapter 6, Access Plates and Panels Identification Description and Operation. Install lower wing dry bay panels 501AB, 501BB, 501CB, 501DB, 501EB, 503AB,503BB, 503CB, 503DB, 503EB, 503FB, 503GB, 503HB, 503JB, 511AB, 525AB, 525BB, 525CB, 525DB, 525EB, 525FB, 525GB, 551AB, and 575AB left, and 601AB, 601BB, 601CB, 601DB, 601EB, 603AB, 603BB, 603CB, 603DB, 603EB, 603FB, 603GB, 603HB, 603JB, 611AB, 621CB, 623AB, 625AB, 625BB, 625CB, 625DB, 625EB, 625FB, 625GB, 651AB, 675AB right. Refer to Chapter 6, Access/Inspection Plates - Description and Operation. Refuel the airplane. Refer to Chapter 12, Fuel – Servicing. Examine the fuel bay panels for leaks.
Restore Access NOTE:
(1) End of task
The lower wing fuel access panels and lower wing dry bay panels are installed before the inspection step to do a leak check of the panels.
None
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MODEL 208 MAINTENANCE MANUAL Task 57-10-00-250 3.
Wing to Carry - Thru Front Spar Attachment Fittings Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing to carry - thru front spar attachment fittings in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing from the airplane. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing to Carry - Thru Front Spar Attachment Fittings. (1) Do a nondestructive testing (NDT) inspection for cracks in the front wing-to-carry-thru spar attach fitting lug. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing to Carry - Thru Spar Attachment Fittings - Description And Operation. (2) Do a NDT inspection for cracks in the front wing-to-carry-thru spar attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing to Carry - Thru Spar Attachment Fittings - Description And Operation. (3) If no cracks are found, restore access. (4) If cracks are found, replace the wing-to-carry-thru spar attach fitting. Refer to Wings - Removal/ Installation.
E.
Restore Access (1) Install the wing. Refer to Wings - Removal/Installation. End of task Task 57-10-00-251 4.
Wing to Carry - Thru Rear Spar Attachment Fittings Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing to carry - thru rear spar attachment fittings in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing from the airplane. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing to Carry - Thru Rear Spar Attachment Fittings. (1) Do a nondestructive testing (NDT) inspection for cracks in the rear wing-to-carry-thru spar attach fitting lug. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing to Carry - Thru Spar Attachment Fittings - Description And Operation. (2) Do a NDT inspection for cracks in the rear wing-to-carry-thru spar attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing to Carry - Thru Spar Attachment Fittings - Description And Operation. (3) If no cracks are found, restore access. (4) If cracks are found, replace the wing-to-carry-thru spar attach fitting. Refer to Wings - Removal/ Installation.
E.
Restore Access (1) Install the wing. Refer to Wings - Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL Task 57-10-00-252 5.
Front Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the front spar lower cap inboard of WS 141.20 in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the applicable access panels on the bottom of the wing to get access to the front spar. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do a Special Detailed Inspection of the Front Spar Lower Cap Inboard of WS 141.20. (1) Do a visual inspection for cracks in the wing front spar lower cap inboard of WS 141.20. (2) Do a nondestructive testing (NDT) inspection for cracks in the wing front spar lower cap between the wing attach fittings and WS 141.20. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Forward Spar Lower Cap Inboard of WS 141.20 - Description And Operation. (3) If no cracks are found, restore access. (4) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures.
E.
Restore Access (1) Installed the access panels that were removed on the bottom of the wing to get access to the front spar. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task Task 57-10-00-253 6.
Rear Spar Lower Cap Inboard of WS 141.20 Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the rear spar lower cap inboard of WS 141.20 in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the applicable access panels on the bottom of the wing to get access to the rear spar. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation.
D.
Do a Special Detailed Inspection of the Rear Spar Lower Cap Inboard of WS 141.20. (1) Do a visual inspection for cracks in the wing rear spar lower cap inboard of WS 141.20. (2) Do a nondestructive testing (NDT) inspection for cracks in the wing rear spar lower cap between the wing attach fittings and WS 141.20. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Aft Spar Lower Cap Inboard of WS 141.20 - Description And Operation. (3) If no cracks are found, restore access. (4) If cracks are found, contact Cessna Propeller Aircraft Product Support for repair procedures.
E.
Restore Access (1) Installed the access panels that were removed on the bottom of the wing to get access to the rear spar. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation. End of task
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MODEL 208 MAINTENANCE MANUAL Task 57-10-00-254 7.
Center Flap Track and Inboard Flap Track Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the center flap track and inboard flap track in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the center and inboard flaps. Refer to Flaps - Maintenance Practices.
D.
Do a Special Detailed Inspection of the Center Flap Track and Inboard Flap Track. (1) Do a nondestructive testing (NDT) inspection for cracks in the inboard flap track at WS 53.00. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Flap Tracks Description And Operation. (2) Do a NDT inspection for cracks in the center flap track at WS 126.50. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Flap Tracks - Description And Operation. (3) If no cracks are found, restore access. (4) If cracks are found, replace the flap track(s). Refer to Flaps - Maintenance Practices.
E.
Restore Access (1) Install the center and inboard flaps. Refer to Flaps - Maintenance Practices. End of task Task 57-10-00-255 8.
Outboard Flap Track Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the outboard flap track in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the outboard flap. Refer to Flaps - Maintenance Practices.
D.
Do a Special Detailed Inspection of the Outboard Flap Track. (1) Do a nondestructive testing (NDT) inspection for cracks in the outboard flap track at WS 214.30. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Flap Tracks Description And Operation. (2) If no cracks are found, restore access. (3) If cracks are found, replace the flap track(s). Refer to Flaps - Maintenance Practices.
E.
Restore Access (1) Install the outboard flap. Refer to Flaps - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL WING STRUT ATTACH FITTING - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives the procedure for the wing strut channel modiÞcation.
Tools and Materials NOTE:
You can use equivalent substitutes for the following items:
NAME
NUMBER
MANUFACTURER
USE
Solvent
Desoto 110
DeSoto, Inc 1700 S. Mt. Prospect Des Plaines, IL 60016
To clean the Þttings before the inspection.
Tip
HTS-5 Tip 170-5/16
APEX Division of Cooper Industries Box 952 Dayton, OH 45401
To remove the strut attach Þttings from the strut.
Nut
MS21042L5
Cessna Aircraft Company Box 7706 Wichita, KS 67277
To use if the replacement of the attach Þtting is necessary.
3.
Wing Strut Channel ModiÞcation A.
Change the Wing Strut Channel (Refer to Figure 201). (1) Remove the cracked wing strut attach Þtting. (a) Remove the wing strut attach Þtting from the strut. Refer to Wings - Removal/Installation. 1 Keep the bolts for later installation. 2 Discard the nuts. (2) Change the wing strut channel. NOTE:
(3)
The wing strut channel modiÞcation is not necessary unless the attach wing strut Þtting(s) are being replaced.
(a) Make a radius on the strut channel. (b) Use Desoto 110 Solvent to clean the strut channel. (c) Apply the epoxy primer, or equivalent, to all the bare metal surfaces. Use the bolts and new nuts to install the wing strut attach Þtting to the strut.
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MODEL 208 MAINTENANCE MANUAL
Wing Strut Fitting Figure 201 (Sheet 1)
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Wing Strut Fitting Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL WING STRUT ATTACH FITTING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the wing strut attach fittings in a serviceable condition.
Task 57-10-01-250 2.
Wing Strut Fittings Special Detailed Inspection (Typical Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing strut fittings in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. (2) Remove the wing struts. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing Strut Fittings. (1) Do a nondestructive testing (NDT) inspection for cracks in the wing strut upper attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (2) Do a NDT inspection for cracks in the wing strut upper attach fitting extrusion radii. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (3) Do a NDT inspection for cracks in the wing strut upper attach fitting lugs. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (4) Do a NDT inspection for cracks in the wing strut lower attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (5) Do a NDT inspection for cracks in the wing strut lower attach fitting extrusion radii. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (6) Do a NDT inspection for cracks in the wing strut lower attach fitting lugs. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (7) If no cracks are found, restore access. (8) If cracks are found, replace the damaged parts. Refer to Chapter 57, Wings - Removal/ Installation.
E.
Restore Access (1) Install the wing struts. Refer to Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. End of task Task 57-10-01-251 3.
Wing Strut Fittings Special Detailed Inspection (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing strut fittings in a serviceable condition.
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MODEL 208 MAINTENANCE MANUAL B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. (2) Remove the wing struts. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing Strut Fittings. (1) Do a nondestructive testing (NDT) inspection for cracks in the wing strut upper attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (2) Do a NDT inspection for cracks in the wing strut upper attach fitting extrusion radii. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (3) Do a NDT inspection for cracks in the wing strut upper attach fitting lugs. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (4) Do a NDT inspection for cracks in the wing strut lower attach fitting holes. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (5) Do a NDT inspection for cracks in the wing strut lower attach fitting extrusion radii. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (6) Do a NDT inspection for cracks in the wing strut lower attach fitting lugs. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Fittings - Description And Operation. (7) If no cracks are found, restore access. (8) If cracks are found, replace the damaged parts. Refer to Chapter 57, Wings - Removal/ Installation.
E.
Restore Access (1) Install the wing struts. Refer to Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. End of task Task 57-10-01-252 4.
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Typical Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing strut attachment to the front spar in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. (2) Remove the wing struts. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing Strut Attachment to Front Spar. (1) Do a nondestructive testing (NDT) inspection for cracks in the forward spar wing strut attach fitting. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation. (2) Do a NDT inspection for cracks in the forward spar wing strut attach fitting lug. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation. (3) If no cracks are found, restore access.
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If cracks are found, replace the forward spar wing strut attach fitting. Refer to Chapter 57, Wings - Removal/Installation.
E.
Restore Access (1) Install the wing struts. Refer to Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. End of task Task 57-10-01-253 5.
Wing Strut Attachment to Front Spar Special Detailed Inspection (Nominal/Standard Bolt Size) (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing strut attachment to the front spar in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. (2) Remove the wing struts. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing Strut Attachment to Front Spar. (1) Do a nondestructive testing (NDT) inspection for cracks in the forward spar wing strut attach fitting. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation. (2) Do a NDT inspection for cracks in the forward spar wing strut attach fitting lug. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation. (3) If no cracks are found, restore access. (4) If cracks are found, replace the forward spar wing strut attach fitting. Refer to Chapter 57, Wings - Removal/Installation.
E.
Restore Access (1) Install the wing struts. Refer to Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. End of task Task 57-10-01-254 6.
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/64 Inch Oversize Bolt Size) (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing strut attachment to the front spar in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. (2) Remove the wing struts. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing Strut Attachment to Front Spar. (1) Do a nondestructive testing (NDT) inspection for cracks in the forward spar wing strut attach fitting. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation.
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Do a NDT inspection for cracks in the forward spar wing strut attach fitting lug. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation. If no cracks are found, restore access. If cracks are found, replace the forward spar wing strut attach fitting. Refer to Chapter 57, Wings - Removal/Installation.
E.
Restore Access (1) Install the wing struts. Refer to Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. End of task Task 57-10-01-255 7.
Wing Strut Attachment to Front Spar Special Detailed Inspection (1/32 Inch Oversize Bolt Size) (Severe Inspection Compliance) A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the wing strut attachment to the front spar in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. (2) Remove the wing struts. Refer to Wings - Removal/Installation.
D.
Do a Special Detailed Inspection of the Wing Strut Attachment to Front Spar. (1) Do a nondestructive testing (NDT) inspection for cracks in the forward spar wing strut attach fitting. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation. (2) Do a NDT inspection for cracks in the forward spar wing strut attach fitting lug. Refer to the Model 208 Nondestructive testing Manual, Part 6, Eddy Current, Wing Strut Attachment to Front Spar - Description And Operation. (3) If no cracks are found, restore access. (4) If cracks are found, replace the forward spar wing strut attach fitting. Refer to Chapter 57, Wings - Removal/Installation.
E.
Restore Access (1) Install the wing struts. Refer to Wings - Removal/Installation. (2) Install the wing strut-to-wing fairings. Refer to Wings - Removal/Installation. End of task
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MODEL 208 MAINTENANCE MANUAL WING PLATES/SKINS - REMOVAL/INSTALLATION 1.
General A.
2.
The wing spars, ribs, and stringers are covered with formed alloy aluminum sheet skins. Upper and lower skins, between front and rear spars from Wing Stations 55.00 to 214.30, form an integral fuel bay. To minimize fuel tank sealing in this area, spanwise, stringers are bonded to upper and lower skins. All rivets on leading edge from upper front spar cap to lower front spar cap are ßush while all other rivets aft of front spar are universal head. Access holes and cover plates are provided in lower skin panels, lower leading edge, and lower trailing edge to provide access to various system components. Access holes and covers are typical for left and right wings.
Wing Plates/Skins Removal/Installation A.
Remove Plates and Covers (1) Index-mark plate or cover to attaching skin or structure and identify plate location to ensure that plate or cover can be installed in same position and location as removed. (2) Remove attaching screws and remove plate or cover. NOTE:
B.
Remove sealed plates under fuel cell carefully to avoid damaging skins.
Install Plates and Covers (1) Select and verify correct cover plate for applicable opening. (2) Position cover plate and install and attach snugly, but not tightly. Torque screws to 20 to 25 inch-pounds. (3) Seal fuel bay cover plates in accordance with Chapter 28.
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MODEL 208 MAINTENANCE MANUAL VORTEX GENERATORS - MAINTENANCE PRACTICES 1.
General A.
2.
Tools and Equipment A.
3.
The model 208B airplane with the TKS installed, uses vortex generators (VG's) to improve air flow on control surfaces. There are ten equally spaced VG's on the upper surface of each wing. This section gives maintenance practices for the wing VG's that includes removal and installation.
For a list of tools and equipment, refer to Wings- General.
Vortex Generator Removal/Installation NOTE:
Removal and installation of the VG'is typical for the left and right wing leading edges.
NOTE:
It is recommended that you replace all used VGs with new VGs when you complete this VG removal/installation procedure.
A.
Remove the Vortex Generator (Refer to Figure 201). (1) Use a toluene solvent cleaning compound to lift the edge of the VG, and remove the VG. (2) When a sufficient piece of the VG is lifted, use hands or pliers and more removal solvent to pull the VG off the wing surface. (3) Discard the used VG. (4) Clean the remaining adhesive from the wing surface with a toluene solvent cleaning compound. (a) Use a cloth in one hand wetted with the toluene solvent and a clean dry cloth in the other hand. (b) Wipe the solvent immediately from the cleaned surface before it evaporates. NOTE: (5)
B.
Cloth should be folder each time surfaced is wiped to present a clean area and avoid redepositing of grease.
Visually inspect the wing for corrosion where the VG was installed. (a) If you find corrosion it must be repaired. Refer to Model 208, Structural Repair Manual, Corrosion - Repair, 51-11-00.
Install the Vortex Generator (Refer to Figure 201). (1) Find the location on the wing where the VG is to be installed.
(2)
(3)
NOTE:
The VG (with the paper backing intact) can be used to outline the area to be prepared.
NOTE:
The vortex generators are installed 27.00 inches center to center, parallel to the wing station, with the peak pointing aft.
To locate the alignment for the points of the VG's do the steps that follow: (a) At W.S.39.00 measure from the top of the porious panel to a point on the upper wing surface 5.75 inches (146.05 mm) Along Contour (A.C.). (b) At W.S.304.00 measure from the top of the porious panel to a point on the upper wing surface 5.31 inches (134.87 mm )Along Contour (A.C.). (c) Snap a line between the two points. (d) Put the peak of each VG at the line, parallel to the wing station. Apply masking tape to the area around the installation surface on the boot. Leave approximately a 0.15 inch, (3.81 mm) distance between the tape and the edges of the VG..
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Vortex Generator Installation Figure 201 (Sheet 1)
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CAUTION: For best results, do not install if temperature is below 50 °F (10 °C). The vortex generators are temperature sensitive and will bond best when then temperature is 50 °F (10 °C) or above. (4) (5) (6) (7) (8)
Apply EC-1300L adhesive to the wing surface where the VG will be installed, and to the VG bond surface. Refer to Chapter 20, Adhesive and Solvent Bonding - Maintenance Practices. Put the VG in the correct position on the wing surface. Mask the VG before painting or sealing. Use the adhesive and apply a fillet seal around the base of the vortex generator. Apply primer and paint to agree with the exterior finish of the airplane. Refer to Chapter 20, Interior and Exterior Finish - Cleaning/Painting.
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MODEL 208 MAINTENANCE MANUAL LEADING EDGE - REMOVAL/INSTALLATION 1.
General A.
2.
The wing leading edge consists of three sections incorporating riveted nose ribs and skins. The leading edge assembly can be removed and replaced as a single unit.
Leading Edge Removal/Installation A. Remove Leading Edge (Refer to Figure 401). (1) Turn off all electrical power. (2) Remove wing tip in accordance with Wings - Removal/Installation, Wing Tip Removal/ Installation. (3) Remove lower cover plates (4) and covers (6). (4) Remove air inlet cover (7) and lift strut upper fairing (8). (5) Remove wing de-ice boot (if installed) in accordance with Chapter 30. (6) Disconnect aileron control system carry-thru cables at turnbuckles above headliner in cabin. (7) Disconnect aileron and spoiler push-pull rods at outboard bellcrank. Refer to Chapter 27. (8) In right wing only, disconnect aileron trim tab cable at quick-disconnects in leading edge root area. (9) In left wing only, disconnect pitot/static lines at wing root connections. Cap open lines to prevent entrance of foreign materials. (10) Remove wires from inboard and outboard fuel level transmitters and disconnect electrical cable connector in root area. (11) Drill out and remove ßush rivets attaching leading edge skins to upper and lower spar caps. (12) Working through lower access openings (5), drill out and remove universal head rivets attaching nose ribs (2) to spar attach brackets (3). (13) Remove leading edge assembly with electrical wire bundle, pitot/static lines, and aileron/aileron trim cables intact. B.
Install Leading Edge (Refer to Figure 401 ). (1) Trial Þt wing leading edge assembly (1) to wing front spar by installing cleco fasteners to temporarily attach leading edge. (2) Working through access holes (5) in lower side of leading edge, rivet nose ribs (2) to nose rib attach brackets (3). Use MS20470AD4 universal head rivets. (3) Rivet leading edge skins to upper and lower spar caps using NAS1097AD5 and NAS1097AD4 rivets (4) Connect electrical wires to outboard and inboard fuel level transmitters. Reconnect electrical cable in wing root area. (5) In left wing only, connect pitot/static lines in wing root area. (6) In right wing only, connect aileron trim cables at quick- disconnects in wing root area of leading edge. (7) Reconnect aileron control system interconnect cables at turnbuckles above headliner in cabin. (8) Connect aileron and spoiler push-pull rods at outboard bellcrank. (9) Check aileron, aileron trim tab, and spoiler surfaces for proper travels and adjust cable tensions per procedures in Chapter 27. Safety turnbuckles and push-pull rod pivot bolts. (10) Install wing de-ice boot (if removed) in accordance with Chapter 30. (11) Install lift strut upper fairing (8) and air inlet cover (7). (12) Replace lower cover plates (4) and access hole covers (6). (13) Replace wing tip in accordance with Wings - Removal/Installation, Wing Tip Removal/ Installation. (14) Turn on electrical power and check electrical components for proper functioning (e.g., fuel quantity system, pitot/static system heaters, stall warning system, landing/taxi lights, and navigation lights).
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Wing Leading Edge Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FLAPS - MAINTENANCE PRACTICES 1.
General A.
2.
Tools, Equipment and Materials A.
3.
4.
The wing trailing edge section consists of riveted ribs, stringers, gussets and skins. A flap is attached to each wing. On the Model 208B and 208 (with TKS installed), four vortex generator boots are attached to the leading edge of the flap assembly. The wing trailing edge structure contains fittings for flap attachment. Fittings are shown in Figure 201
For a list of required tools, equipment and materials, refer to Wings - General.
Flap Removal/Installation A.
Remove Flap Assembly (Refer to Figure 201). (1) Lower the flaps to the fully extended (30 degrees) position. (2) Disconnect the flap push-pull rods by removing the bolts at the inboard and center hinge positions. Keep the spacers. (3) At the outboard hinge position, disconnect the cable by removing the cotter pin and clevis pin. (4) At the outboard hinge position, remove the nut, bolt, and roller from the flap track attachment. (5) At the inboard and center hinge positions, remove the nut, bolt, bushing, and roller, from each forward flap track attachment. (6) Support the flap assembly and remove the nut, bolt, and roller, from each inboard rear and center rear flap track attachment. Remove the flap assembly.
B.
Install the Flap Assembly (Refer to Figure 201.) (1) Position the flap assembly on the wing flap tracks and install the roller, bolt, and nut at each inboard rear and center rear flap track attachment. (2) At the inboard and center hinge positions, install the roller, bushing, bolt, and nut at each forward flap track attachment. (3) At the outboard hinge position, install the roller, bolt, and nut. Connect the cable to the flap by installing the clevis pin and cotter pin. (4) With spacers positioned on each side of the flap push-pull rod end, install the bolt and nut at each inboard and center hinge position. (5) Check the flap control system cable tensions and the flap travels in accordance with Chapter 27. Rig again as required.
Adjustment of Fuselage-Mounted Flap Roller NOTE: A.
This procedure applies to Airplanes 208 and 208B thru 208B0194 Not Incoporating SK208-71.
Procedure (Refer to Figure 201). (1) After the installation of the flaps, check the clearance between the lower surface of the roller and the upper surface of the flap track channel. Do this check at five different flap settings from 30 degrees down. Maintain an 1/8-inch clearance if possible. (2) Do not allow the top of the roller to run above the top of the flange on the flap track channel. It is possible that the 1/5-inch clearance will have to be decreased in order to prevent the top of the roller from running above the top of the flap track channel. (3) Install the spacer as required between the roller support assembly and the fuselage to make sure the roller engages with the flap track properly. NOTE:
5.
If interference occurs with the washer and roller support assembly as a result of the extreme wing spar eccentric adjustment, grind the washer flat on the outside diameter to eliminate the interference.
Flap Vortex Generator Boots Removal/Installation A.
Remove Flap Vortex Generator Boot (Refer to Figure 201 ).
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Flap Installation Figure 201 (Sheet 1)
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Flap Installation Figure 201 (Sheet 2)
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WARNING: Cement and solvent vapors are toxic and extremely flammable. Use only in a well ventilated area away from sparks or vapors. Excess exposure could cause injury or death. If dizziness or nausea occur, obtain fresh air immediately. Avoid contact with skin or eyes. Use solvent resistant gloves to minimize skin exposure. Use safety glasses to minimize chance of eye contact. If eye contact occurs, flush eyes with water for 15 minutes and see a physician. If skin contact occurs, wash thoroughly with soap and water. If swallowed, do not induce vomiting. See a physician immediately. WARNING: Confirm that the aircraft is electrically grounded to prevent static sparks which could ignite solvent vapors. (1) (2) (3) (4) (5) (6) B.
Remove flap. Refer to Flap Removal/Installation. Using a pressure handle squirt can filled with methyl n-propyl ketone and starting at one corner of the upper trailing edge of the boot, apply a minimum amount of methyl n-propyl ketone to the seam line while tension is applied to peel back the corner of the boot. Using methyl n-propyl ketone, separate boot from flap for a distance of four inches all the way along the upper trailing edge. If boot is to be preserved, continue to use methyl n-propyl ketone to soften the adhesion line and pull down and toward the lower trailing edge with uniform tension. Remove installation cement using BF Goodrich KE 9002 paint remover or equivalent. Clean area thoroughly using methyl n-propyl ketone.
Install Flap Vortex Generator Boots (Refer to Figure 201). (1) Using one inch masking tape, mask off area to be covered by Vortex Generator boot. Allow onehalf inch extra on each side. If an adhesion test is to be made, allow one inch on end to install adhesion test strip. (2) Thoroughly clean the metal surfaces with cleaning solvent at least twice. (3) Remove all paint and primer within the masked area. (4) For final cleaning, swab with clean solvent and quickly wipe dry with a clean, dry cloth to avoid leaving a film. (5) Apply GACO N-700-A neoprene coating to the leading edge skin to avoid fuel degradation of the boot or installation cement. (6) Fill gaps of skin splices that lead under boots with GACO N-700-A. (7) Moisten a lint free cloth with cleaning solvent and carefully clean the rough, back surface of the boot at least twice. Change cloths frequently to avoid recontamination of the cleaned areas.
WARNING: EC1300L cement contains methyl n-propyl ketone and is extremely flammable. Extinguish all open flames. Avoid sparks. Use only in well ventilated areas. Avoid prolonged breathing of vapors. Avoid skin contact. (8)
Thoroughly mix the EC1300L installation cement. If necessary, the cement may be thinned with Toluene or methyl n-propyl ketone (up to 5% by volume). (9) Apply one even brush coat to the cleaned back surface of the Vortex Generator boot and to the cleaned installation surface. (10) Allow the cement to dry a minimum of one hour at 50°F (10°C) or above when the relative humidity is less than 75%. If the humidity is 75% to 90%, allow additional drying time. NOTE:
Do not apply the cement if the relative humidity is higher than 90% or if the temperature is below 50°F (10°C).
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MODEL 208 MAINTENANCE MANUAL (11) Snap a chalk line along the centerline of the leading edge of the flap. Intensify chalk line on leading edge with a ball point pen. (12) Stir cement again and apply a second coat to both surfaces and allow to air dry a minimum of one hour. NOTE:
Vortex Generator boot and leading edge may be cemented for a maximum of 48 hours before actual installation if cemented parts are covered and kept clean.
(13) Position boot so its centerline is against the centerline leading edge of the flap. NOTE:
Installation may best be accomplished using two people: one to hold and guide the boot during installation, the other to reactivate the cement and roll the boot down.
(14) Using a clean, lint free cloth dampened with Toluene, reactivate a 3-inch by 18-inch long section of cement on the flap leading edge. (15) Reactivate a matching section on the boot cemented surface. Vortex Generator boot will adhere only where cement is reactivated. (16) When cement is tacky, press boot to flap making sure that centerlines coincide. Then rubber roll boot firmly against flap skin in the tacky area. (17) If boot should attach off course (centerline not coinciding with leading edge centerline), apply methyl n-propyl ketone with a small brush or squirt can to soften the bond line. Apply only a minimum of solvent to bond line while applying sufficient tension to peel back boot. (18) Remove slowly enough to allow solvent to soften cement, thus preventing removal of cement coat or injury to boot. (19) Do not use excess quantities of solvent. (20) To avoid boot damage, avoid twisting, bending boot sharply, or jerking boot loose from bonded area. (21) Allow to dry thoroughly before continuing with application. Reapply cement if any has pulled loose. (22) After boot is fastened in place along its centerline, begin to reactivate cement on either upper or lower surface. (23) Start at inboard end and wipe with Toluene moistened cloth, first along cemented aircraft surface in one direction, and return to start by wiping corresponding cemented surface of boot (approximately three inches wide). Too much wiping will remove cement. (24) Hold boot back to reveal bond line and begin reactivating. (25) Keep moistened cloth tight into fold of bond line of the boot to flap skin. (26) To avoid trapping air, do not allow boot to touch reactivated cement until desired time. (27) Roll down boot with rubber roller, starting at bond line, and roll spanwise while working toward trailing edge. (28) Work carefully to avoid trapping air. Let the roller do the work of mating the two surfaces. NOTE:
If boot lifts after rolling, and cement cobwebs, the cement is too wet. Wait until cement is tacky, and roll again. Keep bond line as straight as possible. This helps to better monitor where the bond line is and eliminate pockets where air can be trapped.
(29) When boot is being installed in a recess, install boot up to edges of recess, then, using a hook knife, trim boot edges and/or ends, as applicable to fit recessed leading edge skin or to butt against adjacent boot or aircraft structure. (30) Rubber roll spanwise over entire surface of boot, applying pressure to ensure good bond. (31) Remove all masking tapes. (32) Apply masking tape to boot edges at trimmed ends or at gaps between sections. (33) Apply GACO N-700-A neoprene coating to protect cut edges of a boot and fair it to adjacent surface. (34) Apply masking tape to boot surface approximately 1/4-inch forward from trailing edges and any trimmed edges. (35) On aircraft surface, apply masking tape approximately 1/8-inch back from area initially cleaned and cemented, forming a neat, straight line.
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MODEL 208 MAINTENANCE MANUAL (36) Apply a heavy brush coat of GACO N-700-A neoprene coating to surfaces between tapes (including trimmed edges). Ensure sealer is continuous. (No voids). (37) Remove masking tape immediately after applying coating (before coating dries). (38) Install flap. Refer to Flap Removal/Installation. C.
Adhesion Test. (1) Using excess boot material, prepare test specimen one inch wide and four or more inches long. (2) Cement specimen to installation surface adjacent to installed boot, following the identical procedure used for boot installation. (3) Leave one inch of the strip uncemented to attach a clamp. (4) Four hours or more after boot installation, attach a spring scale to uncemented end of each strip and measure force required to remove the strip at a rate of one inch per minute. The pull shall be applied 180 degree to the surface. (Strip doubled back on itself). (5) A minimum of five pounds tension (pull) shall be required to remove test strip. NOTE:
Required acceptability of the boot adhesion shall be based on carefully lifting one corner of the boot in question sufficiently to attach a spring clamp and attaching a spring scale to this clamp. Pull with force 180 degree to the surface, and in such a direction that the boot tends to be removed on the diagonal. If a force of five pounds per inch of width can be exerted under these conditions, the installation shall be considered satisfactory. Width increases as corner peels back.
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MODEL 208 MAINTENANCE MANUAL AILERONS - MAINTENANCE PRACTICES 1.
General A.
2.
An aileron is attached to the trailing edge of each wing. Both the right and left ailerons have an aileron trim tab that is attached to the aileron by piano-type hinges.
Aileron Assembly Removal/Installation A.
Remove the Aileron Assembly (Refer to Figure 201). (1) Remove the wing tip. Refer toWing Tip - Removal/Installation. (2) Remove the attach bolt and aileron push-pull rod. (3) Remove the aileron trim-tab-hinge pin retaining screw, the screws, and the cover. (4) Disconnect the aileron trim-tab-push-pull rods at the attach bracket. NOTE: (5)
B.
3.
If the right aileron is to be removed, the push-pull rods are attached to the aileron trim-tab-actuator push-pull rods.
Remove the aileron hinge bolts and remove the aileron. Be careful not to cause damage to the ground strap connector.
Install the Aileron Assembly (Refer to Figure 201). (1) Put the aileron assembly in position until the inboard and outboard hinge holes align with the wing hinge bearings. Install the outboard and inboard aileron bolts into the wing hinge bearings. Install one nut with a cotter pin on the inboard bolt. install safety-wire on the outboard bolt as shown in Chapter 20, Safetying - Maintenance Practices. (2) Connect the aileron trim tab push-pull rods to the attach bracket. (3) Install the cover with the attach screws. (4) Install the trim tab-hinge-pin retaining screw through the loop in the outboard end of the hinge pin. (5) Connect the aileron push-pull rod to the aileron with the attach bolt. Torque the attach bolt as shown in Chapter 20, Torque Data - Maintenance Practices. (6) Install the wing tip. Refer to Wing Tip - Removal/Installation. (7) Examine the aileron and aileron-trim-tab cable tensions and travels in accordance with Chapter 27, Aileron and Spoiler System - Adjustment/Test. Adjust the rigging as necessary.
Aileron Trim Tab Removal/Installation A.
Remove the Aileron Trim Tab (Refer to Figure 201). (1) Remove the wing tip. Refer to Wing Tip - Removal/Installation. (2) Remove the trim tab hinge pin retaining screw, the screws, and the cover. (3) Disconnect the trim tab push-pull rods with the removal of the cotter pins, castellated nuts, and bolts. Keep the bushings. (4) Remove the hinge pin from the hinge and remove the trim tab.
B.
Install the Aileron Trim Tab (Refer to Figure 201). (1) Put the trim tab into the aileron cutout, align the hinge half holes and install the hinge pin. (2) Align the holes in push-pull rod clevises with the holes in the trim tab horns. Install the bushings, bolts, and nuts. Safety the nuts with cotter pins. (3) Install the cover with screws. (4) With the loop of the hinge pin under the head of the retaining screw, tighten the screw. (5) Install the wing tip. Refer to Wing Tip Removal/Installation.
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MODEL 208 MAINTENANCE MANUAL
Aileron Installation Figure 201 (Sheet 1)
57-60-00 © Cessna Aircraft Company
Page 202 Apr 1/2010
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MODEL 208 MAINTENANCE MANUAL
Aileron Installation Figure 201 (Sheet 2)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SPOILERS - MAINTENANCE PRACTICES 1.
General A.
2.
A slot-lip spoiler is attached to the trailing edge of each wing.
Spoiler Removal/Installation A.
Remove Spoiler (Refer to Figure 201). (1) Disconnect the push-pull rod from the spoiler at the outboard hinge by removing the cotter pin, nut, and bolt. (2) Remove the hinge pin retaining screw at the outboard hinge, center hinge, and inboard hinge. (3) Remove the hinge pins, from the outboard hinge, center hinge, and inboard hinge, respectively. (4) Remove the spoiler.
B.
Install the Spoiler (Refer to Figure 201). (1) Position the spoiler near the wing so that the hinge halves mate and the hinge pin holes align at the outboard hinge, center hinge, and inboard hinge. (2) Put the hinge pins into the hinge pin holes, and secure with the screw at all three locations. (3) Connect the push-pull rod to the spoiler at the outboard hinge location. Install the bolt, washer, nut, and cotter pin. (4) Examine the spoiler travel in accordance with Chapter 27. Rig again as required.
57-70-00 © Cessna Aircraft Company
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MODEL 208 MAINTENANCE MANUAL
Spoiler Installation Figure 201 (Sheet 1)
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CHAPTER
PROPELLERS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
61-00-00
Pages 1-2
Dec 1/2006
61-00-01
Pages 101-106
Aug 1/1995
61-10-00
Pages 1-2
Aug 1/1995
61-10-00
Pages 201-212
Jun 3/2002
61-10-00
Pages 501-506
Apr 1/2010
61-10-00
Pages 601-606
Jun 1/2011
61-11-00
Pages 1-3
Mar 1/2012
61-11-00
Pages 201-208
Mar 1/2012
61-11-00
Pages 501-506
Mar 1/2012
61-11-00
Pages 601-607
Mar 1/2012
61-20-00
Pages 201-203
Jun 1/2011
61-40-00
Page 1
Sep 2/1997
61-40-00
Pages 201-203
Mar 1/2000
61-60-00
Page 701
Mar 1/2000
61-Title 61-List of Effective Pages 61-Record of Temporary Revisions 61-Table of Contents 61-List of Tasks
61 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Mar 1/2012
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS PROPELLER - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-00-00 61-00-00 61-00-00 61-00-00
Page 1 Page 1 Page 1 Page 2
PROPELLER - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-00-01 Page 101 61-00-01 Page 101
PROPELLER (HARTZELL) - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-10-00 Page 1 61-10-00 Page 1 61-10-00 Page 1
PROPELLER (HARTZELL) - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment/Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-10-00 Page 201 61-10-00 Page 201 61-10-00 Page 201 61-10-00 Page 208
PROPELLER (HARTZELL) - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing Requirements. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Initial Weight Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Final Weight Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-10-00 Page 501 61-10-00 Page 501 61-10-00 Page 501 61-10-00 Page 501 61-10-00 Page 505
PROPELLER (HARTZELL) - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hartzell Propeller Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-10-00 Page 601 61-10-00 Page 601 61-10-00 Page 601
PROPELLER (MCCAULEY) - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-11-00 Page 1 61-11-00 Page 1 61-11-00 Page 1
PROPELLER (McCAULEY) - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment/Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Blade Damage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Corrosion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Grease or Oil Leakage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-11-00 Page 201 61-11-00 Page 201 61-11-00 Page 201 61-11-00 Page 206 61-11-00 Page 207 61-11-00 Page 207 61-11-00 Page 207
DYNAMIC BALANCING (McCAULEY) - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing Requirements. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Initial Weight Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Final Weight Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-11-00 Page 501 61-11-00 Page 501 61-11-00 Page 501 61-11-00 Page 501 61-11-00 Page 504
PROPELLER (McCAULEY) - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . McCauley Propeller Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-11-00 Page 601 61-11-00 Page 601 61-11-00 Page 601
PROPELLER CONTROL - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Governor Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Overspeed Governor Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Overspeed Governor Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-20-00 Page 201 61-20-00 Page 201 61-20-00 Page 201 61-20-00 Page 201 61-20-00 Page 201 61-20-00 Page 202
PROPELLER BETA INDICATING SYSTEM - DESCRIPTION AND OPERATION . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-40-00 Page 1 61-40-00 Page 1 61-40-00 Page 1
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MODEL 208 MAINTENANCE MANUAL PROPELLER BETA INDICATING SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Beta Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment/Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-40-00 Page 201 61-40-00 Page 201 61-40-00 Page 201 61-40-00 Page 201
COMPOSITE PROPELLER - CLEANING/PAINTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Composite Propeller Painting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61-60-00 Page 701 61-60-00 Page 701 61-60-00 Page 701
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 61-10-00-720
Hartzell Propeller Functional Check
61-10-00 Page 601
61-11-00-720
McCauley Propeller Functional Check
61-11-00 Page 601
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Page 1 of 1 Mar 1/2012
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL PROPELLER - GENERAL 1.
Scope A.
2.
This chapter contains information on the propeller and propeller governors. For speciÞc information on Hartzell or McCauley propellers, refer to the applicable maintenance manual listed in the introduction - List of Vendor Publications.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Polane Black
F63B8
Sherwin-Williams 101 Prospect Ave. Cleveland, OH
To paint propeller, both face and camber.
Catalyst
V66V29
Sherwin- Williams
To harden paint.
Accelerator
V66VB11
Sherwin-Williams
To shorten paint cure time.
Polane Primer Sealer
E65A4
Sherwin-Williams
To paint propeller, both face and camber.
White Plane
Z99WB612
Sherwin-Williams
To paint propeller tip.
Spraylat Copper Lightning Guard
599SA-A8574-1
Sherwin-Williams
To paint propeller.
Polane primer
E65A4
Sherwin-Williams
To paint propeller, both face and camber.
Primer
D61-A-23
Sherwin-Williams
To paint propeller.
Catalyst
V66V27
Sherwin-Williams
To harden paint.
Wrench Extension (5/8×12 point)
AST-2877
Hartzell Propeller 350 Washington Ave. Piqua, OH 45356
To torque propeller retaining bolts.
Beta System Compressor
CT-2834
Hartzell Propeller
To retract beta feedback ring.
Propeller Dome Nut Wrench
B-842
Hartzell Propeller
To torque nut on propeller dome.
Propeller Blade Angle Protractor
C-2820
Hartzell Propeller
To measure blade angle.
Commercially Available
To check propeller beta feedback ring (collar) axial runout.
Dial Indicator
Measuring Device
C-088
Dyer Company P.O. Box 4966 Lancaster, PA
To measure composite blade surface damage.
Composite Blade Repair/Paint Kit
A-2328-3
Hartzell Propeller
To repair composite blade minor surface damage.
Torque Wrench
509006-19
Cessna Aircraft Co.
To tighten nuts.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Torque Wrench Adapter
B-5588
McCauley Propeller Systems PO Box 7704 Wichita, KS 67277-7704
To tighten nuts.
Feedback Collar Retainer
D-5945
McCauley Propeller Systems
To retract propeller feedback collar.
Files
Commercially Available
Propeller damage repair.
120 Grit Sanding Disk (for rotary grinder)
Commercially Available
Propeller damage repair.
600 Grit Sand paper
Commercially Available
Propeller damage repair.
Crocus Cloth
Commercially Available
Propeller damage repair.
Emery Cloth (coarse grain)
Commercially Available
Propeller damage repair.
Rotary Grinder
Commercially Available
Propeller damage repair.
Hinkel Technologies 32100 Stephenson Highway Madison Heights, MI 48071
Treat exposed metal after propeller repair.
Commercially Available
Cleaning.
Commercially Available
To lubricate prop studs, nut threads and spacers.
Alodine
MIL-PRF-83483
Methyl n-Propyl Ketone Lubricant
3.
MIL-G-T-83483
DeÞnition A.
This chapter is divided into sections and subsections to assist maintenance personnel in locating speciÞc systems and information. For locating speciÞc information within the chapter, refer to the Table of Contents at the beginning of the chapter.
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MODEL 208 MAINTENANCE MANUAL PROPELLER - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in propeller system troubleshooting. Refer to Figure 101.
61-00-01 © Cessna Aircraft Company
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Propeller Troubleshooting Chart Figure 101 (Sheet 1)
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Propeller Troubleshooting Chart Figure 101 (Sheet 2)
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Propeller Troubleshooting Chart Figure 101 (Sheet 3)
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Propeller Troubleshooting Chart Figure 101 (Sheet 4)
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Propeller Troubleshooting Chart Figure 101 (Sheet 5)
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MODEL 208 MAINTENANCE MANUAL PROPELLER (HARTZELL) - DESCRIPTION AND OPERATION 1.
General A.
2.
The Hartzell propeller installation consists of a Hartzell Model HC-B3MN-3/M10083 three-bladed, constant-speed, full-feathering, reversible, governor-regulated propeller equipped with composite blades. A propeller control lever on the control quadrant in the cockpit establishes a setting in the propeller governor through a linkage to the engine compartment. This setting (of the governor pilot valve) establishes propeller speed by balancing governor-boosted oil pressure/ßow against a servo piston in the propeller hub with the action of return springs in the hub and centrifugal counter-weights on the blade shanks acting to drive the servo piston in the opposite direction. Since the servo piston is linked to the blades, its position thus governs their setting or blade angle and hence determines propeller speed. Increasing oil pressure against the piston drives the blades toward low pitch (high RPM) and into reverse while the return springs and the counterweights acting against the piston, drives the blades toward high pitch (low RPM) and into feather. The source of propeller system oil is the engine pressure lubrication system boosted to a higher pressure by the propeller governor gear pump.
Description A.
The propeller assembly consists of a hollow steel spider hub which supports three propeller blades and also houses an internal oil pilot tube and feather return springs. Movement of propeller blades is controlled by a hydraulic piston mounted at the front of propeller spider hub. The servo piston is connected by a link to the trailing edge root of each blade. Centrifugal counterweights on each blade and feathering springs in servo piston tend to drive servo piston into the feather or high pitch position. This movement is opposed by the propeller governor oil pressure. The governor oil pressure is applied to servo piston via passages in governor body, an oil transfer tube, and oil transfer housing on propeller shaft, and via the hollow centerbore of propeller shaft and propeller hub. An increase in governor oil pressure moves blades toward low pitch position (increased RPM). A decrease in governor oil pressure allows the blades to move toward high pitch position (decreased RPM) under the inßuence of feathering springs and blade counterweights (Refer to Figure 1).
B.
The servo piston is also connected by three spring-loaded sliding rods to a feedback ring mounted at rear of propeller. Movement of feedback ring is transmitted by a carbon block through the propeller reversing lever to Beta valve on propeller governor. This movement is used to control propeller blade angle from the normal forward low pitch stop to full reverse position.
C.
The high-strength, lightweight composite blades consist of an aluminum blade shank retention section, into which is molded a high-density foam Þller, varying layers of Kevlar material covering blade foam section, and a metal cap molded into blade leading edge. Completing the assembly is a Kevlar Þlament winding which creates blade primary retention. Secondary retention is an integral part of the assembly and is retained by blade clamp of propeller assembly. Blade balance is achieved by the incorporation of a balance tube centrally located in blade foam core and retained by aluminum blade shank. It is essential that the propeller blades be properly maintained in accordance with Hartzell Propeller Products recommended service procedures. These procedures are detailed in Hartzell Composite Blade Inspection and Repair Procedures Manual No. 135-E.
61-10-00 © Cessna Aircraft Company
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Hartzell Propeller Assembly Cutaway Figure 1 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL PROPELLER (HARTZELL) - MAINTENANCE PRACTICES 1.
General A.
2.
Maintenance practices for the propeller consist of removal, installation, adjustment/tests, and blade clamp lubrication. Adjustment/checks include the following: Beta feedback ring axial runout check, feather blade angle check and adjustment, and low pitch stop check and adjustment.
Propeller Removal/Installation A.
Remove Propeller (Refer to Figure 201). (1) Make sure all electrical power switches are in OFF position. (2) Open upper right cowling door and remove right nose cap. (3) Remove propeller reversing lever (1) and carbon block (3) from propeller Beta feed back ring (4). Refer to Pratt & Whitney Engine Maintenance Manual for removing the propeller reversing lever. (4) Remove spinner (15) by removing screws and Þber washers (5) securing spinner to spinner bulkhead (22). NOTE:
Mark an index on spinner and spinner bulkhead to ensure reinstallation of spinner as originally installed.
CAUTION: Do not use masking or other adhesive tapes to secure brushes as adhesive will degrade conductivity. (5)
If propeller has electric anti-ice system installed, loosen nuts securing propeller anti-ice system brush block assembly (32) to engine reduction gearbox (33) and carefully insert a length of safety-wire between brushes and slip ring (31). Tie safety-wire around brush holder to secure brushes in holder and remove brush block, brushes, and bracket.
CAUTION: Ensure beta system compressor tool is not cocked. Do not forcibly pull the feedback ring against the guide which limits forward travel. Position Beta system compressor tool at forward portion of propeller. Attach tool ßanges to rod end ring of propeller servo piston (16). Tighten tool until propeller feedback ring (4) is pulled forward to allow access to 12 point bolts (25). (7) Attach lifting sling to hoist and position hoist forward of airplane. Attach sling to propeller by positioning blades at 10 o'clock and 2 o'clock positions. (8) Position drip pan under propeller to catch residual oil which will drain from propeller when removed. (9) Cut safety-wire from bolts (25) and using proper wrench extension, loosen and remove bolts. (10) With propeller supported in sling, remove propeller from airplane and O-ring (19) from propeller shaft. Discard O-ring. (11) To remove spinner bulkhead (22), loosen jam nuts (26) and using ßats of low stop rods (17), turn low stop rods one-third turn each in sequence to evenly back rods out of Beta follow-up ring (4).
(6)
NOTE: B.
If any single low stop rod is turned more than one-third turn at a time, remaining two stop guides will bind up in follow up ring.
Install Propeller (Refer to Figure 201 and Figure 202). (1) If propeller is equipped with electric anti-ice system, install slip ring (31). (2) Install spinner bulkhead (if removed) by installing guide lugs (28), spinner bulkhead support (21), retaining rings (27), and spinner bulkhead (22) over low pitch stop rods (17) at back of propeller. Install bolts (29).
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Hartzell Propeller Installation Figure 201 (Sheet 1)
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Hartzell Propeller Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Hartzell Propeller Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL
Hartzell Propeller Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL
Propeller Mounting Bolt Torque Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (3)
Place jam nuts (26) onto low pitch stop rods. Position Beta follow-up ring (4) to low pitch stop rods (17) and using ßats on stop rods thread into follow-up ring one-third turn at a time on each stop rod. NOTE:
(4) (5)
If any single low stop rod is turned more than one-third turn at a time, remaining two stop guides will bind up in follow up ring.
Measure distance from forward side of follow-up ring (4) to aft side of spinner bulkhead (22) at each low pitch stop rod location to ensure follow-up ring alignment. Adjust stop rods as required to make all three dimensions equal and tighten jam nuts (26). Install Beta system compressor tool on propeller.
CAUTION: Ensure beta system compressor tool is not cocked. Do not forcibly pull feedback ring against guide which limits forward travel. (6) (7) (8) (9) (10) (11) (12) (13)
Operate Beta system compressor tool to pull feedback ring forward while observing for smooth movement without binding or interference. Leave feedback ring pulled forward during installation for access to engine propeller ßange (23). Lightly lubricate O-ring (19) with engine oil and install on engine propeller ßange (23). Lift propeller into position at front of airplane using hoisting sling and hoist. Lubricate threads and bolt washer face of 12-point bolts (25) with Hartzell A-3338-1 lubricant or equivalent lubricant conforming to AMS 2518 or MIL-T-83483. Very carefully position propeller onto engine shaft, using extreme caution to avoid damaging feedback ring (4). Install washers (24) onto bolts (25) with chamfered ID of washer under bolt head. Install bolts (25) and torque to 40 foot-pounds using torque sequence A shown in Figure 202. Repeat sequence A but torque to 80 foot-pounds. Final torque all bolts from 100 to 105 footpounds using sequence B. Safety-wire bolts. Remove Beta system compressor tool. Check that carbon block (3) will slide freely in groove of Beta feedback ring (4) at all points without binding or excessive friction. NOTE:
The carbon block initially supplied with each propeller has been preÞt. If a different carbon block is being installed, it may be necessary to sand it to obtain a total clearance between carbon block and side of groove of 0.001 to 0.002 inch at the tightest point.
(14) Install carbon block (3) onto reversing lever (1). NOTE:
The lower end of the propeller reversing lever is machined with a stepped notch.
CAUTION: Make sure the stepped notch at the end of the propeller reversing lever (1) is under the guide pin (16) in the reversing lever guide pin bracket (15). (15) Install reversing lever (1) to Beta valve clevis (2) and follow-up ring (4). Refer to Pratt & Whitney Engine Maintenance Manual for installing the propeller reversing lever. (16) Position spinner (15) to spinner bulkhead (22) as indexed during removal procedure and secure with screws and Þber washers (5). (17) Install right nose cap. (18) Check clearance between spinner (15) and nose cap, clearance should be 0.32 inch, +0.10 or -0.10 inch.
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MODEL 208 MAINTENANCE MANUAL
3.
Adjustment/Checks A.
Beta Feedback Ring Axial Runout Check. NOTE: (1) (2) (3) (4)
Checking adjustment of the Beta feedback ring axial runout is not required unless there is reason to believe linkage settings have been tampered with or feedback ring is bent.
Open right upper cowling door. Remove right nosecap. Clamp dial indicator in position to check axial runout of forward inside face of Beta feedback ring groove. Rotate propeller by hand and check that axial runout does not exceed 0.010 inch total indicator reading and that there is no binding between carbon block and feedback ring. NOTE:
(5)
If Beta feedback ring runout is excessive proceed as follows: (a) Index mark location of spinner to spinner bulkhead and remove spinner. (b) Mark one of the three low pitch stop rods. Do not change setting on this rod. (c) On the other two rods, loosen jam-nuts at Beta feedback ring and jam-nuts aft of rod end ring. (d) Adjust runout by carefully screwing the two rods into or out of Beta feedback ring using ßats on the rods. (e) When feedback ring axial runout is within tolerance, torque jam-nuts at feedback ring to 180 inch-pounds, +18 or -15 inch-pounds.
(f) (g) (h) (i) (j) B.
The carbon block initially supplied with each propeller has been preÞt. If a different carbon block is being installed, it may be necessary to sand it to obtain a total clearance between carbon block and side of groove of 0.001 to 0.002 inch at the tightest point.
NOTE:
Inability to obtain satisfactory runout adjustment is reason to suspect feedback ring is bent or warped which can usually be conÞrmed by visual inspection. If this is the case, propeller must be removed and repaired on a propeller test stand in accordance with Hartzell Turbine Propeller Overhaul Instructions Manual No. 118-E or returned to Hartzell for repair.
NOTE:
Rotation of the low pitch stop rods will not cause low pitch stop nut adjustment to change. Do not change position of stop nuts.
On rod that was marked and not adjusted, measure distance between low stop nut and propeller piston boss with a precision measuring instrument, such as an inside micrometer or vernier or dial calipers. Adjust other two low stop nuts to match. Adjust elastic stop nuts forward of rod end ring so that ring is an equal distance from the ends of all three rods. Torque jam-nuts aft of ring to 180 inch-pounds, +18 or -18 inchpounds. Remove dial indicator. Install spinner as indexed in removal procedure. Install right nose cap.
Feather Blade Angle Check and Adjustment (Refer to Figure 203). (1) Position airplane out of the wind. NOTE: (2)
Airplane must remain in a stable position throughout this procedure.
Remove spinner by removing screws and Þber washers securing spinner to spinner bulkhead. NOTE:
Mark an index on spinner and spinner bulkhead to ensure reinstallation of spinner as originally installed.
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MODEL 208 MAINTENANCE MANUAL (3)
(4) (5) (6) (7)
Zero propeller protractor on propeller servo piston by using a parallel bar. Recheck by turning parallel bar over. If readings are different, parallel bar is not parallel and rod end ring should be removed to zero the propeller protractor directly on the servo piston. This method should be used if a parallel bar is not available. Rotate each blade to a horizontal position and measure each feather blade angle on the back of blade at the 42-inch station. The average of the three individual blade angles must be 78.4 degrees, +0.2 or -0.2 degrees. If adjustment is not required, replace rod end ring if it was removed, and torque jam-nuts to 180 inch-pounds, +18 or -18 inch-pounds. Install spinner using Þber washers and screws as indexed during removal. If adjustment of the feather blade angles is required, proceed as follows: (a) Place drip pan under propeller to catch residual oil. (b) Using precision measuring instrument, such as inside micrometer or dial or vernier calipers, measure distance between one low pitch stop nut and servo piston boss, and record dimension. (c) Remove forward rod end ring (7) and low pitch stop nuts (8) from all three low pitch stop rods (6). (d) Remove Flexlock nut (4) from servo piston (5). Remove link pin units (9) from side of servo piston (5). (e) Remove piston (5) from cylinder. (f) Remove safety-wire from the four feather adjustment screws (3). (g) Equally adjust the four feather adjustment screws (3) on front of spring cup (2) to provide required feather angle. One turn equals 1.5 degrees of blade angle change with blade angle increasing when screws are turned in and decreasing when screws are turned out. Record number and direction of turns to nearest eighth turn. (h) Safety-wire feather adjustment screws (3). (i) Slide piston (5) onto cylinder. (j) Install link arms (10) into slots on piston (5) and install link pin units (9). (k) Install Flexlock nut (4) and torque to 120 foot-pounds, +12 or -12 foot-pounds. Ensure that piston does not rotate against low pitch stop rods (6) and cause binding. (l) Recheck blade feather angles per steps B.(4) and B.(5). If angles are not satisfactory, readjust per steps B.(7)(d) through (k). (m) Reinstall low pitch stop nuts (8). Adjust nuts using a precision measuring instrument so that distance between low pitch stop nuts and piston boss is equal to original distance per step B.(7)(b), corrected for feather adjustment as follows: Each turn IN of feather adjustment screws, ADD 0.031 inch to original distance. Each turn OUT of feather screws, SUBTRACT 0.031 inch from original distance.
EXAMPLE + 2.683 inches
Original stop nut distance = First feather adjustment correction:
+ 0.043 inch
1 3/8 turns IN x 0.031 inch = Second feather adjustment correction: 1/2 turn OUT x 0.031 inch =
- 0.016 inch
Corrected low pitch stop nut distance:
+ 2.710 inches
(n) (o) C.
Reinstall forward rod end ring (7). Adjust three elastic nuts forward of the ring so ring is an equal distance from end of all three rods. Torque three jam-nuts to 180 inch-pounds, +18 or -18 inch-pounds. Install spinner as indexed during removal and secure with screws and Þber washers.
Low Pitch Stop Check and Adjustment (Refer to Figure 203). (1) Position airplane out of wind. NOTE:
Airplane must remain in a stable position throughout this procedure.
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MODEL 208 MAINTENANCE MANUAL
Feather and Low Pitch Blade Angle Adjustment Cutaway Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2)
Remove spinner by removing screws and Þber washers securing spinner to spinner bulkhead. NOTE:
(3)
(4) (5) (6)
(7)
(8) (9)
Mark an index on spinner and spinner bulkhead to ensure reinstallation of spinner as originally installed.
Measure and record distance between piston bosses and low pitch stop nuts using a precision measuring instrument, such as an inside micrometer or dial or vernier calipers. All three distances should be equal unless tampering has occurred. If not equal, adjust any two nuts so that distances are equal to the third. Place drip pan under propeller to catch residual oil which is lost during next step. Remove Flexlock nut (4) from front of piston (5). Grasp counterweights and pull forward to rotate blades to a lower angle. Pull alternate counterweights to allow piston to slide forward as evenly as possible without excessive cocking. Pull forward until piston bosses are Þrmly and squarely in contact with all three low stop nuts, but feedback linkage is not pulled forward. Zero propeller protractor on horizontal portion of servo piston, except that use of parallel bar or removal of rod end retaining ring is not required. Check zero as propeller is rotated to three equally spaced positions. Different readings indicate that servo piston is excessively cocked due to not being Þrmly and squarely in contact with all three low pitch stop nuts per step 6. Rotate each blade to a horizontal position and measure each low pitch blade angle on back of blade at 42-inch station. Record each angle. The average of the three individual blade angles must be 9.0 degrees, +0.5 or -0.5 degrees. NOTE:
The propeller low pitch blade angle speciÞcation is 9.0 degrees, +0.1 or -0.1 degree; however, this tight tolerance can only be achieved accurately with propeller installed on a test bench. Do not attempt adjustment of an installed propeller if blade angle is within the +0.5 or -0.5 degree tolerance.
(10) Rotate blades back to feathered position using caution when re-engaging the hole in front of piston with threaded pilot. (11) Install Flexlock nut (4) and torque to 120 foot-pounds, +12 or -12 foot-pounds. Ensure that piston (5) does not rotate against low pitch stop rods (6) causing binding.
CAUTION: If there is any doubt as to how to measure, calculate, or adjust required low pitch stop nut distance, seek assistance before attempting adjustment. (12) If adjustment of low pitch blade angle is required, adjust the pitch stop nuts so that distance from piston bosses to low pitch stop nuts is equal to original recorded distance per step C(3) corrected for required blade angle change as follows: (a) If measured blade angle is too LOW, SUBTRACT 0.035 inches for each degree of difference between 9.0 degrees and measured blade angle. (b) If measured blade angle is too HIGH, ADD 0.035 inches for each degree of difference between measured blade angle and 9.0 degrees. EXAMPLE 1: Blade angle too LOW - Subtract correction Measured blade angle: (7.8 degrees) Original low stop nut distance:
2.683 inches
Correction: (9.0-7.8) x 0.035 inch =
-0.042 inch
Corrected low stop nut distance =
2.641 inches
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MODEL 208 MAINTENANCE MANUAL EXAMPLE 2: Blade angle too HIGH - Add correction Measured blade angle: (9.8 degrees) Original low stop nut distance:
2.683 inches
Correction: (9.8-9.0) x 0.035 inch =
+0.028 inch
Corrected low stop nut distance =
2.711 inches
D.
Propeller Blade Clamp Lubrication. (1) The propeller manufacturer recommends that propeller blade clamps be lubricated each 100 hours. Blade clamps are greased through zerk Þttings (two each per blade clamp). Lubricate as follows:
CAUTION: Care must be taken to avoid blowing out blade clamp gaskets. This is accomplished by removing one of two zerks of each blade clamp. (2) (3) (4) (5) (6)
Remove spinner. Remove one each of the two zerk Þttings from each blade clamp. Using grease conforming to speciÞcation shown in Chapter 12 and a pressure type grease gun, pump grease into zerk Þtting not removed from each blade clamp. Stop pumping when new grease ßows from hole where zerk was removed. Replace removed zerk Þtting. Reinstall spinner.
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MODEL 208 MAINTENANCE MANUAL PROPELLER (HARTZELL) - ADJUSTMENT/TEST 1.
2.
General A.
This section gives information necessary to do a functional test for the dynamic balancing of the propeller. This data is applicable to all Hartzell HB-B3MN-3 series propellers installed on the Model 208.
B.
The correctable propeller imbalance is a result of variations in the respective component weight and installation. The installation of balance weights in the correct position is an effective method to reduce the effects of these physical variations.
C.
The unknown factors in this problem are the amount and location of weight to be added. The balancer equipment will indicate the amount of imbalance in velocity (inches per second - IPS). This is translated into the amount of weight to be added and the angular location of the weight. The amount and location of weight added to the propeller is found with the balancer equipment. Refer to the manufacturer manual provided with the equipment for the balancing procedures.
D.
Applicable Documents. (1) Applicable Caravan l Pilot's Operating Handbook. (2) Applicable Caravan l Pilot's Checklist.
Balancing Requirements A.
Balancing Requirements for the Dynamic Balancing and Functional Test Procedure (Refer to Figure 501). (1) Balance the Hartzell HB-B3MN-3 series propellers installed on the Model 208 after all the engine rigging is satisfactorily completed. Refer to Engine Control Rigging - Adjustment and Test.
(2) (3) (4) (5)
NOTE:
All propellers balanced by this procedure must have logged a minimum of three hours of installed running time to adequately seat the seals and distribute the lubricating grease.
NOTE:
The maximum allowable vibration for Hartzell HB-B3MN-3 series propellers is 0.07 IPS. If the vibration value is greater than 2.00 IPS, the propeller must be rejected for excessive unbalance.
Remove all the dynamic balance weights from the propeller. Remove the propeller shaft oil seal cavity drain plug. Install the adapter and vibration sensor in the propeller shaft oil seal drain port. Position the airplane into the wind and away from buildings and blast fences. NOTE:
(6) 3.
Do not balance the propellers when it is raining or when the wind gusts are 5 knots over any prevailing, steady wind.
Use the Aces equipment, or equivalent, to dynamically balance the propeller. Refer to the manufacturer manual provided with the equipment for the balancing procedures.
Initial Weight Installation A.
Install the Initial Weight (Refer to Figure 502). (1) Use the propeller balancing equipment to find the initial weight installation location. Refer to the manufacturer manual provided with the equipment for the balancing procedures. (2) Remove the screw from the spinner identified for the initial weight location. NOTE:
For initial weight installation, the weights can be installed on the outside of the spinner. The initial weight installation uses MS24694-XX countersunk screws in the holes located between the nutplates.
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MODEL 208 MAINTENANCE MANUAL
Vibration Sensor Figure 501 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Balance Weight Location Figure 502 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Balance Weight Location Figure 502 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (3)
Find the number of AN960-10 washers necessary to equal the initial correction weight found in the dynamic propeller balancing procedure plus the weight of the MS24694 screw removed from the spinner. Refer to the manufacturer manual provided with the equipment for the balancing procedures. (a) Use the scale listed in the manufacturer manual to weigh the hardware used for the initial weight installation. NOTE:
(4) (5) (6) 4.
The total weight of the washers and screw must not exceed 25.5 grams at any one location. If the necessary weight exceeds 25.5 grams, it must be equally distributed between adjacent locations.
Change the number of washers to adjust for the MS27039-1-XX screw necessary to attach the initial weights on the outside of the spinner. Install the MS27039-1-XX screw and AN970-3/AN960-10 washers in the spinner attachment screw hole. Use the propeller balancing equipment to find the weight and angle correction. Refer to the manufacturer manual provided with the equipment for the balancing procedures.
Final Weight Installation A.
Install the Final Weight (Refer to Figure 502 and Figure 503). NOTE:
(1) (2) (3)
For final weight installation, the spinner must be removed. The final weights are installed on the inner side of the spinner bulkhead. The final weight installation uses MS24694-XX countersunk screws in the holes between the nutplates.
Remove the MS27039-1-XX screw and washers installed during the initial weight installation. Remove the screws and fiber washers from the spinner. Put index marks on the spinner and the spinner bulkhead. NOTE:
(4) (5) (6)
Index marks are used to make sure the spinner is installed on the spinner bulkhead in its initial position.
Remove the spinner from the airplane. Change the total weight of the removed washers and MS27039-1-XX screw, to adjust for the weight of the removed MS24694 screw, and the attaching MS24694-XX screw and MS21044-N3 nut. Install the final weight hardware inside the spinner bulkhead in the countersunk hole nearest to the location found in the weight and angle correction. NOTE:
The total weight of washers, screw and nut must not exceed 25.5 grams at any one location. If the necessary weight exceeds 25.5 grams, it must be equally located between adjacent locations.
(7) (8) (9) (10)
Put the spinner in position against the spinner bulkhead. Use the index marks to make sure the spinner is in its initial position. Install the screws and fiber washers in the spinner bulkhead. After the final installation of the weights, do a test of the propeller balance to make sure the propeller is within permitted balance limits. Refer to the manufacturer manual provided with the equipment for the balancing procedures. (11) Remove the adapter and vibration sensor from the propeller shaft oil seal drain port. (12) Install the propeller shaft oil seal cavity drain plug.
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MODEL 208 MAINTENANCE MANUAL
Hartzell Spinner Figure 503 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PROPELLER (HARTZELL) - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the Hartzell propeller in a serviceable condition. NOTE:
For different views of the propeller and the spinner installation that are not included in this section, refer to Figure 201, in Propeller (Hartzell) - Maintenance Practices.
Task 61-10-00-720 2.
Hartzell Propeller Functional Check A.
General (1) This section gives the information needed to do the functional check of the Hartzell propeller.
B.
Special Tools (1) Mild Soap and Water. (2) Age Master No. 1. (3) ICEX. (4) Stoddard Solvent or equivalent. (5) Isopropyl Alcohol.
C.
Access NOTE: (1) (2)
D.
The propeller spinner is removed after the propeller is washed for the inspection.
Remove the nose cap to get access to the propeller governor. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Remove the upper left cowling door to get access to the overspeed governor. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
Do the Hartzell Propeller Detailed Inspection.
CAUTION: Moisture of any type must never touch exposed Kevlar composite material. (1)
(2) (3) (4) (5) (6) E.
Examine the propeller blades for any openings in the Kevlar composite material before you wash the blades. (a) If you find openings in the Kevlar composite material, apply paint to the exposed areas. Refer to Composite Propeller - Cleaning/Painting. Wash the propeller blades and the boots with mild soap and water before you start the inspection. (a) Do not let the soap solution come into contact with the blade clamps. Put a mark on the spinner and the bulkhead to record the alignment for the next installation. (a) Do not use a lead pencil. Remove the propeller spinner. Refer to Propeller (Hartzell) - Maintenance Practices. Be careful to not remove the spinner index mark when you clean the spinner and the bulkhead. (a) Clean the spinner and the bulkhead with Stoddard solvent to remove all grease before you start the inspection. Clean the slip ring and the deice brush block with isopropyl alcohol, Stoddard solvent, or equivalent.
Examine the Spinner and the Bulkhead. (1) Examine the accessible surface of the bulkhead and the inner and outer spinner surface for condition, cracks, corrosion, and fractures. (2) Examine the spinner bulkhead, spinner bulkhead support, spinner attach screws, and spinner attach nutplates for condition, corrosion, and wear. (3) Examine the attach holes in the spinner for cracks and hole elongation.
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MODEL 208 MAINTENANCE MANUAL (4) (5)
(6) (7) F.
Examine the balance weights for condition, corrosion, security, and correct installation. Refer to Final Weight Installation found in Propeller (Hartzell) - Adjustment/Test. Visually examine the spinner dome surface and the bulkhead for burned spots, pits, or other signs of a lightning strike. (a) If there are signs of a lightning strike, refer to Chapter 5, Unscheduled Maintenance Checks, Lightning Strike. Examine the attach screws for condition. Make sure that there is a fiber washer installed on each attach screw. Examine the viewable area of the engine propeller shaft seal just aft of the spinner bulkhead for leaks and condition.
Examine the Blades. (1) Examine all blades for condition, gouges, scratches, leading edge looseness, erosion, debonds, delaminations, cracks, and exposed composite materials. (2) If installed, examine the anti-ice boots for abrasions, exposed heating elements, cuts, nicks, and security of attachment. Refer to Chapter 30, Propeller Anti-Ice - Maintenance Practices, Figure 201. (a) Examine the wiring from the boots to the terminal strips on the spinner bulkhead for condition, chafing, correct routing, and security of attachment at all clamps. (b) Examine the connector between the boot and the wire harness for security of attachment. (c) Examine the wire harness connectors at the terminal strips for condition and security of attachment. (d) Examine the boot edge dressing for condition. 1 If necessary, touch-up damaged or exposed areas. (3) Examine the terminal strips for condition and security of attachment to spinner bulkhead. (4) Use the “Coin Tap” procedure to examine for debond damage adjacent to any crack in the paint between the erosion shield and the composite material. Refer to the Hartzell Propeller Owners Manual 146, Maintenance Practices. NOTE:
(5)
Paint cracks can occur along the line at which the erosion shield contacts the blade surface. Any crack in the paint of a composite blade finish is considered minor damage. Circumferential cracks can occur in the paint and the resin on the primary retention windings because of resin build-up during manufacture. Refer to the Hartzell Propeller Owners Manual.
Examine the blades and the blade clamps for condition, cracks, corrosion, evidence of lightning strikes, and security. Make sure that all hardware is correctly safetied. NOTE:
Lightning strike damage normally shows by burned spots on the blade clamps and the leading and trailing edges of the blades.
(a)
If there are signs of a lightning strike, refer to Chapter 5, Unscheduled Maintenance Checks, Lightning Strike. (6) Examine all blade clamp counter weights for condition and security. (7) Examine all blade clamp static balance weights (if installed) for condition and security. (8) Examine the red alignment marks on the blades and the clamps to make sure that the blades have not slipped in the blade clamps. (9) Move the counterweights back and forth to examine if there is freedom of blade movement on the hub pilot tube. (10) If the blade(s) are possibly tight (will not turn slightly), remove the link arm(s) from servo piston and turn each blade individually with your hand.
CAUTION: Make sure that you do not scratch or damage the link arms. NOTE:
Examine the blades for play. Radial play must not be more than 0.5 degrees. End play and fore and aft movement cannot be more than +/- 0.06 (1.5 mm).
(11) If the blade(s) are tight, rough, or binding, return the propeller to an approved repair facility.
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MODEL 208 MAINTENANCE MANUAL (12) If the blades are serviceable, connect the link arms. (13) Lubricate the blade clamps. Refer to Chapter 12, Propeller (Hartzell) - Servicing. G.
Examine the Hub (Refer to Figure 601).
CAUTION: Oil leaks from the propeller or the engine can get on the wing, wing struts, and/or the horizontal stabilizer deice boots and cause damage. (1) (2) (3) (4)
Examine the exposed area for condition, cracks, corrosion, and security of the components to the hub. Visually examine all three link arms for condition and security. Examine the hub servo piston and the blade clamps for oil and grease leaks. Visually examine the propeller for security of installation. NOTE:
(5)
Examine the exterior area of the servo piston for condition, corrosion, and security of the flex lock nut. NOTE:
H.
If the safety wire installation is correct, the propeller is secured.
The flex lock nut installation is correct if the torque putty on nut and the shaft is not broken.
Examine the Beta Feedback Ring (Refer to Figure 601). (1) Examine the feedback ring for condition, corrosion, and security of installation. (2) Clean the feedback ring and the brush holder with isopropyl alcohol, Stoddard solvent, or equivalent. (3) Examine the wires between the feedback ring and the terminal strip for condition, chafing, and security. (4) Examine the reversing lever for condition and security. (5) Examine the reversing lever for free play. (a) If there is free play at the beta valve, remove the bolt and examine the sleeve bushing for signs of wear at the attach location. 1 If there is wear, replace the bushing. (b) If the lever has free play at the beta cable clevis, remove bolt at the clevis and examine the sleeve spacer for signs of wear at the attach point. 1 If there is wear, replace the spacer. (6) Examine the carbon brush for wear and signs of damage. (7) To examine the carbon brush for wear, do the steps that follow: (a) Hold the carbon brush against the feedback ring. (b) Turn the feedback ring and measure the clearance between the carbon brush and the feedback ring around the full circumference of the feedback ring. NOTE:
The clearance between the brush and the feedback ring must not be more than 0.010 inch (0.254 mm) at any area around the full circumference of the ring.
CAUTION: Do not turn the elastic low pitch stop nuts installed on the low pitch stop rods. (8)
Examine the low pitch stop rods (3 each) for condition and security. (a) The low pitch stop rod locknuts are installed correctly if the torque putty has not been disturbed. (9) Examine the varistor installed near the center at the top of the forward side of the firewall for condition and security of installation. (10) Examine the electrical connections for condition, routing, signs of chaffing, and security.
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Propeller and Spinner Installation - Hartzell Figure 601 (Sheet 1)
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Propeller and Spinner Installation - Hartzell Figure 601 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (11) Discoloration of the varistor or the electrical leads, or a failure can be a sign that a lightning strike has occurred. (a) If you think that there was a lightning strike to the airplane, refer to Chapter 5, Unscheduled Maintenance Checks, Lightning Strike. (12) Apply Age Master No. 1 to the de-ice boots in accordance with the manufactures recommendations. Refer to Chapter 30, Pneumatic Surface Deice - Maintenance Practices. (13) If the operating conditions make it necessary, apply ICEX II to the boots. Refer to Chapter 30, Pneumatic Surface Deice - Maintenance Practices. I.
Examine the Propeller Governor (Refer to Figure 601). (1) Examine the propeller governor for condition, oil leaks, and security. (2) Examine the speed adjuster return spring for condition and security. (3) Examine the air bleed link for corrosion, condition, and security. (4) Examine the governor interconnecting rod for corrosion, condition, security and wear. (a) Make sure that the rod end bearings turn freely and do not bind. (5) Examine all hardware for corrosion, condition, and correct safety. NOTE:
It is not necessary to safety wire the four self locking mounting nuts.
J.
Examine the Propeller Cable Terminal Rod End (Refer to Chapter 76, PT6A-114/-114A Engine Rigging - Adjustment/Test, Figure 510). (1) Disconnect the rod end from the propeller speed adjusting lever. Refer to Chapter 76, Propeller Control - Maintenance Practices. (2) Wipe the rod end clean using a clean lint-free cloth. (3) Examine the rod end for corrosion, pitting, and cleanliness. (4) Lubricate the rod end ball with MIL-L-7870. (5) Connect the rod end to the adjusting lever. Refer to Chapter 76, Propeller Control - Maintenance Practices.
K.
Examine the Overspeed Governor (Refer to Propeller Control - Maintenance Practices, Figure 201). (1) Examine the overspeed governor for condition, oil leaks, and security. (a) Make sure that the hardware is safety wired except for the four self-locking attach nuts. (2) Examine the electrical wiring and the electrical connector at the governor reset test solenoid for signs of damage, correct wire routing, and security. (3) Examine the governor reset test solenoid for condition and security. (4) Install the upper left cowling door. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. (5) Install the nose cap. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (6) Install the propeller spinner. Refer to Propeller (Hartzell) - Maintenance Practices. (7) Do the Propeller Overspeed Governor Functional Check. Refer to Propeller Control Maintenance Practices.
L.
Restore Access NOTE:
(1) End of task
The propeller spinner, nose cap, and upper left cowling door were installed before the functional check.
None
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MODEL 208 MAINTENANCE MANUAL PROPELLER (MCCAULEY) - DESCRIPTION AND OPERATION 1.
General A.
2.
Airplanes 20800189 and On, 208B0218 and On, and airplanes that have incorporated CAB 90-20, are equipped with a McCauley Model 3GFR34C703/106GA-0 three-bladed, constant-speed, fullfeathering, reversible, governor-regulated propeller.
Description and Operation A.
A propeller governor setting is established by a propeller control lever located on the cockpit control quadrant through linkage to engine compartment. This governor pilot valve setting establishes propeller speed by balancing governor-boosted oil pressure/ßow against a propeller hub servo piston with the action of return springs in the hub and centrifugal counterweights on blade shanks acting to drive servo piston in opposite direction. Since the servo piston is linked to the blades, its position thus governs their setting or blade angle and hence determines propeller speed. Increasing oil pressure against the piston drives the blades toward low pitch (high RPM) and into reverse while the return springs and counterweights, acting against the piston, drives the blades toward high pitch (low RPM) and into feather. Source of propeller system oil is engine pressure lubrication system boosted to a higher pressure by the propeller governor gear pump. Refer to Figure 1. NOTE:
For information and procedures not contained in this Chapter, refer to the McCauley MPC700 Propeller Overhaul Manual, MPC-26 McCauley Owner/Operator Information Manual, and McCauley Technical Report No. 722. (refer to List of Vendor Publications)
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McCauley Propeller Assembly Cutaway Figure 1 (Sheet 1)
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McCauley Propeller Assembly Cutaway Figure 1 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL PROPELLER (McCAULEY) - MAINTENANCE PRACTICES 1.
General A.
2.
Maintenance practices for the propeller consists of removal, installation and adjustment/test. Adjustment/test include the beta feedback collar axial runout check.
Propeller Removal/Installation A.
Remove Propeller (Refer to Figure 201).
CAUTION: Do not forcibly pull the feedback collar against the guide which limits the forward travel. (1) (2) (3) (4) (5) (6) (7)
(8) (9) (10) (11) (12) (13)
B.
Ensure airplane electrical power is OFF. Open upper right hand cowling and remove right nose cap. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Disconnect propeller reversing lever (26) from control cable and beta valve clevis (25). Remove propeller reversing lever (26) and carbon block (24) from propeller feedback collar (23). Refer to Pratt & Whitney Engine Maintenance Manual for removing the propeller reversing lever. Remove spinner (17) by removing screws (19) and Þber washers (18). If propeller anti-ice is installed, loosen nuts securing anti-ice brush block assembly (22) and carefully insert an electrical tie strap between brushes and slip ring (15). Secure brushes in holder and remove brush block assembly and bracket. Install assembly tool, D-5945, on forward end of beta rods (5). NOTE:
Do not disturb beta rod nuts. The position of beta rod nuts, with respect to the beta rod, determines the low pitch setting.
NOTE:
Adjustment of low pitch setting may only be performed by an approved propeller repair station.
Pull beta rods forward until roll pin on tool bottoms against the plate. This will position feedback collar forward to make mounting nuts accessible. Attach lifting sling to hoist and position hoist forward of the airplane. Attach sling to propeller by positioning blades at 10 O'clock and 2 O'clock. Position drip pan under propeller to catch residual oil which will drain from the propeller when removed. Remove mounting nuts (9) and spacers (8). With propeller supported by the sling, remove propeller from engine ßange (10). If removal of the spinner bulkhead (13) is required on standard propeller, remove screws (11) and washers (12) and remove bulkhead. On propeller with anti-ice installation, remove screws securing anti-ice leads (21) to slip ring (15) and screws securing lead clamps to bulkhead.
Install Propeller (Refer to Figure 201). (1) Ensure airplane electrical power is OFF. (2) If spinner bulkhead (13) was removed, position spinner bulkhead on propeller and install washers (12) and screws (11). Torque screws (11) 20 to 25 inch-pounds. (3) On propeller with anti-ice installation, install screws securing anti-ice leads (21) to slip ring (15) and secure leads to bulkhead using screws and clamps removed. (4) Install D-5945 tool. (5) Apply a light coating of engine oil to O-ring (14) and install in the propeller hub. (6) Inspect stud and nut threads for cleanliness and absence of nicks, burrs or other damage. (7) Apply MIL-PRF-83483C lubricant liberally to propeller studs, nut threads and both faces of spacers (8).
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McCauley Propeller Installation Figure 201 (Sheet 1)
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McCauley Propeller Installation Figure 201 (Sheet 2)
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McCauley Propeller Installation Figure 201 (Sheet 3)
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McCauley Propeller Installation Figure 201 (Sheet 4)
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CAUTION: It is important that propeller be seated against engine ßange with a straight push. Rotation, cocking or wiggling will damage the o-ring groove and oil leakage may result. (8)
With propeller supported by a hoist and sling position propeller on engine ßange (10) and install spacers (8) and nuts (9). Using the B-5588 torque wrench adapter or equivalent, torque nuts 68 to 72 foot-pounds. (9) On propeller with anti-ice installation, install anti-ice brush block assembly (22). Clearance between anti-ice brush block and slip ring is 0.064 inch, +0.015 or -0.015 inch. Torque the nuts that attach the brush block bracket assembly to the engine from 145 to 165 inch-pounds (16.38 to 18.64 N-m). (10) Remove D-5945 tool. NOTE:
The lower end of the propeller reversing lever is machined with a stepped notch.
CAUTION: Make sure the stepped notch at the end of the propeller reversing lever (26) is under the guide pin (37) in the reversing lever guide pin bracket (36). (11) Install propeller reversing lever (26) and carbon block (24) in propeller feedback collar (23). Refer to Pratt & Whitney Engine Maintenance Manual for installing the propeller reversing lever. (12) Connect propeller reversing lever (26) to control cable and beta valve clevis (25). (13) Slide spinner support (1) on feathering spring housing (2).
CAUTION: Preform the following procedure exactly as written to prevent damage. (14) Lightly press spinner (17) against spinner support (1) and check alignment of spinner holes with spinner bulkhead holes. Spinner holes should be approximately 1/2 hole diameter forward from alignment with bulkhead holes. If not add or remove shims (16) to obtain this alignment. (15) Once shimming is complete, push hard on front of spinner to align holes and install screws (19) and washers (18). (16) Install right nose cap half and close cowling. 3.
Adjustment/Test A.
Beta Feedback Collar Axial Runout Check. NOTE: (1) (2) (3) (4)
Checking adjustment of the beta feedback collar axial runout is not required unless there is reason to believe linkage settings have been tampered with or feedback ring is bent.
Open right upper cowling door. Remove right nose cap half. Clamp dial indicator in position to check axial runout of forward face of beta feedback collar groove. Rotate propeller by hand and check that axial runout does not exceed 0.010 inch total indicator reading and that there is no binding between carbon block and feedback collar. NOTE:
The carbon block initially supplied with each propeller has been preÞtted. If a different carbon block is being installed, it may be necessary to sand it to obtain a total clearance between the carbon block and side of the groove of 0.001 inch to 0.002 inch at the tightest point. If clearance between feedback carbon block and groove of feedback collar exceeds 0.010 inch, replacement of feedback carbon block assembly is required.
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4.
Propeller Blade Damage A.
5.
Propeller Corrosion A.
6.
A propeller blade is highly stressed. The fact that propeller blades are likely to be subjected to damage such as nicks, gouges, scratches, corrosion pits, etc. demands frequent inspection and maintenance. (1) Refer to McCauley MPC26 Owner/Operator Information Manual for propeller blade inspection and repair information (refer to List of Vendor Publications). (2) Repair of small nicks and scratches may be performed by qualiÞed mechanics in the Þeld in accordance with procedures speciÞed in McCauley MPC26 Owner/Operator Information Manual also FAA Advisory Circular 43.13-1A. However, whenever a signiÞcant amount of metal is removed, or in the case of previously reworked blades which may be at or near the minimum width and thickness limits, the propeller shall be inspected by a McCauley FAA approved propeller repair station to determine if the minimum allowable blade width and thickness limits have not been exceeded. If the limits have been exceeded, blade replacement is required. If not, after Þling and polishing, the damaged area should be inspected by ßuorescent dye penetrant method to verify all damage has been removed and the blade is not cracked. The area should then be protected by localized application of chemical Þlm per MlL-C-5541 (e.g. Alodine) and repainted per manufacturers instructions as necessary. (3) Large nicks or scratches or other damage involving such things as bent blades, balance, diameter reduction etc. should be corrected only by a McCauley FAA approved propeller repair station. (4) Damage to blade anti-ice boots may conceal blade damage. Damage must be given careful inspection, anti-ice boot elasticity may obscure blade damage. If boot is damaged or cut completely through to the blade, or if blade damage is suspected, the boot must be removed for blade inspection/repair. A damaged boot may result in an electrical open or short circuit in the boot heating element. Boot replacement is required. A damaged heating element may also cause arcing to the blade surface. Damage of this type may also require blade replacement. Refer to Chapter 30 for anti-ice boot replacement.
Aluminum alloys used in propeller blades are susceptible to corrosion. The degree of corrosion likely to occur is largely dependent upon environmental exposure. If a painted or anodic blade surface is penetrated by stone damage etc., the material exposed may corrode if not reprotected. Corrosion may be accelerated if the propeller is operated in industrial or coastal areas. (1) Preventative measure should be taken. Damaged or blistered paint should be removed and repainted. Blades can be wiped with a cloth dampened with oil or waxed with an automotive type paste wax on a regular basis to minimize corrosion. (2) If corrosion develops, it should be removed as soon as possible. This can be accomplished by ensuring that blades are frequently inspected for evidence of corrosion. In the early stages, a light polishing is all that is required. However, if the corrosion is deep-seated, further material removal will be required. If a signiÞcant amount of metal is removed, or in the case of previously reworked blades which may be at or near the minimum width and thickness limits, the propeller shall be inspected by a McCauley FAA approved propeller repair station to make sure the minimum allowable blade width and thickness limits have not been exceeded. If the limits have been exceeded, propeller blade replacement is required. After the corrosion has been removed, the area should then be protected by localized application of chemical Þlm per MlL-C-5541 (e.g. Alodine) and repainted per the paint manufacturers instructions as necessary.
Propeller Grease or Oil Leakage NOTE: A.
The presence of oil or grease deposits on a nacelle does not necessarily indicate the propeller is leaking; leakage can come from the engine.
On new propellers, slight grease leakage during the Þrst several hours of operation is no cause for concern. Lubricants used during the assembly of shims at the propeller blade shank are liberally applied. Even though they are cleaned prior to shipment, centrifugal forces during the Þrst hours of operation can result in grease streaks on the blades. Such leakage will normally cease within the Þrst ten hours of operation.
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MODEL 208 MAINTENANCE MANUAL B.
A propeller oil leak can come from (1) the engine lubricating system or (2) the propeller hub cavity which holds approximately 2 quarts of turbine oil and is independent of engine oil. (1) Leakage of engine oil would normally come from piston O-rings and be observed around the tool attachment holes on the forward end of the propeller cylinder. Repair of such leaks can sometimes be performed, (by a propeller repairman) without removing the propeller from the engine. (2) Loss of small amounts of oil from the hub cavity need not be immediately replenished. The hub contains more oil than is needed for propeller lubrication. If leak persists the source must be determined and corrected, and oil replenished per McCauley MPC700 Propeller Overhaul Manual (refer to List of Vendor Publications). (3) Leakage from the hub cavity normally requires repair by an approved propeller repair station. However, if leakage is determined to be coming from a blade shank O-ring (blade to hub seal), before removing the propeller for repair, wipe off residue, run engine and cycle propeller pitch. After shutdown inspect for leaks; if no leakage is observed propeller may be returned to service (blade O-rings sometimes have a friction/stretch problem causing a leak which can be cured by recycling propeller pitch).
C.
If propeller leakage is suspected, but the source is not readily apparent, before removing propeller from airplane. (1) Remove propeller spinner. (2) Wipe clean all propeller, ßange and spinner bulkhead parts. (3) Use White "Dy-Check" developer or prepare a solution of alcohol and chalk dust to coat the hub and blade shank areas.
CAUTION: Do not attempt engine run up without spinner installed unless spinner bulkhead Þllets and anti-ice leads have been removed. (4) (5) (6) D.
After solution drys, reinstall spinner and run engine for at least Þve minutes. Shut down engine and examine coated surfaces. The sources of any leakage will show as a stain on the coated surfaces. If it is deÞnitely established that propeller is leaking, remove propeller and mark so proper inspection can be made during disassembly by an authorized propeller repair station.
If leakage is shown only through engine shaft and hub ßanges, it is not necessary to overhaul propeller. Remove propeller and carefully inspect end of engine shaft and propeller hub to determine cause of O-ring damage. After correction, install new O-ring and reinstall propeller.
CAUTION: Under no circumstances should an additional or oversize o-ring be used. E.
Internal leakage. An internal O-ring failure could allow propeller oil from the hub cavity into the engine lubricating system. The propeller uses turbine oil which is compatible with engine oil. An indication of propeller internal leakage would be an unexplainable increase in engine oil level. If this occurs the propeller hub oil level should be checked, if abnormally low, remove the propeller for repair by an approved McCauley propeller repair station.
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MODEL 208 MAINTENANCE MANUAL DYNAMIC BALANCING (McCAULEY) - ADJUSTMENT/TEST 1.
2.
General A.
This section gives information necessary to do a functional test for the dynamic balancing of the propeller. This data is applicable to all McCauley 3GFR34C703/106 GA-0 series propellers installed on the Model 208.
B.
The correctable propeller imbalance is a result of variations in the respective component weight and installation. The installation of balance weights in the correct position is an effective method to reduce the effects of these physical variations.
C.
The unknown factors in this problem are the amount and location of weight to be added. The balancer equipment will indicate the amount of imbalance in velocity (inches per second - IPS). This is translated into the amount of weight to be added and the angular location of the weight. The amount and location of weight added to the propeller is found with the balancer equipment. Refer to the manufacturer manual provided with the equipment for the balancing procedures.
D.
Applicable Documents. (1) Applicable Caravan l Pilot's Operating Handbook. (2) Applicable Caravan I Pilot's Checklist. (3) McCauley MPC700 Propeller Overhaul Manual and MPC26 Owner/Operator Information Manual (refer to List of Vendor Publications).
Balancing Requirements A.
Balancing Requirements for the Dynamic Balancing and Functional Test Procedure (Refer to Figure 501). (1) Dynamic balance the McCauley 3GFR34C703/106 GA-0 series propellers installed on the Model 208 after all the engine rigging is satisfactorily completed. Refer to Engine Control Rigging Adjustment and Test.
(2) (3) (4) (5)
NOTE:
If after Þve attempts the dynamic balance of the propeller is still not satisfactorily balanced, refer to MPC26 Owner/Operator Information Manual, Vibration Troubleshooting to help Þnd the source of the vibration. If it is determined the propeller needs to be statically balanced, remove the propeller from the airplane and have it statically balanced by an authorized McCauley repair station. After the propeller static balance has been completed, do the dynamic balance procedure again.
NOTE:
The maximum permitted vibration of the McCauley 3GFR34C703/106 GA-0 series propellers is 0.07 IPS.
Remove all the dynamic balance weights from the propeller. Remove the propeller shaft oil seal cavity drain plug. Install the adapter and vibration sensor in the propeller shaft oil seal drain port. Position the airplane into the wind and away from buildings and blast fences. NOTE:
(6) 3.
Do not balance the propellers when it is raining or when the wind gusts are 5 knots over any prevailing, steady wind.
Use the Aces equipment, or equivalent, to dynamically balance the propeller. Refer to the manufacturer manual provided with the equipment for the balancing procedures.
Initial Weight Installation A.
Install the Initial Weight (Refer to Figure 502). (1) Use the propeller balancing equipment to Þnd the initial weight installation location. Refer to the correct manufacturer manuals for dynamic propeller balancing procedures.
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Vibration Sensor Figure 501 (Sheet 1)
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Balance Weight Location Figure 502 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (2)
Remove the screw from the spinner identiÞed for the initial weight location. NOTE:
(3)
For initial weight installation, the weights can be installed on the outside of the spinner. The initial weight installation uses AN502-10-XX screws.
Find the number of AN970-3 washers necessary to equal the initial correction weight found in the dynamic propeller balancing procedure plus the weight of the A-1635-133 screw removed from the spinner. Refer to the correct manufacturer manuals for dynamic propeller balancing procedures. (a) Use the scale listed in the manufacturer manual to weigh the hardware used for the initial weight installation. NOTE:
(4) (5) (6) 4.
The total number of AN970-3 washers (4.1 grams per washer) cannot exceed 6 at any one location. If more than 6 washers are necessary, they must be equally located between adjacent locations. One AN970-3 washer must be placed forward of the spinner bulkhead for stress relief. All the other AN970-3 washers must be put aft of the spinner bulkhead. The AN502-10-XX screw may not project past the MS21083-N3 nut more than 0.125 inches.
Change the number of washers to adjust for the AN502-10-XX screw necessary to attach the initial weights on the outside of the spinner. Install the AN502-10-XX screw and AN970-3 washers in the spinner attachment screw hole. Use the propeller balancing equipment to Þnd the weight and angle correction. Refer to the manufacturer manual provided with the equipment for the balancing procedures.
Final Weight Installation A.
Install the Final Weight (Refer to Figure 502 and Figure 503). NOTE:
(1) (2) (3) (4) (5)
For Þnal weight installation, the spinner must be removed. The Þnal weights are installed on the spinner bulkhead. The Þnal weight installation uses an AN502-10-XX screw, an NAS1149F0332P washer, and a MS21083-N3 nut.
Remove the AN502-10-XX screw and AN970-3 washers installed during the initial weight installation. Remove the screws and washers from the spinner. Remove the spinner from the airplane. Change the total weight of the removed washers and AN502-10-XX screw, to adjust for the weight of the removed A-1635-133 screw, and the attaching NAS1149F0332P washer and MS21083-N3 nut. Install the Þnal weight hardware inside the spinner bulkhead in the hole nearest to the location found in the weight and angle correction. NOTE:
The total number of AN970-3 washers (0.145 ounces or 4.1 grams per washer) cannot exceed six at any one location. If more than six washers are necessary, they must be equally located between adjacent locations. One AN970-3 washer must be put forward of the spinner bulkhead for stress relief. All of the other AN970-3 washers must be put aft of the spinner bulkhead. The AN502-10-XX screw must not extend past the MS21083-N3 nut more than 0.125 inches.
CAUTION: Make sure you obey the procedure that follows as written. This will help prevent damage to the equipment. (6) (7)
Lightly push the spinner against the spinner support. Make sure that the spinner holes align with the spinner bulkhead holes. NOTE:
The spinner holes must be approximately 1/2 hole diameter forward from the bulkhead holes.
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McCauley Spinner Figure 503 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (8) (9) (10) (11) (12) (13)
If the holes do not align correctly, add or remove shims as necessary. When the shims are correct, push hard on the front of the spinner to align the spinner holes with the spinner bulkhead holes. Install the screws and washers in the spinner bulkhead. After the Þnal installation of the weights, do a test of the propeller balance to make sure the propeller is within permitted balance limits. Refer to the correct manufacturer manuals for dynamic propeller balancing procedures. Remove the adapter and vibration sensor from the propeller shaft oil seal drain port. Install the propeller shaft oil seal cavity drain plug.
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MODEL 208 MAINTENANCE MANUAL PROPELLER (McCAULEY) - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the McCauley propeller in a serviceable condition. NOTE:
For different views of the propeller and the spinner installation that are not included in this section, refer to Figure 201, in Propeller (McCauley) - Maintenance Practices.
Task 61-11-00-720 2.
McCauley Propeller Functional Check A.
General (1) This task gives the information needed to do the functional check of the McCauley propeller.
B.
Special Tools (1) Mild Soap and Water. (2) Stoddard Solvent or equivalent. (3) Isopropyl Alcohol.
C.
Access (1) Remove the nose cap to get access to the propeller governor. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (2) Remove the upper left cowling door to get access to the overspeed governor. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a McCauley Propeller Detailed Inspection. (1) Examine the propeller blades for any damage before you wash the blades.
CAUTION: Do not let the soap solution come into contact with the hub. The soap solution can contaminate the O-ring that is installed in the hub. (2)
Wash the propeller blades and the boots with mild soap and water before you start the inspection. NOTE:
(3) (4) (5) (6) E.
The propeller spinner is removed after the propeller is washed for the inspection.
Put a mark on the spinner and the bulkhead to record the alignment for the next installation. (a) Do not use a lead pencil. Remove the propeller spinner. Refer to Propeller (McCauley) - Maintenance Practices. (a) Make sure that you keep the front spinner support spacers for the next installation of the spinner. Be careful to not remove the spinner index mark when you clean the spinner and the bulkhead. (a) Clean the spinner and the bulkhead with Stoddard solvent to remove all oil and grease before you start the inspection. If installed, clean the de-ice slip ring assembly and the de-ice brush block with isopropyl alcohol, Stoddard solvent, or equivalent.
Examine the Spinner and Bulkhead (1) Examine the accessible surface of the bulkhead and the inner and outer spinner surface for condition, cracks, corrosion, and fractures. (2) Examine the spinner bulkhead, spinner bulkhead support, spinner attach screws, and spinner attach nutplates for condition, corrosion, and wear. (3) Examine the attach holes in the spinner for cracks and hole elongation. (4) Examine the spinner Þllets for condition, cracks, corrosion, and security. (5) Examine the balance weights for condition, corrosion, security, and correct installation. Refer to Final Weight Installation found in Dynamic Balancing (McCauley) - Adjustment/Test.
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MODEL 208 MAINTENANCE MANUAL (6)
(7) (8)
F.
Visually examine the spinner dome surface and the bulkhead for burned spots, pits, or other signs of a lightning strike. (a) If there are signs of a lightning strike, refer to Chapter 5, Unscheduled Maintenance Checks, Lightning Strike. If installed, examine the deice leads for condition, chaÞng, and security. For airplanes with TKS ,examine the feed shoes, slinger ring, propeller feed nozzle, propeller nozzle bracket, Þtting, and propeller hose assembly for condition, corrosion, security, and correct installation. Refer to Chapter 30, TKS Anti-Ice Propeller (McCauley) - Maintenance Practices. (a) Make sure that the feed nozzle is extended into the slinger ring channel with an edge distance of 0.10 to 0.15 inches (2.54 to 3.81 mm) from the slinger ring. If necessary adjust. Refer to Chapter 30, TKS Anti-Ice Propeller (McCauley) - Maintenance Practices. (b) Turn the propeller slowly by hand and make sure that the distance between the slinger ring and the feeder tube stays in an alignment tolerance of 0.10 to 0.15 inches (2.54 to 3.81 mm). Refer to Chapter 30, TKS Anti-Ice Propeller (McCauley) - Maintenance Practices. (c) While turning the propeller, make sure that the propeller feeder nozzle is spraying on the second groove of the adjacent feed shoe when the propeller is in full Þne pitch and that each tube has a 3/16 inch (4.76 mm) clearance from the propeller boot. Refer to Chapter 30, TKS Anti-Ice Propeller (McCauley) - Maintenance Practices.
Examine the Blades. (1) Examine all blades and blade surfaces for condition, gouges, scratches, corrosion, erosion, cracks, nicks, evidence of lightning strikes, and security. (a) If a propeller blade is found to have damage, refer to the McCauley MPC26 Owner/Operator Information Manual for repair procedures (refer to List of Vendor Publications). (b) If there are signs of a lightning strike, refer to Chapter 5, Unscheduled Maintenance Checks, Lightning Strike. (2) Examine all blade attachment points for oil leaks.
CAUTION: Oil leaks from the propeller or the engine can get on the wing, wing struts, and/or the horizontal stabilizer deice boots and cause damage. (3)
(4)
(5) (6) (7)
(8)
Examine the cylinder attachment point for oil leaks. (a) If oil is coming from the area of the beta spring housing, the piston seal is possibly leaking. Remove the propeller from service and return it to a McCauley authorized repair facility. Refer to Propeller (McCauley) - Maintenance Practices. Examine the area around the beta rod (3 each) bushings for oil leaks. NOTE:
The propeller hub is Þlled with turbine oil of the same type that is used in the engine. There are NO grease Þttings on this propeller.
NOTE:
Oil leaks found around the propeller mounting ßange can or can not come from the ßange. Other items such as the governor beta valve, or prop shaft seal can cause the oil leaks.
Examine the propeller mounting area for oil leaks. Examine the viewable area of the engine propeller shaft seal just aft of the spinner bulkhead. If installed, examine the anti-ice boots for abrasions, exposed heating elements, cuts, nicks, and security of attachment. (a) Examine the wiring from the boots to the terminal strips on the spinner bulkhead for condition, chaÞng, correct routing, and security of attachment at all clamps. (b) Examine the connector between the boot and the wire harness for security of attachment. (c) Examine the wire harness connectors at the terminal strips for condition and security of attachment. (d) Examine the boot edge dressing for condition. If necessary, touch-up damaged or exposed areas. 1 Examine the terminal strips for condition and security of attachment to spinner bulkhead.
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MODEL 208 MAINTENANCE MANUAL (9) G.
For airplanes with TKS, examine the feeder boots for abrasions, cuts, nicks, and security of attachment.
Examine the Hub (Refer to Figure 601). (1) Examine the exposed area for condition, cracks, corrosion, and security of the components to the hub. (2) Examine the hub for oil leaks at the blade butts and the mount ßange. (3) Examine the feathering spring housing for condition, cracks, corrosion, and security. (4) Examine the cylinder for condition, oil leaks at mount ßange, and security of attachment. (5) Visually examine the propeller for security of installation. (6) Examine the attach nuts for condition and that each stud has a spacer under the elastic attach nut. (7) Visually examine the nuts for security. NOTE: (a)
H.
The nut installation is correct if the torque putty on the nuts is not broken.
If you are not sure that the installation is correct, torque the nuts again, and apply new torque putty. Refer to Propeller (McCauley) - Maintenance Practices.
Examine the Beta System Feedback Collar (Refer to Figure 601). (1) Examine the beta feedback collar for condition, corrosion, and security of installation. (2) Examine the reversing lever for condition and security. (3) Examine the reversing lever for free play. (a) If there is free play at the beta valve, remove the clevis pin and examine the sleeve bushing for signs of wear at the attach location. If there is wear, replace the bushing. 1 (b) If the lever has free play at the beta cable clevis, remove bolt at the clevis and examine the sleeve spacer for signs of wear at the attach point. If there is wear, replace the spacer. 1 (4) Examine the alignment pin for condition and security. (5) Examine the carbon brush for wear and signs of damage. (6) To examine the carbon brush for wear, do the steps that follow: (a) Hold the carbon brush against the feedback collar. (b) Turn the feedback collar and measure the clearance between the carbon brush and the feedback collar around the full circumference of the feedback collar. NOTE:
The clearance between the brush and the feedback collar must not be more than 0.010 inch (0.254 mm) at any area around the full circumference of the collar.
CAUTION: Do not turn the elastic low pitch stop nuts installed on the beta rods. (7)
Examine the beta rods (3 each) for condition and security. (a) The beta rod locknuts are installed correctly if the torque putty has not been disturbed. (8) Examine the varistor installed near the center at the top of the forward side of the Þrewall for condition and security of installation. (9) Examine the electrical connections for condition, routing, signs of chafÞng, and security. (10) Discoloration of the varistor or the electrical leads, or a failure can be a sign that a lightning strike has occurred. (a) If you think that there was a lightning strike to the airplane, refer to Chapter 5, Unscheduled Maintenance Checks, Lightning Strike. I.
Examine the Propeller Governor (Refer to Figure 601). (1) Examine the propeller governor for condition, oil leaks, and security. (a) Make sure that the hardware is safety wired except for the four self-locking attach nuts. (2) Examine the speed adjusting lever return spring for condition and security. (3) Examine the air bleed link for corrosion, condition, and security. (4) Examine the governor interconnecting rod for corrosion, condition, security and wear. (5) Make sure that the rod end bearings turn freely and do not bind.
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Propeller and Spinner Installation - McCauley Figure 601 (Sheet 1)
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Propeller and Spinner Installation - McCauley Figure 601 (Sheet 2)
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Propeller and Spinner Installation - McCauley Figure 601 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (6)
Examine all hardware for corrosion, condition, and correct safety.
NOTE:
It is not necessary to safety wire the four self locking mounting nuts.
J.
Examine the Propeller Cable Terminal Rod End (Refer to Chapter 76, PT6A-114/-114A Engine Rigging - Adjustment/Test, Figure 510). (1) Disconnect the rod end from the propeller speed adjusting lever. Refer to Chapter 76, Propeller Control - Maintenance Practices. (2) Wipe the rod end clean with a clean lint-free cloth. (3) Examine the rod end for corrosion, pitting, and cleanliness. (4) Lubricate the rod end ball with MIL-L-7870. (5) Connect the rod end to the adjusting lever. Refer to Chapter 76, Propeller Control - Maintenance Practices.
K.
Examine the Overspeed Governor (Refer to Figure 601). NOTE: (1) (2) (3) (4) (5) (6)
(7) L.
Verify part number of governor to complete engine run portion of functional check.
Examine the overspeed governor for condition, oil leaks, and security. (a) Make sure that the hardware is safety wired except for the four self-locking attach nuts. Examine the electrical wiring and the electrical connection at the test solenoid for signs of damage, correct wire routing, and security. Examine the governor reset test solenoid for condition and security. Install the upper left cowling door. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. Install the nose cap. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Install the propeller spinner. Refer to Propeller (McCauley) - Maintenance Practices. (a) Make sure that the correct number of spacers between feathering spring housing and spinner support are installed at the locations that were recorded during the removal of the spinner. Do the Propeller Overspeed Governor Functional Check. Refer to Propeller Control Maintenance Practices.
Restore Access NOTE:
(1) End of task
The propeller spinner, nose cap, and upper left cowling door were installed before the functional check.
None
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MODEL 208 MAINTENANCE MANUAL PROPELLER CONTROL - MAINTENANCE PRACTICES 1.
General A.
2.
3.
A propeller governor is furnished with the engine for use with either Hartzell or McCauley propeller.
Description and Operation A.
The propeller governor combines the function of a normal propeller governor, a beta (reversing) valve, and a power turbine governor. The governor, mounted on the top front of the engine, provides connection for propeller control cable. A beta valve, mounted to the propeller governor, is connected to fuel control unit by adjustable linkage. The beta valve is utilized to reverse pitch angle of propeller blades in order to provide reverse thrust.
B.
A propeller overspeed governor is mounted on the left side of the engine reduction gearbox and prevents propeller overspeed if the primary propeller governor fails. The overspeed governor regulates the flow of oil to the propeller pitch change mechanism with a fly weight and speeder spring arrangement similar to that of primary governor. The overspeed unit governs at 104 percent Ng speed (approximately 1976 RPM) for Governor A210507, and 106 percent Ng speed (approximately 2014) for Governor D210507. Since it has no mechanical controls, the overspeed governor has a testing solenoid that resets the governor below its normal overspeed setting for ground test. The overspeed test switch is on left side of instrument panel.
Propeller Governor Removal/Installation A.
Remove Propeller Governor. (1) Remove right nose cap. (2) Disconnect propeller reversing lever from beta control valve clevis. Refer to Pratt & Whitney Engine Maintenance Manual for propeller reversing lever connection instructions. (3) Disconnect propeller governor interconnect rod from propeller governor air bleed link. (4) Remove propeller air pressure tube (Py) line at propeller governor. (5) Disconnect propeller control linkage from propeller speed-adjusting lever on governor. (6) Remove four self-locking nuts and plain washers securing propeller governor to engine reduction gearbox, and withdraw governor and gasket from mounting pad.
B.
Install Propeller Governor. (1) Install new gasket, with raised side of gasket facing upward, over studs on propeller governor mounting pad, located on reduction gearbox.
CAUTION: Make sure the drive splines of the governor are correctly engaged by verifying that the flange of the governor rests on the gasket squarely with no gap. Rotate the propeller to assist engagement, if necessary. (2) (3) (4) (5) (6) 4.
Install propeller governor over studs, onto gasket. (a) Secure with four washers and self-locking nuts. (b) Torque nuts to 170 to 190 inch-pounds. Connect beta control valve clevis to propeller reversing lever with clevis pin and secure with washers and cotter pin. Refer to Pratt & Whitney Engine Maintenance Manual for propeller reversing lever connection instructions. Connect propeller governor interconnect rod to propeller governor air bleed link. Connect propeller control linkage to propeller speed adjusting lever on propeller governor. (a) Connect propeller air pressure line (Py) to the propeller governor. Install right nose cap.
Propeller Overspeed Governor Removal/Installation A.
Remove Propeller Overspeed Governor (Refer to Figure 201). (1) Disengage PROP O-SPD TEST circuit breaker. (2) Open left upper cowling door to gain access.
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MODEL 208 MAINTENANCE MANUAL (3) (4) B.
Remove safety wire and disconnect electrical connector from overspeed governor reset test solenoid. Remove self-locking nuts and washers securing overspeed governor to engine reduction gearbox mounting pad and remove governor. Discard gasket.
Installation of Propeller Overspeed Governor (Refer to Figure 201). (1) Wipe engine mounting pad clean and place new gasket over engine mounting pad studs.
CAUTION: Do not use sealing compounds of any kind on gasket. (2)
Install overspeed governor on studs, engaging spines on governor drive shaft of engine.
CAUTION: Do not install the overspeed governor on studs with a force-fit. (3) (4) (5) (6) 5.
Install washers and self-locking nuts and torque nuts to 20 ft lbs. Connect the electrical connector to the governor reset test solenoid and safety wire the connector. Close left upper cowling door. Reset PROP O-SPD TEST circuit breaker.
Propeller Overspeed Governor Functional Check A.
Do a Propeller Overspeed Governor Functional Check. (1) Start engine in accordance with Pilot's Operating Handbook and Approved Airplane Flight Manual. (2) Advance propeller control lever to MAX. (3) Position power lever below 1500 RPM. (4) Press and hold propeller overspeed governor test switch, located on left side of instrument panel. NOTE:
(5) (6) (7) (8) (9)
The A210507 Governor effectivity is for Airplanes 20800001 thru 20800136 and 208B0001 thru 208B0105. The D210507 Governor effectivity is for airplanes 20800137 and On, and 208B0106 and On.
Move the power lever setting and make sure that the propeller RPM becomes stable at 1750 RPM, +20 or -20 RPM for Governor A210507, and 1785 RPM, +20 or -20 RPM for Governor D210507. Observe ITT and the torque limits. (a) Make sure the torque limits do not become more than 800 ft lbs. Make sure the propeller RPM (Np) does not increase more than 1770 RPM for Governor A210507, and 1805 RPM for Governor D210507. Reduce power lever setting to 1500 RPM. Release overspeed governor test switch to normal.
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MODEL 208 MAINTENANCE MANUAL
Propeller Overspeed Governor Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PROPELLER BETA INDICATING SYSTEM - DESCRIPTION AND OPERATION 1.
General NOTE: A.
2.
The propeller beta indication system is applicable to all Brazilian and British certiÞed airplanes.
The beta indicating system is utilized to advise the pilot that the engine power lever has been positioned past the ßight Idle position on the power quadrant, and that engine propeller is in the reverse position as indicated by the illumination of the BETA light on the annunciator panel.
Description A.
The BETA light is located in the annunciator panel and the beta switch is mounted on the engine. The beta switch is actuated by the propeller reversing lever connected to the propeller governor, carbon block and cam. The carbon block rides in the propeller feedback ring of the propeller assembly. The feedback ring movement is controlled by the movement of the propeller servo piston, and the servo piston is controlled by the amount of oil allowed to ßow to the propeller by the governor.
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MODEL 208 MAINTENANCE MANUAL PROPELLER BETA INDICATING SYSTEM - MAINTENANCE PRACTICES 1.
General NOTE: A.
2.
3.
The propeller Beta indication system is applicable to all Brazilian and British certiÞed airplanes.
The beta indicating system consists of the BETA annunciator light and engine-mounted switch. The light is located on instrument panel. The switch is located on left side of engine above propeller overspeed governor. Access to switch is gained by removing the engine cowling.
Beta Switch Removal/Installation A.
Remove Beta Switch. (Refer to Figure 201.) (1) Ensure electrical power is OFF. (2) Remove upper engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. (3) Identify and disconnect electrical wiring from switch. (4) Remove attaching hardware securing switch to bracket. (5) Remove switch.
B.
Install Beta Switch. (Refer to Figure 201.) (1) Position switch on bracket plate and secure with screws, washers, and nuts. (2) Identify and connect electrical wiring to switch. (3) Adjust switch. Refer to Adjustment/Test. (4) Install upper engine cowling. Refer to Engine Cowling and Nose Cap - Maintenance Practices.
Adjustment/Test A.
Adjust Beta Switch. (Refer to Figure 201) (1) Apply electrical power to airplane. (2) Place power lever in idle position. (3) Loosen screw holding switch mounting plate. Disengage switch from cam and ensure annunciator BETA light is on. (4) Actuate switch by depressing roller lever on switch. Ensure annunciator BETA light is out. (5) Ensure propeller is in feathered position, loosen jamnut and adjust cam on plunger so distance between forward edge of cam and aft edge of block is 0.350 inch with plunger spring-loaded against propeller reversing lever. Tighten jamnut. (6) Adjust switch mounting plate fore and aft so centerline of switch roller is 0.125-inch forward of sloped portion of cam. (7) Adjust switch mounting plate toward cam 0.040 to 0.060 inch further after switch is actuated on cam. Tighten mounting plate screws.
B.
Adjust Beta Switch (Engine). (Refer to Figure 201) (1) Start engine. Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. (2) Place propeller control lever at MAX RPM position. (3) Place power lever at IDLE position (at detent gate). Check annunciator BETA light is off. (4) Slowly move power lever toward REVERSE until annunciator BETA light comes on. Ensure this point is within one lever width of idle gate. (5) Advance the power lever to the IDLE position and check that the annunciator BETA light is out. (6) If adjustment is necessary, proceed as follows: (a) Verify power lever idle position is properly rigged (rear linkage). Refer to Chapter 76, Engine Control Rigging - Adjustment/Test. (b) Readjust microswitch mounting plate per Step A. Maintain position of plate toward the cam but shift fore and aft position per steps (c) and (d). (c) If BETA light goes off too early (i.e., power lever at idle), move plate forward 0.05 inch.
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Propeller Beta Indicating System Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (d)
If the BETA light comes on too late (i.e., more than one lower width aft of the idle gate), move the plate aft 0.05 inch. NOTE:
The fore and aft adjustment above should be sufÞcient if position of mounting plate toward cam is correct. If this position is incorrect, no adjustment may be necessary.
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MODEL 208 MAINTENANCE MANUAL COMPOSITE PROPELLER - CLEANING/PAINTING 1.
General A.
2.
Composite propeller cleaning/painting consists of repainting the propeller.
Composite Propeller Painting A.
Paint Propeller. (1) Roll on one coat of 13 parts spray Þll D61-A-23 and 1 part V66V27 catalyst. Allow to dry. (2) Sand with vibrator type sander with appropriate grit paper as required, leaving minimum thickness for purpose of Þlling pits and blemishes only. (3) Spray one coat of 13 parts E65A4 Polane primer sealer with 1 part V66V29 catalyst, (and V66VB11 accelerator as needed). Wipe with acetone soaked cheese cloth to smooth down any pronounced Kevlar Þber. Sand with appropriate grit paper as required (dry sand). (4) Spray face and camber with one light coat of 13 parts E65A4 Polane primer sealer with 1 part V66V29 catalyst, (and V66VB11 accelerator as needed). Sand with No. 400 grit sandpaper until desired Þnish is achieved, leaving no more than 4 mils total Polane primer sealer thickness. (5) Tape nickel leading edge, both face and camber, 0.75 inch aft from leading edge radius. (6) Spray with Spraylat (copper) Lightning Guard (P/N 599SA-A8574-1) 8 parts (A) with 1 part (B). Minimum thickness of 1 mil recommended. (7) Remove tape and feather edge of paint line. (8) Allow to dry 30 minutes and sand with variable grit sandpaper until smooth. (9) Spray light coat of 13 parts E65A4 Polane primer sealer with 1 part V66VB29 catalyst (and V66VB11 accelerator as needed) on both face and camber. Repeat second light coat on both face and camber, to achieve full coverage, but no lead or trail buildup. Total thickness not to exceed 2 mils dry. (10) Spray one light coat of 6 parts Polane black F63B8 with 1 part V66V29 catalyst, (and V66VB11 accelerator as needed) on both face and camber. Repeat second light coat on both face and camber so as to have full coverage but no lead or trail buildup. Total thickness of approximately 4 mils is recommended. (11) Tip (1), spray with 6 parts white plane 299WB612 with 1 part V66V29 catalyst, on camber side only.
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71 CHAPTER
POWER PLANT
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
71-00-00
Page 1
Aug 1/1995
71-00-01
Pages 1-3
Mar 1/1999
71-00-01
Pages 201-206
Jul 1/2010
71-00-01
Pages 501-512
Mar 1/2001
71-00-01
Pages 601-602
Jun 1/2011
71-00-02
Page 601
Mar 1/2001
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Jun 1/2011
71-20-00
Pages 201-206
Jul 1/2010
71-20-00
Pages 601-602
Jun 1/2011
71-40-00
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Mar 1/1999
71-41-00
Pages 401-402
Aug 1/1995
71-41-00
Pages 601-602
Jun 1/2011
71-42-00
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Apr 1/2010
71-43-00
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Aug 1/1995
71-50-00
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Aug 1/1995
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Oct 15/1999
71-60-00
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Jun 1/2011
71-70-00
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Mar 1/1999
71-Title 71-List of Effective Pages 71-Record of Temporary Revisions 71-Table of Contents 71-List of Tasks
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS POWERPLANT - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-00-00 71-00-00 71-00-00 71-00-00
POWER PLANT - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-00-01 Page 1 71-00-01 Page 1
POWER PLANT - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Build-Up Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Build-Up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-00-01 Page 201 71-00-01 Page 201 71-00-01 Page 201 71-00-01 Page 202 71-00-01 Page 206
POWER PLANT - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 600 SHP Engine (PT6A-114) Acceleration Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 675 SHP Engine (PT6A-114A) Acceleration Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-00-01 Page 501 71-00-01 Page 501 71-00-01 Page 501 71-00-01 Page 503 71-00-01 Page 510
POWER PLANT - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Compartment Zonal Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-00-01 Page 601 71-00-01 Page 601 71-00-01 Page 601
HOT SECTION - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hot Section Inspection/Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-00-02 Page 601 71-00-02 Page 601 71-00-02 Page 601
ENGINE COWLING AND NOSE CAP - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Cowling Door Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rub Strip Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cowling Center Panel Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lower Cowl Panel Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Cap Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-10-00 Page 201 71-10-00 Page 201 71-10-00 Page 201 71-10-00 Page 201 71-10-00 Page 201 71-10-00 Page 201 71-10-00 Page 206
ENGINE MOUNT - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Mount Truss Assembly Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Mount Elastomers Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Mount Bracket to Engine Mount Ring Bolt Removal/Installation. . . . . . . . . . . Engine Mount Bracket to Engine Bolt Removal/Installation . . . . . . . . . . . . . . . . . . . . . .
71-20-00 Page 201 71-20-00 Page 201 71-20-00 Page 201 71-20-00 Page 204 71-20-00 Page 205 71-20-00 Page 205
ENGINE MOUNT - INSPECTION/CHECK. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Mounts and Firewall Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Truss and Ring Assembly Special Detailed Inspection . . . . . . . . . . . . . . . . . . .
71-20-00 Page 601 71-20-00 Page 601 71-20-00 Page 601 71-20-00 Page 602
ENGINE EQUIPMENT ATTACH BRACKETS - MAINTENANCE PRACTICES . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Control Cable Bracket Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Control Cable Bracket Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Condition Control Cable Bracket Removal/Installation. . . . . . . . . . . . . . . . . . . . . . Emergency Power Cable Bracket Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . Standby Alternator Adjuster Arm Assembly/Mounting Bracket Assembly Removal/ Installation (Airplanes 20800001 Thru 20800079). . . . . . . . . . . . . . . . . . . . . . . . . . . . Standby Alternator Adjuster Arm Assembly/Mounting Bracket Assembly Removal/ Installation (Airplanes 20800080 and On, and 208B0001 and On) . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL ENGINE WASH RING - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Wash Ring Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-41-00 Page 401 71-41-00 Page 401 71-41-00 Page 401
ENGINE WASH RING - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Wash Ring, Air Plenum, and Thermocouple (T1) Detailed Inspection . . . . .
71-41-00 Page 601 71-41-00 Page 601 71-41-00 Page 601
COMPRESSOR BLADE WASH - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Motoring Wash . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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TURBINE BLADE WASH - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbine Blade Wash . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-43-00 Page 201 71-43-00 Page 201 71-43-00 Page 201
ELECTRICAL WIRING HARNESS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wiring Harness Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-50-00 Page 201 71-50-00 Page 201 71-50-00 Page 201
INERTIAL AIR SEPARATOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inertial Air Separator Control Linkage Removal/Installation. . . . . . . . . . . . . . . . . . . . . . Inertial Air Separator Control Linkage Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-60-00 Page 201 71-60-00 Page 201 71-60-00 Page 201 71-60-00 Page 201
INERTIAL AIR SEPARATOR - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inertial Air Separator Detailed Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71-60-00 Page 601 71-60-00 Page 601 71-60-00 Page 601
ENGINE DRAIN LINES - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Drain Lines Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 71-00-01-210
Engine Compartment Zonal Inspection
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71-20-00-220
Engine Mounts and Firewall Detailed Inspection
71-20-00 Page 601
71-20-00-240
Engine Truss and Ring Assembly Special Detailed Inspection
71-20-00 Page 602
71-41-00-220
Engine Wash Ring, Air Plenum, and Thermocouple (T1) Detailed Inspection
71-41-00 Page 601
71-60-00-220
Inertial Air Separator Detailed Inspection
71-60-00 Page 601
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MODEL 208 MAINTENANCE MANUAL POWERPLANT - GENERAL 1.
Scope A.
2.
This chapter contains maintenance information on the powerplant and associated components. For engine related information not found in this chapter, refer to applicable Pratt & Whitney maintenance manuals, listed in Introduction - List of Vendor Publications.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Engine Hoist Sling
CPWA 32327
Kellstrom Co., Inc. WeatherÞeld, CT
Used to aid in engine removal.
3.
DeÞnition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief deÞnition of the sections incorporated in this chapter is as follows: (1) The section on powerplant provides removal and installation instructions for the engine. (2) The section on engine cowlings provides removal and installation instructions for the engine cowlings and nose cap. (3) The section on mounts provides removal, installation and repair procedures for the engine mount. (4) The section on attach Þttings provides removal and installation procedures for attach brackets used on the engine and engine mount. (5) The section on compressor wash provides instructions using when washing the engine. (6) The section on inertial air separator provides removal, installation and inspection instructions for the inertial air separator.
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MODEL 208 MAINTENANCE MANUAL POWER PLANT - DESCRIPTION AND OPERATION 1.
General A.
Power plant installed in Model 208 Series airplanes is a Pratt and Whitney Aircraft of Canada, Ltd., Model PT6A-114 (600 SHP) gas turbine engine. Airplanes 20800277 and On, Airplanes 208B0179 and On, and 208B0001 thru 208B0178, incorporating SK208-80, have a PT6A-114A (675 SHP) gas turbine engine is installed. Engines employ three-stage axial, single-stage centrifugal compressor, driven by a single-stage turbine (free turbine). A second single-stage turbine, counter rotating with Þrst, drives propeller through a reduction gearbox. Fuel is sprayed into an annular combustion chamber by 14 individually removable fuel nozzles mounted around the gas generator case. An ignition unit and two spark igniter plugs are used to start combustion. A hydro-pneumatic fuel control unit (FCU) schedules fuel ßow to maintain power setting selected by the power control lever. Propeller speed remains constant at any selected propeller control lever position through action of a propeller governor. When engine power lever is moved aft into beta range (reverse), maximum propeller speed is limited by pneumatic section to propeller governor.
B.
Refer to Figure 1 for an illustration of engine components. Refer to Figure 2 for an engine air ßow diagram.
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MODEL 208 MAINTENANCE MANUAL
Engine Components Figure 1 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Engine Air Flow Figure 2 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL POWER PLANT - MAINTENANCE PRACTICES 1.
General A.
2.
Powerplant maintenance practices include engine removal/installation and engine build-up.
Engine Removal/Installation A.
Remove Engine (Refer to Figure 201 and Figure 202).
CAUTION: Chock main wheels and place a tailstand under tailcone before attempting engine removal. (1) (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13)
(14) (15) (16) (17) (18) (19)
Turn electrical power off. Pull fuel firewall shutoff control out (off). Remove upper cowling doors and lower cowling panels. Drain residual fuel from lines and fuel filter using filter drain. Remove fuel supply hose at fuel heater. Remove fuel motive flow hose at fuel control unit. Remove right nose cap and oil cooler. Remove top cowl center panel assembly and nosecap. Remove propeller. Disconnect and remove propeller speed control cable. Remove the left nose cap/induction air duct/inertial air separator. Disconnect cabin heater bleed air line at flow control valve and bleed air hose at mixing air valve. Remove starter/generator cooling air hose from starter/generator. Remove engine fire detector wiring harness. Disconnect electrical wiring connectors and ground wires at the following equipment locations: (a) Propeller overspeed governor and ITT harness (left front of engine). (b) Propeller tachometer generator (right front of engine). (c) Cabin bleed air heater flow control valve (lower right side of engine). (d) Oil temperature bulb (right side of engine). (e) Fuel control heater (right rear of engine). (f) Gas generator section tachometer generator (lower right side of engine). (g) Starter/generator (center top of engine accessory case). (h) Ignition exciter high tension leads at ignition exciter (right engine mount truss). Disconnect engine power control cables at fuel control unit. Remove torquemeter pressure and vent lines at forward upper right side of engine mount truss. Connect hoist sling to forward and aft lifting brackets and connect sling to engine hoist. Raise hoist to just support weight of engine and remove nuts and bolts at each of four corners of engine mounting ring. Ensure all wiring and lines are free, then carefully move hoist and engine forward to clear engine mount truss. If engine is to be returned for overhaul or replaced, remove the following items: (a) Engine induction air plenum. Refer to Chapter 71, Engine cowling and Nose Cap Maintenance Practices. (b) Engine mount ring, elastomers, and engine mount brackets. Refer to Chapter 71, Engine mount - Maintenance Practices (c) Propeller overspeed governor. Refer to Chapter 61, Propeller Control - Maintenance Practices. (d) Propeller tachometer generator. Refer to Chapter 77, Propeller RPM Indicator Maintenance Practices. (e) Oil temperature sensing bulb. Refer to Chapter 79, Oil Indicating - Maintenance Practices. (f) Oil cooler bracket and pressure/return hoses. Refer to Chapter 79, Oil Distribution Maintenance Practices. (g) Standby alternator (if equipped). Refer to Chapter 24, Standby Electrical System Maintenance Practices. (h) Torque sensing line and fittings.
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MODEL 208 MAINTENANCE MANUAL B.
3.
Install Engine (Refer to Figure 201 and Figure 202). (1) Install engine mount brackets, elastomers, and engine mount ring. (2) Connect lifting hoist sling to forward and aft lifting brackets on engine and lift engine into position forward of engine mount. (3) Make sure that all engine lines and equipment are clear. (4) Put lubrication on the engine mount bolts before you install them to prevent corrosion. (5) Move the hoist and engine aft to align the engine mount ring holes with the holes in the engine mount. (6) Install the mount bolts and torque the bolt/nuts to 480 to 690 inch-pounds. Remove the hoist and sling (7) Connect torquemeter pressure and vent lines at upper left firewall. Bleed torquemeter indicating system. (8) Connect engine power controls at fuel control unit. Rig controls. (9) Connect the electrical leads of the following items of electrical equipment: (a) Ignition exciter high tension leads at ignition exciter (right engine mount truss). (b) Starter/generator (center top of engine accessory case). (c) Gas generator section tachometer generator (lower right side of engine). (d) Fuel control heater (right rear of engine). (e) Oil temperature bulb (right rear of engine). (f) Cabin bleed air heater flow control valve (lower right side of engine). (g) All engine to engine mount ground straps. (h) Propeller overspeed governor and ITT harness (left front of engine). (i) Propeller tachometer generator (right front of engine). (10) Install engine fire detector warning harness. (11) Connect starter/generator cooling air hose to starter/generator. (12) Connect engine bleed air line to cabin bleed air heater flow control valve. Connect engine bleed air hose to cabin bleed air heater mixing air valve. (13) Install left nosecap/induction air duct/inertial air separator, if not previously installed. (14) Install propeller, if not previously installed. (15) Install and connect propeller governor control cable. (16) Install left and right nosecap bulkhead assemblies and top cowling center panel. (17) Install oil cooler and right nosecap. (18) Connect fuel supply hose at fuel heater and fuel motive flow hose at fuel control unit. (19) Push fuel firewall shutoff control fully in. (20) With fuel line disconnected at fuel manifold below engine, motor engine with starter to purge fuel lines. (21) Start engine and perform operational check. Refer to Pilot's Operating Handbook and FAAApproved Airplane Flight Manual. (a) Use the Pratt and Whitney PT6A-114/-114A/-135/-135A Engine Maintenance Manual with the Pilot's Operating Handbook and FAA-Approved Airplane Flight Manual to do the operational check of the different components on the engine. (22) Shut down engine and check for fluid leaks, connections or hardware, etc. (23) Replace engine cowling.
Engine Build-Up Precautions A.
The following precautions should be followed throughout the build-up of an engine. (1) Take extreme care to prevent dirt, hardware, tools or other foreign material from entering engine. (2) Do not remove packings and gaskets from their packages until needed for assembly purposes. (3) Clean packings and gaskets, if necessary, prior to installation with dry air under pressure or with clean, lint-free rags Do not use solvents. (4) Visually inspect all packings and gaskets for cuts, nicks, and other flaws prior to installation. In no case shall packings and gaskets that are damaged or altered be used. (5) Lubricate gaskets, packings, and back-up rings with the appropriate system fluid before installation. (6) Handle fuel and oil lines carefully to avoid denting or scratching them. Be especially careful not to damage the threads of fittings and line coupling nuts.
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MODEL 208 MAINTENANCE MANUAL
Engine Installation Figure 201 (Sheet 1)
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Engine Installation Figure 201 (Sheet 2)
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Engine Hoisting Sling Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (7)
(8) (9) (10) (11)
(12) (13) 4.
Caps should not be removed from lines until immediately before installation. If lines are disconnected for any reason, they should be recapped until ready for connection. Also, all installed lines, ducts, and electrical connectors that terminate with open ends should be capped or covered in a suitable manner to exclude the entrance of dirt and foreign objects. Before installing any part, be sure it is thoroughly clean. Brushes used in cleaning should not mar or scratch the metal surface. When making fuel and oil connections, apply anti-seize compound (JAN-A- 669) to male threads sparingly, being careful not to permit entry into lines. Do not twist hose assemblies when installing. The stripe on the sides of the hose will show if any twist exists. A twisted hose under pressure may fail or loosen itself. It is important to use correct size and type of clamps when securing various hoses, tubing, and wire bundles directly to engine or to engine via brackets. If clamps of insufficient size are used and tightened excessively, the line may be damaged by not being able to slip through the clamps when the engine grows because of thermal expansion. Route and clamp all lines as shown. This will aid in keeping line chafing to a minimum after engine installation. All electrical bonding, grounding, and mating surfaces shall be clean metal surfaces free of anodic films, oxides, grease, paint, or other high-resistance film. Whenever paint has been removed to make connections, the connections shall be refinished to prevent corrosion. Use lockwire to secure bolts and fittings as required.
Engine Build-Up A.
Install the following items of equipment on the engine before proceeding to install on the airplane. (1) Standby alternator (if equipped). Refer to Chapter 24, Standby Electrical System - Maintenance Practices. (2) Oil cooler pressure and return hoses. Refer to Chapter 79, Oil Distribution - Maintenance Practices. (3) Primary exhaust stack. Refer to Chapter 78, Primary and Secondary Exhaust Duct Maintenance Practices. (4) Starter/generator. Refer to Chapter 80, Starter/Generator - Maintenance Practices. (5) Oil temperature sensing bulb. Refer to Chapter 79, Oil Indicating - Maintenance Practices. (6) Propeller tachometer generator. Refer to Chapter 77, Propeller RPM Indicator - Maintenance Practices. (7) Propeller overspeed governor. Refer to Chapter 61, Propeller Control - Maintenance Practices. (8) Engine induction air plenum. Refer to Chapter 71, Engine cowling and Nose Cap - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL POWER PLANT - ADJUSTMENT/TEST 1.
2.
General A.
This adjustment/test procedure outlines individual procedures to guide maintenance personnel for operating and adjusting an engine. Procedures described are not necessarily in maintenance sequence; select an individual or group of procedures to meet the maintenance requirement.
B.
For engine power lever and control rigging, refer to Chapter 76, Engine Control Rigging - Adjustment/ Test.
Engine Operating Limits A.
The following limitations shall be observed during testing. If at any time the limits are exceeded, immediately shut down the engine by placing the throttle at ßight idle and the fuel condition lever in cutoff. (1) For limits during engine adjustment and testing, refer to Figure 501, Engine Operating Limits, and Table 501 and Table 502.
Table 501. PT6A-114 Engine Operating Limits POWER SETTING
TORQUE FOOTPOUNDS (8)
MAXIMUM ITT °C
GAS GENERATOR RPM % Ng (1) (9)
PROPELLER RPM
OIL PRESSURE PSIG (2)
OIL TEMP °C (5)
SHAFT HORSEPOWER (7)
Takeoff
1980
805
101.6
1900
85 to 105
10 to 99
600
Maximum Climb
1980
765
101.6
1900
85 to 105
0 to 99
600
Maximum Cruise
1980
740
101.6
1900
85 to 105
0 to 99
600
685
52 to 54 (Minimum)
40 (minimum)
-40 to 99
805
101.6
85 to 105
0 to 99
600
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Idle Maximum Reverse (3)
1980
1825
© Cessna Aircraft Company
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL Table 501. PT6A-114 Engine Operating Limits (continued) POWER SETTING
TORQUE FOOTPOUNDS (8)
MAXIMUM ITT °C
GAS GENERATOR RPM % Ng (1) (9)
PROPELLER RPM
OIL PRESSURE PSIG (2)
OIL TEMP °C (5)
Transient
2400 (10)
900 (4)
102.6 (4)
2090 (11)
85 to 105
104 (12)
Starting Maximum Continuous (6) 1. 2.
3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.
1090 (4) (13) 1980
805
SHAFT HORSEPOWER (7)
-40 (minimum) 101.6
1900
85 to 105
10 to 99
600
For every 10°C (18°F) below -30°C (-22°F) ambient temperature, reduce maximum allowable Ng by 2.2%. Normal oil pressure is 85 to 105 PSI at gas generator speeds above 72% with oil temperature between 60°C and 70°C (140°F and 185°F). Oil pressure below 85 PSI is undesirable and should be tolerated only for completion of the ßight, preferably at a reduced power setting. Oil pressure below normal should be reported as an engine discrepancy and should be corrected before the next takeoff. Oil pressures below 40 PSI are unsafe and require that either the engine be shut down or a landing be made as soon as possible using the minimum power required to sustain ßight. Minimum oil pressure above 27,000 Ng is 85 PSI. Reverse power operation is limited to one minute. These values are time limited to Þve seconds. For increased oil service life, an oil temperature below 80°C (176°F) is recommended. A minimum oil temperature of 55°C (130°F) is recommended for fuel heater operation at takeoff power. Use of this rating is intended for abnormal situations (i.e., maintain altitude or climb out of extreme icing or windshear conditions). The maximum allowable SHP is 600. Less than 600 SHP is available under certain temperature and altitude conditions as reßected in the takeoff, climb and cruise performance charts. If maximum torque is used, set Np so as to not exceed power limitations. 100% Ng is 37,500 RPM. These values are limited to 20 seconds. If propeller governor fails toward overspeed, permissible to complete a ßight with propeller control via overspeed governor (on engines so equipped) provided this limit is not exceeded. Maximum permissible transient oil temperature is 104°C (219°F) for 10 minutes. Investigate starting temperatures above 850°C (1562°F) for cause.
Table 502. PT6A-114A Engine Operating Limits POWER SETTING
TORQUE FOOTPOUNDS (8)
MAXIMUM ITT °C
GAS GENERATOR RPM % Ng (1) (9)
PROPELLER RPM
OIL PRESSURE PSIG (2)
OIL TEMP °C (5)
SHAFT HORSEPOWER (7)
Takeoff
1980
805
101.6
1900
85 to 105
10 to 99
675
Maximum Climb
1980
765
101.6
1900
85 to 105
0 to 99
675
Maximum Cruise
1980
740
101.6
1900
85 to 105
0 to 99
675
685
52 to 54 (Minimum)
40 (minimum)
-40 to 99
Idle
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MODEL 208 MAINTENANCE MANUAL Table 502. PT6A-114A Engine Operating Limits (continued) POWER SETTING
TORQUE FOOTPOUNDS (8)
MAXIMUM ITT °C
GAS GENERATOR RPM % Ng (1) (9)
PROPELLER RPM
OIL PRESSURE PSIG (2)
OIL TEMP °C (5)
SHAFT HORSEPOWER (7)
Maximum Reverse (3)
1980
805
101.6
1825
85 to 105
0 to 99
675
Transient
2400 (10)
900 (4)
102.6 (4)
2090 (11)
85 to 105
104 (12)
Starting Maximum Continuous (6) 1. 2.
3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 3.
-40 (minimum)
1090 (4) (13) 1980
805
101.6
1900
85 to 105
10 to 99
675
For every 10°C (18°F) below -30°C (-22°F) ambient temperature, reduce maximum allowable Ng by 2.2%. Normal oil pressure is 85 to 105 PSI at gas generator speeds above 72% with oil temperature between 60°C and 70°C (140°F and 185°F). Oil pressure below 85 PSI is undesirable and should be tolerated only for completion of the ßight, preferably at a reduced power setting. Oil pressure below normal should be reported as an engine discrepancy and should be corrected before the next takeoff. Oil pressures below 40 PSI are unsafe and require that either the engine be shut down or a landing be made as soon as possible using the minimum power required to sustain ßight. Minimum oil pressure above 27,000 Ng is 85 PSI. Reverse power operation is limited to one minute. These values are time limited to Þve seconds. For increased oil service life, an oil temperature below 80°C (176°F) is recommended. A minimum oil temperature of 55°C (130°F) is recommended for fuel heater operation at takeoff power. Use of this rating is intended for abnormal situations (i.e., maintain altitude or climb out of extreme icing or windshear conditions). The maximum allowable SHP is 675. Less than 675 SHP is available under certain temperature and altitude conditions as reßected in the takeoff, climb and cruise performance charts. If maximum torque is used, set Np so as to not exceed power limitations. 100% Ng is 37,500 RPM. These valves are limited to 20 seconds. If propeller governor fails toward overspeed, permissible to complete a ßight with propeller control via overspeed governor (on engines so equipped) provided this limit is not exceeded. Maximum permissible transient oil temperature is 104°C (219°F) for 10 minutes. Investigate starting temperatures above 850°C (1562°F) for cause.
600 SHP Engine (PT6A-114) Acceleration Check A.
Acceleration Check (Refer to Figure 502). (1) Before any adjustments are made to the acceleration adjuster dome. Mark acceleration dome and fuel control unit with a marker pen to establish an initial reference point. (2) Start engine in accordance with Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. Operate engine at idle for Þve minutes to allow temperatures to stabilize. (3) Slowly advance power lever to obtain take off power (1900 RPM and 1658 foot-pounds torque). Record percent Ng at takeoff power and mark power lever position on pedestal. (4) Reduce power to idle. (5) Compute 97.5 percent Ng recorded previously. (6) Set power lever to obtain 63 percent Ng. (7) Move power lever rapidly from 63 percent Ng to position marked on pedestal cover for takeoff power, and record time to obtain 97.5 percent takeoff Ng as previously computed. As soon as 97.5 percent of takeoff Ng is achieved, retard power lever to idle to preclude and overtorque condition.
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Engine Operating Limits Figure 501 (Sheet 1)
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Engine Operating Limits Figure 501 (Sheet 2)
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Engine Operating Limits Figure 501 (Sheet 3)
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Engine Operating Limits Figure 501 (Sheet 4)
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Engine Operating Limits Figure 501 (Sheet 5)
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Fuel Control Linkage Figure 502 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (8)
4.
Acceleration time should fall within limits shown in Figure 503. If not, rotate acceleration adjuster dome one click at a time until requirement is met. Rotate dome clockwise to increase acceleration rate. Do not exceed three clicks. Lockwire adjuster dome. (Refer to Figure 502.)
675 SHP Engine (PT6A-114A) Acceleration Check A.
Acceleration Check (Refer to Figure 502 and Figure 503). (1) Before any adjustments are made to the acceleration adjuster dome. Mark acceleration dome and fuel control unit with a marker pen to establish an initial reference point. (2) Start engine in accordance with Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. Operate engine at idle for Þve minutes to allow temperatures to stabilize. (3) Slowly advance power lever to obtain take off power (1900 RPM and 1865 foot-pounds torque). Record percent Ng at takeoff power and mark power lever position on pedestal. (4) Reduce power to idle. (5) Compute 97.5 percent Ng recorded in step (2). (6) Set power lever to obtain 63 percent Ng. (7) Move power lever rapidly from 63 percent Ng to position marked on pedestal cover for takeoff power, and record time to obtain 97.5 percent takeoff Ng as previously computed. As soon as 97.5 percent of takeoff Ng is achieved, retard power lever to idle to preclude an overtorque condition. (8) Acceleration time shall fall within limits shown in Figure 503. If not, rotate acceleration adjuster dome one click at a time until requirement is met. Rotate dome clockwise to increase acceleration rate. Do not exceed three clicks. Lockwire adjuster dome.
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Engine Acceleration Time Versus Temperature Chart Figure 503 (Sheet 1)
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Engine Acceleration Time Versus Temperature Chart Figure 503 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL POWER PLANT - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the engine compartment in a serviceable condition.
Task 71-00-01-210 2.
Engine Compartment Zonal Inspection A.
General (1) The Zonal Inspection Program (ZIP) includes a series of General Visual Inspection (GVI) tasks. This section gives ZIP procedures for an zonal inspection of the engine compartment. NOTE:
An engine compartment zonal inspection is a general visual examination that includes all systems and structural components in the engine compartment area, installation, or assembly. This includes checks for evidence of degradation such as corrosion, cracks, chafing of tubing, loose duct support, wiring damage, cable wear, fluid leaks, insufficient drainage, and for other conditions that could cause corrosion/damage. This level of inspection is completed during normal lighting conditions such as daylight, hangar light, drop-light, or flashlight at approximately "arm length" inspection distance to the object. It can be necessary to remove and/or open access panels or doors to do an engine and engine compartment zonal inspection. A mirror can be necessary to enhance visual access to all exposed surfaces in the inspection area. Stands, ladders, or platforms can be necessary to get access to the area that is checked.
B.
Special Tools (1) None
C.
Access (1) Remove the engine cowlings. Refer to Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do the Engine Compartment Zonal Inspection. NOTE: (1) (2) (3) (4) (5)
This inspection is from the forward tip of the nose spinner to FS 100.00.
Examine the engine compartment for damage and signs of overheating. Refer to Chapter 20, High Intensity Radiated Fields (HIRF) - Inspection/Check, External Zonal Visual Inspection of Lightning and High Intensity Radiated Fields. Examine all of the systems and structural components for damage, corrosion, cracks, loose fasteners, loose/misalignment of linkage, and correct installation. Examine all tubing, hose, and fluid fittings for signs of leaks, damage, chafing, and correct clamp installation. Examine all placards and markings for security of installation, legibility, and correct location. (a) For the correct placards and placard locations. Refer to the Pilots Operating Handbook or Chapter 11, of the Model 208 Illustrated Parts Catalog. Examine for contamination and look carefully for quantities of combustible material. (a) Remove all of the combustible material that has collected. NOTE:
Combustible material can be fuel vapor, engine oil, and/or dust or lint that has collected.
NOTE:
An inspection for contamination and combustible material meets the requirements of the Enhanced Zonal Inspection Program.
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Restore Access (1) Install the engine cowlings. Refer to Engine Cowling and Nose Cap - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL HOT SECTION - INSPECTION/CHECK 1.
General A.
2.
Hot section inspection in conducted with power section only removed from airplane.
Hot Section Inspection/Check A.
Inspect Hot Section. (1) Remove engine cowling and nosecap. (2) Place suitable drip pan under front section of engine to catch residue oil lost during following steps. (3) Remove propeller. Refer to Chapter 61. (4) Disconnect pressure and scavenge oil lines to oil cooler and remove cooler. Refer to Chapter 79. (5) Remove primary exhaust duct. Refer to Chapter 78. (6) Disconnect torque pressure and vent lines at top right front of engine. Refer to Chapter 77. (7) Remove Þre detector loop. Refer to Chapter 26. (8) Disconnect electrical leads at propeller tachometer generator. (9) Remove propeller reduction gearbox drain plug/chip detector and drain oil from gearbox. Refer to Chapter 79. (10) At propeller governor, disconnect propeller speed control cable, propeller governor interconnect rod, and front clevis at Beta follow-up cable from reversing lever. Remove Beta follow-up cable from front attach bracket at "C" ßange. (11) Disconnect Py reference line at propeller governor. Cap line and port. (12) Disconnect electrical cable at propeller overspeed governor. (13) Position hoist in front of engine and attach sling to propeller mounting ßange with four socket head capscrews. Tighten capscrews Þngertight. (14) Remove power section from engine. Refer to Pratt and Whitney Maintenance Manual. (15) Perform hot section inspection in accordance with procedures in Pratt and Whitney Maintenance Manual. (16) Install power section to engine. Refer to Pratt and Whitney Maintenance Manual. (17) Remove hoist and sling from engine. (18) Connect electrical leads to propeller overspeed governor. (19) Connect Py reference line at propeller governor. (20) Install Beta follow-up cable to mounting bracket "C” ßange. (21) Connect front clevis of Beta follow-up cable and propeller governor interconnect rod to propeller reversing lever. Install propeller speed control cable to propeller governor speed lever. Check propeller control cable rigging per Chapter 77. (22) Install propeller reduction gearbox chip detector in sump of propeller reduction gearbox. Refer to Chapter 77. (23) Install electrical leads at propeller tachometer generator. (24) Install Þre detector loop. Refer to Chapter 76. (25) Connect torque indicator system pressure and vent lines at top right front of engine. Bleed system as required. Refer to Chapter 77. (26) Install primary exhaust duct. Refer to Chapter 78. (27) Install oil cooler. Refer to Chapter 79. (28) Install propeller. Refer to Chapter 61.
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MODEL 208 MAINTENANCE MANUAL ENGINE COWLING AND NOSE CAP - MAINTENANCE PRACTICES 1.
General A.
2.
3.
4.
5.
Engine cowling and nose cap maintenance practices consist of upper cowling door removal/ installation, cowling center panel removal/installation and lower cowl panel removal/installation.
Upper Cowling Door Removal/Installation A.
Remove Upper Cowling Door (Refer to Figure 201). (1) Depress latch handle (8) and raise cowling door. (2) Support cowling door section and depress lock button of hinge pin (7). Remove hinge pin (7) at forward and aft locations. (3) Remove cowling door. (4) Do a primary and secondary exhaust duct general visual inspection (alignment check). Refer to Task 78-10-00-210.
B.
Install Upper Cowling Door (Refer to Figure 201). (1) Support cowling door and align hinge pin hole in hinge bearing (5) with hinge pin hole in center engine mount truss (6). Insert hinge pin (7) at both forward and aft hinge locations. (2) Ensure hinge pins (7) are locked in place.
Rub Strip Removal/Installation NOTE:
This procedure is intended for Airplanes 20800130 and On and 20800001 thru 20800129 incorporating SK208-47 and Airplanes 208B0068 and On and 208B0001 thru 208B0067 incorporating SK208-47.
NOTE:
Removal and installation procedures are typical for all rub strips.
A.
Remove Rub Strip (Refer to Figure 201). (1) Remove existing rivets. Refer to Model 208 Series Structural Repair Manual, Chapter 51 Standard Practices and Structures - General. (2) Remove rub strip.
B.
Install Rub Strip (Refer to Figure 201). (1) Clean existing adhesive from upper cowling door. Refer to Chapter 20, General Solvents/Cleaners - Maintenance Practices. (2) Secure new rub strip to upper cowling door using temporary sheet metal fasteners. (3) Match drill Number 30 (0.125 inch diameter) holes through rub strip. (4) Countersink rub strip hole locations to 0.125 inch diameter X 100 degrees. (5) Remove rub strip and debur holes. (6) Apply EA9309.3NA Adhesive to rub strip. (7) Install rub strip using MS20426AD4 rivets as required. (8) Clean excess adhesive from parts. Refer to Model 208 Series Structural Repair Manual, Chapter 51 - Standard Practices and Structures - General.
Cowling Center Panel Removal/Installation A.
Remove Cowling Center Panel (Refer to Figure 201). (1) Remove cowling center panel (2) by disengaging quarter-turn fasteners (32) at forward and aft ends of panel. Disengage quarter-turn fasteners (32) by rotating counterclockwise.
B.
Install Cowling Center Panel. (Refer to Figure 201). (1) Position cowl center panel to airplane and secure fasteners (32) at forward and aft ends of panel.
Lower Cowl Panel Removal/Installation A.
Remove Lower Cowl Panel (Refer to Figure 201). (1) Remove left (1) and right (3) upper cowling doors. Refer to Upper Cowling Door Removal/ Installation.
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Cowling and Nose Cap Installation Figure 201 (Sheet 1)
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Cowling and Nose Cap Installation Figure 201 (Sheet 2)
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Cowling and Nose Cap Installation Figure 201 (Sheet 3)
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Cowling and Nose Cap Installation Figure 201 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4)
Remove quick-release pin (21), washer (20), and clevis pin (18) to release lower cowl support (19) and remove lower left cowl panel (16). Disconnect oil breather hose and starter/generator oil drain line at exhaust stack connections on lower right cowl panel (15). Remove drag link spring fairing right removable panel. Refer to Chapter 32, Nose Landing Gear - Maintenance Practices. NOTE:
(5)
Ensure lower left cowl panel is supported, then disengage by rotating fasteners (32) counterclockwise at forward and aft ends of lower left cowl panel (16). Remove lower left cowl panel (16) from airplane. NOTE:
(6) B.
Secondary exhaust stack is attached to the lower right cowl panel (15) with screws (forward) and attach brackets (aft). Exhaust stack can remain attached to lower right cowl panel during removal.
Ensure lower right cowl panel (15) is supported. Disengage quarter- turn fasteners (32) at forward and aft ends of lower right cowl panel (15). Remove lower right cowl (15) from airplane.
Install Lower Cowl Panels. NOTE: (1) (2) (3) (4) (5) (6)
6.
Drag link spring fairing removable panel is integral with lower left cowl panel (16) and removed as a unit.
Lower right cowl panel (15) must be installed prior to installing lower left cowl panel (16).
Support lower right cowl panel (15) and install by rotating quarter-turn fasteners (32) clockwise on forward and aft ends of panel. Support lower left cowl panel (16) and install by rotating quarter-turn fasteners (32) clockwise on forward and aft ends of panel. Align lower cowl support (19) holes with mounting bracket (28) and secure using clevis pin (18), washer (20), and quick-release pin (21). Connect oil breather hose and starter/generator oil drain line at exhaust stack connections on lower right cowl panel (15).Tighten retaining clamps. Install drag link spring fairing right removable panel. Refer to Chapter 32, Nose Landing Gear Maintenance Practices. Install left (1) and right (3) upper cowling doors. Refer to Upper Cowling Door Removal/ Installation.
Nose Cap Removal/Installation A.
Remove Nose cap. (Refer to Figure 201). NOTE:
(1) (2) (3) (4) (5) (6) (7)
Right nose cap half (30) can be removed from airplane without further equipment removal; however, left nose cap half (29), induction air duct (26) and inertial separator (25) are assembled as a single unit. Removal of this assembly requires removal of left lower cowl section (16) and left cowl bulkhead (23).
Disengage right nose cap half (30) quarter-turn fasteners (32) on upper and lower centerline of nose cap. Remove right nose cap half (30) from airplane. Remove left (1) and right (3) upper cowling doors. Refer to Upper Cowling Door Removal/ Installation. Remove cowling center panel (2). Refer to Cowling Center Panel Removal/Installation. Remove lower left (16) and lower right (15) cowl panels. (Refer to Figure 201). Disengage quarter-turn fasteners (32) securing right cowl bulkhead (22) to left cowl bulkhead (23). Remove bulkheads from airplane. Remove bolt (33), washers (34) and (35), and nut (36) to detach actuator tube (37) from actuating link (38). Remove bolts (39) and (41), washers (40) and (42), and nut (44) to detach actuator plate (43) from induction air plenum.
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Secure actuator plate (43) to inertial separator (25) using tape or other suitable method.
CAUTION: Ensure actuator plate (43) is secured to inertial separator (25) to avoid the possibility of damage during removal. (9)
B.
Remove nuts and bolts (27) securing forward section of left nose cap half (29) to mounting bracket (28). With assembly supported, remove hinge pins (24) from forward, aft, and top surfaces of inertial separator (25). Remove assembly from airplane.
Install Nose cap (Refer to Figure 201). (1) With induction air duct and inertial separator (25) installed, support left nose cap half (20) in position. Align nose cap bracket mounting holes with bracket (28) and hinge halves of inertial separator (25). Align with hinge halves of induction air duct (26) and insert hinge pins (24). (2) Install bolts and nuts (27) and safety nuts. (3) Attach actuator plate (43) to induction air duct (26) using bolts (39) and (41), washers (40) and (42), and nuts (44). (4) Attach actuator tube (37) to actuator link (38) using bolt (33), washers (34) and (35), and nut (36).
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MODEL 208 MAINTENANCE MANUAL ENGINE MOUNT - MAINTENANCE PRACTICES 1.
General A.
2.
Engine maintenance practices consist of the engine mount truss removal and installation, the engine mount elastomer removal and installation, the engine mount bracket to engine mount ring bolt removal and installation, and the engine mount bracket to engine bolt removal and installation.
Engine Mount Truss Assembly Removal/Installation A.
Remove the Engine Mount Truss Assembly (Refer to Figure 201). (1) Remove the upper cowling doors. Refer to Upper Cowling Door Removal/Installation. (2) Remove the lower cowling sections. Refer to Lower Cowl Panel Removal/Installation. (3) Remove the upper center cowling panel section. Refer to Cowling Center Panel Removal/ Installation. (4) Remove the right nosecap. Refer to Nose Cap Removal/Installation. (5) Remove the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation. (6) Remove the engine. Refer to Engine Removal/Installation. (7) Remove the nose gear. Refer to Chapter 32, Nose Landing Gear - Maintenance Practices. (8) Remove the clamps that attach the electrical wire bundles, lines, and hoses to the engine mount truss assembly. (9) Seal and stow the electrical wire bundles, lines, and hoses. (10) Hold the engine mount truss assembly and remove the bolts that attach the upper engine mount truss assembly, the lower engine mount truss assembly, and the center engine mount truss assembly to the firewall structure. (11) Keep for installation the nuts and barrel nuts that attach the upper engine mount truss assembly, the lower engine mount truss assembly, and the center engine mount truss assembly to the firewall structure.
B.
Install the Engine Mount Truss Assembly (Refer to Figure 201). (1) Put the engine mount truss assembly in position against the firewall structure.
CAUTION: You must use a countersunk washer with an internal wrenching bolt. This will help prevent damage to the bolt. (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13) (14) (15) (16)
Put the countersunk washers with the countersunk face next to the bolt head radius on the bolts that attach the upper engine mount truss assembly and the lower engine mount truss assembly to the firewall structure. Refer to Chapter 20, Torque Data - Maintenance Practices. Install the bolt, countersunk washer, washer, and nut that attach the upper engine mount truss assembly to the firewall structure. Install the bolt, washers, special washer (if necessary), and nut that attach the center engine mount truss assembly to the firewall structure. Install the bolt, countersunk washer, and barrel nut that attach the lower engine mount truss assembly to the firewall structure. Torque the bolts that attach the upper engine mount truss assembly and the lower engine mount truss assembly to the firewall structure to a range of 450 to 500 inch-pounds. Torque the bolt that attaches the center engine mount truss assembly to the firewall structure to a range of 160 to 190 inch-pounds. Install the nose gear. Refer to Chapter 32, Nose Landing Gear - Maintenance Practices. Install the engine. Refer to Engine Removal/Installation. Make sure the routing of the wire bundles, the lines, and the hoses is correct. Install the clamps on the lines and hoses on the engine mount truss assembly. Install the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation. Install the right nosecap. Refer to Nose Cap Removal/Installation. Install the upper center cowling section. Refer to Cowling Center Panel Removal/Installation. Install the lower cowling panel sections. Refer to Lower Cowl Panel Removal/Installation. Install the upper cowling doors. Refer to Upper Cowling Door Removal/Installation.
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Engine Mount Figure 201 (Sheet 1)
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Engine Mount Figure 201 (Sheet 2)
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3.
Engine Mount Elastomers Removal/Installation A.
Remove the Engine Mount Elastomers (Refer to Figure 201). (1) Remove the upper cowling doors. Refer to Upper Cowling Door Removal/Installation. (2) Remove the lower cowling sections. Refer to Lower Cowl Panel Removal/Installation. (3) Remove the upper center cowling panel section. Refer to Cowling Center Panel Removal/ Installation. (4) Remove the right nosecap. Refer to Nose Cap Removal/Installation. (5) Remove the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation.
CAUTION: Before you remove any engine mount bracket to engine mount ring bolts, make sure you hold the engine with a hoist and a sling. This will help prevent engine movement and damage. (6)
Hold the engine with a hoist and a sling. Refer to Powerplant - General.
CAUTION: Make sure you remove and install only one engine mount bracket to the engine mount ring bolt at a time. This will help prevent engine movement and damage. (7)
Remove the cotter pin, nut, washer, countersunk washer, and bolt that attach the engine mount bracket to the engine mount ring. (8) Remove the aft elastomer. (9) If necessary, adjust the hoist to make enough clearance between the engine mount bracket and the engine mount ring to free the forward elastomer. (10) Remove the forward elastomer. (11) Keep for installation the spacer and the pins. B.
Install the Engine Mount Elastomers (Refer to Figure 201). (1) If necessary, adjust the hoist to make enough clearance between the engine mount bracket and the engine mount ring to install the forward elastomer. (2) Put the spacer, pins, and forward elastomer in position.
CAUTION: You must use a countersunk washer with an internal wrenching bolt. This will help prevent damage to the bolt. (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13)
Put a countersunk washer with the countersunk face next to the bolt head radius on the bolt that attaches the engine mount bracket to the engine mount ring. Refer to Chapter 20, Torque Data - Maintenance Practices. Put the aft elastomer in position and install the bolt, with the countersunk washer under the bolt head, that attaches the engine mount bracket to the engine mount ring. Install the washer and nut on the bolt that attaches the engine mount bracket to the engine mount ring. Torque the nut that attaches the engine mount bracket to the engine mount ring to a range of 480 to 690 inch-pounds. Install the cotter pin that attaches the engine mount bracket to the engine mount ring. Remove the hoist and the sling from the engine. Install the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation. Install the right nosecap. Refer to Nose Cap Removal/Installation. Install the upper center cowling section. Refer to Cowling Center Panel Removal/Installation. Install the lower cowling panel sections. Refer to Lower Cowl Panel Removal/Installation. Install the upper cowling doors. Refer to Upper Cowling Door Removal/Installation.
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4.
Engine Mount Bracket to Engine Mount Ring Bolt Removal/Installation A.
Remove the Engine Mount Bracket to Engine Mount Ring Bolt (Refer to Figure 201). (1) Remove the upper cowling doors. Refer to Upper Cowling Door Removal/Installation. (2) Remove the lower cowling sections. Refer to Lower Cowl Panel Removal/Installation. (3) Remove the upper center cowling panel section. Refer to Cowling Center Panel Removal/ Installation. (4) Remove the right nosecap. Refer to Nose Cap Removal/Installation. (5) Remove the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation.
CAUTION: Before you remove any engine mount bracket to engine mount ring bolts, make sure you hold the engine with a hoist and a sling. This will help prevent engine movement and damage. (6)
Hold the engine with a hoist and a sling. Refer to Powerplant - General.
CAUTION: Make sure you remove and install only one engine mount bracket to the engine mount ring bolt at a time. This will help prevent engine movement and damage. (7) B.
Remove the cotter pin, nut, washer, countersunk washer, and bolt that attach the engine mount bracket to the engine mount ring.
Install the Engine Mount Bracket to Engine Mount Ring Bolt (Refer to Figure 201).
CAUTION: You must use a countersunk washer with an internal wrenching bolt. This will help prevent damage to the bolt. (1)
Put a countersunk washer with the countersunk face next to the bolt head radius on the bolt that attaches the engine mount bracket to the engine mount ring. Refer to Chapter 20, Torque Data - Maintenance Practices. (2) Install the bolt, with the countersunk washer under the bolt head, that attaches the engine mount bracket to the engine mount ring. (3) Install the washer and nut on the bolt that attaches the engine mount bracket to the engine mount ring. (4) Torque the nut that attaches the engine mount bracket to the engine mount ring to a range of 480 to 690 inch-pounds. (5) Install the cotter pin that attaches the engine mount bracket to the engine mount ring. (6) Remove the hoist and the sling from the engine. (7) Install the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation. (8) Install the right nosecap. Refer to Nose Cap Removal/Installation. (9) Install the upper center cowling section. Refer to Cowling Center Panel Removal/Installation. (10) Install the lower cowling panel sections. Refer to Lower Cowl Panel Removal/Installation. (11) Install the upper cowling doors. Refer to Upper Cowling Door Removal/Installation. 5.
Engine Mount Bracket to Engine Bolt Removal/Installation A.
Remove the Engine Mount Bracket to Engine Bolt (Refer to Figure 201). (1) Remove the upper cowling doors. Refer to Upper Cowling Door Removal/Installation. (2) Remove the lower cowling sections. Refer to Lower Cowl Panel Removal/Installation. (3) Remove the upper center cowling panel section. Refer to Cowling Center Panel Removal/ Installation. (4) Remove the right nosecap. Refer to Nose Cap Removal/Installation. (5) Remove the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation.
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MODEL 208 MAINTENANCE MANUAL (6)
Remove the safety wire from the bolt that attaches the engine mount bracket to the engine.
CAUTION: Do not remove more than two of the 12 total engine mount bracket to engine bolts at one time. This will help prevent engine movement. (7) B.
Remove the washer and bolt that attach the engine mount bracket to the engine.
Install the Engine Mount Bracket to Engine Bolt (Refer to Figure 201).
CAUTION: Make sure you use the correct washer for the type of bolt used to connect the engine mount bracket to the engine. You must use a countersunk washer with an internal wrenching bolt. This will help prevent damage to the bolt. (1)
If necessary, put a countersunk washer with the countersunk face next to the bolt head radius on the bolt that attaches the engine mount bracket to the engine. Refer to Chapter 20, Torque Data - Maintenance Practices. NOTE:
(2) (3) (4) (5) (6) (7) (8) (9)
Although most airplanes have an internal wrenching bolt with a countersunk washer that attaches the engine mount bracket to the engine, some airplanes have a hex head bolt with a flat washer.
Install the bolt, with the washer under the bolt head, that attaches the engine mount bracket to the engine. Torque the bolt that attaches the engine mount bracket to the engine to a range of 275 to 300 inch-pounds. Install safety wire on the bolts that attach the engine mount bracket to the engine. Refer to Chapter 20, Safetying - Maintenance Practices. Install the left nosecap, induction air duct, and inertial air separator as a single assembly. Refer to Nose Cap Removal/Installation. Install the right nosecap. Refer to Nose Cap Removal/Installation. Install the upper center cowling section. Refer to Cowling Center Panel Removal/Installation. Install the lower cowling panel sections. Refer to Lower Cowl Panel Removal/Installation. Install the upper cowling doors. Refer to Upper Cowling Door Removal/Installation.
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MODEL 208 MAINTENANCE MANUAL ENGINE MOUNT - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the engine mounts in a serviceable condition.
Task 71-20-00-220 2.
Engine Mounts and Firewall Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the engine mounts, ground straps, and firewall.
B.
Special Tools (1) None
C.
Access (1) Remove the engine cowlings. Refer to Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a Detailed Inspection of the Tubular Mounts (Refer to Figure 201 found in Engine Mount Maintenance Practices). (1) Examine the areas that follow for condition, cracks at welds, distortion, corrosion, and security of attachment: • Engine truss tube surface • Areas around the engine truss welds • Right and left side landing gear junctions • Engine truss to firewall attach bolts • ECTM instrument box (if applicable) attachment area at lower right side of truss • Lower truss tube and ignition exciter box attachment area at the right side of truss. (2) Examine the ground straps, attach bolts (5 ea.), and mount bolts (4 ea.) for security of installation. (a) The mount bolts to engine mount ring must have the nuts installed against the mount ring (bolt heads aft) with cotter pins installed.
E.
Do a Detailed Inspection of the Engine Shock Mounts. (1) Examine the engine upper, right, and left side mount brackets, ground straps, engine mounts, mount bolts and engine mount to engine ring attach bolts for condition, security, cracks, corrosion, and deterioration of elastomers. NOTE:
(2) (3)
Mount brackets on airplanes 2080001 Thru 20800187 and 208B0001 Thru 208B0216 must be modified per CAB90-7 Rev.1 for improved drainage and mount bolt lubrication to aid in preventing corrosion.
Examine the elastomer closely where the rubber material is bonded to the plate for signs of separation. (a) If separation is found, replace the elastomer. If replacement of the elastomer is necessary, do the following with the components removed. (a) Examine the mount bracket for condition, cracks, and corrosion. (b) Examine the attaching bolt and nut for condition and corrosion. (c) Examine the spacer and shims for condition and wear. (d) When installing, make sure that the chamfer of the special washer is against the shoulder of the bolt head.
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CAUTION: The mount bolt is an internal wrenching type. Where the shank meets the head there is a radius shoulder that must mate with the chamfer of the special washer. Incorrect installation of these parts could cause the head of the bolt to break off. F.
Do a Detailed Inspection of the Firewall Structure. (1) Examine the forward surface of the firewall for corrosion, condition, cracks, missing rivets, and signs of damage. (2) Examine the sealant for overall condition at the fittings where items pass through the firewall skin. (a) If sealant is found unserviceable, replace the sealant. Refer to Chapter 20, Fuel, Weather and High-Temperature Sealing - Maintenance Practices.
CAUTION: The firewall is sealed at skin overlaps and joints with a fire resistant sealant. Where items pass through the firewall such as control cables and wire harnesses, fittings are sealed with a white ablative type fire resistant sealant (DAPCO U000117). (3)
Examine the brackets and fittings attached to the forward side of the firewall for condition, corrosion, and security.
G.
Restore Access (1) Install the engine cowlings. Refer to Engine Cowling and Nose Cap - Maintenance Practices. End of task Task 71-20-00-240 3.
Engine Truss and Ring Assembly Special Detailed Inspection A.
General (1) This task includes the Supplemental Inspection Document (SID) requirements necessary to keep the engine truss and ring assembly in a serviceable condition.
B.
Special Tools (1) None
C.
Access (1) Remove the engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a Special Detailed Inspection of the Engine Truss and Ring Assembly. (1) Do a nondestructive testing (NDT) inspection for cracks in the engine mount ring assembly. Refer to the Model 208 Nondestructive testing Manual, Part 8, Magnetic Particle, Engine Truss and Ring Assembly - Description And Operation. (2) Do a NDT inspection for cracks in the engine mount assembly at the engine mount ring assembly. Refer to the Model 208 Nondestructive testing Manual, Part 8, Magnetic Particle, Engine Truss and Ring Assembly - Description And Operation. (3) Do a NDT inspection for cracks in the engine mount assembly at the firewall attachments. Refer to the Model 208 Nondestructive testing Manual, Part 8, Magnetic Particle, Engine Truss and Ring Assembly - Description And Operation. (4) If no cracks are found, restore access. (5) If cracks are found, repair or replace the damaged part(s). Refer to Chapter 71, Engine Mount - Maintenance Practices or the Model 208 Structural Repair Manual.
E.
Restore Access (1) Install the engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL ENGINE EQUIPMENT ATTACH BRACKETS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
4.
Engine equipment attach brackets maintenance practices consist of attach brackets removal/ installation.
Power Control Cable Bracket Removal/Installation A.
Remove Power Control Cable Bracket (Refer to Figure 201). (1) Open or remove upper right cowling door. (2) Remove cotter pin (3), nut (4), washers (5), retaining washer (6), spacer (8) and lever arm bolt (10) from rod end (15B). Discard cotter pin, but retain remaining hardware for reinstallation. (3) Remove and retain jamnuts (15A), rod end (15B) and washer (15C) from power control cable. (4) Withdraw power control cable (1). (5) Remove screws attaching power control cable bracket (2) to wiring convolute and retain for reinstallation. (6) Remove power control cable bracket (2) from engine.
B.
Install Power Control Cable Bracket (Refer to Figure 201). (1) Position power control cable bracket (2) to wiring convolute and secure using previously retained screws. (2) Install power control cable (1) through hole in power control cable bracket (2). (3) Install jamnuts (15A), rod ends (15B) and washer (15C). (4) Install new MS24665-86 cotter pin (3) and previously retained nut (4), washers (5), retaining washer (6), spacer (8) and lever arm bolt (10) to power control cable (1). (5) Rig power control cable. Refer to Chapter 76, Engine Control Rigging - Adjustment/Test.
Propeller Control Cable Bracket Removal/Installation A.
Remove Propeller Control Cable Bracket (Refer to Figure 201). (1) Remove cotter pin (3), nut (4), washers (5), retaining washers (6), spacer (8) and bolt (10) from propeller control cable (14) at lever arm (9). Discard cotter pin, but retain remaining hardware for reinstallation. (2) Remove and retain jamnuts (15A), washer (15C), and rod ends (15B) at propeller control cable bracket. (3) Withdraw propeller control cable (14) from propeller control cable bracket (15). (4) Remove three nuts and bolts attaching propeller control cable bracket (15) to engine and retain hardware for reinstallation.
B.
Install Propeller Control Cable Bracket (Refer to Figure 201 ). (1) Install three bolts and nuts at holes at bottom of propeller control cable bracket (15) attaching it to engine. (2) Insert propeller control cable (14) through propeller control cable bracket (15). (3) Install jamnuts (15A), washer (15C), and rod ends (15B). (4) Install new MS 24665-86 cotter pin (3), nut (4), washers (5), retaining washer (6), spacer (8) and lever arm bolt (10) to propeller control cable (14). (5) Rig propeller control cable. Refer to Chapter 76, Engine Control Rigging - Adjustment/Test.
Fuel Condition Control Cable Bracket Removal/Installation A.
Remove Fuel Condition Control Cable Bracket (Refer to Figure 201). (1) Remove screws (11) and washers (15C) retaining fuel condition cable (13) to fuel condition cable bracket (12). (2) Remove fuel condition control cable bracket (12) from engine and retain existing hardware for reinstallation.
B.
Install Fuel Condition Control Cable Bracket (Refer to Figure 201). (1) Install fuel condition cable bracket (12) to engine using existing hardware. (2) Install screws (11) and washers (15C) retaining fuel condition cable (13) to fuel condition cable bracket (12).
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Engine Equipment Attach Brackets Figure 201 (Sheet 1)
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Engine Equipment Attach Brackets Figure 201 (Sheet 2)
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Engine Equipment Attach Brackets Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (3) 5.
6.
Rig fuel condition cable. Refer to Chapter 76, Engine Control Rigging - Adjustment/Test.
Emergency Power Cable Bracket Removal/Installation A.
Remove Emergency Power Control Cable Brackets (Refer to Figure 201 ). (1) Remove safety wire from nuts (4) attaching power cable brackets (17) and bracket (18) to screws (11). (2) Remove nuts (4), spacers (8) and screws (11), releasing emergency power cable (19). Retain all hardware for reinstallation. (3) Remove safety wire from nuts (4) securing emergency power cable brackets (17) to U-bolt (16). (4) Remove nuts (4) and U-bolt (16) and retain for reinstallation. (5) Remove emergency power cable brackets (17) and brackets (18) from fuel control unit.
B.
Install Emergency Power Cable Brackets (Refer to Figure 201). (1) Install emergency power cable brackets (17) using U-bolt (16), screws (11) spacers (8) and nuts (4). Safety wire nuts. (2) Install bracket (18) retaining emergency power control cable (19). Secure with nuts (4). Safety wire nuts. (3) Rig emergency power control cable. Refer to Chapter 76, Engine Control Rigging - Adjustment/ Test.
Standby Alternator Adjuster Arm Assembly/Mounting Bracket Assembly Removal/Installation (Airplanes 20800001 Thru 20800079) A.
Remove Standby Alternator Adjuster Arm Assembly (Refer to Figure 201). (1) Remove alternator (22). Refer to Chapter 24, Standby Electrical System - Maintenance Practices. (2) Loosen clamp (29) and remove drain hose (30) and elbow (28A) if installed. (3) Remove adapter (28). (4) Disconnect line (23C) and union (23B) from drive housing (23). Discard packing (23A). (5) Remove ground strap (36) from adjuster arm assembly (25) by removing bolt (37), nut (39) and washer (38). Retain bolt, nut and washer for reinstallation. (6) Remove nuts (26) and washers (40) securing drive housing (23) to engine. Discard gasket (24), but retain nuts and washers for reinstallation. (7) Remove adjuster arm assembly (25) from engine and discard gasket (24A).
B.
Install Standby Alternator Adjuster Arm Assembly. (Refer to Figure 201.) (1) Using new gasket (24A), installed on aft face of scavenge pump, position adjuster arm assembly (25) over four studs of pump and install second new gasket (24) over studs. Refer to 208 Illustrated Parts Catalog for gasket part numbers. (2) Slide drive housing (23) on studs, making sure drive splines are engaged. (3) Install nuts (26) and washers (40). (4) Using bolt (37), washer (38) and nut (39), install ground strap (36) to adjuster arm assembly (25). (5) Using new packing (23A), install union (23B) and connect line (23C). Refer to 208 Illustrated Parts Catalog for packing part number. (6) Install adapter (28). (7) Connect drain hose (30) to elbow (28A), if installed, and secure with clamp (29) (as needed). (8) Install alternator (22). Refer to Chapter 24, Standby Electrical System - Maintenance Practices
C.
Remove Standby Alternator Mounting Bracket Assembly (Refer to Figure 201) (1) Remove alternator (22). Refer to Chapter 24, Standby Electrical System - Maintenance Practices. (2) Remove nut (35), three bolts (34) and washers (33). Retain for reinstallation. (3) Remove two bolts, nuts, and washers attaching mounting bracket assembly (21) to drain valve bracket mount and retain hardware for reinstallation. Refer to Engine Drain Lines - Maintenance Practices. (4) Remove mounting bracket assembly (21) from engine.
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MODEL 208 MAINTENANCE MANUAL D.
7.
Install Standby Alternator Mounting Bracket Assembly (Refer to Figure 201 ). (1) Position mounting bracket assembly (21) aligning three holes at bottom over three matching holes in base of engine. (2) Install nut (35), three bolts (34) and washers (33). (3) Position mounting bracket assembly (21) to drain valve mount and secure using previously retained bolts, nuts and washers. Refer to Engine Drain Lines - Maintenance Practices. (4) Install alternator (22). Refer to Chapter 24, Standby Electrical System - Maintenance Practices.
Standby Alternator Adjuster Arm Assembly/Mounting Bracket Assembly Removal/Installation (Airplanes 20800080 and On, and 208B0001 and On) A.
Remove Standby Alternator Adjuster Arm Assembly (Refer to Figure 201). (1) Remove alternator (22). Refer to Chapter 24, Standby Electrical System Maintenance Practices. (2) Loosen ground strap (36) from by removing bolt (37), washer (38) and nut (39). Retain bolt, washer and nut for reinstallation. (3) Remove nuts (26) and washers (40) securing drive housing (23) to engine. Discard gasket (24) but retain nuts and washer for reinstallation. (4) Remove adjuster arm assembly (25) from engine and discard gasket (24A).
B.
Install Standby Alternator Adjuster Arm Assembly (Refer to Figure 201). (1) Using new gasket (24A) installed on aft face of scavenge pump, position adjuster arm assembly (25) over four studs of pump and install second new gasket (24) over studs. Refer to 208 Illustrated Parts Catalog for gasket part numbers. (2) Slide drive housing (23) on studs making sure drive splines are engaged. NOTE: (3) (4) (5)
Install drive housing (23) with chamfered edge of housing adjacent to adjuster arm assembly (25).
Install washers (40) and nuts (26). Using bolt (37), washer (38) and nut (39), install ground strap (36) to adjuster arm assembly (25). Install alternator (22). Refer to Chapter 24, Standby Electrical System - Maintenance Practices.
C.
Remove Standby Alternator Mounting Bracket Assembly (Refer to Figure 201) (1) Remove alternator (22). Refer to Chapter 24, Standby Electrical System - Maintenance Practices. (2) Remove nut (35), three bolts (34) and washers (33). Retain hardware for reinstallation. (3) Remove two bolts, nuts and washer attaching mounting bracket assembly (21) to drain valve mount and retain for reinstallation. Refer to Engine Drain Lines - Maintenance Practices. (4) Remove mounting bracket assembly (21) from engine.
D.
Install Standby Alternator Mounting Bracket Assembly (Refer to Figure 201). (1) Position mounting bracket assembly (21) aligning three holes at bottom over matching holes at base of engine. (2) Install nut (35), three bolts (34) and washers (33). (3) Position mounting bracket assembly (21) to drain valve mount and secure using previously retained bolts, nuts and washers. Refer to Engine Drain Lines - Maintenance Practices. (4) Install alternator (22). Refer to Chapter 24, Standby Electrical System - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL ENGINE WASH RING - REMOVAL/INSTALLATION 1.
General A.
2.
An engine wash ring is installed inside the induction air plenum assembly to aid in performing periodic compressor washes as recommended by the engine manufacturer. The assembly consists of a curved tube with 19 drilled holes for discharge of cleaning and rinse solutions. A capped exterior connection is provided to connect wash ring to solution source.
Engine Wash Ring Removal/Installation A.
Remove Engine Wash Ring (Refer to Figure 401). (1) Remove upper cowling doors. (2) Remove upper center cowl panel section. (3) Remove lower cowling panel sections. (4) Remove left nosecap/induction air duct/inertial air separator. NOTE:
It may be necessary to remove induction air plenum top panel to gain access to engine wash ring.
CAUTION: Cover engine air inlet to prevent foreign objects from dropping into engine. (5) (6) (7) B.
Remove nuts (7) and washers from screws (1) securing clamps (4) to wash ring (8). Carefully remove clamps (4) from wash ring (8). Slide wash ring out of plenum through grommet (10) and remove.
Install Wash Ring (Refer to Figure 401). (1) Insert wash ring through grommet (10) into plenum.
CAUTION: Ensure clamps (4) do not obstruct wash ring spray holes when installing clamps. (2) (3) (4) (5) (6) (7)
Install clamps (4) and secure with screws (1), washers, and nuts (7). Install induction air plenum top panel, if removed. Remove cover from engine inlet. Install nosecap/induction air duct/inertial air separator. Install lower cowling panel sections. Install upper center cowl panel section and upper cowling doors.
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Engine Compressor Wash Ring Installation Figure 401 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ENGINE WASH RING - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the engine wash ring, air plenum, and trim thermocouple (T1) in a serviceable condition.
Task 71-41-00-220 2.
Engine Wash Ring, Air Plenum, and Thermocouple (T1) Detailed Inspection A.
General (1) This task gives the procedures to do an inspection of the engine wash ring, air plenum, and trim thermocouple (T1).
B.
Tools and Equipment (1) None
C.
Access (1) Remove the left and the right upper cowling doors. Refer to Engine Cowling and Nose Cap Maintenance Practices. (2) Remove the right and the left nose gear fairings. Refer to Chapter 32, Nose Gear Fairing Maintenance Practices. (3) Remove the left and the right lower cowl panels. Refer to Engine Cowling and Nose Cap Maintenance Practices. (4) Remove the inspection panel assembly from the air plenum.
D.
Do an Engine Wash Ring, Air Plenum, and Thermocouple (T1) Detailed Inspection.
CAUTION: Security of the air plenum attaching hardware is very important. Ingestion of loose hardware during engine startup can cause damage to the airplane or injury to personnel. (1)
(2) (3) (4)
(5)
Examine the air plenum assembly and all attaching brackets and hardware for condition, cracks, corrosion, chafing, excessive wear, and security of installation. (a) If you find that there is missing hardware, it is necessary to do a more detailed inspection of the engine compressor section. Examine the upper and the lower air plenum band assemblies for condition, cracks, corrosion, and security. Examine the induction air bleed tube assembly for condition, cracks, corrosion, and security. Examine the engine wash ring tube for condition and security. (a) Examine the tube attach brackets for condition and security. (b) Examine the tube attach clamps for condition and security. (c) Examine the tube cap and chain for condition and security. Examine the trim thermocouple (T1) for condition, bends, and security of attachment at the T5 boss. Refer to the Pratt & Whitney Engine Maintenance Manual, Section 77-20-01.
CAUTION: If using a wrench to examine the security of the trim thermocouple attach bolts, use an additional wrench for the backup nut that is attached to the terminal lug. If you do not obey these instructions, damage can occur to the T5 lugs. (6) (7) (8) (9)
Examine the connection at the T5 boss for condition, corrosion, and security. Examine the exterior of the gas generator case for general condition. cracks, distortion, corrosion, and signs of overheating. Examine the air inlet screen wire mesh for condition, cleanliness, and damage. Examine the rims and the flanges of the screen for security and damage.
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MODEL 208 MAINTENANCE MANUAL E.
Restore Access (1) Install the inspection panel assembly to the air plenum. (2) Install the left and the right lower cowl panels. Refer to Engine Cowling and Nose Cap Maintenance Practices. (3) Install the right and the left nose gear fairings. Refer to Chapter 32, Nose Gear Fairing Maintenance Practices. (4) Install the left and the right upper cowling doors. Refer to Engine Cowling and Nose Cap Maintenance Practices. End of task
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MODEL 208 MAINTENANCE MANUAL COMPRESSOR BLADE WASH - MAINTENANCE PRACTICES 1.
General A.
2.
Compressor blade wash is accomplished to remove deposit buildup accumulated on compressor blades during normal operation.
Engine Motoring Wash NOTE:
Refer to Pratt and Whitney Engine Maintenance Manual for more information about the engine compressor wash procedures, wash schedules, and cleaning solution mixture. These documents recommend the use of demineralized water for motoring washes, and recommend drinking-quality water as an alternative. The Pratt and Whitney Engine Maintenance Manual gives specified information about drinking-quality water and demineralized water.
NOTE:
The engine shall be washed by the following method (motoring wash only) using the starter. Cleaning solution flow rate should be 2-3 GPM.
CAUTION: Observe starter cycle limitations outlined in the pilot's operating handbook and faa approved airplane flight manual to prevent damage to the starter/generator when motoring engine. A.
Desalination or Performance Recovery Wash. NOTE:
(1) (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13)
Removal and capping off of P3 pneumatic lines, during compressor blade wash is not necessary on airplanes utilizing the newer Pratt and Whitney engine PT6A-114A. This change is due to the incorporation of an improved P3 air filter drain adapter, which eliminates the possibility of fuel contamination and damage to the P3 lines due to mishandling, during this maintenance procedure.
Open upper left cowling door and connect wash ring, refer to Pratt and Whitney Engine Maintenance Manual. Place suitable catch pan under engine. Open right upper cowling door. Disconnect the flexible hose from the heater diverter air valve and move hose away from valve. Disconnect compressor duct from the flow control valve tee and move the duct away from tee, refer to Chapter 21, Compressor Bleed Air Heater - Maintenance Practices. Locate and remove the P3 line from the engine to the P3 filter assembly. Cap the open port at the P3 filter assembly (this step applies only to the older engine configuration). Ensure ignition and airplane bleed air is OFF. Perform the desalination or performance recovery wash, refer to the Pratt and Whitney Engine Maintenance Manual. Reconnect hose to heater diverter air valve. Reconnect the compressor duct at the flow control valve and secure, refer to Chapter 21, Compressor Bleed Air Heater - Maintenance Practices Remove the cap at the P3 filter assembly and reconnect the P3 line from the engine to the P3 filter assembly (this step applies only to the older engine configuration). Start engine and run at 80 percent Ng for one minute to dry engine. Shut down engine and close upper cowling doors.
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MODEL 208 MAINTENANCE MANUAL TURBINE BLADE WASH - MAINTENANCE PRACTICES 1.
General A.
2.
Turbine blade wash is accomplished to remove deposit build- up accumulated on turbine blades during normal operation.
Turbine Blade Wash A.
Wash Turbine Blades (Refer to Figure 201). NOTE:
Refer to Pilot's Operating Handbook for starter motor cycle limitations.
NOTE:
A minimum cool-down period of 40 minutes should be observed after engine running and prior to injecting rinse ßuid.
NOTE:
Compressor turbine blade washing is to be accomplished using water of drinking quality (potable) only at ambient temperatures of 36°F and above, and a potable water/methanol solution at ambient temperatures below 36°F. Consult applicable Engine Maintenance Manual for solution strengths according to ambient temperature.
(1) (2)
Remove one of the two igniter plugs. Install spray tube P/N PWC32271 in igniter plug boss and tighten spray tube Þnger tight. Ensure that arrow on tab of spray tube is directed toward engine reduction gearbox and is parallel with centerline of engine.
CAUTION: Delivery hose should be supported so as to prevent damage to the spray tube. (3) (4)
Connect a suitable cleaning rig hose to the spray tube adapter. Water/solution supply delivery pressure is 40 psi. Ensure engine ignition is off and aircraft bleed air is shut off.
CAUTION: Observe starter motor limitations. (5)
Carry out a 30 second motoring cycle, introducing the wash solution when the compressor attains a speed of approximately 5% Ng. NOTE:
Approximately one- half gallon of rinse solution will be passed through the compressor turbine during a 30 second cycle.
CAUTION: When using a water/methanol solution, perform an additional 30 seconds dry motoring cycle after each washing cycle to purge engine of volatile fumes. Ensure prescribed starter cooling periods are observed. (6) (7) (8)
Repeat washing cycle as required to remove contaminants from turbine blades. Remove spray tube assembly and reinstall igniter plug. Perform two consecutive engine dry motoring runs, ensuring that starter running time limitations are not exceeded.
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Trubine Blade Wash Spray Tube Assembly Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL WIRING HARNESS - MAINTENANCE PRACTICES 1.
General A.
2.
The engine wiring harness consists of cables to connect the following items of equipment: propeller overspeed governor and ITT harness (left front of engine), propeller tachometer generator (right front of engine), cabin bleed air heater ßow control valve (lower right side of engine), oil temperature sender (right side of engine), fuel control heater (right rear of engine), gas generator section tachometer generator (lower right side of engine), starter/generator (center top of engine accessory case), and ignition exciter leads (right engine mount truss).
Wiring Harness Removal/Installation A.
Remove Wiring Harness. (1) Disconnect all quick-disconnect connectors at the items of equipment mentioned above, tag connectors, cut sta-ties or loosen clamps, and remove wiring harness.
B.
Install Wiring Harness. (1) Reconnect all quick-disconnect connectors at the tagged locations of the items of equipment mentioned above and reinstall sta-ties and clamps.
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MODEL 208 MAINTENANCE MANUAL INERTIAL AIR SEPARATOR - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Inertial air separator maintenance practices consist of separator control linkage removal/installation and separator control rigging.
Inertial Air Separator Control Linkage Removal/Installation A.
Remove Inertial Air Separator Control Linkage (Refer to Figure 201). (1) Open upper cowling doors and remove lower cowling panels. (2) In engine compartment, disconnect actuating handle tube rod end (47) from lever arm (59). (3) Remove actuating handle tube rod end (47) from swivel (49). (4) In cabin, remove setscrew (55) from bezel (52), loosen clamp (66), and withdraw actuating handle tube (56) from instrument panel. (5) In engine compartment, disconnect actuating tube forward rod end (30) from actuating link (31). (6) Remove pivot bolt (41), actuating tube (42) and lever arm (59). (7) Remove bolt (20) and (33) from forward vane bellcrank (16) and aft vane bellcrank (34). Remove forward vane push-pull rod (21).
B.
Install Inertial Air Separator Control Linkage (Refer to Figure 201). (1) If removed, install inertial separator assembly. (2) Install bushings (62) in each end of lever arm (59), insert spacer (63) through bushings. (3) Align pivot holes in lever arm (59) with holes in mounting brackets and install pivot bolt (41). Secure with washer (18) and nut (64). (4) Connect actuating tube forward rod end (30) of actuating tube (42) to actuating link (31). (5) In cabin, insert actuating handle tube (56) through bezel (52), clamp (66), vapor shield (65), and bushing (57) into engine compartment. Tighten setscrew (55). (6) In engine compartment, install actuating handle tube rod end (47) on swivel (49) and connect to lever arm (59). (7) Install forward vane push-pull rod (21) to forward vane bellcrank (16) and to aft vane bellcrank (34). (8) Rig inertial air separator control linkage. (9) Install cowling panel sections and close upper cowl doors.
Inertial Air Separator Control Linkage Rigging A.
Rig Inertial Separator Control Linkage (Refer to Figure 201). (1) With aft vane (12) in NORMAL (closed) position and actuating handle tube (56) and actuating handle (53) fully forward and locked (handle vertical), install actuating handle tube rod end (47) at top of lever arm (59). Leave nut (44) loose. (2) Install actuating tube forward rod end (30) and actuating tube aft rod end (43) on actuating tube (42) and install tube between lever arm (59) and actuating link (31). Leave nut (46) loose. (3) Rotate eccentric bushing (24) so that actuating link (31) bottoms against slot in stop link (22) before going beyond centerline of the two links. (4) Bend tabs on lock washer (23) to prevent rotation of eccentric bushing (24). (5) Holding forward vane (13) in fully open position against stop block, attach forward vane push-pull rod (21) between forward vane bellcrank (16) and aft vane bellcrank (34). Ensure clevis (19) attaches to forward vane bellcrank (16) and aft rod end (36) aligns with inside lever arm hole in aft vane bellcrank. (6) In cabin, pull out actuating tube handle to BYPASS position and check that forward vane (13) seats Þrmly against airfoil assembly (7) inside inertial separator. If forward vane (13) does not contact airfoil (7), disconnect clevis (19) from eyebolt in forward vane bellcrank (16) and screw in eyebolt to shorten bellcrank arm. Recheck for vane-to-airfoil contact. (7) Conversely, if forward vane (13) contacts airfoil (7) in separator before aft vane (12) opens fully against its stop (BYPASS position), screw eyebolt out of forward vane bellcrank (16) to lengthen its lever arm and recheck. (8) Tighten all nuts on ends of adjustable push rods.
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Inertial Separator and Control Linkage Figure 201 (Sheet 1)
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Inertial Separator and Control Linkage Figure 201 (Sheet 2)
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Inertial Separator and Control Linkage Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL INERTIAL AIR SEPARATOR - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the inertial air separator in a serviceable condition.
Task 71-60-00-220 2.
Inertial Air Separator Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the inertial air separator.
B.
Special Tools (1) None
C.
Access (1) Open the upper cowling doors. Refer to Engine Cowling and Nose Cap - Maintenance Practices. (2) Remove the lower cowling panels. Refer to Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do the Inertial Air Separator Detailed Inspection (Refer to Figure 601). (1) Put the inertial separator vanes in the bypass mode. (2) Use a flashlight and a mirror to do an inspection of all rivets on both sides of the forward and the aft inertial separator vanes to find if rivets are loose or missing. (a) If there are loose or missing rivets, do the Rivet Replacement in this section. (b) If there are no loose or missing rivets, do the Restore Access in this section.
E.
Rivet Replacement. (1) Remove the inertial separator and forward and/or aft vanes. (2) Examine the inertial air separator vanes for loose or missing rivets. (3) If there are missing rivets, do the steps that follow: (a) Examine the induction air plenum and the engine inlet areas for the missing rivet(s) and debris. (b) Examine the engine for foreign object damage (FOD). Refer to the Pratt & Whitney PT6A114 Engine Maintenance Manual. (4) Replace all loose or missing rivets with same type rivet. If necessary, it is possible to use the next size larger diameter rivet. (5) Reinstall the forward and/or aft vanes. (6) Thoroughly clean any debris from the inside of the inertial separator. (7) Reinstall the inertial separator.
F.
Restore Access (1) Install the lower cowling panels. Refer to Engine Cowling and Nose Cap - Maintenance Practices. (2) Close the upper cowling doors. Refer to Engine Cowling and Nose Cap - Maintenance Practices. End of task
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Inertial Air Separator Detailed Inspection Figure 601 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ENGINE DRAIN LINES - MAINTENANCE PRACTICES 1.
General A.
2.
A series of drain lines is installed to vent ßuids overboard or to catch and retain for later disposal. Drains provided are: propeller shaft (1), forward (2) and aft (3) combustion chamber, fuel manifold (7), fuel control (14), EPA can (6), oil collection can and drain valves.
Engine Drain Lines Removal/Installation A.
Due to the simplicity of the engine drains installation, refer to Figure 201 for a guide to removal and installation.
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Engine Drains Installation Figure 201 (Sheet 1)
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Engine Drains Installation Figure 201 (Sheet 2)
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Engine Drains Installation Figure 201 (Sheet 3)
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CHAPTER
ENGINE FUEL AND CONTROL
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
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73-Title 73-List of Effective Pages 73-Record of Temporary Revisions 73-Table of Contents
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
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Issue Date
By
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS ENGINE FUEL AND CONTROL - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-00-00 Page 1 73-00-00 Page 1 73-00-00 Page 1 73-00-00 Page 1
FUEL DISTRIBUTION - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-10-00 Page 1 73-10-00 Page 1
FUEL SCAVENGE SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EPA Can Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-10-10 Page 201 73-10-10 Page 201 73-10-10 Page 201 73-10-10 Page 201
FUEL PUMP - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Pump Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-11-00 Page 601 73-11-00 Page 601 73-11-00 Page 601
FUEL FLOW INDICATING - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-30-00 Page 101 73-30-00 Page 101
FUEL FLOW INDICATING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Flow Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Flow Transmitter Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Conical Seal Installation (Fuel Transmitter Line Assembly Leakage) . . . . . . . . . . . . .
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FUEL TOTALIZER INDICATING SYSTEM - TROUBLESHOOTING. . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-32-00 Page 101 73-32-00 Page 101
FUEL TOTALIZER INDICATING SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Signal Conditioner Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Totalizer Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-32-00 Page 201 73-32-00 Page 201 73-32-00 Page 201 73-32-00 Page 201 73-32-00 Page 201
SHADIN MINI-FLOW FUEL TOTALIZER - DESCRIPTION AND OPERATION . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-33-00 Page 1 73-33-00 Page 1 73-33-00 Page 1
SHADIN MINI-FLOW FUEL TOTALIZER - REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shadin Mini-Flow Fuel Totalizer Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . Signal Interface Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73-33-00 Page 401 73-33-00 Page 401 73-33-00 Page 401 73-33-00 Page 401
73 - CONTENTS © Cessna Aircraft Company
Page 1 of 1 Apr 1/2010
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MODEL 208 MAINTENANCE MANUAL ENGINE FUEL AND CONTROL - GENERAL 1.
Scope A.
2.
This chapter describes those units, components and associated mechanical systems or electrical circuits which furnish or control fuel to the engine and fuel indicating system.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Conical Seal
VSF1015N8B 0.500 (-8 Tube Size)
Voi-Shan VSI Corporation 8463 Hiquera Street P.O. Box 515 Culver City, CA 90230
To seal fuel transmitter Þttings.
3.
DeÞnition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief deÞnition of the sections incorporated in this chapter is as follows: (1) The section on distribution covers maintenance practices for that portion of the system used to distribute fuel to the engine. (2) The section on indicating covers maintenance practices for that portion of the system used to indicate fuel ßow and fuel quantity.
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MODEL 208 MAINTENANCE MANUAL FUEL DISTRIBUTION - DESCRIPTION AND OPERATION 1.
Description and Operation A.
The engine fuel system consists of an oil-to-fuel heater, an engine-driven fuel pump, a fuel control unit, a ßow divider and dump valve, a dual fuel manifold with 14 simplex nozzles, a fuel ßow indicator system, and two fuel drain lines. The system provides fuel ßow to satisfy the speed and power demands of the engine. Fuel from the airplane fuel reservoir is delivered to the oil-to-fuel heater which is essentially a heat exchanger which utilizes heat from the engine lubricating oil system to preheat fuel before it is delivered to the fuel control unit. A fuel temperature-sensing oil bypass valve regulates fuel temperature by either allowing oil to ßow through heater circuit or bypass it to engine oil tank.
B.
Fuel from oil-to-fuel heater then enters engine-driven fuel pump chamber through a 74-micron inlet screen. The inlet screen is spring-loaded and should it become blocked, the increase in differential pressure will overcome the spring and allow unÞltered fuel to ßow into pump chamber. The pump increases fuel pressure and delivers it to fuel control unit via a 10-micron Þlter in pump outlet. A bypass valve and cored passages in pump body enables unÞltered high pressure fuel to ßow to fuel control unit in the event outlet Þlter becomes blocked.
C.
The fuel control unit consists of a fuel metering section, a temperature compensating section, and a gas generator (Ng) pneumatic governor. The fuel control unit determines the proper fuel schedule to provide power required as established by power lever input. This is accomplished by controlling speed of compressor turbine. The temperature compensating section alters the acceleration fuel schedule to compensate for fuel density differences at different fuel temperatures, especially during engine start. The temperature compensator alters the acceleration fuel schedule of fuel control unit to compensate for variations in compressor inlet air temperature. Engine characteristics vary with changes in inlet air temperature, and the acceleration fuel schedule must, in turn, be altered to prevent compressor stall and/or excessive turbine temperatures. The power turbine governor, located in propeller governor housing, provides power turbine overspeed protection in the event of propeller governor failure. This is accomplished by limiting fuel to gas generator. During reverse thrust operation, maximum power turbine speed is controlled by power turbine governor.
D.
The ßow divider schedules the metered fuel, from fuel control unit, between primary and secondary fuel manifolds. The fuel manifold and nozzle assemblies deliver fuel to combustion chamber through ten primary and four secondary fuel nozzles. During engine start, metered fuel is delivered initially to the primary nozzles, with secondary nozzles cutting in above a preset valve. All nozzles are operative at idle and above.
E.
When fuel cutoff valve in fuel control unit closes during engine shutdown, both primary and secondary manifolds are connected to a dump valve port and residual fuel in manifolds is allowed to drain into EPA fuel reservoir can attached to Þrewall where it can be drained daily. NOTE:
For more information pertaining to the airplane engine fuel components, not covered in this chapter, refer to the Pratt & Whitney maintenance manual listed in the List of Publications in the front of this manual. For proper fuel grades and speciÞcations, refer to Chapter 12, Fuel - Servicing.
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MODEL 208 MAINTENANCE MANUAL FUEL SCAVENGE SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Description A.
3.
This section covers removal and installation procedures for the fuel scavenging system.
The fuel scavenge system consists of combustion chamber drain lines, Environmental Protection Agency (EPA) fuel reservoir can, a drain valve and overboard vent tube assembly.
EPA Can Removal/Installation A.
Remove EPA Fuel Reservoir Can (Refer to Figure 201). (1) Open left upper cowling door to gain access to EPA fuel reservoir can (6). (2) Disconnect vent tube assembly (1) from EPA fuel reservoir can (6). (3) Disconnect combustion chamber drain line (8) from EPA fuel reservoir can (6). (4) Remove (four each) bolts (4) and washers (5) securing EPA fuel reservoir can (6). (5) Remove EPA fuel reservoir can (6) from brackets (2) and (3).
B.
Install EPA Fuel Reservoir Can (Refer to Figure 201). (1) Position EPA fuel reservoir can (6) to brackets (2) and (3). (2) Install (four each) bolts (4) and washers (5) securing EPA fuel reservoir can (6). (3) Connect combustion chamber drain line (8) from EPA fuel reservoir can (6). (4) Connect vent tube assembly (1) to EPA fuel reservoir can (6). (5) 0n completion of maintenance, close left-hand upper cowling door.
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Fuel Scavenge System Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FUEL PUMP - INSPECTION/CHECK 1.
General A.
2.
This section provides inspection criteria for the fuel pump.
Fuel Pump Inspection A.
Fuel (1) (2) (3) (4) (5) (6) (7)
Pump Inspection (Refer to Figure 601). Remove the drain line and fitting from the fuel pump seal drain port. Insert a suitable cotton swab into the port, wiping the internal passage wall above the threads. Remove the swab and examine it for presence of a reddish brown (iron-oxide) stain. If none is evident, the unit may continue in service. Install fitting and drain line. If stain is found, remove pump from engine. Examine the pump face, input coupling side, for residue from fretting corrosion (iron-oxide deposits) in the area shown in Inspection Area B. If none is present the pump may be reinstalled. If a fretting corrosion residue is present, this indicates excessive spline wear and the pump must be forwarded to an approved facility for overhaul/repair. Install a replacement pump in accordance with paragraph B.
CAUTION: Do not remove the coupling as this may cause chipping of the splines and/or loss of circlip retention with detrimental effect. The circlip allows very limited movement of the coupling. B.
Fuel Pump Installation. (1) Prior to installation of a time-continued or replacement pump, accomplish the following: (a) Check for presence of oil on the coupling splines of the removed pump, and ensure that the hole in the gearbox drive shaft (Inspection Area C) is free of obstruction. This hole may be cleaned with a length of MS9226-03 lockwire, or equivalent. (b) Ensure that the shaft and coupling splines are free of grease or similar lubricant. (2) The shaft and coupling splines are to be liberally coated with engine oil only. Install pump on the engine. (3) This coupling inspection should be entered in the airplane log book.
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MODEL 208 MAINTENANCE MANUAL
Engine Driven Fuel Pump Inspection Figure 601 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FUEL FLOW INDICATING - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in troubleshooting the fuel ßow indicating system. Refer to Figure 101.
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MODEL 208 MAINTENANCE MANUAL
Fuel Flow Indicator Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FUEL FLOW INDICATING - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
This section provides removal and installation instructions for the fuel ßow indicating system and installation procedures for conical seals in the fuel transmitter line.
The fuel ßow indicating system measures rates of ßow by means of an inline transmitter, which provides an electrical signal, and a fuel ßow indicator, which displays the fuel ßow information.
Fuel Flow Indicator Removal/Installation
CAUTION: Handle fuel ßow indicator with care during removal/installation to prevent damage to indicator.
4.
A.
Remove Fuel Flow Indicator (Refer to Figure 201). (1) Turn all electrical power OFF. (2) Loosen screw (2) only. (3) Slide fuel ßow indicator (1) aft, out of instrument panel. (4) Disconnect electrical connector (3). (5) Remove fuel ßow indicator (1).
B.
Install Fuel Flow Indicator (Refer to Figure 201). (1) Position fuel ßow indicator (1) in instrument panel. (2) Connect electrical connector (3) to fuel ßow indicator (1). (3) Position fuel ßow indicator (1) in instrument panel and tighten screw (2).
Fuel Flow Transmitter Removal/Installation
CAUTION: Handle fuel ßow transmitter with care during removal/installation to prevent damage to transmitter. A.
Remove Fuel Flow Transmitter (Refer to Figure 201, Sheet 2). (1) Turn all electrical power OFF. (2) Open upper right cowling door to gain access to fuel ßow transmitter (2). (3) Disconnect electrical connector (5) at fuel ßow transmitter (2). (4) Remove safety wires (1) and fuel lines (3) and (4). (5) Remove fuel ßow transmitter (2). NOTE:
B.
Cap all open fuel lines and Þttings.
Install Fuel Flow Transmitter (Refer to Figure 201, Sheet 2). (1) Position fuel ßow transmitter (2) between fuel lines (3) and (4).
CAUTION: Ensure arrow on fuel ßow transmitter is pointing in proper direction. Arrow should indicate direction of ßow from pump to ßow divider. (2) (3) (4) (5) (6)
Remove caps and connect fuel lines (3) and (4). Safety wire fuel lines (3) and (4) to fuel ßow transmitter (2). Connect electrical connector (5) to fuel ßow transmitter (2). Close cowling door. Perform engine operational check. Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual.
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Fuel Flow Indicator Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Fuel Flow Indicator Installation Figure 201 (Sheet 2)
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5.
Conical Seal Installation (Fuel Transmitter Line Assembly Leakage)
CAUTION: Do not install seal into line assembly. Seal shall be installed on male Þtting only. Do not lubricate the seal or threads. NOTE:
A.
Conical seals are allowed only at the fuel transmitter line connections and are only to be used if the cause of the leak is a damaged male Þtting (slight deformation) of the fuel transmitter. Any leakage caused by other damage, such as cracked Þttings or deformed line assembly ßares, cannot be corrected with conical seals, but must be corrected by replacement of the torque transducer and/or line assemblies.
Installation of Conical Seal (Refer to Figure 201, Sheet 2). (1) Disconnect airplane battery. (2) Remove safety wire (1) from Þttings. (3) Disconnect fuel lines (3) and (4). (4) Inspect line ßares and fuel transmitter Þttings for cracks or deformities. If cracks or deformities are found, replace lines/transmitter as required. (5) lf lines or transmitter are replaced connect lines and check for leaks in accordance with step (10). If no leak is found, it will not be necessary to install conical seals. (6) If leak is found install conical seal onto male ßare portion of Þtting. The ßats on conical seals are designed to provide proper positioning of seal onto the male Þtting to prevent cocking of the seal that could cause ßow restriction. (7) Thread line assembly nut onto the male ßared Þtting several turns with Þngers until joint is snug. If line assembly nut cannot be tightened snugly with Þnger torque, disassemble and correct problem to prevent damage to conical seal. (8) Conical seals are subject to cold creep, therefore, a double tightening procedure is required. Tighten line assembly nuts from 575 to 625 inch-pounds. Torque values are higher than torque values for lines assembly nuts without conical seals installed. Allow 5 minutes elapsed time for cold creep to occur and then recheck torque after 30 minutes, and retighten if required. (9) Safety wire Þttings. (10) Check for leaks by running the engine momentarily to full power to conÞrm that operation is normal, and that no leakage is evident upon making a post shutdown inspection.
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MODEL 208 MAINTENANCE MANUAL FUEL TOTALIZER INDICATING SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in troubleshooting the fuel totalizer system. Refer to Figure 101.
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MODEL 208 MAINTENANCE MANUAL
Fuel Totalizer Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL FUEL TOTALIZER INDICATING SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
3.
4.
This section contains description, operation, removal and installation instructions for the fuel totalizer indicating system.
Description and Operation A.
The fuel totalizer system is available to aid the pilot in monitoring total fuel consumed each ßight. The totalizer system utilizes fuel ßow indicator system circuitry and an additional signal conditioner and totalizer indicator to display pounds of fuel consumed. The indicator is located on right side of instrument panel and is labeled FUEL CONSUMED, POUNDS. A Þve-digit display is centered in the indicator. The display can be reset by means of display reset pushbutton mounted to the left of display. A display pushbutton lock, located below pushbutton, can be rotated to engage the button and prevent inadvertent zeroing of display. The fuel ßow indicator and fuel totalizer systems are protected by pull-off type circuit breakers, labeled FUEL FLOW and FUEL TOTAL, respectively.
B.
In operation, the fuel ßow transducer in the standard fuel system generates an electrical signal which is proportional to fuel ßow rate and transmits this signal to standard fuel ßow indicator, where it is registered in pounds per hour. The voltage output from fuel ßow indicator is then sent to a totalizer signal conditioner where it is conditioned and sent to the fuel totalizer indicator, which displays its value to the pilot in total pounds of fuel consumed. On Airplanes 20800001 Thru 20800034, the totalizer signal conditioner is located on the cowl deck lower surface, forward of left and right fuel quantity indicators. On Airplanes 20800035 and On and 208B0001 and On, the totalizer signal conditioner is located on the aft side of the Þrewall.
Signal Conditioner Removal/Installation A.
Remove Signal Conditioner (Refer to Figure 201). (1) Ensure electrical power is OFF. (2) Remove nine screws (3) securing cowl deck access cover (2) to cowl deck. (3) Remove cowl deck access cover (2). (4) Disconnect electrical connector (6) from signal conditioner (5). (5) On Airplanes 20800001 Thru 20800034, remove (four each) screws (1) and nuts (4) securing signal conditioner (5) to lower surface of cowl deck. (6) On Airplanes 20800035 and On and 208B0001 and On, remove (two each) screws (13) securing conditioner (5) to Þrewall stiffener (14). (7) Remove signal conditioner up through cowl deck access hole.
B.
Install Signal Conditioner (Refer to Figure 201). (1) On Airplanes 20800001 Thru 20800034, place signal conditioner through cowl deck access hole and position it to lower surface of cowl deck. Install (four each) screws (1) and nuts (4) securing signal conditioner (5). (2) On Airplanes 20800035 and On and 208B0001 and On, place signal conditioner through cowl deck access hole and position it to mounting nutplates located in Þrewall stiffener (14). Install (two each) screws (13) securing signal conditioner (5) to Þrewall stiffener (14). (3) Connect electrical connector (6) to signal conditioner (5). (4) Position cowl deck access cover on upper surface of cowl deck. (5) Install nine screws (3) securing cowl deck access cover (2) to cowl deck.
Fuel Totalizer Removal/Installation A.
Remove Fuel Totalizer (Refer to Figure 201). (1) Ensure electrical power is OFF. (2) Remove nine screws (12) securing right removable ßight panel and slide panel aft to gain access to nuts (8). (3) Loosen lower clamp screws and remove and identify electrical leads (10) and (11). (4) Remove (two each) screws (7) and nuts (8) securing fuel totalizer indicator (9) to ßight panel. (5) Slide fuel totalizer indicator (9) aft out of ßight panel.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Fuel Totalizer System Installation Figure 201 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL B.
Install Fuel Totalizer (Refer to Figure 201). (1) Slide fuel totalizer indicator (9) forward through right ßight panel. (2) Install (two each) screws (7) and nuts (8) securing fuel totalizer indicator to ßight panel. (3) Identify and install electrical leads (10) and (11), and secure by tightening lower clamp screws. (4) Slide right removable ßight panel forward and secure with nine screws (12).
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SHADIN MINI-FLOW FUEL TOTALIZER - DESCRIPTION AND OPERATION 1.
2.
General A.
The Shadin Mini-Flow Fuel Totalizer system is a highly accurate digital fuel management system.
B.
System components include a microprocessor based, panel mounted indicator and a signal interface connected to the existing fuel ßow transducer and indicator.
Description and Operation A.
The Shadin MIni-Flow Fuel Totalizer provides complete fuel management information under real ßight conditions. Manual data entry is not required except for initial input of fuel on board information. It is connected to a GPS or Loran-C receiver for navigation data such as ground speed and estimated time enroute.
B.
Fuel management functions, fuel ßow, fuel used, fuel remaining and time remaining are independent of the navigation function and will continue to function without navigation information.
C.
This system provides the following functions. (1) SpeciÞc Range - Fuel consumption is calculated in Nautical Miles per Pound (NM/Lb.) or in Nautical Miles per 10 pound (NM/10 Lb.) of fuel burned. This provides an indication of cruise speed efÞciency. (2) Fuel to Destination - Fuel needed to reach destination is calculated under real wind conditions by using destination information provided by Loran-C or GPS waypoints. (3) Fuel Reserve - Indicates the amount of fuel remaining after reaching destination as selected on Loran-C or GPS receiver. (4) Endurance - The system calculates the time in hours and minutes based on the fuel onboard and fuel consumption. (5) Fuel Remaining - Indicates the amount of fuel remaining onboard. (6) Fuel Used - Displays amount of fuel used since last fuel entry. (7) Not Enough Fuel - Display digits will ßash when the rotary switch is in Fuel to Destination position and the Fuel to Destinations is more than the Fuel Remaining. The Fuel Remaining digits will show a negative sign followed by the amount of fuel short to reach the destination. (8) Fuel Reserve Will Be Used - Display digits will ßash when the rotary switch is in either Fuel to Destination or Reserve Fuel. This warning alerts the pilot to prevailing conditions requiring the use of the 45 minute Fuel Reserve or part of it. (9) Fuel Flow - The system provides a digital readout of the fuel ßow per hour to the nearest pound.
D.
Fuel ßow is always displayed in a dedicated window. All other functions are displayed in a separate window.
E.
A non volatile memory keeps basic setting, fuel remaining and fuel used information stored during power shut down.
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Page 1 Oct 15/1999
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SHADIN MINI-FLOW FUEL TOTALIZER - REMOVAL/INSTALLATION 1.
2.
3.
General A.
This section contains removal and installation procedures for the Shadin Mini-Flow Fuel Totalizer system.
B.
System components include the microprocessor based indicator mounted in the upper right hand side of the instrument panel and a signal interface mounted on the inboard side of the glove box.
Shadin Mini-Flow Fuel Totalizer Removal/Installation A.
Remove Fuel Totalizer (Refer to Figure 401). (1) Disengage circuit breaker FUEL TOTL on left circuit breaker panel. (2) Remove screws from right removable ßight panel. (3) Gain access to fuel totalizer by moving removable ßight panel aft. (4) Remove screws holding fuel totalizer to removable ßight panel. (5) Slide fuel totalizer aft out of removable ßight panel. (6) Disconnect electrical connector. (7) Remove fuel totalizer from airplane.
B.
Install Fuel Totalizer (Refer to Figure 401). (1) Connect electrical connector to fuel totalizer. (2) Place fuel totalizer on aft side of removable ßight panel and secure with screws. (3) Secure removable removable ßight panel with screws. (4) Engage circuit breaker FUEL TOTL on left circuit breaker panel.
Signal Interface Removal/Installation A.
Remove Signal Interface (Refer to Figure 401). (1) Disengage circuit breaker FUEL TOTL on left circuit breaker panel. (2) From inside glove box, remove screws holding signal interface to inboard side of glove box. (3) Reach under instrument panel and remove signal interface and mounting bracket by bringing them downward and outward. (4) Identify and disconnect electrical connectors. (5) Remove signal interface from airplane.
B.
Install Signal Interface (Refer to Figure 401). (1) Identify and connect electrical connectors to signal interface. (2) Reach under the instrument panel to position the signal interface and mounting bracket against inboard side of glove box. (3) From inside glove box, secure signal interface to glove box with screws. (4) Engage circuit breaker FUEL TOTL on left circuit breaker panel.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Shadin Mini-Flow Fuel Totalizer System Figure 401 (Sheet 1)
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Page 402 Oct 15/1999
CHAPTER
IGNITION
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
74-00-00
Page 1
Aug 1/1995
74-10-00
Pages 201-203
Aug 1/1995
74-20-00
Pages 201-203
Aug 1/1995
74-21-00
Page 201
Mar 1/2012
74-Title 74-List of Effective Pages 74-Record of Temporary Revisions 74-Table of Contents
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By
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MODEL 208 MAINTENANCE MANUAL
CONTENTS IGNITION - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
74-00-00 74-00-00 74-00-00 74-00-00
Page 1 Page 1 Page 1 Page 1
IGNITION EXCITERS - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ignition Exciter Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
74-10-00 Page 201 74-10-00 Page 201 74-10-00 Page 201 74-10-00 Page 201
IGNITION CABLES - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ignition Cable Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
74-20-00 Page 201 74-20-00 Page 201 74-20-00 Page 201 74-20-00 Page 201 74-20-00 Page 201
SPARK IGNITERS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
74-21-00 Page 201 74-21-00 Page 201
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL IGNITION - GENERAL 1.
Scope A.
2.
This chapter describes those components which supply ignition to the engine.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Digital Voltmeter
Model 8100A
John Fluke Mfg. Co. 6920 Seaway Blvd. Everett, WA 98206
To check electrical continuity in electrical circuits.
Fluorocarbon Spray Lubricant
MS-122
Miller-Stevenson Chemical Co. Danbury, CT 06810
To coat threads of ignition exciter connectors.
Petroleum Solvent
ASM3160
Commercially available
To clean spark igniters.
Methyl Alcohol
TT-I-735
Commercially available
To clean spark igniters.
3.
DeÞnition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief deÞnition of the topics incorporated in this chapter is as follows: (1) The section on ignition exciters provides description, operation and removal/installation procedures for the ignition exciters. (2) The section on ignition cables provides description, operation and removal/installation procedures for the ignition cables. (3) The section on spark igniters provides description, operation and removal/installation procedures for the spark igniters.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL IGNITION EXCITERS - MAINTENANCE PRACTICES 1.
2.
General A.
The ignition system consists of an ignition exciter box, two high tension leads, two spark igniters, an ignition monitor light on the annunciator panel, an ignition switch and a starter switch. Electrical energy from the exciter box, mounted on the right engine mount truss, is transmitted via two high tension leads to two igniters, at four and nine o'clock positions on the gas generator case adjacent to the fuel manifold. The ignition system is normally energized only during engine start.
B.
Ignition is controlled by two switches, located on left switch and circuit breaker panel, labeled IGNITION and STARTER. The ignition switch has two positions, ON and NORMAL. The NORMAL position arms ignition so that ignition will be obtained when the starter switch is activated. The NORMAL position is used during all ground starts and during air starts with starter assist. The ON position of switch provides continuous ignition, regardless of the position of starter switch. This position is used for air starts without starter assist and during encounters with heavy precipitation before induction system inertial separator is placed in bypass position.
C.
The starter switch has three position, OFF, START, and MOTOR. The OFF position shuts off ignition system current and is the normal position for all operations except engine start. The START position energizes engine ignition system provided ignition switch is in the NORMAL position. After the engine starts during a ground or air start, starter switch must be manually positioned to OFF.
D.
A green annunciator panel light, labeled IGNITION ON, will illuminate when ignition energy is being applied to igniters. A Þve amp push-to-reset type circuit breaker is provided to protect the ignition system primary wiring circuit.
E.
The ignition exciter is a sealed unit containing electronic components encased in an epoxy resin. The unit is energized during the engine starting sequence to initiate combustion in the combustion chamber and as desired during ßight. The exciter transforms 28 VDC input to a high voltage output through solid state circuitry, a transformer, and diodes.
Tools, Equipment and Materials A.
3.
Refer to Ignition - General for a list of required tools, equipment and materials.
Ignition Exciter Removal/Installation A.
Remove Ignition Exciter (Refer to Figure 201 ).
WARNING: Residual voltage in ignition exciter may be dangerously high. Ensure ignition is switched off and system has been inoperative for at least six minutes before removing any ignition components. Always disconnect coupling nuts at ignition exciter end Þrst. Always use insulated tools to remove cable coupling. (1) (2)
Ensure battery switch is OFF. Remove bus voltage supply cable connector (12) from input connector on ignition exciter (13).
CAUTION: Do not allow ignition cable braiding or ferrule to rotate when removing coupling nuts. (3) (4) B.
Remove two high-tension lead coupling nuts (11) from output connectors on ignition exciter (13) and remove leads. Remove four nuts (10) and bolts (14) securing exciter to bracket and remove exciter.
Install Ignition Exciter (Refer to Figure 201). (1) Secure ignition exciter (13) to bracket (4) with four bolts (14) and nuts (10). Ensure input connector is facing up.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Ignition Exciter Installation Figure 201 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CAUTION: Do not allow any lubricant to come in contact with central conductors of exciter connectors. Contact with conductors may result in a high resistance path, which could generate heat and oxidation. (2)
Lightly coat threads of ignition exciter connectors with ßuorocarbon spray lubricant.
CAUTION: Do not allow ignition cable braiding ferrules to rotate when screwing on coupling nuts. (3) (4)
Install high-tension leads to exciter connectors and torque high-tension lead coupling nuts (11) Þnger tight plus 45 degrees. Safety wire nuts. Connect bus voltage supply cable connector (12) to ignition exciter (13) and torque coupling nut Þnger tight plus 45 degrees. Safety wire nut.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL IGNITION CABLES - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
The two individual ignition cable assemblies carry electrical energy output from ignition exciter to engine-mounted spark igniters. Each lead assembly consists of an electrical lead contained in a ßexible metal braiding. Coupling nuts at each end of the assembly facilitate connection to respective connectors on ignition exciter and spark igniters. Mounting ßanges for attachment to engine Þreseals are brazed onto ßexible braiding.
Tools, Equipment and Materials A.
4.
Maintenance practices for the ignition cable consist of removal and installation.
For a list of required tools, equipment and materials, refer to Ignition - General.
Ignition Cable Removal/Installation A.
Remove Ignition Cables (Refer to Figure 201). (1) Remove engine cowling Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (2) Remove inertial air separator. Refer to Chapter 71, Inertial Air Separator - Maintenance Practices. (3) Isolate electrical power supply from ignition system.
CAUTION: When unscrewing coupling nuts (5) of ignition cables, do not allow braiding ferrules or igniter to turn at same time. (4) (5) (6) B.
Remove safety wire from coupling nuts (5) and remove nuts from left and right ignition cables (9) and ignition exciter (3). Remove nuts and bolts securing Þreseal mounting ßanges (6) to Þreseals. Remove nuts and bolts from clamps (10) and withdraw ignition cables (9) from induction air box. Clamps should remain on cables when same cables are to be reinstalled.
Install Ignition Cables (Refer to Figure 201). (1) Thread ignition cables (9) through clearance holes in Þre seals (7) and (8) (six o’clock position) and install clamps (10). Secure with nuts and bolts. (2) Secure Þreseal mounting ßanges (6) to rear Þreseal (7) and front Þreseal (8) with mounting bolts and nuts. NOTE:
Bolt heads should be located on induction air inlet side of Þreseal.
CAUTION: Under no circumstances is any lubricant containing grease or silicone or lubricants such as petrolatum to be used on any ignition components. CAUTION: Do nto allow any lubricant to come in contact with central conductors of cables. Contact with conductors may result in a high resistance path which could generate heat and oxidation. CAUTION: Do not apply lubricant on any cables having teßon insulated sleeves. (3)
Lightly spray insulated end of cables having rubber insulated sleeves with ßuorocarbon spray lubricant.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Ignition Cables Installation Figure 201 (Sheet 1)
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Page 202 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CAUTION: When screwing on coupling nuts (5) of ignition cables, do not allow braiding ferrules or igniter to turn at same time. (4) (5) (6) (7) (8)
Connect coupling nuts (5) at ends of ignition cables to spark igniters (1) and high-tension output connectors (2) on ignition exciter (3). Screw couplings onto mating threads by hand, ensuring that no binding occurs between coupling nut and cable. Tighten coupling nuts and torque Þnger tight plus 45 degrees. Safety wire coupling nuts. Install inertial air separator. Refer to Chapter 71, Inertial Air Separator - Maintenance Practices. Install engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
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Page 203 Aug 1/1995
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL SPARK IGNITERS - MAINTENANCE PRACTICES 1.
General A.
For the engine spark ignitors procedures, use the Pratt and Whitney Engine Maintenance Manual. Refer to the Introduction, Supplier Publication List.
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76 CHAPTER
ENGINE CONTROLS
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
76-10-00
Pages 1-2
Dec 3/2001
76-10-00
Pages 601-603
Jun 1/2011
76-10-01
Pages 201-214
Apr 1/2010
76-10-01
Page 601
Jun 1/2011
76-10-02
Pages 501-509
Jun 1/2011
76-10-03
Pages 501-508
Mar 1/2001
76-Title 76-List of Effective Pages 76-Record of Temporary Revisions 76-Table of Contents 76-List of Tasks
76 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY MAINTENANCE MANUAL
RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS ENGINE CONTROLS - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
76-10-00 76-10-00 76-10-00 76-10-00
Page 1 Page 1 Page 1 Page 1
ENGINE CONTROLS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Controls Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
76-10-00 Page 601 76-10-00 Page 601 76-10-00 Page 601
QUADRANT ASSEMBLY AND CONTROLS - MAINTENANCE PRACTICES . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Quadrant Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Control Cable Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Condition Control Cable Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Speed Control Cable Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Power Control Cable Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Power Lever Frangible/Shear Wire Removal/Installation . . . . . . . . . . . . .
76-10-01 Page 201 76-10-01 Page 201 76-10-01 Page 201 76-10-01 Page 201 76-10-01 Page 209 76-10-01 Page 210 76-10-01 Page 210 76-10-01 Page 212
QUADRANT ASSEMBLY AND CONTROLS - INSPECTION/CHECK . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Power Lever Annunciator Light (EPL) Operational Check . . . . . . . . . . . .
76-10-01 Page 601 76-10-01 Page 601 76-10-01 Page 601
ENGINE CONTROL RIGGING - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Power Control Annunciator Light Switch Adjustment . . . . . . . . . . . . . . . . Emergency Power Control Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Control Lever Reverse Gas Generator N g Pickup Adjustment . . . . . . . . . . . Fuel Control Lower Idle Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Condition Control Lever High Idle Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller Speed Control Lever Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
76-10-02 Page 501 76-10-02 Page 501 76-10-02 Page 501 76-10-02 Page 501 76-10-02 Page 504 76-10-02 Page 506 76-10-02 Page 506 76-10-02 Page 509 76-10-02 Page 509
PT6A-114/-114A ENGINE RIGGING - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . Rig Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
76-10-03 Page 501 76-10-03 Page 501
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MODEL 208 MAINTENANCE MANUAL
LIST OF TASKS 76-10-00-720
Engine Controls Functional Check
76-10-00 Page 601
76-10-01-710
Emergency Power Lever Annunciator Light (EPL) Operational Check
76-10-01 Page 601
76 - LIST OF TASKS © Cessna Aircraft Company
Page 1 of 1 Jun 1/2011
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ENGINE CONTROLS - DESCRIPTION AND OPERATION 1.
Scope A.
2.
DeÞnition A.
3.
This chapter describes those controls which govern operation of the engine. These include the power control lever, fuel condition control lever, propeller speed control lever and emergency power control lever.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief deÞnition of the topics incorporated in this chapter is as follows: (1) Quadrant assembly provides description, operation and removal/installation procedures for the throttle quadrant. (2) Engine control rigging provides instructions for rigging the airplane.
Description and Operation A.
Engine controls for the Model 208 airplane consist of a power control lever, fuel condition control lever, and propeller speed control lever. Power and fuel condition control levers are engine controls, while the third controls propeller speed. A fourth lever, emergency power control, is available for use in the event of primary power control system failure within the hydro-pneumatic metering and governing system. (1) A power control lever, located in the cockpit, is connected through an airframe linkage to a cambox assembly mounted on the front of the Fuel Control Unit (FCU) at the rear of the engine. The power control lever controls the engine power through full range from maximum takeoff power, back through idle, then to maximum reverse power. The power control lever schedules fuel in the forward power range, while directly controlling propeller blade angle and fuel in the reverse power range. (a) Control of reverse blade angles is accomplished through and externally grooved back ring provided with the propeller. Feedback ring motion is proportional to propeller blade angle, and picked up by a carbon block running in the feedback ring. The relationship between the axial position to the feedback ring and propeller blade angle is used to maintain control of blade angle from idle to full reverse. (2) A fuel condition control lever, located in the cockpit, is connected through an airframe linkage to a combined lever and stop mechanism at the top of the fuel control unit. The fuel condition control lever is then connected by an FCU linkage to a fuel shutoff lever on the side of the unit. Additionally, the control lever and stop function as an idle stop for the FCU control rod. (a) The fuel condition control lever consists of three positions: high idle, low idle and cutoff. When the cockpit control lever is in high idle position, the FCU is at 65 percent gas generator speed (Ng). When in low idle position, the FCU is at 52 percent Ng. (3) A propeller speed control lever, located in the cockpit, is connected through an airframe linkage to the propeller governor mounted on the top forward section of the engine. The cockpit propeller speed control lever regulates propeller blade angle from maximum RPM position (1900 RPM) to full feather. Propeller speed control is achieved through a propeller governor which controls blade angle. (4) An Emergency Power Lever (EPL), located in the cockpit, is connected through an airframe linkage to a lever mechanism on the aft side of the FCU. The EPL schedules fuel in the event of primary control failure. A red warning light on the annunciator panel, EMERGENCY POWER LEVER, indicates the EPL is not in normal stowed position. (5) The following airplanes incorporate frangible/shear wire, which is installed from the EPL to the pedestal cover: • Airplanes 20800351 and On. • Airplanes 208B0920 and On. • Airplanes 20800001 thru 20800350 Incorporating SK208-142. • Airplanes 208B0001 thru 208B0919 Incorporating SK208-142.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL (6)
Use of the EPL is indicated by breakage of this frangible/shear wire. After use of the EPL, appropriate engine maintenance actions must be accomplished in accordance with the Pratt & Whitney Maintenance Manual, Chapter 71-00.
B.
Propeller control is provided by a standard governor, Beta control valve, and overspeed governor. The propeller governor performs two functions. Under normal ßight conditions, the governor acts as a Constant Speed Unit (CSU), maintaining cockpit selected propeller speed by varying blade angle to match the load to the engine torque in response to changing ßight conditions. During low airspeed operation, the propeller governor is used to select the required propeller blade angle. When the engine is operating in the Beta control range, engine power is adjusted by the FCU and power turbine governor to limit power turbine speed (N1) at a speed approximately 4 percent lower than that set on the propeller governor.
C.
During constant speed operation, desired propeller speed is set by the propeller speed control lever located in the cockpit. Lever movement is translated by an externally mounted control lever on the propeller governor, into a change in compression in the governor speeder spring. Increased compression results in increased propeller speed. (1) Propeller underspeed condition is sensed by governor ßyweights, allowing the pilot valve to move down under the inßuence of the speeder spring, thus permitting high pressure oil to ßow to the propeller servo piston, decreasing propeller blade angle. If, during constant speed operations, the blade angle decreases below a speciÞed positive blade angle, the propeller feedback ring moves forward, actuating the Beta control valve and preventing the ßow of high pressure oil to the propeller servo piston. (2) Constant speed operation allows the Beta control valve to act as a hydraulic Þne pitch stop for the propeller, maintaining positive propeller blade angles during ßight. (3) Propeller overspeed condition results in ßyweights forcing the pilot valve to move up against the speeder spring tension, allowing propeller oil to bleed to the reduction gearbox. Propeller feather spring and blade counterweights then cause the propeller pitch to coarsen, decreasing propeller speed.
D.
Propeller low blade angles and reverse angles are scheduled by a cam-box and cable system connected to the power control lever and controlled from the cockpit. During low airspeed and low engine power, propeller blades are at hydraulic Þne pitch stop. The propeller will exhibit an underspeed condition, relative to speed selection on propeller speed control lever. Movement of the power control lever, below idle position into the Beta control range, moves the cam arrangement in the cam-box. Through the Beta control linkage, the Beta control valve opens, permitting oil to ßow to the servo piston, thus decreasing propeller blade angle. (1) As the propeller blade angle approaches the control lever angle selection, the feedback ring causes the propeller reversing lever to close the Beta control valve. Oil ßow is terminated to the propeller and further pitch change is prevented. Resulting blade angle is proportional to the selected power control lever position. Low blade angles and reverse may only be selected when the propeller is underspeeding, relative to the speed selected by the propeller speed control lever. When a high forward ßight speed is achieved, low blade angles and reverse cannot be selected, as the propeller will windmill at selected speed, even when the gas generator is at IDLE.
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MODEL 208 MAINTENANCE MANUAL ENGINE CONTROLS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the engine controls in a serviceable condition.
Task 76-10-00-720 2.
Engine Controls Functional Check A.
General (1) This task gives the procedures to do a functional check of the engine controls.
B.
Special Tools (1) MIL-L-7870 lubricant, or equivalent.
C.
Access (1) None
D.
Do a Engine Power Lever Detailed Inspection. For illustrations of the power lever, refer to Quadrant Assembly and Controls - Maintenance Practices, Figure 202. (1) Examine the power control cable from the power lever to the cam box input lever for security of installation, wear, corrosion, routing, evidence of damage, and deterioration. (a) Examine the cable for security at the firewall jam nut. (b) Examine all cable attach brackets for condition and security. (2) Examine the rubber seals at the end of the flex cable for condition, security, and deterioration. (3) Disconnect the rod end from the lever arm. Refer to Quadrant Assembly and Controls Maintenance Practices. (a) Wipe the rod end clean using a clean lint-free cloth. (b) Examine the rod end bearing for condition, corrosion, pitting, security, and freedom of movement. (c) Lubricate the rod end ball with MIL-L-7870 oil or an equivalent. (4) Examine the lever arm for condition and security. (5) Connect the rod end to the lever arm. Refer to Quadrant Assembly and Controls - Maintenance Practices. (6) Examine the rod end clevis at the engine power control lever for condition, security, and freedom of movement. (7) Adjust the friction lock to ON. (a) Examine the control for positive locking action. (8) Adjust the friction lock OFF. (a) Move the control from the IDLE position to the FULL POWER position. (b) Make sure that there is freedom of operation.
E.
Do a Fuel Condition Control Lever Detailed Inspection. For illustrations of the fuel condition control lever, refer to Quadrant Assembly and Controls - Maintenance Practices, Figure 202. (1) Examine the control cable from the cockpit lever to the governor lever for security of installation, wear, corrosion, routing, evidence of damage, and deterioration. (a) Examine the cable for security at the firewall jam nut. (b) Examine all cable attach brackets for condition and security. (2) Examine the rubber seals at the end of the flex cable for condition, security, and deterioration. (3) Disconnect the rod end from the lever arm. Refer to Quadrant Assembly and Controls Maintenance Practices. (a) Wipe the rod end clean using a clean lint-free cloth. (b) Examine the rod end bearing for condition, corrosion, pitting, security, and freedom of movement. (c) Lubricate the rod end ball with MIL-L-7870 oil or an equivalent. (4) Examine the lever arm for condition and security.
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MODEL 208 MAINTENANCE MANUAL (5) (6) (7) (8)
Connect the rod end to the lever arm. Refer to Quadrant Assembly and Controls - Maintenance Practices. Examine the rod end clevis at the fuel condition control lever for condition, security, and freedom of movement. Adjust the friction lock to ON. (a) Examine the control for positive locking action. Adjust the friction lock OFF. (a) Move the control from the CUTOFF position to the HIGH IDLE position. (b) Make sure that there is freedom of operation. (c) Make sure that the lever on the fuel control contacts the HIGH IDLE stop.
F.
Propeller Speed Control Lever Detailed Inspection. For illustrations of the propeller speed control lever, refer to Quadrant Assembly and Controls - Maintenance Practices, Figure 202. (1) Examine the propeller speed control cable from the cockpit lever to the governor lever for security of installation, wear, corrosion, routing, evidence of damage, and deterioration. (a) Examine the cable for security at the firewall jam nut. (b) Examine all cable attach brackets for condition and security. (2) Examine the rubber seals at the end of the flex cable for condition, security, and deterioration. (3) Disconnect the rod end from the lever arm. Refer to Quadrant Assembly and Controls Maintenance Practices. (a) Wipe the rod end clean using a clean lint-free cloth. (b) Examine the rod end bearing for condition, corrosion, pitting, security, and freedom of movement. (c) Lubricate the rod end ball with MIL-L-7870 oil or an equivalent. (4) Connect the rod end to the lever arm. Refer to Quadrant Assembly and Controls - Maintenance Practices. (5) Examine the rod end clevis at the cable connection to the propeller speed control lever for condition, security, and freedom of movement. (6) Adjust the friction lock to ON. (a) Examine the control for positive locking action. (7) Adjust the friction lock OFF. (a) Move the control from the FEATHER position to the HIGH RPM position. (b) Make sure that there is freedom of operation. (c) Make sure that the lever arm contacts the HIGH RPM stop.
G.
Do a Functional Check of the Engine Power Control Lever. (1) Start the engine. Refer to the Model 208 Pilot’s Operating Handbook and Approved Airplane Flight Manual. (2) Operate the engine at IDLE for five minutes to let the temperatures stabilize. (3) Put the propeller speed control lever to the MAX forward position. (4) Move the power control lever from IDLE, then slowly aft to the REVERSE position. (5) Make sure that the propeller RPM increases to peak, then decreases 10 RPM to 15 RPM before the gas generator (Ng) begins to increase from idle. (a) If necessary, do the Power Control Lever Reverse Gas Generator (Ng) Pickup Adjustment. Refer to Engine Control Rigging - Adjustment/Test.
H.
Do a Functional Check of the Fuel Condition Control Lever. (1) Make sure that the engine temperature is stabilized. (2) Make sure that the power control lever is at IDLE. (3) Make sure that the fuel condition control lever is at LOW IDLE. (4) Put the generator switch to the ON position. (a) Adjust the electrical load to 40 Amperes. (5) Put the BLEED AIR HEAT switch to the ON position. (6) Turn the CABIN HEAT TEMP control knob to the full HOT position. (7) Make sure that the Ng is from 52 percent to 55 percent. (a) If the Ng is not at the specified range, do the Fuel Control Lower Idle Adjustment. Refer to Engine Control Rigging - Adjustment/Test. (8) Put the fuel condition control lever to the HIGH IDLE position.
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MODEL 208 MAINTENANCE MANUAL (9)
Make sure that the Ng is from 64 percent to 66 percent. (a) If the Ng is not at the specified range, do the Fuel Control High Idle Adjustment. Refer to Engine Control Rigging - Adjustment/Test. (10) Move the fuel condition control lever to the LOW IDLE position. I.
Do a Functional Check of the Propeller Control Lever Reverse Gas Generator Ng Pickup. (1) Make sure that the engine temperature is stabilized. (2) Put the propeller speed control lever to the MAX RPM position. (3) Put the power lever at the IDLE position (at the detent gate). (4) Slowly move the power control lever to the REVERSE position. (a) Make sure that the propeller RPM increases to peak; then decreases 10 to 15 RPM before the gas generator (Ng) begins to increase from idle. (b) If necessary, do the Propeller Speed Control Lever Adjustment. Refer to Engine Control Rigging - Adjustment/Test. (5) Move the power lever to the IDLE position. (6) Move the propeller speed control lever to the MIN RPM position. (7) Shut down the engine. Refer to the Model 208 Pilot’s Operating Handbook and Approved Flight Manual.
J.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL QUADRANT ASSEMBLY AND CONTROLS - MAINTENANCE PRACTICES 1.
General A.
2.
3.
This section provides removal and installation instructions for the quadrant assembly, power control cable, fuel condition control cable, propeller speed control cable and emergency power control cable. For rigging procedures, refer to Engine Controls Rigging - Adjustment/Test.
Control Quadrant Removal/Installation A.
Remove Control Quadrant (Refer to Figure 201 and Figure 202). (1) Make sure the electrical power is OFF. (2) Remove the nut and washer from the center of the elevator trim control wheel. (3) Remove the elevator trim control wheel from the airplane. (4) Remove the nut, screw, washer and support, from the pedestal structure and quadrant cover. (5) Remove screws that attach the emergency power control knob, propeller speed control lever knob, fuel condition lever knob and the flap control lever knob to their respective levers. (6) Remove each knob from its lever. (7) Remove the flap control lever knob. (a) Rotate the knob clockwise. (b) Remove knob from the airplane. (8) Remove the quadrant cover from the pedestal by removing the screw and washer. (9) Remove the friction knob, flap control lever and associated parts. (10) Disconnect the electrical wires from the go-around switch. (11) Disconnect the fuel condition control clevis, propeller speed control clevis, power control clevis and emergency power control cable by removing cotter pins, washers and clevis pins. (a) Retain all hardware, except cotter pins, for reinstallation. (12) Remove the quadrant attach screws, washers, nuts that secure the quadrant assembly. (13) Remove the quadrant assembly from the airplane.
B.
Install Control Quadrant (Refer to Figure 201 and Figure 202). (1) Set the quadrant assembly onto the pedestal and secure using quadrant attach screws, washers and nuts. (2) Connect the fuel condition control clevis, propeller speed control clevis, power control clevis and emergency power control cable clevis to control lever arms using nuts, washers and new cotter pins. (3) Connect the electrical wires to the go-around switch. (4) Install friction knob, flap control lever, and associated parts. Refer to Chapter 27, Elevator Trim - Maintenance Practices. (5) Install the quadrant cover on the pedestal and secure using screws and washers. (6) Install the flap lever knob by rotating knob clockwise. (7) Attach the emergency power control knob, propeller speed control lever knob, fuel condition lever knob and the flap control lever knob to their respective levers using screws. (8) Set the support to the pedestal structure and quadrant cover and secure using screw, washer and nut. (9) Attach the elevator trim control wheel using washer and nut.
Power Control Cable Removal/Installation A.
Remove Power Control Cable (Refer to Figure 202). (1) Open the upper cowling doors. (2) In the cockpit. remove the carpet and upholstery panels as required to access the cable retention area aft of the control pedestal and at the cabin side of the firewall where the control cables route. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery - Maintenance Practices. (3) Get access to power control cable connection at power control lever located inside control pedestal. Refer to Control Quadrant Removal/Installation. (4) In engine compartment, loosen jam nuts at power control cable bracket and fuel control unit. (5) Remove the cotter pin, washer and clevis pin from power control clevis and disconnect cable.
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MODEL 208 MAINTENANCE MANUAL
Control Quadrant Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Control Quadrant Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Control Quadrant Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL
Engine Control Cable Installation Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL
Engine Control Cable Installation Figure 202 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL
Engine Control Cable Installation Figure 202 (Sheet 3)
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Engine Control Cable Installation Figure 202 (Sheet 4)
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MODEL 208 MAINTENANCE MANUAL (6) (7) (8) (9) B.
4.
Remove power control clevis and jam nut. Cut and remove the safety wire, cotter pin, nut, washers, retaining washer, spacer and lever arm bolt, from lever arm. (a) Retain hardware, except cotter pin, for reinstallation. Cut and remove the tie straps and remove the cabin side power control jam nut. Withdraw and remove the power control cable from the cabin by pulling through to the engine compartment.
Install Power Control Cable (Refer to Figure 202). (1) Insert the power control cable through the hole in the firewall and route to the pedestal in the cabin. (a) Install jam nut over power control cable in cabin area. (2) Put the power control cable through the power control cable bracket and connect it and the cable rod end to the lever arm using bolt, spacer, washers, retaining washer, and nut. Do not install cotter pin, or safety wire at this time. (3) Tighten the jam nuts using your fingers. (4) In the cabin, route power control cable up through pedestal structure to the control quadrant. (5) Connect the power control cable clevis to the power lever arm in the quadrant assembly and secure using clevis pin, washer, and new cotter pin. (6) Tighten jam nut and install tie-straps. (7) Adjust the power control linkage. Refer to Engine Control Rigging - Adjustment/Test. (8) Tighten jam nuts, install new cotter pin and safety wire nut. (9) Install previously removed carpet and upholstery panel. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery Maintenance Practices. (10) Close upper cowling.
Fuel Condition Control Cable Removal/Installation A.
Remove Fuel Condition Control Cable (Refer to Figure 202). (1) Open the upper cowling doors. (2) In the cockpit, remove the carpet and upholstery panel as required to access the cable retention area aft of control pedestal and at cabin side of firewall where control cables route. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery - Maintenance Practices. (3) Cut the tie straps and remove the cabin side fuel condition control cable jam nut. (4) Get access to the fuel condition control cable connection at the fuel condition control lever inside control pedestal. Refer to Control Quadrant Removal/Installation. (5) Remove the cotter pin, washer and clevis pin from the fuel condition control clevis. (6) Disconnect the fuel condition control cable and remove the fuel condition control clevis and jam nut. (7) In engine compartment, remove the screw and washer securing fuel condition control cable to the fuel condition control bracket. (8) Cut and remove the safety wire and remove the cotter pin, nut, washers, retaining washer, and lever arm bolt, to disconnect the fuel condition control cable rod end from the lever arm. (a) Retain all hardware, except the cotter pin, for reinstallation. (9) Remove the fuel condition control cable through to engine compartment.
B.
Install Fuel Condition Control Cable (Refer to Figure 202). (1) Insert the fuel condition control cable, through the hole in firewall and route to the pedestal in the cabin. Install jam nut over cable in cabin area. (2) Attach the fuel condition control cable to the fuel condition control cable bracket, using washer and screw. (3) Connect the fuel condition control cable and rod end to the lever arm, using bolt, washers, retaining washer, and nut. Do not install cotter pin or safety wire at this time. (4) In the cabin, route fuel condition control cable up through pedestal structure to quadrant. (5) Connect the fuel condition control clevis to the fuel condition control lever arm in quadrant assembly and secure using a clevis pin, washer, and new cotter pin. (6) Tighten jam nut and install tie straps.
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MODEL 208 MAINTENANCE MANUAL (7) (8) (9)
Adjust the fuel condition control linkage. Refer to Engine Control Rigging - Adjustment/Test. Tighten jam nuts, install new cotter pin and safety wire nut. Install previously removed carpet and upholstery panel. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery Maintenance Practices. (10) Close upper cowling. 5.
6.
Propeller Speed Control Cable Removal/Installation A.
Remove Propeller Speed Control Cable (Refer to Figure 202). (1) Open the upper cowling doors. (2) In the cockpit, remove carpet and upholstery panels as required to access the cable retention area aft of control pedestal and at cabin side of firewall where the control cables route through. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery - Maintenance Practices. (3) Cut tie straps and remove the cabin side propeller speed control cable jam nut. (4) Get access to propeller speed control cable connection, at the propeller speed control lever inside the control pedestal. Refer to Control Quadrant Removal/Installation. (5) Remove the cotter pin, washer and clevis pin from the propeller speed control clevis and disconnect the cable. (6) Remove the power control clevis and jam nut. (7) In the engine compartment, loosen the jam nuts on the propeller speed control cable bracket at the propeller governor. (8) Cut and remove the safety wire and remove the cotter pin, nut, retaining washers and, washers, spacer and lever arm bolt, to disconnect propeller speed control cable, rod end, from lever arm. (a) Retain hardware, except cotter pin, for reinstallation. (9) Remove the remaining clamps securing the propeller speed control cable to the engine as required. (10) Withdraw and remove the propeller speed control cable from the cabin by pulling through the firewall to the engine compartment.
B.
Install Propeller Speed Control Cable (Refer to Figure 202). (1) Put the propeller speed control cable through the hole in the firewall and route to the pedestal in the cabin. Install a jam nut over the propeller speed control cable, in the cabin area. (2) Put the propeller speed control cable, through the propeller speed control bracket and connect the propeller speed control cable rod end to the lever arm using the lever arm bolt, spacer, retaining washer, washers, and nut. Do not install the cotter pin or safety wire at this time. (3) Tighten the jam nuts with finger pressure. (4) In the cabin, route the propeller speed control cable, up through the pedestal structure to the quadrant. (5) Connect the propeller speed control cable clevis, to the propeller speed lever arm in the quadrant assembly and secure with a clevis pin, washer, and new cotter pin. (6) Tighten the jam nut and install the tie straps. (7) Adjust the propeller speed control linkage. Refer to Engine Control Rigging - Adjustment/Test. (8) Install previously removed clamps. Tighten jam nuts, install new cotter pin and safety wire nut. (9) Install previously removed carpet and upholstery panel. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery Maintenance Practices. (10) Close upper cowling.
Emergency Power Control Cable Removal/Installation A.
Remove the Emergency Power Control Cable (Refer to Figure 202). (1) Open the upper cowling doors. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. (2) In the cockpit, remove carpet and upholstery panels as necessary to access the cable retention area aft of the control pedestal and at the cabin side of the firewall where the control cables are found. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery - Maintenance Practices.
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MODEL 208 MAINTENANCE MANUAL (3) (4)
Cut the tie straps, then remove the cabin side emergency power control cable jam nut. Get access to the emergency power control cable connection at the emergency power control lever inside control pedestal. Refer to Control Quadrant Removal/Installation. (5) Remove cotter pin, washer and clevis pin from the emergency power control clevis, then disconnect the emergency power control cable. (6) Remove the emergency power control clevis and jam nut. (7) For Airplanes 20800396 and On and 208B1171 and On, do the steps that follow: (a) In the engine compartment, remove the nut, washer and bolt from the mount that attaches the emergency power control cable to the airplane. (8) For Airplanes 20800001 thru 20800395 and 208B0001 thru 208B1170, do the steps that follow: (a) In the engine compartment, cut the safety wire from the nuts that attach the inner and outer brackets. (b) Remove the nuts that attach the inner and outer brackets, then remove the inner bracket and outer bracket that attach the emergency power control cable from the airplane. (9) For Airplanes 20800131 and On and 208B0087 and On that have the bushing installed in the manual override arm, do the steps that follow: (a) Cut and remove the safety wire. (b) Remove the cotter pin, nut, retaining washer, washers, lever arm bolt and bushing. (c) Keep all of the hardware other than the cotter pin for installation. (10) Disconnect the emergency power control cable and rod end from the manual override arm. (11) Pull the emergency power control cable through the firewall toward the engine compartment to remove the emergency power control cable from the airplane. B.
Install the Emergency Power Control Cable (Refer to Figure 202). (1) Put the emergency power control cable through the hole in the firewall, and route it to the pedestal in the cabin. (2) Install the jam nut over the emergency power control cable in the cabin area. (3) For Airplanes 20800131 and On and 208B0087 and On that have the bushing installed in the manual override arm, do the steps that follow: (a) Install the lever arm bolt, retaining washer, washers, nut and install bushing. (b) Connect the emergency power control cable and rod end to the manual override arm. (c) Do not install the cotter pin or wire at this time. (4) For Airplanes 20800001 thru 20800395 and 208B0001 thru 208B1170, do the steps that follow: (a) Attach the emergency power control cable, to inner bracket and outer bracket using nuts. (b) Safety the nuts with wire. (5) For Airplanes 20800396 and On and 208B1171 and On, do the steps that follow: (a) In the engine compartment, install the nut, washer and bolt in the mount that attaches the emergency power control cable to the airplane. (6) In the cabin, install the emergency power control cable up through the pedestal structure to the control quadrant. (7) Connect the emergency power control cable clevis to the emergency power control lever arm in the control quadrant assembly, then attach it with the clevis pin, washer, and new cotter pin. (8) Tighten the jam nut. (9) Install the tie straps. (10) Tighten the jam nuts, then install the new cotter pin and safety the nut with wire. (11) Install the carpet and upholstery panel. Refer to Chapter 6, Access Plates and Panels Identification - Description and Operation and Chapter 25, Cabin Upholstery - Maintenance Practices. (12) Adjust the emergency power control linkage. Refer to Engine Control Rigging - Adjustment/Test. (13) Close the upper cowling door.
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MODEL 208 MAINTENANCE MANUAL
7.
Emergency Power Lever Frangible/Shear Wire Removal/Installation A.
Remove Frangible/Shear Wire (Refer to Figure 203). NOTE:
(1)
(2)
(3) (4) B.
If the frangible/shear wire is missing or broken, determine if engine limitations have been exceeded. If so, perform all engine maintenance actions required by the Pratt & Whitney Maintenance Manual, Chapter 71-00.
The following airplanes incorporate frangible/shear wire, which is installed from the Emergency Power Lever (EPL) to the pedestal cover: • Airplanes 20800351 thru 20800395 • Airplanes 208B0920 thru 208B1170 • Airplanes 20800001 thru 20800350 Incorporating SK208-166 • Airplanes 208B0001 thru 208B0919 Incorporating SK208-142. The following airplanes do not incorporate frangible/shear wire installed from the Emergency Power Lever (EPL) to the pedestal cover: • Airplanes 20800396 and On • Airplanes 208B1171 and On • Airplanes 20800001 thru 20800395 Incorporating SK208-166 • Airplanes 208B0001 thru 208B1170 Incorporating SK208-166. Make sure the engine is not running. Cut and remove the frangible/shear wire.
Install Frangible/Shear Wire (Refer to Figure 203). (1) If the wire was found to be broken and engine limitations have been exceeded, then perform all engine maintenance actions required by the engine maintenance manual. Refer to the Pratt & Whitney Maintenance Manual, Chapter 71-00. (2) Make sure the EPL is in the normal position. (3) Cut to length one MS20995CY15 or C489003 frangible/shear wire. (4) Install the wire from the EPL to the pedestal cover screw, using the double-twist method.
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Emergency Power Lever Frangible/Shear Wire Figure 203 (Sheet 1)
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Emergency Power Lever Frangible/Shear Wire Figure 203 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL QUADRANT ASSEMBLY AND CONTROLS - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the quadrant assembly and controls in a serviceable condition.
Task 76-10-01-710 2.
Emergency Power Lever Annunciator Light (EPL) Operational Check A.
General (1) This task gives the procedures to do an operational check of the emergency power lever (EPL) annunciator light.
B.
Special Tools (1) None
C.
Access (1) None
D.
Do an Operational Check of the Emergency Power Lever Annunciator Light. (1) Apply electrical power to the airplane. (2) If installed, cut and remove the frangible/shear wire from the EPL to the pedestal cover screw. Refer to Emergency Power Lever Frangible/Shear Wire Removal/Installation in Quadrant Assembly and Controls - Maintenance Practices. (3) Move the emergency power control lever through its full range of travel forward of the NORM gate, then back to the NORM gate. (4) Make sure that the EMERGENCY POWER LEVER annunciator light stays on. NOTE: (5) (6)
Move the emergency power control lever aft of the NORM gate. Make sure that the EMERGENCY POWER LEVER annunciator light goes off. NOTE:
(7) (8) (9)
The IDLE stop position is forward of NORM gate.
The normal stowed position is aft of the NORM gate.
If adjustments are necessary, do the Emergency Power Control Annunciator Light Switch Adjustment. Refer to Engine Control Rigging - Adjustment/Test. If removed, install the frangible/shear wire from the EPL to the pedestal cover. Refer to Emergency Power Lever Frangible/Shear Wire Removal/Installation in Quadrant Assembly and Controls - Maintenance Practices. Remove electrical power from the airplane.
E.
Restore Access (1) None End of task
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MODEL 208 MAINTENANCE MANUAL ENGINE CONTROL RIGGING - ADJUSTMENT/TEST 1.
General A.
2.
This section provides adjustment and rigging procedures for engine controls and associated annunciators.
Emergency Power Control Annunciator Light Switch Adjustment A.
Adjust Emergency Power Control Annunciator Light Switch. (1) Verify engines are powered OFF. (2) Verify electrical power is applied to airplane. (3) Verify EMERGENCY POWER LEVER annunciator light remains illuminated when moving emergency power control lever through full travel range, forward of NORM gate, then back to NORM gate. NOTE: (4)
Move emergency power control lever aft of NORM gate and verify EMERGENCY POWER LEVER annunciator light extinguishes. NOTE:
(5)
3.
IDLE stop position is forward of NORM gate.
Normal stowed position is aft of NORM gate.
Adjust EMERGENCY POWER LEVER annunciator light switch, if required. (a) Remove applicable quadrant components. Refer to Control Quadrant Removal/ Installation. (b) Loosen screw securing EMERGENCY POWER LEVER annunciator switch to switch plate and adjust per steps 2.A.(3) and (4). (c) Reinstall control quadrant components. Refer to Control Quadrant Removal/Installation.
Emergency Power Control Rigging A.
Emergency Power Control Rigging Procedures (Refer to Figure 501 and Figure 502). NOTE: (1)
(2) (3)
(4)
Emergency power control cable linkages must allow sufficient travel to permit fuel control unit manual override arm full travel from OFF position to MAX position.
The following airplanes incorporate frangible/shear wire, which is installed from the Emergency Power Lever (EPL) to the pedestal cover: • Airplanes 20800351 and On. • Airplanes 208B0920 and On. • Airplanes 20800001 thru 20800350 Incorporating SK208-142. • Airplanes 208B0001 thru 208B0919 Incorporating SK208-142. If installed, cut and remove the frangible/shear wire from the EPL to the pedestal cover screw. Refer to Emergency Power Lever Frangible/Shear Wire Removal/Installation in Chapter 76, Quadrant Assembly and Controls - Maintenance Practices. Place emergency power control lever at maximum power position by moving emergency power control lever forward until lever stops. (a) Visually check in engine compartment and verify fuel control unit manual override arm is against maximum speed adjustment screw. Place emergency power control lever at normal position by moving emergency power control lever aft of NORM gate until lever stops. NOTE:
(5)
Aft emergency power control lever travel is limited by fuel control internal stop. Normal position is aft of NORM gate.
Rig emergency power control to minimum ineffective range. (a) Verify EMERGENCY POWER LEVER annunciator light is functional prior to rigging.
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Emergency Power Lever Adjustment Figure 501 (Sheet 1)
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Fuel Control Unit Adjustment Figure 502 (Sheet 1)
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CAUTION: The emergency power control lever must be at the normal stowed position during engine start. If it is not at the stowed position, there will be an over temperature condition. (b) (c)
Verify emergency power control lever is in normal stowed position. Perform a normal engine start. Refer to Pilot’s Operating Handbook and Approved Airplane Flight Manual. (d) Beginning at IDLE position, move emergency power control lever forward of NORM gate, slowly advancing to takeoff power. 1 Verify minimum 0.25 inch ineffective range forward of NORM gate before Ng begins advancing from idle. (6) Adjust emergency power control travel at fuel control unit, if required. (a) Working from within engine compartment, cut safety wire, remove cotter pin and loosen rod end jam nut at fuel control unit manual override arm. (b) Adjust emergency power control travel at fuel control unit to achieve minimum 0.25 inch ineffective range forward of NORM gate. Repeat rigging of emergency power control to minimum ineffective range. (c) Tighten rod end jam nut, install new cotter pin through bolt and nut on fuel control unit manual override arm and safety wire. Refer to Model 208 Series Illustrated Parts Catalog for cotter pin part number. (7) Return emergency control power lever securely to IDLE side of NORM gate and verify EMERGENCY POWER LEVER annunciator illuminates. (8) Return emergency power control lever to NORM position and verify EMERGENCY POWER LEVER annunciator extinguishes. (9) Shut down engine. Refer to Pilot's Operating Handbook and Approved Airplane Flight Manual. (10) If required, install MS20995CY15 or C489003 frangible/shear wire from the EPL to the pedestal cover. Refer to Emergency Power Lever Frangible/Shear Wire Removal/Installation in Chapter 76, Quadrant Assembly and Controls - Maintenance Practices. 4.
Power Control Lever Reverse Gas Generator Ng Pickup Adjustment A.
Adjust Power Control Lever Reverse Gas Generator (Ng) Pickup (Refer to Figure 503). (1) Start engine, observing all operating limitations. Refer to Pilot’s Operating Handbook and Approved Airplane Flight Manual. (2) Operate engine at IDLE for five minutes, allowing temperatures to stabilize. (3) Place propeller speed control lever in MAX forward position. (4) Move power control lever from IDLE, then slowly aft to REVERSE position. (a) Verify propeller RPM increases to peak, then decreases 10 RPM to 15 RPM before gas generator (Ng) begins increasing from idle. Adjust control lever reverse gas generator (Ng) pickup as required. 1 Cut and remove safety wire on reverse gas generator pickup bolt. 2 Loosen jam nut while securing generator pickup bolt.
CAUTION: Reverse gas generator pickup bolt adjustment is sensitive and shall be adjusted in increments of one-eighth turn between adjustments. 3
(b)
Rotate reverse generator pickup bolt clockwise or counterclockwise in increments of one-eighth turn to achieve a minimum torque of 900 foot-pounds at MAX REVERSE. 4 Torque jam nut while securing generator pick up bolt. 5 Safety wire reverse gas generator pickup bolt. Refer to Chapter 20, Safetying Maintenance Practices. Shut down engine. Refer to Pilot’s Operating Handbook and Approved Flight Manual.
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Engine Control Rigging Figure 503 (Sheet 1)
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5.
Fuel Control Lower Idle Adjustment A.
Adjust Fuel Control Lower Idle (Refer to Figure 504). (1) Start engine, observing all operating limitations. Refer to Pilot's Operating Handbook and Approved Flight Manual.
(2) (3) (4) (5) (6) (7)
NOTE:
Low idle maximum has been approved at 55 percent for all engine configurations.
NOTE:
Do not allow Ng to drop below 52 percent. Advance power lever as required.
(a) Operate engine at idle for five minutes, allowing temperatures to stabilize. (b) Advance power lever as required to achieve 52 to 55 percent Ng. Position generator switch to ON and adjust electrical load to 40 Amperes. Position BLEED AIR HEAT switch to ON. Rotate CABIN HEAT TEMP control to full HOT. Position fuel condition control lever to LOW IDLE. Position power control lever to IDLE position, forward and against detent gate. Verify Ng is 52 percent to 55 percent. (a) If 52 percent to 55 percent Ng is not achieved, adjust idle adjusting screw. Cut and remove safety wire on idle adjusting screw. 1 2 Using and Allen key, hold idle adjusting screw securely to prevent movement and release torque on nut.
CAUTION: Idle speed adjustment is sensitive and shall be adjusted in increments of one-eighth turn between idle speed checks. 3
(8) (9) 6.
Rotate idle speed adjusting screw clockwise or counterclockwise in increments of one-eighth turn to increase or decrease idle speed. Tighten jam nut, but do not safety wire at this time. a If idle speed remains unchanged during adjustment, FCU arm is at pickup point and must be rerigged. Refer to Power Control Forward Linkage Rigging. Shut down engine. Refer to Pilot's Operating Handbook and Approved Airplane Flight Manual. Torque jam nut from 20 inch-pounds to 25 inch-pounds and safety wire nut. Refer to Chapter 20, Safetying - Maintenance Practices.
Fuel Condition Control Lever High Idle Adjustment A.
Adjust Fuel Condition Control Lever High Idle (Refer to Figure 504). (1) Start engine, observing all operating limitations. Refer to Pilot's Operating Handbook and Approved Airplane Flight Manual. (2) Operate at idle for five minutes, allowing temperatures to stabilize. (3) Position power control lever to IDLE. (4) Position generator to ON and adjust electrical load to 40 Amperes. (5) Position BLEED AIR HEAT switch to OFF. (6) Position fuel control lever to HIGH IDLE. (a) If 64 to 66 percent Ng is not achieved, adjust high idle stop bolt. 1 Cut and remove safety wire on high idle stop bolt.
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Fuel Control Linkage Figure 504 (Sheet 1)
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Propeller Speed Linkage Adjustment Figure 505 (Sheet 1)
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CAUTION: Idle adjustment is sensitive and shall be adjusted in increments of one-eighth turn between idle checks. 2
(7) (8) 7.
Rotate high idle stop bolt, in increments of one-eighth turn, clockwise to increase or counterclockwise to decrease idle speed. Finger tighten jam nut, but do not safety wire at this time. NOTE:
Adjusting the high idle stop bolt will also effect the amount of lever cushion.
NOTE:
In order to maintain proper lever cushion, it may be necessary to adjust the nut on the upper cut off and flight idle linkage. Turning the nut clockwise will increase the idle speed, which will allow a higher idle stop and more lever cushion.
Shut down engine. Refer to Pilot's Operating Handbook and Approved Airplane Flight Manual. Torque jam nut from 20 inch-pounds to 25 inch-pounds and safety wire nut. Refer to Chapter 20, Safetying - Maintenance Practices.
Propeller Speed Control Lever Adjustment A.
Adjust Propeller Speed Control Lever (Refer to Figure 505). (1) Start engine, observing all operating limitations. Refer to Pilot's Operating Handbook and Approved Airplane Flight Manual. (2) Operate at idle for five minutes, allowing temperatures to stabilize. (3) Position propeller speed control lever to MAX PROP RPM. (4) Advance power control lever to achieve 1900 RPM. (5) Continue advancing power control lever slowly. Verify propeller governor maintains 1900 RPM, +10 or -10 RPM. (a) If 1900 RPM, +10 or -10 RPM, is not maintained, recheck propeller speed control lever rigging. Refer to Power Control Aft Linkage Rigging. (b) Return power control lever to IDLE position. (c) If propeller RPM exceeds 1900 RPM, +10 or -10 RPM, adjust propeller governor maximum RPM stop.
CAUTION: Propeller governor maximum rpm stop adjustment is sensitive and shall be adjusted in increments of one-eighth turn between adjustments. Cut and remove safety wire on propeller governor maximum RPM stop jam nut. Using an Allen wrench, hold screw securely to prevent movement and release torque on jam nut. 3 Tighten jam nut and safety wire nut. Refer to Chapter 20, Safetying - Maintenance Practices. (d) With power lever at IDLE, move propeller speed control lever to FEATHER position. Verify propeller feathers and propeller governor speed adjusting lever contacts feather stop. Adjust cable bulkhead fittings as required. Shut down engine. Refer to Pilot’s Operating Handbook and Approved Airplane Flight Manual. 1 2
(6) 8.
Engine Operating Limits A.
Engine operating limits are provided for both PT6A-114 and PT6A-114A engines. Refer to Chapter 71, Powerplant - Adjustment/Test.
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MODEL 208 MAINTENANCE MANUAL PT6A-114/-114A ENGINE RIGGING - ADJUSTMENT/TEST 1.
Rig Engine
CAUTION: The propeller reversing linkage will be damaged if the power lever is moved aft of the idle position with the engine not running and the propeller in feather. A.
Power Control Forward Linkage Rigging (1) Refer to Figure 501, disconnect the push-pull cable rear clevis from the propeller reversing cam.
Push-Pull Cable Rear Clevis Figure 501 (2)
Refer to Figure 502, disconnect the propeller governor interconnect rod from the governor air bleed link.
Propeller Governor Interconnect Rod Figure 502 (3)
Refer to Figure 502, disconnect the push-pull cable front terminal from the propeller reversing lever.
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(5) (6)
Pull the propeller push-pull cable forward to insure that the push-pull control terminal is against the internal stop of the low pitch adjuster stop. The collar should be visible through holes in the low pitch adjuster stop. Cable travel should be approximately 1.00 to 1.25 inches. Travel is measured by pushing the push-pull cable aft against the stop. Place a piece of masking tape on the push-pull cable at the low pitch adjuster stop. Pull the push-pull cable forward against the stop and measure the distance from the forward edge of the low pitch adjuster stop to the aft edge of the masking tape. To obtain the approximate 1.00 to 1.25 inches travel on the wire rope terminal, screw the locknut onto the adjuster stop and Þnger tighten only. Adjust at the low pitch adjuster stop (Refer to Figure 502) to obtain correct travel. NOTE:
The locknut must be tightened and fastened with lockwire after Þnal adjustments.
NOTE:
Adjustment is not routinely necessary.
Reconnect front terminal to reversing lever. Ensure small bushing is installed between terminal and lever. The rear of the Beta Valve clevis slot end should be ßush with the front face of the cap nut (Detail A of Figure 502). NOTE: (a)
If surging is encountered during ßight test, adjust the slot end FORWARD approximately 1/32 inch. To adjust, proceed as follows:
To adjust the slot FORWARD, turn the low pitch stop adjuster CLOCKWISE as viewed from the front of the airplane. To adjust the slot AFT turn the low pitch stop adjuster COUNTERCLOCKWISE. NOTE:
(7)
The low pitch adjuster as an assembly, is completely free to rotate, thus by itself makes no adjustment to amount of travel of the cable.
1 With the safety wire cut. 2 Loosen the locknut. 3 Hold the low pitch adjuster stop. (b) Torque low pitch adjuster locknut 150 - 200 inch pounds and lockwire. Pull the propeller reversing lever forward. Refer to Figure 503.
Propeller Reversing Lever Figure 503 (8) (9) (10) (11) (12) (13)
Hold the fuel governor air bleed link on the max forward stop. Adjust the propeller governor interconnect rod until the retaining bolt aligns with the outer hole of the governor air bleed link. Shorten the interconnect rod one-half turn and reconnect. Torque locknut 32 to 36 inch pounds and lockwire. Move the cambox input lever into the forward power range. Pull the propeller reversing cam Þrmly aft.
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MODEL 208 MAINTENANCE MANUAL (14) Maintaining forward pressure on the rear terminal to remove slack, adjust until hole in the rear clevis aligns with the middle hole in the propeller reversing cam. (15) Lengthen clevis one-half turn and reconnect. Do not install washer and cotter pin. Check for thread engagement in witness hole. (16) Move the power lever from IDLE to MAX POWER checking for free movement and cushion. Cushion should be one-eighth to one-quarter inch. B.
Power Control Aft Linkage Rigging (1) Referring to Figure 504, disconnect the power cable rod end at the cambox input lever and push the propeller reversing cam forward Þrmly.
Power Cable Attach Figure 504 (2) (3)
Move the cambox input lever COUNTERCLOCKWISE as far as possible without the reversing cam moving. This should allow the cam follower pin to rest in the track point of the reversing cam. (Detail A) With the setting as in step 2, check that the cambox lever is approximately in the 8 o'clock position. Adjustments should be made by removing the input lever and relocating on the serrated shaft. NOTE:
(4) (5) (6) (7) (8)
Adjustments are not routinely necessary.
Keeping the power lever cable rod end disconnected from the input lever, move the pedestal lever from IDLE to MAX POWER to IDLE assuring that there is no binding. Set the pedestal lever at IDLE and apply the friction lock. Maintaining forward pressure on the propeller reversing cam, align the rod end with the outer hole in the cambox input lever and connect the rod end to the lever. Release the friction lock and move the pedestal lever from IDLE to MAX POWER to IDLE making sure that the cam returns to track point. Place the fuel condition lever in CUTOFF.
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Interconnect Rod Figure 505 (9)
Refer to Figure 505, assure that the forward end of the interconnect rod is installed in the top hole of the FCU actuating lever. (10) Disconnect the interconnect rod from the FCU arm. (11) Lift the FCU arm gently CLOCKWISE until the pickup point is felt. The FCU arm should be approximately 20 degrees below horizontal. NOTE:
To reset FCU arm position, mark the FCU arm and serrated spacer with a marker pen. Loosen the FCU arm extension and adjust the serrated spacer. There are 24 serrations on the inner face of the spacer and 25 on the outer face.
(a)
There should be approximately 0.030 inch clearance between the cam follower pin and the FCU actuating lever. (b) There should be 0.060 to 0.100 inch clearance between the ßat on the interconnect rod and the FCU arm extension. (12) Refer to Figure 505, adjust the interconnect rod until the hole in the rod end aligns with the inner hole of the FCU arm. (13) Lengthen the interconnect rod an additional two turns and connect to the FCU arm. (14) Lift the FCU arm CLOCKWISE until the pick-up point is felt. Check the following: (a) There should be approximately 0.030 inch clearance between the cam follower pin and the FCU actuating lever. (b) There should be 0.060 to 0.100 inch clearance between the ßat on the interconnect rod and the FCU arm extension. NOTE:
If necessary, rotate the rod until the ßat surface is parallel with the FCU arm extension. Adjust the FCU arm per the NOTE SECTION of step 11 and repeat steps 12 through 14.
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Gas Generator Pickup Screw Figure 506 (15) Adjust the reverse gas generator pick-up screw (Refer to Figure 506) until the end is ßush with the FCU actuating lever. (16) Move the pedestal power lever from IDLE to MAX POWER. Check the following: (a) No binding is present throughout travel. (b) Refer to Figure 507, check that the FCU max power stop is contacted with the power lever in MAX POWER. (c) The cam follower pin is not bottomed out in the slot of the propeller reversing cam. (d) One-eighth inch gap (cushion) at the pedestal slot.
FCU Max Power Stop Figure 507
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Rear Clevis/Cam Follower Pin Figure 508 (17) Refer to Figure 508, disconnect the rear clevis. Move the power lever from IDLE to MAX REVERSE and observe the following: (a) No binding is present throughout travel. (b) The cam follower pin is not bottomed out in the slot of the propeller reversing cam. (c) One-eighth inch gap (cushion) at the pedestal slot. (18) Return the power lever to the IDLE position; reconnect the clevis and safety. C.
Fuel condition lever linkage rigging
Fuel Cutoff Lever Figure 509 (1) (2) (3) (4)
Refer to Figure 509, place the fuel cutoff lever in CUTOFF. Set the cockpit fuel condition lever in CUTOFF. Allow one-eighth to one-quarter inch cushion. With the cutoff lever in CUTOFF, align the cable terminal with the hole in the fuel cutoff lever and connect. Move the cockpit fuel condition lever from CUTOFF to HIGH and back to CUTOFF to insure there is no binding.
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Move the cockpit lever to LOW IDLE. The fuel cutoff lever should be approximately vertical as shown in Figure 509. Adjust the output terminal or engine bracket attach bulkhead Þttings as required. NOTE:
D.
Assure that the fuel cutoff lever is against the hard stop on the FCU when the cockpit control lever is in CUTOFF.
Propeller Speed Control Lever Linkage Rigging (1) Place the cockpit propeller speed control lever in the maximum forward position.
Maximum RPM Stop Figure 510 (2) (3) (4) (5)
(6)
Refer to Figure 510, hold the propeller governor speed adjusting lever against the maximum RPM stop. Align the propeller cable terminal rod end with the speed adjusting lever. Connect and safety. Adjust cable mounting bracket Þttings on the reduction gearbox ßange to obtain one-eighth to one-quarter inch cushion at the cockpit propeller speed control lever maximum RPM position. Move the propeller speed control lever through its full range of travel and check the following: (a) Ensure the propeller governor speed adjusting lever contacts the feather stop and the maximum RPM stop. (b) No binding is present. If either or both stops are not contacted, proceed as follows:
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Feather Stop Adjustment Figure 511 (a) (b) (c) (d) (e)
Refer to Figure 511, remove safety-wire from the lever retaining bolt. Remove lever and keep the lower lever in hard contact with the feather stop. Reinstall the speed adjusting lever on the serrated shaft so that it aligns with the boss on top of the governor body. Repeat step 5. If satisfactory, reinstall the lever retaining screw and safety-wire. If unsatisfactory, readjust position and repeat step 5. NOTE:
Serrations will relocate lever position in increments of 5 degrees.
NOTE:
After completion of rigging of the engine, refer to Chapter 71, Powerplant Adjustment/Test for engine run rig checks and additional adjustments.
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ENGINE INDICATING
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LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
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Apr 1/1996
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77-Title 77-List of Effective Pages 77-Record of Temporary Revisions 77-Table of Contents
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
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CONTENTS ENGINE INDICATING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-00-00 Page 1 77-00-00 Page 1 77-00-00 Page 1 77-00-00 Page 2
WET TORQUE INDICATING SYSTEM - DESCRIPTION AND OPERATION . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-11-00 Page 1 77-11-00 Page 1
WET TORQUE INDICATING SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-11-00 Page 101 77-11-00 Page 101
WET TORQUE INDICATING SYSTEM - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Separator Breather Pad Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Indicator Vent Line Leak Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Indicator Pressure Line Leak Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Indicator Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-11-00 Page 201 77-11-00 Page 201 77-11-00 Page 201 77-11-00 Page 201 77-11-00 Page 201 77-11-00 Page 205 77-11-00 Page 205
ELECTRICAL TORQUE INDICATING SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-12-00 Page 101 77-12-00 Page 101
ELECTRICAL TORQUE INDICATING SYSTEM - MAINTENANCE PRACTICES . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Pressure Transmitter Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Transmitter Vent Line Leak Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Transmitter Pressure Line Leak Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Transmitter Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque System Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Torque System Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-12-00 Page 201 77-12-00 Page 201 77-12-00 Page 201 77-12-00 Page 201 77-12-00 Page 201 77-12-00 Page 205 77-12-00 Page 205 77-12-00 Page 206 77-12-00 Page 206
PROPELLER RPM INDICATOR - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-13-00 Page 101 77-13-00 Page 101
PROPELLER RPM INDICATOR - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller RPM Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tach Generator Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propeller RPM Indicator Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-13-00 Page 201 77-13-00 Page 201 77-13-00 Page 201 77-13-00 Page 201 77-13-00 Page 201
GAS GENERATOR RPM INDICATOR - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Gas Generator RPM Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tach Generator Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Gas Generator RPM Indicator Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-14-00 Page 201 77-14-00 Page 201 77-14-00 Page 201 77-14-00 Page 201 77-14-00 Page 201
INTERTURBINE TEMPERATURE INDICATOR - MAINTENANCE PRACTICES . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Interturbine Temperature Indicator Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . .
77-21-00 Page 201 77-21-00 Page 201 77-21-00 Page 201
INTERTURBINE TEMPERATURE INDICATOR - ADJUSTMENT/TEST . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Interturbine Temperature Indicator Adjustment and Test . . . . . . . . . . . . . . . . . . . . . . . . Interturbine Temperature Indicator Bench Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Interturbine Temperature Indicator Airplane Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Insulation Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-21-00 Page 501 77-21-00 Page 501 77-21-00 Page 501 77-21-00 Page 501 77-21-00 Page 509 77-21-00 Page 510
77 - CONTENTS © Cessna Aircraft Company
Page 1 of 3 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ALTAIR ENGINE TREND MONITORING - DESCRIPTION AND OPERATION . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description/Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Manuals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Retrieving Data from ADAS+ Processor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Uploading Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-30-00 Page 1 77-30-00 Page 1 77-30-00 Page 1 77-30-00 Page 1 77-30-00 Page 2 77-30-00 Page 2
ENGINE TREND MONITORING SYSTEM - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Indications (Engine Off) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Start-Up and Communication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Log File Error Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-30-00 Page 101 77-30-00 Page 101 77-30-00 Page 101 77-30-00 Page 101 77-30-00 Page 101
ALTAIR ADAS+ ENGINE TREND MONITORING SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Processor Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outside Air Temperature (OAT) Sensor Removal/Installation . . . . . . . . . . . . . . . . . . . . Annunciator Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ice Vane Installation Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot and Static Transducers Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-30-00 Page 201 77-30-00 Page 201 77-30-00 Page 201 77-30-00 Page 201 77-30-00 Page 201 77-30-00 Page 205 77-30-00 Page 205 77-30-00 Page 205
ALTAIR ADAS+ ENGINE TREND MONITORING SYSTEM - ADJUSTMENT/TEST . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADAS+ Processor Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot and Static Sensor Calibration Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pitot and Static Sensors Calibrating Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OAT Sensor Calibration Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Sensor Calibration Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Sensor Calibrating Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Discrete Switch Tests. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Live Data Sensor Test - Engine Ground Run. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-30-00 Page 501 77-30-00 Page 501 77-30-00 Page 501 77-30-00 Page 501 77-30-00 Page 501 77-30-00 Page 501 77-30-00 Page 503 77-30-00 Page 503 77-30-00 Page 503 77-30-00 Page 504 77-30-00 Page 505
ALTAIR ENGINE TREND MONITORING TORQUE TRANSDUCER - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Transducer Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-30-10 Page 201 77-30-10 Page 201 77-30-10 Page 201 77-30-10 Page 201
ALTAIR ADASd ENGINE TREND MONITORING - DESCRIPTION AND OPERATION. . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description/Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Manuals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Retrieving Data from ADAS d Processor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Uploading Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-31-00 Page 1 77-31-00 Page 1 77-31-00 Page 1 77-31-00 Page 1 77-31-00 Page 2 77-31-00 Page 2
ALTAIR ADASd ENGINE TREND MONITORING - TROUBLESHOOTING. . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Indications (Engine Off) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Start-Up and Communication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-31-00 Page 101 77-31-00 Page 101 77-31-00 Page 101 77-31-00 Page 101 77-31-00 Page 101
ALTAIR ADASd ENGINE TREND MONITORING SYSTEM - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Processor Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-31-00 Page 201 77-31-00 Page 201 77-31-00 Page 201 77-31-00 Page 201
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ALTAIR ADASd ENGINE TREND MONITORING SYSTEM - ADJUSTMENT/TEST . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADAS d Processor Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Live Data Sensor Test - Engine Off . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Live Data Sensor Test - Engine Ground Run. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-31-00 Page 501 77-31-00 Page 501 77-31-00 Page 501 77-31-00 Page 501 77-31-00 Page 501 77-31-00 Page 502
GEA-71 ENGINE/AIRFRAME UNIT - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GEA-71 Engine/Airframe Unit Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
77-40-00 Page 201 77-40-00 Page 201 77-40-00 Page 201
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Page 3 of 3 Apr 1/2010
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL ENGINE INDICATING - GENERAL 1.
Scope A.
2.
This chapter contains information on the systems and components used to monitor and indicate engine conditions.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Tachometer Generator
AG34 (MIL-G26611)
Task Corporation 1009 E. Vermont Ave. Anaheim, CA 92803
Used during functional check of percent RPM indicator.
Frequency Counter
Model 5326
Hewlett Packard 1601 California Ave. Palo Alto, CA 94304
Used during functional check of percent RPM indicator.
Torque System Calibration Tester
5790011-1
Cessna Aircraft Company Cessna Parts Distribution Department 701, CPD 2 5800 East Pawnee Road Wichita, KS 67218-5590 (P.O. Box 7704 Wichita, KS 67277-7704)
Used during functional check of percent RPM indicator.
Pressure Tester
2311F
BarÞeld Instrument Corp. 1478 Central Ave. East Point, GA 30344
To troubleshoot the torque indicating system.
Pressure Gage
304-00101
BarÞeld Instrument Corp.
To troubleshoot the torque indicating system.
Dry Nitrogen Bottle (Alternate Shop Air with Regulator Capable of 40.0 PSI, Pressure and Shutoff Valve)
Regulated to 100 PSI
Commercially available
To leak check system.
Pressure Gages (0 to 100 PSI)
Commercially available
To monitor pressures.
Container (1 quart capacity)
Commercially available
To bleed system.
Air Data Test Set
6520-10
Laverslab, Inc. 10435 Greenbough, Suite 300 Stafford, TX77477
To check for torque indicator vacuum leak.
TT1000A Test Set
2312G-8
BarÞeld Instrument Corp.
To test ITT system.
TT1200
101-00920
BarÞeld Instrument Corp.
To test ITT system.
Thermo Gun
Model 500A
Raychem Corp. 300 Constitution Drive Menlo Park, CA 94825-1111
To heat probes.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
Precision Millivolt Source
MV116
Electronic Development Corp. 11 Hamlin Street Boston, MA 02127-4112
Measure millivolts.
Test Harness
579005-11
Cessna Aircraft Company
To test ITT indicator.
Test Harness
579005-13
Cessna Aircraft Company
To calibrate ITT indicator.
Thermos Container (1 quart)
Commercially available
To test ITT system.
28.0 VDC Power Supply (with less than 0.25 VDC peak-to-peak ripple)
Commercially available
To supply voltage.
Deadweight Tester (See Note 1)
Model 61-10
Chandler Engineering Co. 7707 E. 38th St. Tulsa, OK 74145
To calibrate torque indicating system.
Digital Volt Meter
8100A
John Fluke Manufacturing Co. 6920 Seaway Boulevard Everett, WA 98206
To measure voltage.
NOTE 1: A deadweight tester, capable of 80.0 PSI with a minimum of 0.1 PSI resolution, must be used. The following precautions need to be taken when using a deadweight tester: (1) The deadweight tester must be level to within one-quarter bubble. Device being tested must be at the same height as the deadweight tester with no dips in the connecting hoses. (2) The cylinder must be kept smooth and dust free. (3) The weights must be recalibrated at six-month intervals. (4) The tester must be Þlled with the type and brand of oil with which the engine has been serviced. 3.
DeÞnition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief deÞnition of the sections incorporated in this chapter is as follows: (1) The section on indicating provides description, operation, troubleshooting and removal/ installation instructions for the torque and propeller RPM indicator systems. (2) The section on temperature indicating provides description, operation, troubleshooting, removal/ installation and test instructions for the interturbine temperature indicator systems.
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MODEL 208 MAINTENANCE MANUAL WET TORQUE INDICATING SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
The torque indicator is located on the upper portion of the instrument panel. The torque indicator monitors the engine torque pressure and converts this pressure into an indication of torque in footpounds. Instrument markings indicate that the normal operating range (green arc) is from 0 footpounds to 1658 foot-pounds, the alternate power range (striped green arc) is from 1658 foot-pounds to 1970 foot-pounds, and maximum torque (red line) is 1970 foot-pounds. NOTE:
Additional information pertaining to the torque sensor is contained in the Pratt/ Whitney PT6 Maintenance Manual (refer to the List of Publications in the Introduction Section of this manual).
B.
Two lines enter the back of the torque indicator. One line measuring pressure from the torque sensor, and the other line, reference pressure from the engine case. The power turbine, on a separate shaft from the compressor section, supplies power to the propeller through a gear reduction system; a torque sensor is incorporated into the propeller reduction system indicating power changes by increasing or decreasing oil pressure to a direct reading indicator on the instrument panel. The reference pressure line connects the indicator to an oil separator and vent port on the engine case.
C.
A schematic has been provided to aid the maintenance technician in system understanding. Refer to Figure 1.
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
Wet Torque Sensing Schematic Figure 1 (Sheet 1)
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL WET TORQUE INDICATING SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been included to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
B.
Trouble shooting and repair of malfunctions which occur due to leaks in the pressure portion of the system can be difÞcult. Maintenance actions involving component replacement may seem to offer a cure but this is only temporary if leaks are not isolated and repaired. The following guidelines should be observed in troubleshooting this type of problem. Detection of extremely small leaks by normal procedures involving looking for a pressure loss can be difÞcult. If a leak is suspected, particularly in cases where repeated rebleeding results in periods of normal operation followed by a malfunction, it may be necessary to gain access to all portions of the pressure line to check for leakage. This may be done Þrst by pressurizing a Þlled system at the engine input Þtting and checking for signs of wetness at all connections including the connections at the torque indicator, and second, by pressurizing a dry system with air or nitrogen and checking all connections with a bubble-type leak detector. It is important to trace the pressure line from one end to the other to make sure the correct line is being checked.
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MODEL 208 MAINTENANCE MANUAL
Wet Torque Indicating System Troubleshooting Chart Figure 101 (Sheet 1)
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Wet Torque Indicating System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL WET TORQUE INDICATING SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Wet torque indicating system maintenance practices include torque indicator removal/installation and functional test.
Torque Indicator Removal/Installation A.
Remove Torque Indicator (Refer to Figure 201). (1) Remove screws (1) securing cowl deck cover (2) to cowl deck assembly, to allow access to engine torque indicator lines. (2) Disconnect lines (3) from unions (4). (3) Loosen mounting screws (5) securing torque indicator (6) to instrument panel. (4) Slide indicator (6) out from instrument panel.
B.
Install Torque Indicator (Refer to Figure 201). (1) Slide indicator (6) into instrument panel. (2) Tighten mounting screws (5) securing indicator to instrument panel. (3) Connect lines (3) to unions (4). NOTE: (4) (5)
3.
4.
Start engine in accordance with Model 208 Pilot's Operating Handbook.
Observe indicator needle, if excessive ßuctuation is present, position rag under indicator pressure line and loosen to bleed off air in system. Install cowl deck cover (2) and secure using screws (1).
Oil Separator Breather Pad Removal/Installation A.
Remove Oil Separator Breather Pad (Refer to Figure 201). (1) Open right upper engine cowling. (2) Disconnect line (4) from elbow (5). (3) Remove adapter (6) from oil separator (8). (4) Remove breather pads (7) from oil separator (8), using a wire with end formed into a hook.
B.
Install Oil Separator Breather Pad (Refer to Figure 201). (1) Insert breather pads (7) into oil separator (8). (2) Screw adapter (6) into oil separator and tighten. (3) Connect line (4) from torque indicator to elbow (5). (4) Close and secure right engine cowling.
Torque Indicator Vent Line Leak Test A.
Leak Test Torque Indicator Vent Line (Refer to Figure 201). (1) Disconnect vent line (4) from elbow (5). (2) Connect vacuum portion of pitot-static test set to line (4). Set test set altimeter to zero. (3) Slowly apply vacuum. Check that indicator reading does not start decreasing below zero. If this is the case, connections to indicator are reversed and must be corrected. Continue applying vacuum until altimeter on test set reads 14,000 feet. (4) Shut off vacuum. Check that test altimeter or gage does not show a loss of vacuum for one minute. (5) If leakage is noted, disconnect vent line from indicator and cap. Recheck test set plumbing and vent line leakage and repair as required (with vent line not connected to indicator, positive pressure may be used to isolate leaks using bubble leak detector on Þttings). Check indicator by applying vacuum per steps (2) through (4) directly to reference port Þtting of indicator. If leakage is noted, check Þtting for Teßon tape on pipe threads and proper torque. If Þtting is not leakage source, replace indicator. (6) Connect vent line (4) to elbow (5).
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MODEL 208 MAINTENANCE MANUAL
Wet Torque Indicating System Installation Figure 201 (Sheet 1)
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Wet Torque Indicating System Installation Figure 201 (Sheet 2)
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Wet Torque Indicating System Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL
5.
Torque Indicator Pressure Line Leak Test A.
Leak Test Torque Indicator Pressure Line (Refer to Figure 201). (1) Disconnect pressure line (6) from adapter (7). (2) Connect air pressure source to pressure line (6). (3) Apply two (2) PSIG pressure and check that torque indicator reading increases. If reading decreases, line to torque indicators are reversed and must be corrected. (4) Apply 40 PSIG pressure and shut off pressure source. Record data as follows: (a) Allow system to set for ten minutes to stabilize. Tap test gage and record exact test pressure (PSIG). Record ambient temperature (°F) and true ambient pressure ("Hg) by reading the window of an altimeter set to zero altitude. (b) Allow system to set for one hour. Record test pressure, ambient temperature, and ambient pressure in same manner as step (a) above. (c) Correct for ambient temperature and pressure differences as follows: 1 Ambient Temperature Correction (PSIG) = Final Reading (°F) - Initial Reading (°F) x 0.1. 2 Ambient Pressure Correction (PSIG) = Final Reading ("Hg) - Initial Reading ("Hg) x -0.49. 3 Corrected Test Pressure = Initial Test Pressure + Ambient Temperature + Ambient Pressure Correction. 4 Leakage = Corrected Test Pressure - Observed Final Test Pressure. (d) Leakage per step (3) shall not exceed 0.5 PSIG. Isolate and repair any leaks by using bubble leak detector solution on Þttings. NOTE:
(e) 6.
Allowable leakage of 0.5 PSIG is due to tolerances in pressure gage and correction procedures. Goal is zero leakage. If there is any doubt as to whether leakage exists, check all Þttings with bubble leak detector ßuid.
Relieve pressure and disconnect test equipment. Reconnect pressure line (6) to adapter (7).
Torque Indicator Functional Test A.
Functional Test Torque Indicator (Refer to Table 201).
CAUTION: Improper connection will damage the indicator. (1) (2) (3) (4) (5)
Connect pressure source and gage (or deadweight tester) to pressure port (marked P) of the indicator. Leave reference port (marked V) of the indicator open to atmosphere. Apply pressures in increasing direction. Check indicator reading tolerance after tapping indicator to overcome friction. Repeat with decreasing pressure, except at the 2500 ft-Ib point. Failure of indicator to comply with above shall result in rejection of unit. Apply 8.6 PSIA pressure (14,000-foot altimeter reading with altimeter originally set to zero) to reference port (marked V) of indicator. Shut off vacuum and check that no discernible leakage occurs in one minute.
Table 201. Scale Error INPUT PSI (NOTE 1)
INPUT PSI (NOTE 2) (REF)
TOLERENCE PSI (REF)
TORQUE (FT-LBS)
TOLERANCE (FT-LBS)
-0.71
0
+1.38 or -1.38
0
+50 or -50
13.49
14.20
+1.38 or -1.38
500
+50 or -50
27.68
28.39
+1.02 or -1.02
1000
+35 or -35
41.87
42.59
+0.78 or -0.78
1500
+28 or -28
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MODEL 208 MAINTENANCE MANUAL Table 201. Scale Error (continued) INPUT PSI (NOTE 1)
INPUT PSI (NOTE 2) (REF)
TOLERENCE PSI (REF)
TORQUE (FT-LBS)
TOLERANCE (FT-LBS)
46.36
47.08
+0.56 or -0.56
1658
+20 or -20
55.22
55.93
+0.56 or -0.56
1970
+20 or -20
70.26
70.98
+1.18 or -1.18
2500
+40 or -40
NOTE 1: To be used when indicator and pressure source are at same level. NOTE 2: To be used when indicator is 20 inches higher than pressure source. 1. 2. 3. 4. 5.
Constant = 35.22 ft-Ibs/PSI. Gage will sense 0.71 PSI lower (level ßight) due to gage location approximately 20 inches higher than engine ports. Zero PSI at the gage = 25 ft-Ibs. Difference will disappear when gage is installed and operating. Pressures listed in Column 1 are with reference port open to ambient pressure and no ßuid head difference between source and indicator. Values listed in Column 1 are for calibration of indicator only and are not applicable to installed system. Tap indicator before reading.
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL TORQUE INDICATING SYSTEM - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been prepared to aid the maintenance technician in system understanding. Refer to Figure 101.
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Electrical Torque Indicating System Troubleshooting Chart Figure 101 (Sheet 1)
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Electrical Torque Indicating System Troubleshooting Chart Figure 101 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL ELECTRICAL TORQUE INDICATING SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Maintenance practices include torque indicator removal/installation, torque transmitter removal/ installation and functional test.
Torque Indicator Removal/Installation NOTE:
3.
4.
During component removal and installation, cap all lines and Þttings.
A.
Remove Torque Indicator (Refer to Figure 201). (1) Loosen screw (5) securing indicator (6) to mounting clamp (3) and remove indicator. (2) Cut safety-wire and remove electrical connector (4) from indicator (6).
B.
Install Torque Indicator (Refer to Figure 201 ). (1) Install electrical connector (4) to indicator (6) and safety-wire. (2) Slide indicator (6) through instrument panel into mounting clamp (3) and secure with screw (5). (3) Start engine in accordance with Pilot's Operating Handbook and check torque indicator operation.
Torque Pressure Transmitter Removal/Installation A.
Remove Torque Pressure Transmitter (Refer to Figure 201). (1) Open right upper cowling door to gain access. (2) Ensure airplane electrical power is OFF. (3) Cut safety wire and remove electrical connector (14). (4) Loosen vent line (7) and pressure line (8) at each end of transmitter (11) and remove lines. (5) Remove bolts (13) from clamps (12) securing transmitter to bracket (15) and remove transmitter.
B.
Install Torque Pressure Transmitter (Refer to Figure 201). (1) Secure torque pressure transmitter (11) to bracket (15) by installing bolts (13) through clamps (12). (2) Connect pressure line (8) and vent line (7) to transmitter. (3) Install electrical connector (14) to transmitter and safety-wire. (4) Restore electrical power to airplane, start engine in accordance with Pilot's Operating Handbook and verify proper torque indicator operation. (5) If torque indication is erratic, loosen pressure line connection (8) at transmitter and allow air in lines to bleed from system. Tighten line. (6) Shut down engine in accordance with Pilot's Operating Handbook.
Torque Transmitter Vent Line Leak Test A.
Leak Test Torque Transmitter Vent Line (Refer to Figure 201). NOTE:
Airplanes 20800001 thru 20800083, ensure torque indicator circuit breaker is in and 26 volts 400 Hz AC current is being supplied to the torque indicating system for the following tests.
NOTE:
Airplanes 20800084 and On and 208B0001 and On, ensure torque indicator circuit breaker is engaged and 28 VDC current is being supplied to the torque indicating system for the following tests.
(1) (2) (3) (4)
Disconnect vent line (4) from elbow (5). Connect vacuum portion of pitot-static test set to line (4). Set test set altimeter to zero. Slowly apply vacuum. Check that indicator reading does not start decreasing below zero. If this is the case, connections to transmitter are reversed and must be corrected. Continue applying vacuum until altimeter on test set reads 14,000 feet (or a pressure of 8.6 PSIA). Shut off vacuum. Check that test altimeter or gage does not show a loss of vacuum for one minute.
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Electrical Torque Indicating System Installation Figure 201 (Sheet 1)
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Electrical Torque Indicating System Installation Figure 201 (Sheet 2)
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Electrical Torque Indicating System Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (5)
(6) 5.
If leakage is noted, disconnect vent line from transmitter and cap. Recheck test set plumbing and vent line leakage and repair as required (with vent line not connected to transmitter, positive pressure may be used to isolate leaks using bubble leak detector on Þttings). Check transmitter by applying vacuum per steps (2) through (4) directly to reference port Þtting of transmitter. If leakage is noted, check Þtting for Teßon tape on pipe threads and proper torque. If Þtting is not leakage source, replace transmitter. Connect vent line (4) to elbow (5).
Torque Transmitter Pressure Line Leak Test A.
Leak Test Torque Transmitter Pressure Line (Refer to Figure 201 ). (1) Disconnect pressure line (6) from adapter (7). (2) Connect air pressure source to pressure line. (3) Apply 2.0 PSIG pressure and check that torque indicator reading increases. If reading decreases, line to torque transmitter is reversed and must be corrected. (4) Apply 40 PSIG pressure and shut off pressure source. Record data as follows: (a) Allow system to set for ten minutes to stabilize. Tap test gage and record exact test pressure (PSIG). Record ambient temperature in (°F) and true ambient pressure ("Hg) by reading the window of an altimeter set to zero altitude. (b) AIIow system to set for one hour. Record test pressure, ambient temperature and ambient pressure in same manner as step (a) above. (c) Correct for ambient temperature and pressure differences as follows: 1 Ambient Temperature Correction (PSIG) = Final Reading (°F) - Initial Reading (°F) x 0.1. 2 Ambient Pressure Correction (PSIG) = Final Reading ("Hg) - Initial Reading ("Hg) x -0.49. 3 Corrected Test Pressure = Initial Test Pressure + Ambient Temperature + Ambient Pressure Correction. 4 Leakage = Corrected Test Pressure - Observed Final Test Pressure. (5) Leakage per Step 3. shall not exceed 0.5 PSIG. Isolate and repair any leaks by using bubble leak detector solution on Þttings. NOTE:
(6) 6.
Allowable leakage of 0.5 PSIG is due to tolerances in pressure gage and correction procedures. Goal is zero leakage. If there is any doubt as to whether leakage exists, check all Þttings with bubble leak detector ßuid.
Relieve pressure and disconnect test equipment. Reconnect pressure line to adapter.
Torque Transmitter Functional Test A.
Functional Test Torque Transmitter (Refer to Table 201).
CAUTION: Improper connection will damage the transmitter. (1) (2) (3) (4)
Connect pressure source and gage (or deadweight tester) to pressure port of the transmitter. Leave reference port (marked "VENT") of the transmitter open to atmosphere. Apply pressures in increasing direction. Check indicator reading tolerance after tapping transmitter to overcome friction. Repeat with decreasing pressure except at the 2500 foot-pound point. Failure of transmitter to comply with above shall result in rejection of unit. Apply 8.6 PSIA pressure (14,000-foot altimeter reading with altimeter originally set to zero) to reference port (marked "VENT") of transmitter. Shut off vacuum and check that no discernible leakage occurs in one minute.
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Table 201. Scale Error INPUT PSI (1)
INPUT PSI (2) (REF)
TOLERENCE PSI (REF)
TORQUE (FT-LBS)
TOLERANCE (FT-LBS)
-0.71
0
+1.38 or -1.38
0
+50 or -50
13.49
14.20
+1.38 or -1.38
500
+50 or -50
27.68
28.39
+1.02 or -1.02
1000
+35 or -35
41.87
42.59
+0.78 or -0.78
1500
+28 or -28
46.36
47.08
+0.56 or -0.56
1658
+20 or -20
55.22
55.93
+0.56 or -0.56
1970
+20 or -20
70.26
70.98
+1.18 or -1.18
2500
+40 or -40
NOTE 1: To be used when transmitter and pressure source are at same level. NOTE 2: To be used when transmitter is 20 inches higher than pressure source. Constant = 35.22 Ft-Lbs/PSI. Zero PSI at the gage = 25 Ft-Lbs. Difference will disappear when gage is installed and operating. Pressures listed in column 1 are with reference port open to ambient pressure and no ßuid head difference between source and indicator. Values listed in column 1 are for calibration of indicator only and are not applicable to installed system.
1. 2. 3. 4. 7.
Torque System Functional Test A.
8.
Functional Test Torque System (Refer to Figure 201). (1) Disconnect plumbing between torque transducer and engine. (2) Connect the torque system calibration tester to the pressure port of the transducer. (3) Close the torque indicator circuit breaker. (4) Using dry Nitrogen source, apply 28.39 PSI, +1.02 or -1.02 PSI, to the transducer. The indicator shall read 1000 Ft-Lbs, +35 or -35 Ft-Lbs. (5) Increase pressure to 42.59 PSI, +.78 or -.78 PSI. The torque indicator shall read 1500 Ft-Lbs, +28 or -28 Ft-Lbs. (6) Increase pressure to 47.08 PSI, +.56 or -.56 PSI. The torque indicator shall read 1658 Ft-Lbs, +20 or -20 Ft-Lbs. (7) Increase pressure to 55.93 PSI, +.56 or - .56 PSI. The torque indicator shall read 1970 Ft-Lbs, +20 or -20 Ft-Lbs. (8) Slowly remove pressure from the torque transducer. (9) Remove torque system calibration tester and restore the system to original condition. (10) If indicator readings are out of tolerance, then perform torque system calibration. Refer to Torque System Calibration.
Electrical Torque System Calibration A.
Aircraft System Calibration. (1) Disconnect plumbing between torque transducer and engine. (2) Connect torque system calibration tester to pressure port of the transducer.
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CAUTION: The pressure source to transducer connection must have absolutely zero leakage before calibration is attempted. (3)
Close torque indicator circuit breaker. NOTE:
(4) (5) (6) (7) (8) (9) (10) (11) (12) (13) (14) (15) (16)
The following steps are to be performed with both indicator and transducer at ambient temperature.
Apply pressure P1 (as stamped on the torque transducer), and observe the torque indicator reading. Calculate D1 where D1 = indicator reading - 1000. Supply pressure P2 (as stamped on the torque transducer) and observe the torque indicator reading. Calculate D2 where D2 = indicator reading - 2000. Calculate slope adjustment, direction and magnitude (D3). The following calibration is an algebraic sum, D3 = D1 -D2. (a) A negative result means a downscale adjustment. (b) A positive result means an upscale adjustment. Loosen screw securing indicator to mounting clamp, draw it from the panel far enough to access the adjustments in the rear portion of the indicator cover. Remove both cap screws (slope and offset), to gain access to the adjustments. Turn the slope adjustment to change the indicator in the amount and direction of D3. Apply pressure P1 (as stamped on the torque transducer). Turn the offset adjustment to center the torque indicator pointer on the 1000 Ft-Lb graduation. Repeat steps 4 through 14 until D1 and D2 are both zero. Using the following formula, calculate the value of the slope for pressure versus torque system indication.
(17) Using the following formula, calculate the pressure for the red wedge indicator (1658 Lbs torque) for the PT6A-114 engine and (1865 Lbs torque) fot the PT6A-114A engine, and for the red radial indication (1970 Lbs torque) for both PT6A-114 and PT6A-114A engines.
(18) Using the following formula, calculate the possible overall system error. (a) °C = Ambient temperature in degrees centigrade. (b) ± Error (Ft-Lbs) = [ ([50-°C] 0.03714) + 1.4] 25 NOTE:
°C should be current ambient temperature.
(19) Verify that red wedge indication corresponds to the respective applied pressure within the overall error allowed as calculated in Step 18. (20) Verify that red radial indication corresponds to the respective applied pressure within the overall error allowed as calculated in Step 18. (21) Remove pressure from the torque transducer. (22) Verify the torque indicator shows (0) zero ft-lbs torque within the overall error as calculated in Step 18.
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MODEL 208 MAINTENANCE MANUAL (23) Replace both cap screws covering the slope and offset adjustments. Reinstall the torque indicator. (24) Remove the torque calibration tester and restore the system to original condition. (25) Pull the torque indicator circuit breaker. (26) Verify the torque indicator needle is in the OFF or below zero area.
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MODEL 208 MAINTENANCE MANUAL PROPELLER RPM INDICATOR - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been provided to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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MODEL 208 MAINTENANCE MANUAL
Propeller RPM Indicator Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL PROPELLER RPM INDICATOR - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Propeller RPM indicator maintenance practices include propeller RPM indicator removal/installation, tach generator removal/installation and propeller RPM indicator functional check.
Propeller RPM Indicator Removal/Installation A.
Remove Propeller RPM Indicator (Refer to Figure 201). (1) Loosen mounting screw (7) and slide indicator (8) out of instrument panel. (2) Disconnect electrical connector (9) from indicator.
B.
Install Propeller RPM Indicator (Refer to Figure 201). (1) Position indicator (8) in place, connect electrical connector (9) to indicator. (2) Slide indicator into instrument panel and tighten mounting screw (7).
Tach Generator Removal/Installation A.
Remove Tach Generator (Refer to Figure 201). (1) Open right upper engine cowling, refer to Chapter 71, Cowling and Nose Cap - Maintenance Practices). (2) Remove right nose cap half. (3) Cut safety wire and disconnect electrical connector (1) from tach generator (5). (4) Remove nut (2) and washer (3) securing tach generator (5) and gasket (4) to mounting pad and remove.
B.
Install Tach Generator (Refer to Figure 201).
CAUTION: Ensure splines of tach generator drive shaft are aligned with engine drive during installation. (1) (2) (3) 4.
Install gasket (4) and tach generator (5) and secure using washer (3) and nut (2). Connect electrical connector (1) to tach generator. Install right nose cap half in accordance with Chapter 71, Cowling and Nose Cap - Maintenance Practices). Close right upper engine cowling.
Propeller RPM Indicator Functional Check A.
Functional Check Propeller RPM Indicator (Refer to Figure 202). NOTE:
(1) (2) (3)
The following test is to be conducted at room temperature using a suitable tachometer test stand which incorporates a tach generator per MIL-G-26611 as a signal source to the indicator. The generator shall be operated in the counterclockwise direction as viewed from the drive end. Monitor the tachometer generator output with a frequency counter with an accuracy of not less than 0.001 Hertz in the 0-80 Hz range.
Remove indicator. Use a tachometer test stand and apply tachometer generator signals from column one of indicator calibration table in Figure 202. Tap indicator and check indicator readings. Readings shall be per column three of calibration table within the following limits: (a) +20 RPM from 0 RPM to 599 RPM. (b) +16 RPM from 600 RPM to 1599 RPM. (c) +10 RPM from 1600 RPM to 2000 RPM. (d) +10 RPM at 1900 RPM. NOTE:
Readings shall be taken in both ascending and descending directions.
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Propeller RPM Indicator Installation Figure 201 (Sheet 1)
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Propeller RPM Functional Check Information Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL (4) (5)
If the indicator does not meet the indicator specification, refer to Chapter 77, Propeller RPM Indicator - Troubleshooting. Reinstall indicator.
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MODEL 208 MAINTENANCE MANUAL GAS GENERATOR RPM INDICATOR - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Gas Generator RPM Indicator maintenance practices include gas generator RPM removal/installation, tach generator removal/installation and gas generator RPM indicator functional check.
Gas Generator RPM Indicator Removal/Installation A.
Remove Gas Generator RPM Indicator (Refer to Figure 201 ). (1) Loosen mounting screw (13) and slide.indicator (14) out of instrument panel. (2) Disconnect electrical connector (15) from indicator (14).
B.
Install Gas Generator RPM Indicator (Refer to Figure 201). (1) Connect electrical connector (15) to indicator (14). (2) Slide indicator into instrument panel and tighten mounting screw (13).
Tach Generator Removal/Installation A.
Remove Tach Generator (Refer to Figure 201). (1) Open necessary upper engine cowling to gain access to tach-generator. (2) Cut safety wire and disconnect electrical connector (1) from tach-generator (5). (3) Remove nut (2) and washer (3) securing tach-generator (5) to mounting and remove.
B.
Install Tach Generator (Refer to Figure 201).
CAUTION: Ensure alignment of spines as tach-generator is pulled down with nuts. (1) (2) (3) 4.
Install gasket (4) and tach-generator (5) and secure using washer (3) and nut (2). Connect electrical connector (1) to tach-generator and secure with safety wire. Close upper engine cowling.
Gas Generator RPM Indicator Functional Check A.
Functional Check Gas Generator RPM Indicator (Refer to Figure 202). NOTE:
(1) (2) (3)
(4)
The following test is to be conducted at room temperature using a suitable tachometer test stand which incorporates a tach-generator per MIL-G-26611 as a signal source to the indicator. The generator shall be operated in the counterclockwise direction as viewed from the drive end. Monitor the tachometer generator output with a frequency counter having an accuracy of not less than 0.001 Hertz in the 0-80 Hz range.
Remove indicator. Using tachometer test stand, apply tachometer generator signals listed in column one of indicator Calibration Table. Tap indicator and check indicator readings. Readings shall be per column three of Calibration Table within the following limits: (a) + or -1 percent RPM from 0 to 30 percent RPM (b) + or -0.8 percent RPM from 30 to 80 percent RPM (c) + or -0.5 percent RPM from 80 to 101.6 percent RPM (d) + or -0.5 percent RPM at 101.6 percent RPM NOTE:
The unit "percent RPM" is to be interpreted as one small division of the auxiliary dial or one-half of one small division of the main dial.
NOTE:
Readings shall be taken in both ascending and descending directions.
Reinstall indicator.
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Gas Generator Percent RPM Installation Figure 201 (Sheet 1)
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Gas Generator Percent RPM Functional Check Information Figure 202 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL INTERTURBINE TEMPERATURE INDICATOR - MAINTENANCE PRACTICES 1.
General A.
2.
Interturbine temperature indicator maintenance practices consist of inter-turbine temperature indicator removal/installation.
Interturbine Temperature Indicator Removal/Installation A.
Remove Interturbine Temperature Indicator (Refer to Figure 201). (1) Loosen mounting screw (10) and slide indicator (11) out of instrument panel. (2) Disconnect electrical connector (12) from indicator (11).
B.
Install Interturbine Temperature Indicator (Refer to Figure 201). (1) Connect electrical connector (12) to indicator (11). (2) Slide indicator into instrument panel and tighten mounting screw (10).
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ITT Indicator Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL INTERTURBINE TEMPERATURE INDICATOR - ADJUSTMENT/TEST 1.
General A.
2.
The interturbine temperature (ITT) indicator adjustment and test section has procedures to use a TT1000A or TT1200 tester with the interturbine temperature indicator. Refer to Figure 501, Figure 502, Figure 503, Figure 504, Figure 505 and Figure 506 for the test equipment connections.
Interturbine Temperature Indicator Adjustment and Test
Table 501. Calibration TEST POINT °C
APPLIED VOLTAGE IN MILLIVOLTS
TOLERANCE TEMPERATURE °C
100
4.095
+20 to -20
300
12.207
+20 to -20
660
27.445
+7.5 to -7.5
805
33.480
+2.5 to -2.5
900
37.325
+5.0 to -5.0
1090
44.729
+2.5 to -2.5
1100
45.108
+5.0 to -5.0
A.
Test the Interconnect Interturbine Temperature (ITT) System (Refer to Figure 501). NOTE: (1)
Let the junctions in the ice bath become stable for 5 minutes before any test.
Make sure that the ITT harness and the TT5 trim harness are disconnected from the system. NOTE:
(2)
A large terminal is designated for the alumel connector.
Adjust the applied millivolt source for each test point of the calibration data. Make sure that the temperature indication is in the range of the speciÞed tolerance for each test point. Refer to Table 501. (a) If the ITT indication does not agree with the applied millivolt input, do the following: 1 Examine the connectors and wiring between the thermocouple harness connector and the ITT indicator. Make sure that the wiring is clean, there is proper contact, and there is correct alumel/chromel pin arrangement. 2 Examine the interconnect ITT indicator. NOTE:
Let the junctions in the ice bath become stable for 5 minutes before any test.
a
(b)
3.
Adjust the applied millivolt source to 0.00 mV output and apply power to the unit. Let the unit become warm and stable for approximately one minute. b Use the calibration data to make an analysis of the indicator operation when the applied millivolt source is adjusted for each test point. Refer to Table 501. If the ITT indicator agrees with the millivolt input, but the indication during the engine operation is incorrect, refer to the Pratt & Whitney Maintenance Manual for ITT Thermocouple and Thermocouple Harness Inspection Procedure.
Interturbine Temperature Indicator Bench Test A.
Bench Test the ITT Indicator with the TT1000A Tester (Refer to Figure 504).
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ITT Indicator Functional Test Figure 501 (Sheet 1)
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ITT System Functional Test Figure 502 (Sheet 1)
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Bench Test Connections Figure 503 (Sheet 1)
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TT1000A BarÞeld Test Set Figure 504 (Sheet 1)
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TT1200 BarÞeld Test Set Figure 505 (Sheet 1)
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ITT Functional Test with TT1000A or TT1200 BarÞeld Test Set Figure 506 (Sheet 1)
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CAUTION: Approximately 45 VDC no-load is present on the test leads during resistance measurements. Let the test leads reach ambient temperature before any test. (1)
Remove the ITT indicator from the airplane. Refer to Chapter 77, Interturbine Temperature Indicator - Adjustment/Test. (2) Connect the test harness to the indicator. (3) Complete the TT1000A self test before you connect the test set to the indicator. (a) With the test set leads disconnected, set the RESISTANCE RANGE switch to the BATT position. (b) Set the FUNCTION switch to RESISTANCE MEASURE. (c) Set the power switch to ON and make sure that the display shows 30 VDC to 50 VDC. (4) Set the FUNCTION SWITCH to INDICATOR TEST. (5) Set the RESISTANCE RANGE switch to 2M ohms (0 ohm resistance). (6) Connect the black clip (-) of TT1000A lead to the alumel wire connected to pin B on indicator. (7) Connect the red clip (+) of the TT1000A lead to the chromel wire connected to pin A on the indicator. (8) Turn the TEMP ADJ knob on TT1000A until the tester display matches the test points, and make sure that the indicator display is in the range of the speciÞed tolerance for each test point. Refer to Table 501. (a) If the indicator does not agree with the TT1000A input, make sure that the harness is clean, there is proper contact, and there is correct alumel/chromel pin arrangement. 1 Do a test to make sure that the indicator does not agree with the TT1000A input. (b) If the indicator agrees with the TT1000A, examine the airplane wiring between the engine thermocouple harness connector and the ITT indicator. Make sure that the wiring is clean, there is proper contact, and there is correct alumel/chromel pin arrangement. 1 Examine the ITT thermocouple and the thermocouple harness. Refer to Pratt & Whitney Maintenance Manual procedure. (9) Disconnect the test harness from the ITT indicator. (10) Install the ITT indicator in the airplane. Refer to Chapter 77, Interturbine Temperature Indicator - Maintenance Practices. B.
Bench Test the ITT Indicator with the TT1200 Tester (Refer to Figure 505).
CAUTION: Approximately 45 VDC no-load is present on the test leads during resistance measurements. Let the test leads reach ambient temperature before any test. (1) (2) (3)
(4) (5) (6) (7) (8) (9)
Remove the ITT indicator from the airplane. Refer to Chapter 77, Interturbine Temperature Indicator - Maintenance Practices. Connect the test harness to the indicator. Complete the TT1200 self test before you connect the test set to the indicator. (a) Set the FUNCTION switch to RES MEAS. (b) Set the RANGE switch to 2M ohms and touch the test lead clips together. (c) Push the PTM button. If BAT is shown, replace the batteries. Set the FUNCTION switch to IND TEST. Set the RANGE switch to 2M ohms position (0 ohm system resistance). Put the 1°/.1° switch to either the 1° or .1° position. Set the °C/MV switch to °C or MV. Connect the black clip (-) of the TT1200 lead to the alumel wire connected to pin B on the indicator. Connect the red clip (+) of the TT1200 lead to the chromel wire connected to pin A on the indicator.
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MODEL 208 MAINTENANCE MANUAL (10) Turn the TEMP ADJ knob on the TT1200 tester until the display matches the test points. If necessary, turn the FINE knob to get the exact adjustment, and make sure that the indicator display is in the range of the speciÞed tolerance for each test point. Refer to Table 501. (a) If the indicator does not agree with the TT1200 input, make sure that the harness is clean, there is proper contact, and there is correct alumel/chromel pin arrangement. 1 Do a test to make sure that the indicator does not agree with the TT1200 input. (b) If the indicator matches the TT1200, examine the airplane wiring between the engine thermocouple harness connector and the ITT indicator. Make sure that the wiring is clean, there is proper contact, and there is correct alumel/chromel pin arrangement. 1 Examine the ITT thermocouple and the thermocouple harness. Refer to Pratt & Whitney Maintenance Manual procedure. (11) Disconnect the test harness from the ITT indicator. (12) Install the ITT indicator in the airplane. Refer to Chapter 77, Interturbine Temperature Indicator - Maintenance Practices. 4.
Interturbine Temperature Indicator Airplane Check A.
Check the ITT Indicator with the TT1000A Tester.
CAUTION: Approximately 45 VDC no-load is present on the test leads during resistance measurements. Let the test leads reach ambient temperature before any test. (1) (2) (3) (4) (5) (6)
Complete the TT1000A self test before you connect the test set to the indicator. (a) With the test set leads disconnected, set the RESISTANCE RANGE switch to the BATT position. Set the FUNCTION switch to INDICATOR TEST. Set the power switch to ON. The display should show 30 VDC to 50 VDC. Set the FUNCTION switch to RESISTANCE MEASURE. Set the RESISTANCE RANGE switch to 2M ohms position (0 ohm system resistance). Disconnect the airplane ITT gage thermocouple wires from the engine ITT harness and the TT5 trim harness. NOTE:
The terminals are marked on the engine side of the connection for the alumel/chromel identiÞcation.
Connect the black clip (-) of the TT1000A lead to the alumel wire connected to pin "B" on the indicator. (8) Connect the red clip (+) of the TT1000A lead to the chromel wire connected to pin "A" on the indicator. (9) Apply airplane power to the ITT indicator. (10) Push the TT1000A tester power button to ON. (11) Turn the TEMP ADJ knob on TT1000A until the tester display matches the test points, and make sure that the ITT indicator display is in the range of the speciÞed tolerance for each test point. Refer to Table 501. (a) If the indicator does not agree with the TT1000A input, make sure that the airplane harness is clean, there is proper contact, and there is correct alumel/chromel pin arrangement. 1 Do a test to make sure that the indicator does not agree with the TT1000A input. (b) If the TT1000A and the indicator agree, examine the ITT thermocouple and the thermocouple harness. Refer to Pratt & Whitney Maintenance Manual procedure.
(7)
B.
Check the ITT Indicator with the TT1200 Tester.
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CAUTION: Approximately 45 VDC no-load is present on the test leads during resistance measurements. Let the test leads reach ambient temperature before any test. (1) (2) (3) (4) (5) (6) (7) (8) (9)
Complete the TT1000A self test before you connect the test set to the indicator. Set the FUNCTION switch to RES MEAS. Set the RANGE switch to 2M ohms and touch the test lead clips together. Push the PTM button. If BAT is shown, replace the batteries. Set the FUNCTION switch to IND TEST. Set the RANGE switch to 2M ohms (0 ohm system resistance). Set the 1°/.1° switch to 1° or .1°. Set the °C/MV switch to °C or MV. Disconnect the airplane ITT gage thermocouple wires from the engine ITT harness and the TT5 trim harness. NOTE:
The terminals are marked on the engine side of the connection for the alumel/chromel identiÞcation.
(10) Connect the black clip (-) of the TT1200 lead to the alumel wire connected to pin "B" on the indicator. (11) Connect the red clip (+) of the TT1200 lead to the chromel wire connected to pin "A" on indicator. (12) Apply airplane power to the ITT indicator. (13) Push the TT1200 tester button to ON. (14) Turn the TEMP ADJ knob on the TT1200 tester until the display matches the test points. If necessary, turn the FINE knob to get the exact adjustment, and make sure that the ITT indicator display is in the range of the speciÞed tolerance for each test point. Refer to Table 501. (a) If the indicator does not agree with the TT1200 input, make sure that the harness is clean, there is proper contact, and there is correct alumel/chromel pin arrangement. 1 Do a test to make sure that the indicator does not agree with the TT1200 input. (b) If the TT1200 and the indicator agree, examine the ITT thermocouple and the thermocouple harness. Refer to Pratt & Whitney Maintenance Manual procedure. 5.
Insulation Check. A.
Check the Insulation with the TT1000A Tester (Refer to Figure 506).
CAUTION: Approximately 45 VDC no-load is present on the test leads during resistance measurements. Let the test leads reach ambient temperature before any test. (1) (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13)
Disconnect the airplane ITT harness from the engine ITT and the TT5 trim harness. Disconnect the connector from the ITT indicator. Connect the black clip (-) of the TT1000A to airframe ground. Connect the red clip (+) of the TT1000A to the alumel wire connected to pin B of the ITT connector. Set the RESISTANCE RANGE switch to 2M ohms. Set the FUNCTION switch to RESISTANCE MEASURE. Set the power switch to ON. Push the PUSH TO MEASURE button (black) and make sure that the indication is greater than 1 MEG ohm. Set the power switch to OFF. Connect the red clip (+) of the TT1000A to the chromel wire connected to pin A of the ITT connector. Set the RESISTANCE RANGE switch to 2M ohms. Set the FUNCTION switch to RESISTANCE MEASURE. Set the power switch to ON.
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MODEL 208 MAINTENANCE MANUAL (14) Push the PUSH TO MEASURE button (black) and make sure that the indication is greater than 1 MEG ohm. (15) Set the power switch to OFF. (16) Connect the airplane ITT harness to the engine ITT and the TT5 harness. (17) Install the connector on the ITT indicator. B.
Check the Insulation with the TT1200 Tester.
CAUTION: Approximately 45 VDC no-load is present on the test leads during resistance measurements. Let the test leads reach ambient temperature before any test. (1) (2) (3) (4) (5) (6)
Disconnect the airplane ITT harness from the engine ITT and the TT5 trim harness. Disconnect the connector from the ITT indicator. Connect the black clip (-) of the TT1200 to airframe ground. Connect the red clip (+) of the TT1200 to the alumel wire connected to pin B of the ITT connector. Set the FUNCTION switch to INS MEAS. Set the RANGE switch to 2M ohms. NOTE:
(7) (8) (9) (10) (11) (12)
Set the power switch to ON. Push the PTM button (black) and make sure that the indication is greater than 1 MEG ohm. Set the power switch to OFF. Connect the red clip of the TT1200 to the chromel wire connected to pin A of the ITT connector. Set the FUNCTION switch to INS MEAS. Set the RANGE switch to 2M ohms. NOTE:
(13) (14) (15) (16) (17)
The position of the 1°/.1° and °C/MV switches are irrelevant.
The position of the 1°/.1° and °C/MV switches are irrelevant.
Set the power switch to ON. Push the PTM button (black) and make sure that the indication is greater than 1 MEG ohm. Set the power switch to OFF. Connect the airplane ITT harness to the engine ITT and the TT5 harness. Install the connector on the ITT indicator.
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADAS+ ENGINE TREND MONITORING SYSTEM - DESCRIPTION AND OPERATION 1.
General A.
2.
3.
This section gives a general description and the operation of the altair engine trend monitoring system.
Description/Operation A.
Description. (1) Airplane and engine maintenance procedures are very important to flight safety and to decrease the cost of operation. (2) The Airplane Data Acquisition System Plus, known as the ADAS+, has three functions: (a) Exceedance Event Recording. 1 The ADAS+ monitors important engine parameters and records instances when preset values are exceeded. (b) Engine Trend Monitoring. 1 The ADAS+ records and stores engine data for trend analysis. (c) Cockpit Indication. 1 The ADAS+ will warn the pilot if an exceedance occurs. It can also show prior exceedance on engine start or shutdown. A cockpit self-test can be done. (3) The ADAS+ lets the operator control and schedule the engine maintenance operations. (4) In its data acquisition role, the ADAS+ is a passive receiver of information. (5) It can record trend data either manually or automatically.
B.
Operation. (1) Manual operation. (a) The pilot can record a dataset from all of the sensors with the ENGINE/ETM annunciator/ switch. (2) Automatic Operation. (a) The system can automatically record exceedance events and data samples that can be analyzed for trends. (3) Retrieving data. (a) Data is collected through a download serial port. (b) The Altair Avionics Monitor Link Program (MLP), can be used to download data and upload system configuration files. (c) The MLP is also used for system diagnostics, calibrations, and real-time live sensor display. (4) System configuration. (a) The ADAS+ system is shipped without a configuration. (b) The configuration is downloaded from the ALTAIR website and then uploaded to the processor for correct operation.
Tools, Equipment and Manuals NOTE:
Equivalents are approved.
NAME
NUMBER
Computer Laptop (with internet access) Serial Interface Adapter (Serial Cable)
TREND-C-033-1
MANUFACTURER
USE
Commercially available
To access the ADAS+ to get information and upload configurations.
Altair Avionics Corporation Customer Service Department 106 Access Road Norwood, MA 02062 Web: www.altairavionics.com Web: www.turbinetracker.com Phone: (781) 762-8600
To connect a laptop computer to the Altair Engine Trend Monitoring System.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
MANUFACTURER
USE
ADAS+ Instructions for Continued Airworthiness
ADAS-G-260-1
Altair Avionics Corporation
To do the troubleshooting procedures on the Altair Engine Trend Monitoring System.
MLP Users Guide
GSS-T-301-1
Altair Avionics Corporation
Used to help the customer use the Monitor Link Program.
Air Data Tester
101-00184
Barfield 4101 NW 29th Street Miami, FL 33142-5617
To supply pressure or vacuum for the pitot and static system tests.
Pitot Static Test Adaptor
PS4769
Nav-Aids Ltd. 2955 Diab Street Montreal, Quebec H4S 1M1
To attach portable air data tester to pitot system.
Dead Weight Pressure Tester
231FA
Barfield
Used to calibrate and test the torque sensor.
4.
Retrieving Data from ADAS+ Processor A.
Get access to the processor data. NOTE: (1) (2)
5.
Data collected is accessed through a download serial port under the copilot's instrument panel on the right side.
Use the Altair Avionics Monitor Link Program (MLP) to retrieve data from the ADAS+ processor. For the vendor publication, refer to Altair ADAS+ Engine Trend Monitoring System - Description and Operation.
Uploading Data A.
Upload the data. (1) To upload the data you must have a current Turbine Tracker account with Altair Avionics Corporation. (2) To get an account, contact the Altair Avionics Corporation at telephone number (781) 762-8600.
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADAS+ ENGINE TREND MONITORING SYSTEM - TROUBLESHOOTING 1.
General A.
2.
3.
This section gives troubleshooting information for the Altair ADAS+ Engine Trend Monitoring (ETM) system. The troubleshooting information given in this section refers to the startup of the ETM system. Also, this section provides many troubleshooting charts for errors in the log file that might be found in the Monitor Link Program (MLP). The charts give troubleshooting information for pitot/static sensor, outside air temperature (OAT) sensor, torque sensor, inner-turbine temperature (ITT) signal, Ng, Np, and Wf signal errors.
System Indications (Engine Off) A.
The ETM Lamp Stays On. (1) A sensor fault occurred in the previous flight. Refer to the ADAS+ Instructions for Continued Airworthiness manual found in the Altair ADAS+Engine Trend Monitoring System - Description and Operation.
B.
The ETM Lamp Flashes. (1) An engine parameter was exceeded in the previous flight and the log file was not downloaded from the processor. Refer to the ADAS+ Instructions for Continued Airworthiness manual found in the Altair ADAS+Engine Trend Monitoring System - Description and Operation.
System Start-Up and Communication A.
If the laptop doesn't connect with the processor, do the steps that follow. (1) Make sure that the airplane battery cables are connected to an APU or a battery. NOTE:
(2) (3)
Make sure that both the circuit breakers on the relay box assembly that is forward of the firewall are engaged. Monitor the annunciator lamps during the system start up. (a) Make sure that after approximately five seconds the annunciator lamps go out. NOTE:
(4) 4.
If the battery is not installed or the APU is not connected to the battery cables, and the APU is connected to the external power receptacle, then the annunciator will illuminate but the processor will not have power.
If the annunciator lamps stay on indefinitely, then the processor did not finish the start up process.
If the system starts successfully, but other problems continue, contact Cessna Product Support.
Log File Error Troubleshooting A.
Pitot/Static Sensor Troubleshooting (1) If a pitot/static error occurs in the log file, do the troubleshooting procedure to find if the error is caused by the sensor or the processor. Refer to Figure 101. (a) If the failure is caused by the sensor, do the applicable troubleshooting procedure. NOTE: (2)
If the pitot/static sensor is the cause of the failure, it will be a range failure, or a rate failure.
If an OAT error occurs in the log file, do the troubleshooting procedure to find if the error is caused by the sensor or the processor. Refer to Figure 102. (a) If the failure is caused by the sensor, do the applicable troubleshooting procedure. NOTE:
If the OAT sensor is the cause of the failure, it will be a bit, range, or rate failure.
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MODEL 208 MAINTENANCE MANUAL (3)
If a torque sensor error occurs in the log file, do the troubleshooting procedure. Refer to Figure 103. NOTE:
(4)
If an ITT sensor error occurs in the log file, do the troubleshooting procedure. Refer to Figure 104. NOTE:
(5)
If the Ng sensor is the cause of the failure, it will be a signal, range, or rate failure.
If an Np sensor error occurs in the log file, do the troubleshooting procedure. Refer to Figure 106. NOTE:
(7)
If the ITT sensor is the cause of the failure, it will be an engine temperature BIT failure, a range failure, or a rate failure.
If an Ng sensor error occurs in the log file, do the troubleshooting procedure. Refer to Figure 105. NOTE:
(6)
If the torque sensor is the cause of the failure, it will be a range, rate, or signal failure.
If the Np sensor is the cause of the failure, it will be a signal, range or rate failure.
If a Wf signal error occurs in the log file, do the troubleshooting procedure. Refer to Figure 107.
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Pitot/Static Sensor Troubleshooting Figure 101 (Sheet 1)
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Pitot/Static Sensor Troubleshooting Figure 101 (Sheet 2)
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Pitot/Static Sensor Troubleshooting Figure 101 (Sheet 3)
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Outside Air Temperature Sensor Troubleshooting Figure 102 (Sheet 1)
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Outside Air Temperature Sensor Troubleshooting Figure 102 (Sheet 2)
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Outside Air Temperature Sensor Troubleshooting Figure 102 (Sheet 3)
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Outside Air Temperature Sensor Troubleshooting Figure 102 (Sheet 4)
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Torque Sensor Troubleshooting Figure 103 (Sheet 1)
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Torque Sensor Troubleshooting Figure 103 (Sheet 2)
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Torque Sensor Troubleshooting Figure 103 (Sheet 3)
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Inner-Turbine Temperature Signal Troubleshooting Figure 104 (Sheet 1)
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Inner-Turbine Temperature Signal Troubleshooting Figure 104 (Sheet 2)
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Inner-Turbine Temperature Signal Troubleshooting Figure 104 (Sheet 3)
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Ng Signal Troubleshooting Figure 105 (Sheet 1)
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Ng Signal Troubleshooting Figure 105 (Sheet 2)
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Ng Signal Troubleshooting Figure 105 (Sheet 3)
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Np Signal Troubleshooting Figure 106 (Sheet 1)
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Np Signal Troubleshooting Figure 106 (Sheet 2)
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Np Signal Troubleshooting Figure 106 (Sheet 3)
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Wf Signal Troubleshooting Figure 107 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADAS+ ENGINE TREND MONITORING SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Tools and Equipment A.
3.
4.
This section gives the removal and installation of the Altair ADAS+ Engine Trend Monitoring (ETM) Components. The Altair ADAS+ ETM components are the processor, pitot and static transducers, annunciator, Outside Air Temperature (OAT) sensor torque transducer, and ice vane switch.
For a list of tools and equipment, refer to Altair ADAS+ Engine Trend Monitoring System - General.
Processor Removal/Installation A.
Remove the Processor (Refer to Figure 201). (1) Open the engine cowlings. (2) Disengage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. (3) Remove the straps that are around the wires and the processor. (4) Remove the nut, washer and screw that attach the bonding jumper to the processer. (5) Disconnect the electrical connectors (PB100, PB101, and PA09) from the processor. (6) Remove the nuts and washers from the clamp halves. (7) Remove the processor and bracket from the engine truss. (8) Remove the nuts and shock mounts from the bracket and processor. (9) Disconnect the processor from the bracket.
B.
Install the Processor (Refer to Figure 201). (1) Connect the processor to the bracket with the shock mounts and nuts. (2) Put the processor and bracket onto the engine truss. (3) Install the nuts and washers on the clamp halves of the bracket. (4) Install the screw, washer, and nut that attach the bonding jumper. (5) Connect the electrical connectors (PB100, PB101, and PA09) to the processor . (6) Install the tie straps on the wires to the processor. (7) Engage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. (8) Close the engine cowlings. (9) Complete a function test of the Altair ADAS+ ETM System. Refer to Altair ADAS+ Engine Trend Monitoring System - Adjustment/Test.
Outside Air Temperature (OAT) Sensor Removal/Installation A.
Remove the OAT Sensor (Refer to Figure 201). NOTE: (1) (2) (3) (4) (5) (6)
B.
The OAT sensor is found below the left engine cowling.
Open the left engine cowling. Disengage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. Remove the safety wire from the jam nut. Remove the jam nut, washer, and O-ring from the OAT sensor. Disconnect the electrical connector (PI93) from the OAT sensor (JI93). Remove the OAT sensor from the airplane.
Install the OAT Sensor (Refer to Figure 201). (1) Put the OAT sensor into the airplane. (2) Connect the electrical connector (PI93) to the OAT sensor (JI93). (3) Install the O-ring, washer, and jam nut on the OAT sensor. (4) Safety the jam nut with wire. Refer to Chapter 20 Safetying - Maintenance Practices. (5) Engage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment.
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Altair ADAS+ ETM Component Installation Figure 201 (Sheet 1)
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Altair ADAS+ ETM Component Installation Figure 201 (Sheet 2)
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Altair ADAS+ ETM Component Installation Figure 201 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (6) (7) 5.
Close the left engine cowling. Complete a function test of the Altair ADAS+ ETM System. Refer to Altair ADAS+ Engine Trend Monitoring System - Adjustment/Test.
Annunciator Removal/Installation A.
Remove the Annunciator (Refer to Figure 201). NOTE: (1) (2) (3)
The annunciator is a switch light.
Disengage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. Pull the switch light out of the instrument panel. Pry the face of the switch off of the annunciator NOTE:
(4) (5) (6)
6.
7.
The face of the switch will stay connected to the switch by its wiring.
Loosen the retaining tab screws that hold the backshell to the instrument panel. Disconnect the electrical connector from the switch light (SI505). Remove the switch light from the airplane.
B.
Install the Annunciator (Refer to Figure 201). (1) Put the switch light into the mounting sleeve in the instrument panel. (2) Install the retaining tab screws that hold the backshell to the instrument panel. (3) Connect the electrical connector to the switch light (SI505). (4) Engage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment.
C.
Complete a function test of the Altair ADAS+ ETM System. Refer to Altair ADAS+ Engine Trend Monitoring System - Adjustment/Test.
Ice Vane Installation Removal/Installation A.
Remove the Magnetic Reed Switch (Refer to Figure 201). (1) Open the left engine cowling. (2) Disengage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. (3) Remove the nuts, washers, and screws from the clamps that attach the magnet. (4) Remove the magnet. (5) Remove the washers and screws that attach the magnetic reed switch to the switch plate assembly. (6) Remove the clamps to where the wire splices into the wire bundle. (7) Disconnect the splices and remove the magnetic reed switch from the airplane.
B.
Install the Magnetic Reed Switch (Refer to Figure 201). (1) Connect the magnetic reed switch wire to the splices. (2) Put the wire into the clamps and install the screws into the clamps.. (3) Install the screws and washers that attach the magnetic reed switch to the switch plate assembly. (4) Install the screws into the clamps and attach the magnet. (5) Engage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. (6) Close the left engine cowling. (7) Complete a function test of the Altair ADAS+ ETM System. Refer to Altair ADAS+ Engine Trend Monitoring System - Adjustment/Test.
Pitot and Static Transducers Removal/Installation A.
Remove the Pitot-Static Transducers (Refer to Figure 201). (1) Disengage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment.
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MODEL 208 MAINTENANCE MANUAL (2) (3) (4) (5) (6) (7) (8) B.
Get access to the pitot-static transducer from under and behind the instrument panel. Disconnect the electrical connector (PI101) from the static transducer (PO1002). Disconnect the electrical connector (PI102) from the pitot transducer (PO1001). Loosen the hose adapter on the static transducer and remove the hose. Loosen the hose adapter on the pitot transducer and remove the hose. Remove the screws from the clamps that attach the pitot-static transducer. Remove the pitot-static transducer from the airplane.
Install the Pitot-Static Transducer (Refer to Figure 201). (1) Put the pitot transducer (PO1001) into the airplane. (2) Put the static transducer (PO1002) into the airplane. (3) Install the screws into the clamps that attach the pitot-static transducer. (4) Install the hose and tighten the hose adapter on the pitot transducer. (5) Install the hose and tighten the hose adapter on the static transducer. (6) Connect the electrical connector (PI102) onto the pitot transducer. (7) Connect the electrical connector (PI101) onto the static transducer. (8) Engage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. (9) Do a pitot system inspection and leak test. Refer to Pitot/Static System - Adjustment/Test. (10) Complete a function test of the Altair ADAS+ ETM System. Refer to Altair ADAS+ Engine Trend Monitoring System - Adjustment/Test.
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADAS+ ENGINE TREND MONITORING SYSTEM - ADJUSTMENT/TEST 1.
General A.
2.
Tools and Equipment A.
3.
This section gives the configuration procedures for the ADAS+ processor.
For a list of required tools and equipment, refer to the Altair ADAS+ Engine Trend Monitoring System - Description And Operation.
ADAS+ Processor Test A.
Do a test of the ADAS+ Processor. (1) Make sure that the battery is connected. (2) Put the battery switch to the ON position. (3) Make sure that both fault lamps on the left instrument panel come on. (4) Make sure that after approximately five seconds, both fault warning lamps go off. (5) If the ENGINE warning lamps do not stay off, troubleshoot the system. Refer to Altair ADAS+ Engine Trend Monitoring System - Troubleshooting. NOTE:
(6) 4.
If the processor is not configured after the ENGINE warning lamps are off, the ETM and ENGINE lamps can come on. The lamps only give correct and applicable indications after the processor is configured.
If the processor is configured and the lamps do not operate correctly, troubleshoot the system. Refer to Altair ADAS+ Engine Trend Monitoring System - Troubleshooting.
Pitot and Static Sensor Calibration Test A.
Do the Pitot and Static Sensor Calibration Test. (1) Connect the Air Data Tester (ADT) to the right side of the pitot-static system. (2) Configure the ADT to the airspeed and altitudes shown in Table 501, Pitot-Static Calibration Test Settings. (3) Make sure that the sensor values shown in the Live Data are in the specified range compared to the ADT gages. (4) If the values are in the range, do the OAT Sensor Calibration Test. (5) Click on the STOP LIVE DATA bar on the Monitor Link Program (MLP). (6) When you see the prompt by the MLP, put the RUN/CONF switch in the CONF position and click OK.
Table 501. Pitot-Static Calibration Test Settings
5.
Test Reading Number
Pressure Altitude (Feet)
Max Altitude Deviation (Feet)
Airspeed (Knots)
Max Airspeed Deviation (Knots)
1
1,500
+300 or -300
70
+15 or -15
2
15,000
+300 or -300
170
+15 or -15
Pitot and Static Sensors Calibrating Procedures A.
Calibrate the Pitot and Static Sensors. (1) Apply power to the airplane. (2) Connect the download cable between the communication port and the serial port. NOTE: (3)
The communication port is under the copilot's instrument panel. The serial port is on the backside of the laptop computer.
Apply power to the laptop.
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MODEL 208 MAINTENANCE MANUAL (4)
Make a connection between the laptop and the processor. (a) Open the Monitor Link Program (MLP) and click on EDIT from the pull down menu. (b) Make the COMMUNICATION PORT and AUTO DETECT selections. NOTE:
The MLP will automatically sense to which communication port the download cable is connected. The MLP display will show SUCCESSFULLY DETECTED COMM PORT at the top of the screen.
If the MLP doesn't sense the communication port, do the troubleshooting procedure. Refer to System Start-Up and Communication. Make the MLU selection from the pull down menu, then make the SENSOR CALIBRATION selection. Make the (10) ALTITUDE sensor selection from the pull down menu (c)
(5) (6)
NOTE: (7)
(a) Click OK. Set the ADT output to 1500 feet (29.92 in Hg reference) with no airspeed. NOTE:
(8) (9) (10) (11) (12)
The calibration type will default to TWO POINT.
The ADT’s airspeed indicator will not zero.
Enter the indicated altitude from the ADT in the first blank and click SET POINT 1. Set the ADT output to 15,000 feet (29.92 in Hg reference) with no airspeed. Enter the indicated altitude on the ADT in the second blank and click SET POINT 2. Click OK and the processor will complete the calibration. When you see the prompt, accept the new calibration coefficients. NOTE:
The MLP must have the current configuration file to change so that it can be uploaded to the Turbine Tracker later. It is important that the correct configuration file selection was made.
(13) The MLP will now let you make a selection to calibrate more sensors. (a) If the airspeed does not need calibration, do the OAT Sensor Calibration Test. (b) If necessary, click “YES” to calibrate the airspeed. 1 Make the (9) AIRSPEED sensor selection from the pull down menu NOTE: a
The calibration type will default to TWO POINT.
Click OK.
CAUTION: Do not let the vertical speed indicator show more than 2500 ft/min. This will help prevent damage to the indicator. 2 3 4 5 6 7 8 9
Gradually zero the airspeed output and let the altitude decrease to approximately 1320 feet. Disconnect the pitot and static hoses from the ADT. Enter 0 in the blank for Point 1 and click SET POINT 1. Connect the pitot and static hoses to the ADT. Set the ADT altitude to 15,000 feet and the airspeed to 170 knots. Enter the indicated airspeed on the ADT in the second blank and click SET POINT 2. Click OK and the processor will complete the calibration. When you see the prompt, accept the new calibration coefficients. NOTE:
The MLP must have the current configuration file to change so that it can be uploaded to the Turbine Tracker later. It is important that the correct configuration file selection was made.
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MODEL 208 MAINTENANCE MANUAL 10 The MLP will now let you make a selection to calibrate more sensors. 11 Click NO. (14) Do the Pitot and Static Sensor Calibration Test 6.
OAT Sensor Calibration Test A.
Do the OAT Sensor Calibration Test. (1) Make the LIVE DATA then TEXT VIEW selection from the pull down menu. NOTE: (2) (3) (4)
This selection will start a new Live Data session.
Give the LIVE DATA file a name and make the SAVE selection. When you see the prompt by the Monitor Link Program (MLP), put the RUN/CONF switch in the RUN position and click OK. Compare the ambient air temperature shown by the aircraft OAT gage with the temperature shown in Live Data. NOTE:
The difference between the two temperatures must be no more than +2 or -2°C.
(a)
(5) (6) 7.
If the OAT sensor temperature is more than +2 or -2°C different, do the sensor troubleshooting procedure. Refer to the ADAS+ Instructions for Continued Airworthiness manual, Altair ADAS+ Engine Trend Monitoring System - Description and Operation. Click on the STOP LIVE DATA bar on the MLP. When you see the prompt by the MLP, put the RUN/CONF switch in the CONF position and click OK.
Torque Sensor Calibration Test A.
Do the Torque Sensor Calibration Test. (1) Disconnect and put a cap on the high-pressure oil line. NOTE: (2) (3)
The high-pressure oil line is on the engine side of the tee forward of the engine mount ring.
Connect the hose from the dead weight tester to the tee fitting. Make the LIVE DATA then TEXT VIEW selection from the pull down menu. NOTE:
This selection will start a new Live Data session.
(4) (5)
Give the LIVE DATA file a name and make the SAVE selection. When you see the prompt by the Monitor Link Program (MLP), put the RUN/CONF switch in the RUN position and click OK. (6) Increase the pressure in the dead weight tester until the instrument panel torque gage becomes stable at the 1980 ft-lbs red mark. (7) Compare the sensor value shown in Live Data to the instrument panel reading. (a) If the value is less than +2 or -2 percent of the instrument panel gage then the torque sensor is sufficiently calibrated. (b) If the value is more than +2 or -2 percent of the instrument panel gage, do the Torque Sensor Calibrating Procedures. (8) Stop the current LIVE DATA session. (9) Disconnect the dead weight tester. (10) Remove the cap and connect the high-pressure oil line to the transducer. (11) Do the Discrete Switch Tests. 8.
Torque Sensor Calibrating Procedures A.
Calibrate the Torque Sensor. (1) Apply power to the airplane.
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MODEL 208 MAINTENANCE MANUAL (2)
Connect the download cable between the communication port and the serial port. NOTE:
(3) (4)
The communication port is under the copilot's instrument panel. The serial port is on the backside of the laptop computer.
Apply power to the laptop. Make a connection between the laptop and the processor. (a) Open the Monitor Link Program (MLP) and click on EDIT from the pull down menu. (b) Make the COMMUNICATION PORT and AUTO DETECT selections. NOTE:
The MLP will automatically sense to which communication port the download cable is connected. The MLP display will show SUCCESSFULLY DETECTED COMM PORT at the top of the screen.
If the MLP doesn't sense the communication port, do the troubleshooting procedure. Refer to System Start-Up and Communication. Make the MLU selection from the pull down menu, then make the SENSOR CALIBRATION selection. Make the (3) ENGINE TORQUE sensor selection from the pull down menu. (c)
(5) (6)
NOTE:
The calibration type will default to TWO POINT.
(a) Click OK. Increase the pressure in the dead weight tester until the instrument panel torque gage becomes stable at the 1980 ft-lbs red mark. (8) Enter the instrument panel torque gage reading in the first blank and click SET POINT 1. (9) Decrease the pressure in the dead weight tester until the instrument panel torque gage becomes stable at the 1000 ft-lbs. (10) Enter the instrument panel torque gage reading in the second blank and click SET POINT 2. (11) Click OK and the processor will complete the calibration. (12) When you see the prompt, accept the new calibration coefficients. (7)
NOTE:
The MLP must have the current configuration file to change so that it can be uploaded to the Turbine Tracker later. It is important that the correct configuration file selection was made.
(13) The MLP will now let you make a selection to calibrate more sensors. (a) Click NO. (14) Do the Torque Sensor Calibration Test 9.
Discrete Switch Tests A.
Do the Discrete Switch Tests. (1) Make the LIVE DATA then TEXT VIEW selection from the pull down menu. NOTE: (2) (3) (4)
This selection will start a new Live Data session.
Give the LIVE DATA file a name and make the SAVE selection. When you see the prompt by the Monitor Link Program (MLP), put the RUN/CONF switch in the RUN position and click OK. Complete a check of the BLEED AIR HEAT switch. (a) Set the BLEED AIR HEAT switch to the OFF position. (b) Make sure that the LIVE DATA shows 0d. (c) Set the BLEED AIR HEAT switch to the ON position. (d) Make sure that the LIVE DATA shows 1d. (e) Set the BLEED AIR HEAT switch to the OFF position. (f) Make sure that the LIVE DATA shows 0d.
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MODEL 208 MAINTENANCE MANUAL (5)
(6)
(7) (8) 10.
Complete a check of the Emergency Power Lever (EPL). (a) Make sure that the engine is off. (b) Cut and remove the frangible/shear wire from the EPL. (c) Set the EPL to the NORM position. (d) Make sure that the LIVE DATA shows 1d. (e) Set the EPL to the MAX position. (f) Make sure that the LIVE DATA shows 0d. (g) Set the EPL to the NORM position. (h) Make sure that the LIVE DATA shows 1d. (i) Install the frangible/shear wire. Refer to Chapter 76, Emergency Power Lever Frangible/ Shear Wire Removal/Installation Complete a check of the INERTIAL SEPARATOR handle. (a) Push the INERTIAL SEPARATOR handle to the NORMAL position. (b) Make sure that the LIVE DATA shows 0d. (c) Pull the INERTIAL SEPARATOR handle to the BYPASS position. (d) Make sure that the LIVE DATA shows 1d. (e) Push the INERTIAL SEPARATOR handle to the NORMAL position. (f) Make sure that the LIVE DATA shows 0d. Click on the STOP LIVE DATA bar on the MLP, . When you see the prompt by the MLP, put the RUN/CONF switch in the CONF position and click OK.
Live Data Sensor Test – Engine Ground Run A.
Do the Live Data Sensor Test – Engine Ground Run.
WARNING: Immediately shut down the engine if an exceedence occurs in any of the engine operating limitations, or if any incorrect engine operation occurs during any of the procedures that follow. Refer to the Pratt & Whitney Maintenance Manual for the applicable procedure. NOTE:
(1) (2) (3)
To let the Altair ADAS+ ETM processor have time to complete its own start-up procedure, the battery power must be started at least 5 seconds before the engine start. After approximately 5 seconds, the white ETM lamp and the amber ENGINE lamp will go off. If one or two of the lamps stay on or flash, troubleshooting procedures must be completed.
Do the steps that follow during an engine ground run. Make sure that the engine is stable. Push the ENGINE/ETM annunciator/switch. NOTE:
(4)
Record the cockpit instrument values. NOTE:
(5)
The values are recorded on a generic form equivalent to Table 502, Engine Ground Run Data.
Connect the download cable between the communication port and the serial port. NOTE:
(6) (7)
The airplane engine running condition must stay stable while the white ETM lamp flashes and the processor records the trend data.
The communication port is under the copilot's instrument panel. The serial port is on the backside of the laptop computer.
Apply power to the laptop. Make a connection between the laptop and the processor. (a) Open the Monitor Link Program (MLP) and click on EDIT from the pull down menu.
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MODEL 208 MAINTENANCE MANUAL (b)
Make the COMMUNICATION PORT and AUTO DETECT selections. NOTE:
The MLP will automatically sense to which communication port the download cable is connected. The MLP display will show SUCCESSFULLY DETECTED COMM PORT at the top of the screen.
(c)
(8) (9)
If the MLP doesn't sense the communication port, do the troubleshooting procedure. Refer to System Start-Up and Communication. Make the MLU and RETRIEVE UNIT’S DATA LOG selections from the pull down menu. The MLP will now let you make selections of the name and location for the file. (a) After the file is downloaded the MLP will let you make a selection to rest the log. Click YES. NOTE:
If you click YES, the log will be uploaded to the Turbine Tracker site.
NOTE:
The log file will automatically be shown by the MLP after it has been downloaded.
1 If the log file it is not shown by the MLP, click on VIEW in the pull down menu. 2 Then click on DATA LOG FILE and COMPLETE to see the log file. (10) To find the manual trend data in the log file, look for a paragraph with the heading TRENDING DATA. (11) Make sure that the time/date stamp of the manual trend data agrees with the time/date of the hand-recorded data. (12) Compare the average indications (not the maximum indications) from the log file to the entries recorded in generic form. (a) Make sure that the indications are in the ranges shown on Table 502, Engine Ground Run Data. 1 When you compare the data recorded during the engine run to the log data, make sure that you make an adjustment for the dial read error. (b) If all the sensor indications are not in tolerance, do the steps that follow. (c) If the Ng, Np, ITT, or Wf sensor indications are not in tolerance, contact the Altair Avionics Corporation at telephone number (781) 762-8600. (d) If the torque sensor indication is not in tolerance, calibrate the torque sensor. Refer to Torque Sensor Calibrating Procedures. (e) If the torque sensor was calibrated again, do the Live Data Sensor Test again. Refer to Live Data Sensor Test – Engine Ground Run. Table 502. Engine Ground Run Data Engine Run Ground Test
Cockpit Indicator Reading
Altair ADAS+ Reading
Difference
Maximum Deviation
Ng
+0.3 or -0.3 percent
Np
+3 or -3 percent
ITT
+5 or -5°C
Wf
+5 or -5 percent
Torque
+2 or -2 percent
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MODEL 208 MAINTENANCE MANUAL ALTAIR ENGINE TREND MONITORING TORQUE TRANSDUCER - MAINTENANCE PRACTICES 1.
General A.
2.
Tools and Equipment A.
3.
This section gives the removal and installation procedures for the Altair Engine Trend Monitoring (ETM) System torque transducer.
For a list of tools and equipment, refer to Altair ADAS+ Engine Trend Monitoring System - Description and Operation.
Torque Transducer Removal/Installation A.
Remove the Torque Transducer (Refer to Figure 201). (1) Open the right engine cowling. (2) Disengage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. (3) Disconnect the electrical connector (PD001) from the torque transducer. (4) Disconnect the oil pressure line. (5) Disconnect the elbow of oil vent line. (6) Put a cap on the oil vent line and oil pressure line to prevent the entry of foreign object debris. (7) Remove the nut, washer, and bolt from the bracket that attaches the torque transducer to the tube. (8) Remove the bracket and torque transducer. (9) Remove the bolts from the clamps that attach the torque transducer to the bracket. (10) Remove the torque transducer from the bracket.
B.
Install the Torque Transducer (Refer to Figure 201). (1) Put the torque transducer onto the bracket. (2) Install the screws into the clamps that attach the torque transducer to the bracket. (3) Install the bracket and torque transducer. (4) Install the bolts, washers, and nuts into the bracket. (5) Remove the caps on the oil vent line and oil pressure line. (6) Connect the oil vent line. (7) Connect the oil pressure line. (8) Connect the electrical connector (PD001) to the torque transducer. (9) Engage the ETM POWER (CB77) and ETM CONTINUOUS POWER (CB78) circuit breakers on the left side of the engine compartment. (10) Install the engine cowlings. (11) Complete a functional test of the Altair ETM System. Refer to Altair ADAS+Engine Trend Monitoring System - Adjustment/Test.
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Torque Transducer Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADASd ENGINE TREND MONITORING - DESCRIPTION AND OPERATION 1.
General A.
2.
3.
This section gives a general description and the operation of the Altair engine trend monitoring system.
Description/Operation A.
Description. (1) The Airplane Data Acquisition System Digital, known as the ADASd, has three functions: (a) Exceedance Event Recording. 1 The ADASd monitors important engine parameters and records instances when preset values are exceeded. (b) Engine Trend Monitoring. 1 The ADASd records and stores engine data for trend analysis. (c) Cockpit Indication. 1 The ADASd will warn the pilot if an exceedance occurs. It can also show prior exceedance on engine start or shutdown. A cockpit self-test can be done. (2) The ADASd lets the operator control and schedule the engine maintenance operations. (3) In its data acquisition role, the ADASd is a passive receiver of information. (4) It can record trend data either manually or automatically.
B.
Operation. (1) Manual operation. (a) The pilot can record a dataset from all of the sensors with the TRND/ACK or CAPTURE softkeys. (2) Automatic Operation. (a) The system can automatically record exceedance events and data samples that can be analyzed for trends. (3) Retrieving data. (a) Data is collected through a download serial port. (b) The Altair Avionics Monitor Link Program (MLP) can be used to download data and upload system conÞguration Þles. (c) The MLP is also used for system diagnostics and real-time live sensor display. (4) System conÞguration. (a) The ADASd system is shipped without a conÞguration. (b) The conÞguration is downloaded from the ALTAIR website and then uploaded to the processor for correct operation.
Tools, Equipment and Manuals NOTE:
Equivalents are approved.
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MODEL 208 MAINTENANCE MANUAL
NAME
NUMBER
Computer Laptop (with internet access)
MANUFACTURER
USE
Commercially available
To access the ADASd to get information and upload conÞgurations.
Serial Interface Adapter (Serial Cable)
TREND-C-033-1
Altair Avionics Corporation Customer Service Department 106 Access Road Norwood, MA 02062 Web: www.altairavionics.com Web: www.turbinetracker.com Phone: (781) 762-8600
To connect a laptop computer to the Altair Engine Trend Monitoring System.
Serial Interface Adapter (USB Cable)
TREND-C-053-1
Altair Avionics Corporation Customer Service Department 106 Access Road Norwood, MA 02062 Web: www.altairavionics.com Web: www.turbinetracker.com Phone: (781) 762-8600
To connect a laptop computer to the Altair Engine Trend Monitoring System.
MLP Users Guide
GSS-T-301-1
Altair Avionics Corporation
Used to help the customer use the Monitor Link Program.
4.
Retrieving Data from ADASd Processor A.
Get access to the processor data. NOTE: (1) (2)
5.
Data collected is accessed through a download serial port under the copilot's instrument panel on the right side.
Use the Altair Avionics Monitor Link Program (MLP) to retrieve data from the ADASd processor. For the vendor publication, refer to Altair ADASd Engine Trend Monitoring - Description and Operation.
Uploading Data A.
Upload the data. (1) To upload the data you must have a current Turbine Tracker account with Altair Avionics Corporation. (2) To get an account, contact Cessna Propeller Aircraft Product Support for assistance; (316) 5175800 or Fax (316) 942-9006.
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADASd ENGINE TREND MONITORING - TROUBLESHOOTING 1.
General A.
2.
3.
This section gives troubleshooting information for the Altair ADASd Engine Trend Monitoring (ETM) system. The troubleshooting information given in this section refers to the startup of the ETM system. Also, this section provides many troubleshooting procedures for errors in the log Þle that might be found in the Monitor Link Program (MLP).
System Indications (Engine Off) A.
The ETM FAULT message is shown. (1) Download the log Þle and do the troubleshooting to correct the error. Refer to ALTAIR ADASd Engine Trend Monitoring System - Description and Operation.
B.
The PREV EXCEED message is shown. (1) An engine parameter was exceeded in a previous ßight and the log Þle was not downloaded from the processor. Download the log Þle to clear the event.
System Start-Up and Communication A.
If the laptop doesn't connect with the processor, do the steps that follow. (1) Make sure that the airplane battery cables are connected to an APU or a battery. (2) Make sure that both the circuit breakers on the relay box assembly, that is forward of the Þrewall, are engaged. (3) Monitor the CAS messages while the avionics system starts up. (a) Make sure that after the avionics system starts up, ETM CAPTURE and PREV EXCEED messages are not shown. NOTE: (4)
4.
If ETM CAPTURE and PREV EXCEED messages are shown after the log is cleared, then the processor did not Þnish the start up process.
If the system starts successfully, but other issues continue, contact Cessna Propeller Aircraft Product Support for assistance; (316) 517-5800 or Fax (316) 942-9006.
System Troubleshooting A.
For additional troubleshooting, refer to Figure 101 and Figure 102.
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ARINC Troubleshooting Figure 101 (Sheet 1)
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Sensor Troubleshooting Figure 102 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADASd ENGINE TREND MONITORING SYSTEM - MAINTENANCE PRACTICES 1.
General A.
2.
Tools and Equipment A.
3.
This section gives the removal and installation of the Altair ADASd Engine Trend Monitoring (ETM) processor.
For a list of tools and equipment, refer to Altair ADASd Engine Trend Monitoring - Description and Operation.
Processor Removal/Installation A.
Remove the Processor (Refer to Figure 201). (1) Disconnect electrical power from the aircraft. (2) Remove the end cover from the copilot's avionics rack. NOTE: (3) (4) (5) (6)
B.
The copilot's avionics rack is found in front of the copilot's door below the instrument panel.
Disconnect the electrical connectors. Remove the screws that attach the bracket. Remove the bracket with the processor attached from the airplane. Remove the nuts, washers, ferrules, and shock mounts that attach the processor to the bracket.
Install the Processor (Refer to Figure 201). (1) Install the processor on the bracket with shock mounts, ferrules, washers, and nuts. (2) Install the bracket and processor assembly in the copilot's avionics rack. (3) Connect the electrical connectors to the processor. (4) Install the end cover on the copilots avionics rack. (5) Connect electrical power to the aircraft. (6) Do a Live Data Sensor Test of the Altair ADASd ETM System. Refer to Altair ADASd Engine Trend Monitoring System - Adjustment/Test.
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Altair ADASd ETM Component Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL ALTAIR ADASd ENGINE TREND MONITORING SYSTEM - ADJUSTMENT/TEST 1.
General A.
2.
Tools and Equipment A.
3.
For a list of required tools and equipment, refer to the Altair ADASd Engine Trend Monitoring System - Description And Operation.
ADASd Processor Test A.
4.
This section gives the tests for the ADASd processor.
Do a test of the ADASd Processor. (1) Make sure that the battery is connected. (2) Put the battery switch to the ON position. (3) Put the AVIONICS 1 and AVIONICS 2 switches to the ON position. (4) If a PREV EXCEED message is shown, push the TRND/ACK softkey to clear the message. (a) Download the log Þle to Þnd the cause of the message. (5) If an ETM FAULT message is shown, clear the log Þle. Refer to ALTAIR ADASd Engine Trend Monitoring System - Description and Operation. (6) If the processor is conÞgured and CAS messages do not operate correctly, do the troubleshooting for the system. Refer to Altair ADASd Engine Trend Monitoring Troubleshooting.
Live Data Sensor Test – Engine Off A.
Do the Live Data Sensor Test – Engine Off Tests. (1) Make the LIVE DATA then TEXT VIEW selection from the pull down menu. NOTE: (2) (3) (4)
(5)
(6)
This selection will start a new Live Data session.
Give the LIVE DATA Þle a name and make the SAVE selection. When you see the prompt by the Monitor Link Program (MLP), put the RUN/CONF switch in the RUN position and click OK. Complete a check of the BLEED AIR HEAT switch. (a) Set the BLEED AIR HEAT switch to the OFF position. 1 Make sure that the LIVE DATA shows 0d. (b) Set the BLEED AIR HEAT switch to the ON position. 1 Make sure that the LIVE DATA shows 1d. (c) Set the BLEED AIR HEAT switch to the OFF position. 1 Make sure that the LIVE DATA shows 0d. Complete a check of the Emergency Power Lever (EPL). (a) Make sure that the engine is off. (b) Cut and remove the frangible/shear wire from the EPL. (c) Set the EPL to the NORM position. 1 Make sure that the LIVE DATA shows 0d. (d) Set the EPL to the MAX position. 1 Make sure that the LIVE DATA shows 1d. (e) Set the EPL to the NORM position. 1 Make sure that the LIVE DATA shows 0d. (f) Install the frangible/shear wire. Refer to Chapter 76, Emergency Power Lever Frangible/ Shear Wire Removal/Installation. Complete a check of the INERTIAL SEPARATOR handle. (a) Push the INERTIAL SEPARATOR handle to the NORMAL position. 1 Make sure that the LIVE DATA shows 0d. (b) Pull the INERTIAL SEPARATOR handle to the BYPASS position. 1 Make sure that the LIVE DATA shows 1d.
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MODEL 208 MAINTENANCE MANUAL (c)
(7) (8) (9) 5.
Push the INERTIAL SEPARATOR handle to the NORMAL position. 1 Make sure that the LIVE DATA shows 0d. Make sure all sensor values agree with values shown on PFD1, PFD2, and MFD. (a) Record the values. Click on the STOP LIVE DATA bar on the MLP. When you see the prompt by the MLP, put the RUN/CONF switch in the CONF position and click OK.
Live Data Sensor Test – Engine Ground Run A.
Do the Live Data Sensor Test – Engine Ground Run.
WARNING: Immediately shut down the engine if an exceedance occurs in any of the engine operating limitations, or if any incorrect engine operation occurs during any of the procedures that follow. Refer to the P&W Maintenance Manual for the applicable procedure. (1) (2) (3) (4)
Make sure that the engine is stable. Push the ENGINE softkey. Push the TRND/ACK softkey. Hold engine parameters steady for at least 5 seconds. NOTE:
(5)
Record the cockpit instrument values. NOTE:
(6)
The values are recorded on a generic form equivalent to Table 502, Engine Ground Run Data.
Connect the download cable between the communication port and the serial port. NOTE:
(7) (8)
No message will appear to show that data is recorded.
The communication port is under the copilot's instrument panel. The serial port is on the laptop computer.
Apply power to the laptop. Make a connection between the laptop and the processor. (a) Open the Monitor Link Program (MLP) and click on EDIT from the pull down menu. (b) Make the COMMUNICATION PORT and AUTO DETECT selections. NOTE:
The MLP will automatically sense to which communication port the download cable is connected. The MLP display will show SUCCESSFULLY DETECTED COMM PORT at the top of the screen.
(c)
If the MLP doesn't sense the communication port, do the troubleshooting procedure. Refer to System Start-Up and Communication. (9) Make the MLU and RETRIEVE UNIT’S DATA LOG selections from the pull down menu. (10) The MLP will now let you make selections of the name and location for the Þle. (a) After the Þle is downloaded the MLP will let you make a selection to reset the log. Click YES. NOTE:
If you click YES, the log will be uploaded to the Turbine Tracker site.
NOTE:
The log Þle will automatically be shown by the MLP after it has been downloaded.
1 2
If the log Þle it is not shown by the MLP, click on VIEW in the pull down menu. Then click on DATA LOG FILE and COMPLETE to see the log Þle.
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MODEL 208 MAINTENANCE MANUAL (11) To Þnd the manual trend data in the log Þle, look for a paragraph with the heading TRENDING DATA. (12) Make sure that the time/date stamp of the manual trend data agrees with the time/date of the hand-recorded data. (13) Compare the average indications (not the maximum indications) from the log Þle to the entries recorded in generic form. (a) Make sure that the indications are in the ranges shown on Table 502, Engine Ground Run Data. (b) If all the sensor indications are not in tolerance, do the steps that follow. (c) Make sure wiring to the ADASd processor is correct. (d) Make sure the data on the G1000 is correct. 1 If the incorrect data is shown, do the troubleshooting for the system with the indication error. 2 If the correct data is shown, contact Cessna Propeller Aircraft Product Support for assistance; (316) 517-5800 or Fax (316) 942-9006. Table 501. Engine Ground Run Data Engine Run Ground Test
Cockpit Indicator Reading
Altair ADASd Reading
Difference
Maximum Deviation
Ng
+0.3 or -0.3 percent
Np
+3 or -3 percent
ITT
+5 or -5°C
Wf
+5 or -5 percent
Torque
+2 or -2 percent
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MODEL 208 MAINTENANCE MANUAL GEA-71 ENGINE/AIRFRAME UNIT - MAINTENANCE PRACTICES 1.
2.
General A.
The GEA-71 Engine/Airframe Unit is a microprocessor line replaceable unit (LRU). It is used to monitor sensor inputs and operate annunciator outputs for the airframe and engine systems.
B.
The GEA-71 sensor inputs are the engine oil pressure, engine oil temperature, engine torque transducer, ITT probes, engine thermocouple, chip detectors, and inertial separator switch. The fuel ßow transducer, Ng tachometer sensor, and Np tachometer sensor supply inputs to the GEA-71 through a signal conditioner.
GEA-71 Engine/Airframe Unit Removal/Installation A.
Remove the Engine/Airframe Unit (Refer to Figure 201). (1) Disconnect electrical power from the airplane. (a) Disengage the ENG INTFC circuit breaker. (2) Remove the MFD from the instrument panel. Refer to Chapter 34, Garmin Display Unit Maintenance Practices. (3) Remove the screw from the lock lever. (4) Lift the lock lever to release the unit from the avionics rack. (5) Remove the unit from the airplane.
B.
Install the Engine/Airframe Unit (Refer to Figure 201). (1) Install the unit in the avionics rack. (2) Lower the lock lever. (3) Install the screw in the lock lever. (4) Install the MFD. Refer to Chapter 34, Garmin Display Unit - Maintenance Practices. (5) Engage the ENG INTFC circuit breaker. (6) Connect electrical power to the airplane. (7) If a new unit is installed, load the software and conÞguration. Refer to the Garmin G1000 Line Maintenance Manual. (8) Do a check to make sure the transponder operates correctly. Refer to the Garmin G1000 Line Maintenance Manual.
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GEA-71 Engine/Airframe Unit Installation Figure 201 (Sheet 1)
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78 CHAPTER
EXHAUST
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MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
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By
Date Removed
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CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS EXHAUST - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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PRIMARY AND SECONDARY EXHAUST DUCT - MAINTENANCE PRACTICES . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Primary Exhaust Duct Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Secondary Exhaust Duct Removal/Installation (Without Cargo Pod) . . . . . . . . . . . . . Secondary Exhaust Duct Removal/Installation (With Twisted Exhaust Duct) . . . . . . Primary Exhaust Duct Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Secondary Exhaust Duct Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Field Installation of New Secondary Exhaust Duct. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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PRIMARY AND SECONDARY EXHAUST DUCT - INSPECTION/CHECK . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Primary and Secondary Exhaust Duct General Visual Inspection (Alignment Check) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Primary and Secondary Exhaust Duct General Visual Inspection . . . . . . . . . . . . . . . .
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LIST OF TASKS 78-10-00-210
Primary and Secondary Exhaust Duct General Visual Inspection (Alignment Check)
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Primary and Secondary Exhaust Duct General Visual Inspection
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MODEL 208 MAINTENANCE MANUAL EXHAUST - GENERAL 1.
Scope A.
2.
This chapter provides information on the components used to direct exhaust gases from the engine.
DeÞnition A.
This chapter is divided into two sections. (1) The section on Primary and Secondary Exhaust Duct - Maintenance Practices provides removal and installation information on the exhaust ducts. (2) The section on Primary and Secondary Exhaust Duct - Inspection/Check provides inspection and adjustment procedures for aligning the primary and secondary exhaust ducts.
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MODEL 208 MAINTENANCE MANUAL PRIMARY AND SECONDARY EXHAUST DUCT - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
This section gives procedures for the removal, installation and inspection of the primary and secondary exhaust ducts.
The engine exhaust system has a primary exhaust duct and a secondary exhaust duct. These ducts let the hot exhaust gases flow from the gas generator section of the engine overboard and away from the airplane. (1) The primary duct attaches to the lower right front of the engine and lets the exhaust flow into the secondary exhaust duct. The fit between the primary exhaust duct and the secondary duct has a small clearance, but the two units are not physically attached together. (a) The primary exhaust duct has a welded corrosion-resistant stainless steel assembly which is attached with the primary exhaust flange to the forward right side of the engine. (2) The secondary exhaust duct goes through the lower right section of the cowling. It is attached to the cowling and lets the exhaust flow away from the airframe. An enclosure around the primary exhaust duct lets the cool air that went through the right nose cap and oil cooler flow into the secondary exhaust duct. (a) On airplanes without a cargo pod, the secondary exhaust duct is a welded corrosionresistant stainless steel assembly. A two hanger and bracket assembly attaches the aft portion of duct to the lower right cowling panel. The design of the duct lets the engine exhaust gases flow under the airplane. (b) On airplanes with a cargo pod, with a partial TKS system or fairing TKS system installed, the secondary exhaust duct is a welded corrosion-resistant stainless steel assembly or an inconel assembly. Three hanger and bracket assemblies attach the aft portion of the duct to the lower right cowling panel. On Airplanes 208B0001 thru 208B0249 and Airplanes 20800106 thru 20800199 and Airplanes 20800001 thru 20800105 that include SK208-23 but do not include CAB90-27, two additional hanger and bracket assemblies attach the midportion of the duct to the lower right cowling panel. On Airplanes 20800001 thru 20800197 and Airplanes 208B0001 thru 208B0235 that include CAB90-27, the additional hanger and bracket assemblies are removed. On Airplanes 20800316 and On and Airplanes 208B0800 and On and Airplanes 20800001 thru 20800260 and Airplanes 208B0001 thru 208B0597 that include CAB00-8 and Airplanes 20800261 thru 20800315 and Airplanes 208B0598 thru 208B0799 that include CAB00-9, and airplanes with a partial TKS or fairing TKS system installed, there are three hanger assemblies with rod ends. The design of the duct lets the engine exhaust gases flow around the cargo pod.
Primary Exhaust Duct Removal/Installation A.
Remove the Primary Exhaust Duct (Refer to Figure 201 and Figure 202). (1) Remove the lower right cowling panel. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. (2) Disengage the quarter-turn fasteners that attach the oil cooler shroud to the engine. (3) Remove the oil cooler shroud from the engine. (4) Remove the bolts, washers, and nuts that attach the primary exhaust duct and primary exhaust flange to the engine. (5) Remove the primary exhaust duct and primary exhaust flange from the engine.
B.
Install the Primary Exhaust Duct (Refer to Figure 201 and Figure 202). (1) Put the primary exhaust duct in position to the primary exhaust flange. (2) Install the nuts, washers, and bolts that attach the primary exhaust duct and primary exhaust flange to the engine. (3) Do a Primary and Secondary Exhaust Duct General Visual Inspection (Alignment Check). Refer to Primary and Secondary Exhaust Duct - Inspection/Check. (4) Put the oil cooler shroud in position to the engine. (5) Engage the quarter-turn fasteners that attach the oil cooler shroud to the engine .
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MODEL 208 MAINTENANCE MANUAL (6) 4.
Secondary Exhaust Duct Removal/Installation (Without Cargo Pod) A.
Remove the Secondary Exhaust Duct (Refer to Figure 201). (1) Remove the lower right cowling panel. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices.
(2) (3) (4) B.
5.
Install the lower right cowling panel. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices.
NOTE:
The secondary exhaust duct is attached with bolts to the lower right cowling panel. These assemblies are removed as a single unit.
NOTE:
Some airplanes without a cargo pod have a cargo pod type of secondary exhaust duct. This information is found in Secondary Exhaust Duct Removal and Installation (with Cargo Pod).
Remove the three bolts, washers, and nuts that attach the oil breather hose attach neck to the secondary exhaust duct. Remove the bolts, washers, and nuts (if applicable) that attach the secondary exhaust duct bellmouth to the lower right cowling panel. Move the secondary exhaust duct forward until the hanger brackets are off of the aft hangers.
Install the Secondary Exhaust Duct (Refer to Figure 201). (1) Attach the secondary exhaust duct bellmouth to the lower right cowling panel. (a) Move the secondary exhaust duct aft through the lower right cowling panel. (b) Align the hanger brackets with the aft hangers. (c) Move the secondary exhaust duct aft until the hanger brackets are on the aft hangers. (d) Install the nuts (if applicable), washers, and bolts that attach the secondary exhaust duct bellmouth to the lower right cowling panel. (2) Install the three nuts, washers, and bolts that attach the oil breather hose attach neck to the secondary exhaust duct. (3) Do a Primary and Secondary Exhaust Duct General Visual Inspection (Alignment Check). Refer to Primary and Secondary Exhaust Duct - Inspection/Check. (4) Install the lower right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
Secondary Exhaust Duct Removal/Installation (With Twisted Exhaust Duct) A.
Remove the Secondary Exhaust Duct (Refer to Figure 202). (1) Remove the lower right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices.
(2) (3) (4) (5)
NOTE:
The secondary exhaust duct is attached with bolts to the lower right cowling panel. These assemblies are removed as a single unit.
NOTE:
Some airplanes without a cargo pod have a cargo pod type of secondary exhaust duct.
Remove the three bolts, washers, and nuts that attach the oil breather hose attach neck to the secondary exhaust duct. Remove the bolts, washers, and nuts (if applicable) that attach the secondary exhaust duct bellmouth to the lower right cowling panel. On Airplanes 20800006 thru 20800105 that do not include SK208-23, move the secondary exhaust duct forward out of the lower right cowling panel and off the aft hangers. On Airplanes 20800106 thru 20800199 and Airplanes 208B0001 thru 208B0249 and Airplanes 20800001 thru 20800105 that include SK208-23 and do not include CAB90-27, move the secondary exhaust duct forward out of lower right cowling panel and off the mid support hangers and the aft hangers.
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Engine Exhaust System Installation (Without Cargo Pod) Figure 201 (Sheet 1)
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Engine Exhaust System Installation (Without Cargo Pod) Figure 201 (Sheet 2)
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Engine Exhaust System Installation (With Twisted Exhaust Duct) Figure 202 (Sheet 1)
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Engine Exhaust System Installation (With Twisted Exhaust Duct) Figure 202 (Sheet 2)
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Engine Exhaust System Installation (With Twisted Exhaust Duct) Figure 202 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL (6)
B.
On Airplanes 20800316 and On and Airplanes 208B0800 and On and Airplanes 20800001 thru 20800260 and Airplanes 208B0001 thru 208B0597 that include CAB00-8 and Airplanes 20800261 thru 20800315 and Airplanes 208B0598 thru 208B0799 that include CAB00-9, and airplanes with a partial TKS or fairing TKS system installed, remove the secondary exhaust duct from the lower right cowling panel. (a) Remove the bolt, washers, and nut that attach the inboard hanger assembly to the ejector duct bracket (inboard). (b) Remove the bolt, washers, and nut that attach the outboard hanger assembly to the ejector duct bracket assembly (outboard). (c) Remove the bolt, washers, and nut that attach the center hanger assembly to the ejector duct bracket assembly (outboard). (d) Remove the bolts, washers, and nuts that attach the ejector duct bracket assembly (outboard) and ejector duct bracket (inboard) to the secondary exhaust duct. (e) Move the secondary exhaust duct forward until it is away from the lower right cowling panel.
Install the Secondary Exhaust Duct (Refer to Figure 202). (1) On Airplanes 20800006 thru 20800105 that do not include SK208-23, attach the secondary exhaust duct bellmouth to the lower right cowling panel. (a) Move the secondary exhaust duct aft through the lower right cowling panel. (b) Align the hanger brackets with the aft hangers. (c) Move the secondary exhaust duct aft until the hanger brackets are on the aft hangers. (d) Install the nuts, washers, and bolts that attach the secondary exhaust duct bellmouth to the lower right cowling panel. (2) On Airplanes 20800106 thru 20800199 and Airplanes 208B0001 thru 208B0249 and Airplanes 20800001 thru 20800105 that include SK208-23 and do not include CAB90-27, attach the secondary exhaust duct bellmouth to the lower right cowling panel. (a) Move the secondary exhaust duct aft through the lower right cowling panel. (b) Align the hanger brackets with the mid support hangers and the aft hangers. (c) Move the secondary exhaust duct aft until the hanger brackets are on the mid support hangers and the aft hangers. (d) Install the nuts, washers, and bolts that attach the secondary exhaust duct bellmouth to the lower right cowling panel. (3) On Airplanes 20800316 and On and Airplanes 208B0800 and On and Airplanes 20800001 thru 20800260 and Airplanes 208B0001 thru 208B0597 that include CAB00-8 and Airplanes 20800261 thru 20800315 and Airplanes 208B0598 thru 208B0799 that include CAB00-9, and airplanes with a partial or fairing TKS system installed, attach the secondary exhaust duct bellmouth to the lower right cowling panel. (a) Move the secondary exhaust duct aft through the lower right cowling panel. (b) Install the nuts, washers, and bolts that attach the ejector duct bracket assembly (outboard) and ejector duct bracket (inboard) to the secondary exhaust duct.
CAUTION: Make sure that the hanger assemblies are installed in the correct order. This will help prevent damage to the equipment. (c)
(d) (e)
(f)
Install the nut, washers, and bolt that attach the center hanger assembly to the ejector duct bracket assembly (outboard). 1 Put the thin washer adjacent to the bolt head and the thick washer between the rod end and the bracket. Install the nuts (if applicable), washers, and bolts that attach the bellmouth of the secondary exhaust duct to the lower right cowling panel. Install the nut, washers, and bolt that attach the outboard hanger assembly to the ejector duct bracket assembly (outboard). 1 Put the thin washer adjacent to the bolt head and the thick washer between the rod end and the bracket. Install the nut, washers, and bolt that attach the inboard hanger assembly to the ejector duct bracket (inboard). 1 Put the thin washer adjacent to the bolt head and the thick washer between the rod end and the bracket.
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MODEL 208 MAINTENANCE MANUAL (g)
Make sure that the secondary exhaust duct is correctly adjusted to the lower right cowling panel. 1 Make sure there is a clearance between the secondary exhaust duct and the skin and bulge area of the lower right cowling panel. a If necessary, adjust the hanger assemblies. 2 Make sure that the rod ends of the hanger assemblies do not rub or touch the brackets. a If necessary, add a washer between the rod end and the bracket. b If necessary, adjust the hanger assemblies. 3 Make sure that you can move the hanger assemblies on the ball joints with your hand. NOTE:
(h)
This will make sure that the hanger assemblies are not loaded before the airplane is operated.
If necessary, adjust the hanger assemblies. a If it is necessary to adjust the hanger assemblies, do the steps that follow. NOTE:
The hanger assemblies come assembled at the correct initial lengths. The initial length for the center hanger is 1.86 inches, +0.11 or -0.11 inch. The initial length for the inboard hanger is 7.00 inches, +0.25 or -0.25 inch. The initial length for the outboard hanger is 3.11 inches, +0.25 or -0.25 inch.
CAUTION: Make sure that the center hanger assembly is adjusted first. Then adjust the inboard and outboard hanger assemblies. This will help prevent damage to the equipment. 1
2
3
If necessary, adjust the center hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly. If necessary, adjust the inboard hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly. If necessary, adjust the outboard hanger assemblies. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly.
CAUTION: After the hanger assembly adjustments, make sure that a wire will not go through the rod end depth inspection holes of the hanger assemblies. This will make sure that the rod end has a sufficient number of threads in the hanger assembly. 4
(4) (5) (6)
Make sure that a wire will not go through the rod end depth inspection hole in the hanger assembly. a If a wire will go through the rod end depth inspection hole, adjust the hanger assemblies again. Install the three nuts, washers, and bolts that attach the oil breather hose attach neck to the secondary exhaust duct. Install the lower right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Close the upper right cowling door.
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MODEL 208 MAINTENANCE MANUAL (7)
On Airplanes 20800316 and On and Airplanes 208B0800 and On and Airplanes 20800001 thru 20800260 and Airplanes 208B0001 thru 208B0597 that include CAB00-8 and Airplanes 20800261 thru 20800315 and Airplanes 208B0598 thru 208B0799 that include CAB00-9, and airplanes with a partial TKS or fairing TKS system installed, make sure that the secondary exhaust duct is correctly adjusted to the lower right cowling panel. (a) Make sure there is a clearance between the secondary exhaust duct and the skin and bulge area of the lower right cowling panel. 1 If necessary, adjust the hanger assemblies. (b) Make sure that the rod ends of the hanger assemblies do not rub or touch the brackets. 1 If necessary, add a washer between the rod end and the bracket. 2 If necessary, adjust the hanger assemblies. (c) Make sure that you can move the hanger assemblies on the ball joints with your hand. NOTE:
(d)
This will make sure that the hanger assemblies are not loaded before the airplane is operated.
1 If necessary, adjust the hanger assemblies. If it is necessary to adjust the hanger assemblies, do the steps that follow.
CAUTION: Make sure that the center hanger assembly is adjusted first. Then adjust the inboard and outboard hanger assemblies. This will help prevent damage to the equipment. 1
2
3
If necessary, adjust the center hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly. If necessary, adjust the inboard hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly. If necessary, adjust the outboard hanger assemblies. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly.
CAUTION: After the hanger assembly adjustments, make sure that a wire will not go through the rod end depth inspection holes of the hanger assemblies. This will make sure that the rod end has a sufficient number of threads in the hanger assembly. 4
6.
Make sure that a wire will not go through the rod end depth inspection hole in the hanger assembly. a If a wire will go through the rod end depth inspection hole, adjust the hanger assemblies again.
Primary Exhaust Duct Inspection A.
Inspection of Primary Exhaust Duct. NOTE: (1)
Since exhaust systems are subject to high thermal stresses and vibration, inspection is very important for early detection of damaged components.
Remove primary exhaust duct.
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WARNING: Never use highly flammable solvents on engine exhaust system. Never use wire brushes to clean exhaust system components or mark on them with lead pencils. (2) (3)
7.
Check areas adjacent to welds and seams for evidence of cracks. Discoloration or deposits are evidence of leaks. Weld discontinuities or crack immediately adjacent to welds or seams may be rewelded if they do not exceed four inches in length. Cracks in areas not adjacent to welds or seams are not repairable and part is to be replaced.
Secondary Exhaust Duct Inspection A.
Inspection of Secondary Exhaust Duct. (1) Remove secondary exhaust duct.
WARNING: Never use highly flammable solvents on engine exhaust system. Never use wire brushes to clean exhaust system components or mark on them with lead pencils. (2) (3)
8.
Check areas adjacent to welds and seams for evidence of cracks. Discoloration or deposits are evidence of leaks. Weld discontinuities or cracks immediately adjacent to welds, seams, or areas not adjacent to any welds or seams may be welded or rewelded if they do not exceed four inches in length. Cracks longer than four inches in length are not repairable and part is to be replaced.
Field Installation of New Secondary Exhaust Duct A.
Installation of the Secondary Exhaust Duct (Refer to Figure 202). (1) Move the new secondary exhaust duct aft through the lower right cowling panel. (2) Make sure that the bellmouth area of the secondary exhaust duct is flush with the forward edge of the lower right cowling panel skin. (3) Align the oil breather hose attach neck holes. (4) Attach the top part of the secondary exhaust duct bellmouth to the top part of the cowling shroud with clamps. NOTE: (5) (6)
The clamps are used to keep the secondary exhaust duct aligned with the lower right cowling panel.
Install the nuts, washers, and bolts that attach the ejector duct bracket assembly (outboard) and ejector duct bracket (inboard) to the secondary exhaust duct. Adjust the hanger assemblies. NOTE:
(a)
(b)
(c)
The hanger assemblies come assembled at the correct initial lengths. The initial length for the center hanger is 1.86 inches, +0.11 or -0.11 inch. The initial length for the inboard hanger is 7.00 inches, +0.25 or -0.25 inch. The initial length for the outboard hanger is 3.11 inches, +0.25 or -0.25 inch.
If necessary, adjust the center hanger assembly to 1.86 inches, +0.11 or -0.11 inch. 1 Loosen the jam nuts on the hanger assembly. 2 Adjust the rod ends of the hanger assembly. 3 Tighten the jam nuts on the hanger assembly. If necessary, adjust the inboard hanger assembly to 7.00 inches, +0.25 or -0.25 inch. 1 Loosen the jam nuts on the hanger assembly. 2 Adjust the rod ends of the hanger assembly. 3 Tighten the jam nuts on the hanger assembly. If necessary, adjust the outboard hanger assemblies to 3.11 inches, +0.25 or -0.25 inch. 1 Loosen the jam nuts on the hanger assembly. 2 Adjust the rod ends of the hanger assembly.
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(7)
3 Tighten the jam nuts on the hanger assembly. Install the hanger assemblies.
CAUTION: Make sure that the hanger assemblies are installed in the correct order. This will help prevent damage to the equipment. Install the nut, washers, and bolt that attach the center hanger assembly to the ejector duct bracket assembly (outboard). 1 Put the thin washer adjacent to the bolt head and the thick washer between the rod end and the bracket. (b) Install the nut, washers, and bolt that attach the outboard hanger assembly to the ejector duct bracket assembly (outboard). 1 Put the thin washer adjacent to the bolt head and the thick washer between the rod end and the bracket. (c) Install nut, washers, and bolt that attach the inboard hanger assembly to the ejector duct bracket (inboard). 1 Put the thin washer adjacent to the bolt head and the thick washer between the rod end and the bracket. Temporarily install the lower right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Close the upper right cowling door. (a)
(8) (9)
CAUTION: Make sure that there is a clearance and correct fit between the secondary exhaust duct and the skin and bulge area of the lower right cowling panel. This will help prevent damage to the equipment. (10) Make sure that the secondary exhaust duct is correctly adjusted to the lower right cowling. (a) Make sure there is a clearance between the secondary exhaust duct and the skin and bulge area of the lower right cowling panel. 1 If necessary, adjust the hanger assemblies. (b) Make sure that the rod ends of the hanger assemblies do not rub or touch the brackets. If necessary, add a washer between the rod end and the bracket. 1 2 If necessary, adjust the hanger assemblies. (c) Make sure that you can move the hanger assemblies on the ball joints with your hand. NOTE:
(d)
This will make sure that the hanger assemblies are not loaded before the airplane is operated.
If necessary, adjust the hanger assemblies. 1 If it is necessary to adjust the hanger assemblies, do the steps that follow.
CAUTION: Make sure that the center hanger assembly is adjusted first. Then adjust the inboard and outboard hanger assemblies. This will help prevent damage to the equipment. 1
2
3
If necessary, adjust the center hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly. If necessary, adjust the inboard hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly. If necessary, adjust the outboard hanger assemblies. a Loosen the jam nuts on the hanger assembly.
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MODEL 208 MAINTENANCE MANUAL b c
Adjust the rod ends of the hanger assembly. Tighten the jam nuts on the hanger assembly.
CAUTION: After the hanger assembly adjustments, make sure that a wire will not go through the rod end depth inspection holes of the hanger assemblies. This will make sure that the rod end has a sufficient number of threads in the hanger assembly. Make sure that a wire will not go through the rod end depth inspection hole in the hanger assembly. a If a wire will go through the rod end depth inspection hole, adjust the hanger assemblies again. (e) If you can not get a clearance between the secondary exhaust duct and the skin and bulge area of the lower right cowling panel, refer to SK208-141 for modification procedures. Remove the lower right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. Match drill the holes at the top of the cowling shroud to the bellmouth of the secondary exhaust duct. Install the nuts, washers, and bolts that attach the top part of the secondary exhaust duct bellmouth to the cowling shroud. Remove the clamps from the top part of the secondary exhaust duct bellmouth. Make sure that the secondary exhaust duct is aligned with the lower right cowling panel. Match drill the holes at the bottom of the cowling shroud to the bellmouth of the secondary exhaust duct. Install the nuts, washers, and bolts that attach the bottom part of the secondary exhaust duct bellmouth to the cowling shroud. Install the three nuts, washers, and bolts that attach the oil breather hose attach neck to the secondary exhaust duct. Install the lower right cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. Close the upper right cowling door. Make sure that the secondary exhaust duct is correctly adjusted to the lower right cowling panel. (a) Make sure there is a clearance between the secondary exhaust duct and the skin and bulge area of the lower right cowling panel. 1 If necessary, adjust the hanger assemblies. (b) Make sure that the rod ends of the hanger assemblies do not rub or touch the brackets. If necessary, add a washer between the rod end and the bracket. 1 2 If necessary, adjust the hanger assemblies. (c) Make sure that you can move the hanger assemblies on the ball joints with your hand. 4
(11) (12) (13) (14) (15) (16) (17) (18) (19) (20) (21)
NOTE:
(d)
This will make sure that the hanger assemblies are not loaded before the airplane is operated.
1 If necessary, adjust the hanger assemblies. If it is necessary to adjust the hanger assemblies, do the steps that follow.
CAUTION: Make sure that the center hanger assembly is adjusted first. Then adjust the inboard and outboard hanger assemblies. This will help prevent damage to the equipment. 1
If necessary, adjust the center hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly.
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3
If necessary, adjust the inboard hanger assembly. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly. If necessary, adjust the outboard hanger assemblies. a Loosen the jam nuts on the hanger assembly. b Adjust the rod ends of the hanger assembly. c Tighten the jam nuts on the hanger assembly.
CAUTION: After the hanger assembly adjustments, make sure that a wire will not go through the rod end depth inspection holes of the hanger assemblies. This will make sure that the rod end has a sufficient number of threads in the hanger assembly. 4
Make sure that a wire will not go through the rod end depth inspection hole in the hanger assembly. a If a wire will go through the rod end depth inspection hole, adjust the hanger assemblies again.
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MODEL 208 MAINTENANCE MANUAL PRIMARY AND SECONDARY EXHAUST DUCT - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the primary and secondary exhaust ducts in a serviceable condition.
Task 78-10-00-210 2.
Primary and Secondary Exhaust Duct General Visual Inspection (Alignment Check) A.
General (1) This task gives the procedures to do a general visual inspection of the primary and secondary exhaust duct alignment.
B.
Special Tools (1) None
C.
Access (1) Open the right upper engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices.
D.
Do a Primary and Secondary Exhaust Duct General Visual Inspection (Alignment Check). Refer to Figure 601. NOTE: (1)
(2)
The following procedure is designed to do a check of the clearance between the primary and the secondary exhaust duct.
Examine the gap between the primary and the secondary exhaust ducts. (a) If necessary, remove the oil cooler shroud to get access to see the clearance. (b) Make sure there is a visible clearance between the primary and the secondary exhaust ducts. If there is interference between the primary and the secondary exhaust ducts, adjust the position of the secondary duct as follows:
CAUTION: Before you do a lower cowl hanger length adjustment, make sure that the cowl is supported to prevent cowl damage. (a) (b) (c) (d) (e)
Disconnect the lower cowl hanger from the lower right cowl. Loosen the jam nut on the lower cowl hanger and turn the clevis end of the hanger up or down, as necessary, to raise or lower the secondary duct to the desired position. Tighten the jam nut on the hanger. Attach the lower cowl hanger to the lower right cowl. Examine the clearance again and adjust if necessary.
CAUTION: Do a check for interference between the nose gear fairing and the nose gear trunnion after you adjust the lower right cowl. (f)
After you get the correct clearance, install the oil cooler shroud if it was removed.
E.
Restore Access (1) Close the right upper engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap Maintenance Practices. End of task
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Exhaust Duct Alignment Inspection Figure 601 (Sheet 1)
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Exhaust Duct Alignment Inspection Figure 601 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL Task 78-10-00-211 3.
Primary and Secondary Exhaust Duct General Visual Inspection A.
General (1) This task gives the procedures to do a general visual inspection of the primary and secondary exhaust duct.
B.
Special Tools (1) None
C.
Access (1) Remove the engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a Primary and Secondary Exhaust Duct General Visual Inspection. (1) Examine the primary exhaust duct and attach bolts for condition, cracks and corrosion. (2) Examine the secondary exhaust duct, and attach screws. (3) Examine the exhaust deflector and retainer screws. (4) Examine the inboard, center and outboard secondary exhaust hanger brackets. NOTE:
Remove the right side lower engine cowl to get access to secondary exhaust.
E.
Restore Access (1) Install the engine cowling. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task
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CHAPTER
OIL
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
79-00-00
Page 1
Apr 1/1996
79-20-00
Pages 201-204
Mar 1/2000
79-30-00
Pages 101-104
Apr 1/1996
79-30-00
Pages 201-204
Sep 2/2002
79-30-00
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Apr 1/2010
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Page 201
Mar 1/2012
79-Title 79-List of Effective Pages 79-Record of Temporary Revisions 79-Table of Contents
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
Page Number
Issue Date
By
Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS OIL - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DeÞnition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
79-00-00 79-00-00 79-00-00 79-00-00
Page 1 Page 1 Page 1 Page 1
OIL DISTRIBUTION - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Cooler Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Large Oil Cooler Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vernatherm Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vernatherm Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
79-20-00 Page 201 79-20-00 Page 201 79-20-00 Page 201 79-20-00 Page 201 79-20-00 Page 201 79-20-00 Page 204 79-20-00 Page 204
OIL INDICATING - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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OIL INDICATING - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Filler Cap and Dipstick . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure/Oil Temperature Indicator Removal/Installation . . . . . . . . . . . . . . . . . . . . Oil Temperature Bulb Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Switch Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
79-30-00 Page 201 79-30-00 Page 201 79-30-00 Page 201 79-30-00 Page 201 79-30-00 Page 201 79-30-00 Page 204 79-30-00 Page 204
OIL INDICATING - ADJUSTMENT/TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Temperature Indicating System Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
79-30-00 Page 501 79-30-00 Page 501 79-30-00 Page 501
CHIP DETECTORS - MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
79-31-00 Page 201 79-31-00 Page 201
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MODEL 208 MAINTENANCE MANUAL OIL - GENERAL 1.
Scope A.
2.
This chapter describes those units and components, external to the engine, concerned with storing and delivering lubricating oil to and from the engine.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Sealant
RTV-106 (Red)
General Electric Company Silicone Products Dept. Mechanicville Rd. Waterford, NY 12188
To aid in oil cooler related sealing.
3.
DeÞnition A.
This chapter is divided into sections to aid maintenance personnel in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief deÞnition of the sections incorporated in this chapter is as follows: (1) The section on distribution provides information on those components and lines used to distribute oil from external sources. (2) The section on indicating provides information on those components and systems used to indicate various parameters of the oil system.
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MODEL 208 MAINTENANCE MANUAL OIL DISTRIBUTION - MAINTENANCE PRACTICES 1.
General A.
This section provides removal/installation instructions for the airframe mounted oil cooler and its associated lines and components.
B.
For oil system servicing, including checking engine oil level, draining/changing engine oil, and oil Þlter removal/installation, refer to Chapter 12, Engine Oil System - Servicing.
WARNING: The U.S.. Environmental Protection Agency advises that mechanics and other workers who handle engine oil are advised to minimize skin contact with used oil and promptly remove used oil from the skin. In a laboratory study, mice developed skin cancer after skin was exposed to used engine oil twice a week without being washed off, for most of their life span. Substances found to cause cancer in laboratory mice may also cause cancer in humans. 2.
Description and Operation A.
3.
A remote oil cooler is mounted in right nose cap. Heated oil scavenged from accessory gearbox and reduction gearbox is routed to oil cooler via a hose under engine. A thermal bypass valve opens to circulate oil through cooler if oil temperature is above 150°F to 179°F. Otherwise, oil is routed past collar and returns via a hose, located on top of engine, to engine oil tank.
Oil Cooler Removal/Installation A. Remove Oil Cooler (Refer to Figure 201). (1) Remove right nose cap. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (2) Cut safety-wire from drain plug (12) and remove plug to drain oil from oil cooler (13). Reinstall plug after oil has drained. (3) Disconnect outlet hose (6) and inlet hose (8) from oil cooler (13). Cap hoses, cooler outlet elbow (7), and cooler inlet union (11) to prevent entry of foreign materials into oil system. (4) Remove bolts (4) (ten each) attaching cooler to bafße (1). (5) Remove bolts (14) (four each) attaching cooler to bracket (3) and remove cooler from airplane. NOTE: B.
4.
Check seals (9) and (9A) for condition, and, if damaged, replace. Corner seals (9A) are to be cemented to oil cooler (13) using RTV-106 or equivalent.
Install Oil Cooler (Refer to Figure 201). (1) Align holes in oil cooler attach bolt bosses with holes in bracket (3) and install four bolts (14). (2) Install ten bolts (4) attaching cooler to bafße (1). (3) Remove caps from hoses and cooler and connect inlet hose (8) to union (11).Connect outlet hose (6) to elbow (7) of cooler. (4) Tighten drain plug (12) from 215 to 240 inch pounds and safety. (5) Install right nose cap half. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (6) Service oil system. Refer to Chapter 12, Engine Oil System - Servicing.
Large Oil Cooler Removal/Installation A.
Remove Large Oil Cooler (Refer to Figure 201). (1) Remove right nose cap. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (2) Cut safety-wire from drain plug (16) and remove plug to drain oil from oil cooler (17). Reinstall plug.
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Oil Cooler Installation Figure 201 (Sheet 1)
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Oil Cooler Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL (3) (4) (5)
Disconnect outlet hose (6) and inlet hose (11) from oil cooler (17). Cap hoses, cooler outlet elbow (7), and cooler inlet elbow (12) to prevent entry of foreign materials into oil system. Remove ten bolts (9) attaching cooler to bafße (2). Remove bolts (3) (four each) attaching cooler to bracket (4) and remove cooler from airplane. NOTE:
B.
5.
Install Large Oil Cooler (Refer to Figure 201). (1) Align holes in oil cooler attach bolt bosses with holes in bracket (4) and install four bolts (3). (2) Install ten bolts (9) attaching cooler to bafße (2). (3) Remove caps from hoses and cooler and connect inlet hose (11) to elbow (12). Connect outlet hose (6) to elbow (7) of cooler. (4) Tighten drain plug (16) snug and secure using safety wire. (5) Install right nose cap half. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices (6) Service oil system. Refer to Chapter 12, Engine Oil System - Servicing.
Vernatherm Removal/Installation NOTE:
6.
Check seals (10) and (15) for condition, and, if damaged, replace. Corner seals (15) are to be cemented to oil cooler (17) using RTV-106 or equivalent.
Removal, installation and test of the vernatherm is typical on both sizes of oil coolers.
A.
Remove Vernatherm (Refer to Figure 201). (1) Remove right nose cap half. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. (2) Identify vernatherm (15 or 18) and remove safety wire. (3) Remove vernatherm from oil cooler (13 or 17). (4) Dispose of old gasket (16 or 19). (5) Perform test on vernatherm. Refer to Vernatherm Function Test.
B.
Install Vernatherm (Refer to Figure 201). (1) Install vernatherm (15 or 18) with new gasket (16 or 19) in oil cooler. (2) Safety wire vernatherm to oil cooler (13 or 17).Refer to Chapter 20, Safetying - Maintenance Practices (3) Install right nose cap half. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
Vernatherm Functional Test A.
Test Vernatherm. (1) Check vernatherm (15 or 18) by submerging in water at start to close temperature of 150°F for Þve minutes. (2) Remove and measure length. (3) Submerge vernatherm in water at fully closed temperature of 178°F for Þve minutes. (4) Minimum travel shall be 0.100 inch. (5) If valve being tested does not function properly, replace with new valve.
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MODEL 208 MAINTENANCE MANUAL OIL INDICATING - TROUBLESHOOTING 1.
General A. A troubleshooting chart has been prepared to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
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Oil Indicating System Troubleshooting Chart Figure 101 (Sheet 1)
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Oil Indicating System Troubleshooting Chart Figure 101 (Sheet 2)
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Oil Indicating System Troubleshooting Chart Figure 101 (Sheet 3)
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MODEL 208 MAINTENANCE MANUAL OIL INDICATING - MAINTENANCE PRACTICES 1.
General A.
2.
3.
Description and Operation A.
The engine has separate systems to measure the quantity, temperature and pressure of the oil in the engine. The oil Þller cap and dipstick location is on the aft, top left side of the engine. The dipstick is used to measure the quantity of the oil in the engine oil tank. During engine operation one gage shows both the oil pressure and temperature. This gage is installed on the upper center portion of the instrument panel. In addition, a red annunciator panel light labeled LOW OIL PRESS, comes on when the engine oil pressure decreases below approximately 40 PSI, +2 or -2 PSI.
B.
An optional magnetic chip detector system monitors if ferrous particle contaminates are in the oil sumps of the propeller reduction gearbox and the accessory gearbox of the engine. The system has two chip detector switches, an amber annunciator panel light labeled CHIP DETECTOR, and the wiring necessary to connect the system. When installed, the chip detector switches replace the related gearbox sump drain plugs. The annunciator panel light comes on if the ferrous particles make an electrical connection across the magnetic electrodes of either chip detector switch.
Oil Filler Cap and Dipstick A.
4.
This section has description, operation and removal and installation procedures for systems and components used to measure and show the condition of the engine oil.
The oil Þller cap and dipstick location is on the aft, top left side of the engine. Open the left upper cowling door to access the oil Þller cap and dipstick. The engine oil tank capacity is 9.5 U.S. quarts with Þve quarts measured on the dipstick. For the servicing of the engine oil, refer to Chapter 12, Engine Oil System - Servicing.
Oil Pressure/Oil Temperature Indicator Removal/Installation A.
Remove the Oil Pressure/Oil Temperature Indicator (Refer to Figure 201). (1) Remove the glareshield access covers to get access to the rear top of the instrument panel. (2) Disengage the OIL TEMP circuit breaker. (3) Disconnect the oil pressure line from the nipple on the rear of the oil pressure and temperature indicator case. (4) Remove the screws that attach the clamp to the oil pressure and temperature indicator case. (5) Remove the oil pressure and temperature indicator from the instrument panel to get access to the electrical connector. (6) Remove the safety wire from the electrical connector. (7) Remove the electrical connector from the oil pressure and temperature indicator. (8) Remove oil pressure and temperature indicator from the instrument panel.
B.
Install the Oil Pressure/Oil Temperature Indicator (Refer to Figure 201). (1) Connect the electrical connector to the oil pressure and temperature indicator. (2) Install safety wire on the electrical connector. (3) Insert the oil pressure and temperature indicator through the hole in the instrument panel and through the clamp. (4) Install the clamp to the oil pressure and temperature indicator case. (5) Connect the pressure line to the nipple on the rear of the oil pressure and temperature indicator case; but do not completely tighten the nut. (6) Do an engine motoring procedure and then bleed the air from the pressure line. (7) Tighten the pressure line on the rear of the oil pressure and temperature indicator case. (8) Engage the OIL TEMP circuit breaker. (9) Make sure the oil temperature indicator system operates correctly. Refer to Oil Temperature Indicating System Test. (10) Start the engine. (11) Make sure there are no oil leaks. (12) Install the glareshield access covers.
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Oil Indicating System Figure 201 (Sheet 1)
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Oil Indicating System Figure 201 (Sheet 2)
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5.
6.
Oil Temperature Bulb Removal/Installation A.
Remove the Oil Temperature Bulb (Refer to Figure 201). (1) Disengage the OIL TEMP circuit breaker. (2) Open the upper right cowl door to get access to the oil temperature bulb. (3) Remove the safety wire from the oil temperature bulb electrical connector. (4) Remove the oil temperature bulb electrical connector. (5) Remove the safety wire from the oil temperature bulb. (6) Remove the oil temperature bulb from the adapter. (7) Discard the O-ring from the oil temperature bulb.
B.
Install the Oil Temperature Bulb (Refer to Figure 201). (1) Put a new O-ring on the oil temperature bulb. (2) Install the oil temperature bulb into the adapter. (3) Connect the oil temperature bulb electrical connector to the oil temperature bulb. (4) Install safety wire on the oil temperature bulb and the oil temperature bulb electrical connector. (5) Engage the OIL TEMP circuit breaker. (6) Start the engine. (7) Make sure there are no oil leaks. (8) Make sure the oil temperature indicator system operates correctly. Refer to Oil Temperature Indicating System Test.
Oil Pressure Switch Removal/Installation A.
Remove the Oil Pressure Switch (Refer to Figure 201). (1) Disengage the ANNUN PANEL circuit breaker. (2) Remove the glareshield access covers to get access at the top of the instrument panel. (3) On Airplanes 20800001 thru 20800080 disconnect the electrical wiring from the oil pressure switch. (4) On Airplanes 20800081 and On and 208B0001 and On disconnect the oil pressure switch electrical connector. (5) On Airplanes 20800001 thru 20800100 remove the oil pressure switch from the tee and then put a cap on the tee. (6) On Airplanes 20800101 and On and 208B0001 and On remove the oil pressure switch from the tee and then put a cap on the tee. (7) Discard the O-ring. (8) Remove the oil pressure switch from the airplane.
B.
Install the Oil Pressure Switch (Refer to Figure 201). (1) On Airplanes 20800001 thru 20800100 use a new O-ring and install the oil pressure switch in the tee. (2) On Airplanes 20800101 and On and 208B0001 and On use a new O-ring and install the oil pressure switch in the tee. (3) On Airplanes 20800001 thru 20800080 connect the electrical wiring to the oil pressure switch. (4) On Airplanes 20800081 and On and 208B0001 and On connect the oil pressure switch electrical connector. (5) Engage the ANNUN PANEL circuit breaker. (6) Start the engine. (7) Make sure the low oil pressure light goes off at approximately 40.0 PSI, +2 or -2 PSI. (8) Make sure there are no leaks around the oil pressure switch. (9) Install the glareshield access covers.
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MODEL 208 MAINTENANCE MANUAL OIL INDICATING - ADJUSTMENT/TEST 1.
General A.
2.
You must test the oil temperature indicating system when the system is in question or when an oil temperature indicator is replaced. The test that follows will make sure the oil temperature indicator will show the correct indication within the recommended tolerances.
Oil Temperature Indicating System Test A.
the Oil Temperature Indicating System (Refer to Figure 501). Remove all of the electrical power from the airplane. Open the upper right cowl door. Remove the safety wire from the electrical connector (J14). Disconnect the electrical connector from the oil temperature bulb. Fabricate a harness as follows: (a) Get two pieces of 20 gauge wire approximately nine feet long. (b) Connect a M32029/29-212 connector pin to one end of each wire. (c) Remove approximately 1/2 inch of insulation from the opposite (terminal) ends of the wires. (6) Connect the harness pins to pins A and B of the oil temperature bulb electrical connector. (7) At a convenient location, connect a decade resistor to the terminal ends of the two wires of the fabricated harness.
Test (1) (2) (3) (4) (5)
NOTE: (8) (9) (10) (11)
A 200 ohm potentiometer, capable of 400 milliamperes, may be substituted for the decade resistor.
Make sure the pins and wiring will not touch ground or interfere with the adjacent structure. Apply electrical power to the airplane. Monitor the oil temperature indication. Change the resistance of the decade resistor until the oil temperature indicator shows each of the temperatures on the table. NOTE:
Refer to table 501 on airplanes that have oil temperature indicators with part numbers 2606015-1, 2606015-2, and 2606015-3. Refer to table 502 on airplanes that have oil temperature indicators with part number 2606015-4.
(12) Make sure the decade resistor settings are within the minimum and maximum resistance on the table for each temperature indication. Table 501. Oil Indicator Temperature versus Gage Tolerance (-1, -2, -3) OIL INDICATOR TEMPERATURE INDICATION
GAGE TOLERANCE (-1, -2, -3) MINIMUM RESISTANCE (OHMS)
NOMINAL RESISTANCE (OHMS)
MAXIMUM RESISTANCE (OHMS)
-40°C
76.71
77.29
77.87
10°C
93.15
93.80
94.46
55°C
109.57
110.34
111.11
99°C
127.63
128.42
129.21
110°C
132.40
133.26
134.12
140°C
146.27
147.11
147.95
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Table 502. Oil Indicator Temperature versus Gage Tolerance (-4) OIL INDICATOR TEMPERATURE INDICATION
GAGE TOLERANCE (-4) MINIMUM RESISTANCE (OHMS)
NOMINAL RESISTANCE (OHMS)
MAXIMUM RESISTANCE (OHMS)
-40°C
76.48
77.39
78.30
10°C
92.80
93.80
94.80
55°C
109.18
110.34
111.50
99°C
127.63
128.42
129.21
104°C
129.79
130.61
131.43
115°C
134.91
135.77
136.63
(13) When the test is completed, remove the electrical power from the airplane. (14) Remove the decade resistor, or potentiometer, and harness from the oil temperature bulb electrical connector. (15) Connect the electrical connector (J14) to the oil temperature bulb. (16) Install safety wire on the electrical connector. Refer to Chapter 20, Safetying - Maintenance Practices. (17) Close the upper right cowl door.
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Oil Temperature Indicating System Test Figure 501 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL CHIP DETECTORS - MAINTENANCE PRACTICES 1.
General A.
For the engine chip detector procedures, use the Pratt and Whitney Engine Maintenance Manual. Refer to the Introduction, Supplier Publication List.
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CHAPTER
STARTING
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LIST OF EFFECTIVE PAGES CHAPTER-SECTION-SUBJECT
PAGE
DATE
80-00-00
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Sep 2/1997
80-10-00
Page 1
Aug 1/1995
80-10-00
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Aug 1/1995
80-10-00
Pages 201-205
Jun 1/2011
80-10-00
Pages 601-602
Jun 1/2011
80-10-01
Pages 201-206
Aug 1/1995
80-11-00
Pages 201-202
Aug 1/1995
80-11-01
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Oct 15/1999
80-Title 80-List of Effective Pages 80-Record of Temporary Revisions 80-Table of Contents 80-List of Tasks
80 - LIST OF EFFECTIVE PAGES © Cessna Aircraft Company
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RECORD OF TEMPORARY REVISIONS Temporary Revision Number
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Issue Date
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Date Removed
By
CESSNA AIRCRAFT COMPANY
MODEL 208 MAINTENANCE MANUAL
CONTENTS STARTING - GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scope. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tools, Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
80-00-00 80-00-00 80-00-00 80-00-00
STARTER/GENERATOR - DESCRIPTION AND OPERATION . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
80-10-00 Page 1 80-10-00 Page 1 80-10-00 Page 1
STARTER/GENERATOR - TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
80-10-00 Page 101 80-10-00 Page 101
STARTER/GENERATOR - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter/Generator Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter/Generator Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter/Generator Inspection and Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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STARTER/GENERATOR - INSPECTION/CHECK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter-Generator (Part Number 23081 Series only) Detailed Inspection. . . . . . . . . . Starter/Generator Restoration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter-Generator Splines Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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STARTER/GENERATOR BRUSHES - MAINTENANCE PRACTICES (Lear Siegler/ Lucas Aerospace Model 23081-023 Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Removal/Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Inspection/Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter/Generator Brush Seating and Run-ln . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Course Brush Pre-Seating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Final Brush Seating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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STARTER/GENERATOR COOLING AIR BLAST TUBE - MAINTENANCE PRACTICES General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Blast Tube Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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300 AMP STARTER/GENERATOR COOLING AIR DUCT - MAINTENANCE PRACTICES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cooling Air Duct Removal/Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
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LIST OF TASKS 80-10-00-220
Starter-Generator (Part Number 23081 Series only) Detailed Inspection
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80-10-00-960
Starter/Generator Restoration
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80-10-00-160
Starter-Generator Splines Cleaning
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MODEL 208 MAINTENANCE MANUAL STARTING - GENERAL 1.
Scope A.
2.
This chapter contains maintenance information on starter/generators, troubleshooting of starter/generators, brush replacement procedures for selected starter/generators and cooling air blast tube removal and installation.
Tools, Equipment and Materials NOTE:
Equivalent substitutes may be used for the following items:
NAME
NUMBER
MANUFACTURER
USE
Digital Voltmeter
Model 87
John Fluke Mfg. Co. 6920 Seaway Blvd. Everett, WA 98206
To measure currrent, voltage and resistance.
Davco Aero 28C Cart (GPU)
28C
Davco Industries P.O. Box 273 Paso Robles, CA 93446
To provide auxiliary power for ground starting.
Sponge Rubber Padded V Blocks
Locally fabricate
To support starter/ generator during brush run-in procedures.
Switch SPST (75 amps DC minimum capacity)
Commercially available
To provide on/off control for brush run-in procedures.
Rheostat (0 to 10 ohms, 15 amp minimum)
Commercially available
To provide variable control for brush run-in procedures.
Tachometer or Stroboscope (0 to 15,000 RPM)
Commercially available
To measure RPM.
Resistor (0.1 ohm, 250 watt)
Commercially available
Used for brush run-in.
3.
DeÞnition A.
This chapter is divided into sections to aid maintenance technicians in locating information. Consulting the Table of Contents will further assist in locating a particular subject. A brief deÞnition of the sections incorporated in this chapter is as follows: (1) The section on starter/generators provides description/operation, troubleshooting, removal/ installation and brush run-in for starter/generator brushes. (2) The section on cooling air blast tubes provides removal/installation procedures for the blast tube.
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MODEL 208 MAINTENANCE MANUAL STARTER/GENERATOR - DESCRIPTION AND OPERATION 1.
General A.
2.
Starting system consists of a starter/generator, a starter switch, and a starter annunciator light. Starter/ generator functions as a motor for engine starting and will motor the gas generator section until a speed of 46 percent Ng is reached, at which time start cycle will automatically be terminated by a speed sensing switch located in the starter/generator. Starter/generator is controlled by a three-position starter switch located on left sidewall switch and circuit breaker panel. The switch has OFF, START and MOTOR positions. The OFF position deenergizes ignition and starter circuits and is the normal position at all times except during engine start. START position of the switch motors starter/generator and rotates gas generator portion of engine for starting. Also, the START position supplies ignition to engine, provided ignition switch is in NORMAL position. When engine has started, starter switch must be manually placed in OFF position to de-energize ignition system. MOTOR position of switch motors the engine without having ignition circuit energized and is used for motoring engine when an engine start is not desired. This can be useful for clearing fuel from engine, performing compressor washes, etc. MOTOR position is spring-loaded back to OFF position. Also, an interlock between the MOTOR position of starter switch and ignition switch prevents starter from motoring unless ignition switch is in NORMAL position. This prevents unintentional motoring of engine with ignition on. Starter operation is indicated by an amber light on annunciator panel, labeled STARTER ENERGIZED. NOTE:
Refer to Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual for starting procedures and starter cycle limitations.
NOTE:
Ground power unit starts are required if battery was left connected on an airplane parked longer than ten days and especially in cold weather (temperatures 0°F or lower).
NOTE:
Ground power unit output should conform to the following speciÞcations: 28.0 VDC, capacity 800 amperes minimum, 1700 amperes maximum.
Description and Operation A.
Airplane may be equipped with one of the three following starter/generators: (1) Lucas Model 23081-023 Starter/Generator. (2) APC Starter/Generator. (3) Lucas Model 23081-023A Starter/Generator.
B.
Starter/generator housing encloses all working components and a terminal block is located on top of the housing to attach electrical leads. Starter/generator is mounted to engine accessory gearbox at the 12 o’clock position and drives engine in start mode through a splined shaft. A drive coupling and shear section is incorporated in the starter/generator between drive spline and armature to prevent damage to engine accessory gearbox should a failure occur; drive ratio is 0.2931:1. Starter/generator drive splines require lubrication from engine. Cooling for starter/generator is provided by an integral fan attached to aft side of armature. Supplemental cooling during ßight is provided by ram air from an intake in right nose cap through a blast tube and hose to a fan housing inlet at the rear of starter/ generator.
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MODEL 208 MAINTENANCE MANUAL STARTER/GENERATOR - TROUBLESHOOTING 1.
General A.
A troubleshooting chart has been developed to aid the maintenance technician in system troubleshooting. Refer to Figure 101.
B.
When troubleshooting starter system, it is important to Þrst check airplane’s electrical wiring and termination before replacing major components, such as starter/generator or generator control unit. Past experience has shown that when a problem does occur, it is usually associated with a loose electrical wire connection, open circuit or a misadjusted generator control unit. Since starting circuits and DC generating circuits work in conjunction with each other, both electrical systems must be considered a possible source of a malfunction between starter/generator and generator control unit. To assist in checking the airplane’s starter/generator and DC generating electrical circuits, an analyzer box is available. Complete procedures for troubleshooting and use of analyzer box are provided in Chapter 24, Electrical Power - Adjustment/Test.
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Starter/Generator Troubleshooting Chart Figure 101 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL STARTER/GENERATOR - MAINTENANCE PRACTICES 1.
General A.
2.
This section gives removal and installation information for all starter/generators used on the airplane. In addition, the cleaning and inspection information is given for the Lucas Model 23081-023 Starter/ Generator and the APC 200SGL119Q Starter/Generator. The Aircraft Parts Corporation 300-amp Starter/Generator is a remove-and-replace-only part.
Starter/Generator Removal and Installation A.
Remove the Starter/Generator (Refer to Figure 201). NOTE:
Two mechanics are required to properly remove or install the starter/generator.
CAUTION: Make sure the starter/generator drive shaft is aligned with and concentric to the armature shaft. Slight misalignment and/or binding of the starter/ generator drive can reduce the unit's service life. (1) (2) (3) (4) (5)
Raise the upper cowl doors to get to the starter/generator. Remove all external power from the airplane, make sure the battery switch is in the OFF position, and disconnect the battery from the airplane electrical system. Remove the cover from the terminal block. Put an identification tag on each of the electrical leads for later identification and remove the terminal nuts. Remove the speed sensor circuit connector. NOTE:
(6)
Removal of the A/C drive unit is necessary for access to the starter generator on airplanes before 208000505 and 208B002025 that have a 200 AMP starter generator option installed.
Loosen the clamp that holds the cooling air blast hose on the starter/generator and remove the hose. NOTE:
Two mechanics are required to properly remove or install the starter/generator. One mechanic is to hold the starter/generator in position to keep the mounting surfaces flush with the quick attach/detach (QAD) adapter pad. This keeps the starter/generator aligned while the other mechanic loosens and removes the V-band clamp.
CAUTION: Hold the starter/generator in place to prevent damage to the splined drive shaft before you do the following step. (7) (8)
Loosen the V-band that holds the starter/generator to the quick attach/detach QAD adapter. Carefully remove the starter/generator from the QAD adapter pad so that the starter/generator drive spline is not put into a bind. (9) Remove the QAD adapter as necessary. (a) Remove the nuts that hold the QAD adapter to the engine accessory gearbox and remove the adapter. (b) Discard the gasket. (10) Use a cloth that is damp with MIL-PRF-680 or an equivalent solvent to clean the starter-generator splines. (11) Use a 10X magnifying glass to examine for signs of electrical arcing damage (in the form of pitting). NOTE:
If there are signs of arching on the starter-generator drive splines, refer to Cessna SNL07-16 and P&WC S.I.L NO. Gen-PT6-024 for additional information and inspection requirements.
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Starter/Generator Installation Figure 201 (Sheet 1)
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Starter/Generator Installation Figure 201 (Sheet 2)
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MODEL 208 MAINTENANCE MANUAL B.
Install the Starter/Generator (Refer to Figure 201). (1) Do the following steps before you install the starter/generator. Make sure: (a) There are no burrs or foreign objects on the starter/generator shaft. (b) The starter/generator guide pins are clean and not bent or damaged. (c) The mounting surfaces of the starter/generator and the QAD adapter pad are clean and do not have any burrs. (d) The QAD adapter is fastened to the engine transfer case correctly. (e) The QAD adapter pad guide pin holes does not have any burrs or foreign objects, and that they are in good condition. (2) Install the QAD adapter onto engine accessory gearbox with a new gasket, and install nuts as necessary. (3) Install a new O-ring around the groove on the splined drive shaft. (4) With the T-bolt unlatched, put the V-band on the starter/generator between the mounting flange and the terminal block. NOTE:
Two mechanics are required to properly remove or install the starter/generator. One mechanic is to hold the starter/generator in position to keep the mounting surfaces flush with the quick attach/detach QAD adapter pad. This keeps the starter/generator aligned while the other mechanic installs and tightens the V-band clamp.
CAUTION: The spline drive shaft must stay aligned with and concentric to the armature. If the starter/generator is allowed to be installed with the drive shaft out of position, excessive vibration and damage may develop during operation. (5) (6)
Carefully look at the spline drive shaft and the armature shaft interface plates. If the drive shaft looks to be out of position, lightly tap on the spline drive shaft with a plastic mallet to move it to a full concentric position. Figure 201. Carefully engage the spline drive shaft with the engine spline.
CAUTION: Keep the starter/generator flush up against the adapter during installation. Do not let the unit hang loosely without the V-band being latched because to much load on the drive shaft shear section may cause damage. (7) (8) (9)
Make sure the dowel pins are engaged. Put the V-band over the mating flanges and latch. Tap the V-band at several places with a rubber mallet to make sure that there is correct alignment of the spline drive shaft and the armature shaft, and tighten the T-bolt nut to two-thirds the recommended torque. NOTE:
The correct torque value is stamped on the V-band.
(10) Tap the V-band repeatedly with the rubber mallet and tighten the T-bolt nut to the recommended torque. (11) Install the cooling air blast hose with the clamp on the starter/generator. (12) Tighten the cooling air blast hose clamp. (13) Connect the speed sensor cable connector to the starter/generator. (14) Install the electrical cables in the same relationship to the terminal posts as you tagged them during the removal procedure, and install the nuts. (15) Put the cover in place over the terminal block. (16) Reconnect the battery to the airplane electrical system. (17) Close the upper cowl doors and do an engine motoring run to make sure that the normal starter speed is attained.
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3.
Starter/Generator Cleaning NOTE: A.
4.
This cleaning procedure applies only to the Lucas Model 23081-023 Starter/Generator. The Aircraft Parts Corporation 300-amp Starter/Generator is a remove-and-replace-only part.
Cleaning Procedures. (1) Remove starter/generator. (2) Remove cooling fan and brush cover and using low pressure air supply (40 PSI), blow as much carbon and copper dust as possible out of armature and stator windings, starting first at drive end of unit. (3) Continue blowing out dust in brush holder and commutator area, using care not to direct air jet at bearing in order to avoid forcing dirt into bearing. (4) Using a clean cloth moistened with stoddard solvent, thoroughly clean exterior of unit. Wipe with clean cloth. (5) Check for oil and/or other deposits that may have infiltrated into unit and remove as much of this residue as possible with solvent-moistened cloth.
Starter/Generator Inspection and Checks NOTE:
A.
This inspections and checks procedure applies only to Lucas Model 23081-023 Starter/ Generator. The Aircraft Parts Corporation 300-amp Starter/Generator is a remove-and-replaceonly part.
Inspections and Checks. (1) Remove starter/generator. (2) Visually check commutator for worn, pitted or burned commutator bars. NOTE: (3)
Check damper assembly for looseness and excessive wear. Check for loose rivets. NOTE:
(4)
(7)
Any visible damage to drive shaft or splines requires that starter/generator be returned for overhaul.
Check starter/generator housing for cracks and loose stator poles. Tighten stator poles as necessary and stake screws. NOTE:
(6)
If the damper assembly is excessively worn, reveals looseness or has loose rivets, the starter/generator must be returned for overhaul.
Check drive shaft for worn, chipped, cracked, and/or twisted splines. NOTE:
(5)
If the commutator bars are damaged, starter/generator must be returned for overhaul.
If housing is cracked or damaged, return starter/generator for overhaul.
With brushes removed, rotate armature slowly to feel bearings for roughness. NOTE:
This step does not apply to the 23081-023A, which is not field serviceable.
NOTE:
If armature movement is rough, starter/generator must be returned for overhaul.
Inspect fan for bent or distorted blades, cracks, or looseness.
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MODEL 208 MAINTENANCE MANUAL STARTER/GENERATOR - INSPECTION/CHECK 1.
General A.
This section has the inspections and checks necessary to keep the starter/generator in a serviceable condition.
Task 80-10-00-220 2.
Starter-Generator (Part Number 23081 Series only) Detailed Inspection A.
General (1) This task gives the procedures to do a detailed inspection of the starter-generator (Part Number 23081 Series only).
WARNING: Parts and brushes are not interchangeable between the 23081023 and 23081-023A. NOTE:
The model 23081-023 starter/generator brushes have a diagonal wear mark that shows when the brush replacement is necessary.
B.
Special Tools (1) None
C.
Access (1) Remove the left and right upper cowling doors. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do a Detailed Inspection of the Starter-Generator (Part Number 23081 Series Only). (1) Examine the starter-generator for condition, security of installation, and signs of overheating. (2) Examine the terminal block and the boot for condition, cracks, and security. (3) Examine the electrical connections at terminal block for cleanliness, signs of heat or arcing, and signs of damage. (4) Examine the quick attach/detach mount for condition, cracks, corrosion, and security of installation. (5) Examine the mount clamp for condition, cracks, and security. (6) Remove the starter/generator from the airplane. Refer to Starter/Generator - Maintenance Practices. (7) Examine the starter-generator brushes and brush holders. Refer to Goodrich Component Manual With Illustrated Parts List, DC Starter-Generator 23081 Series I. (8) Clean the starter-generator splines. Refer to Task 80-10-00-160. (9) Install the starter/generator. Refer to Starter/Generator - Maintenance Practices.
E.
Restore Access (1) Install the left and right upper cowling doors. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task Task 80-10-00-960 3.
Starter/Generator Restoration A.
General (1) This task gives the procedures to do the starter/generator restoration.
B.
Special Tools (1) None
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MODEL 208 MAINTENANCE MANUAL C.
Access (1) Remove the left and right upper cowling doors. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Do the Starter/Generator Restoration (Refer to Figure 201). (1) Remove the starter/generator from the airplane. Refer to Starter/Generator - Maintenance Practices. (2) Send the starter/generator to an authorized repair facility for restoration (overhaul). (3) Install the starter/generator. Refer to Starter/Generator - Maintenance Practices.
E.
Restore Access (1) Install the left and right upper cowling doors. Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices. End of task Task 80-10-00-160 4.
Starter-Generator Splines Cleaning A.
General (1) This task provides the procedures to perform a cleaning/inspection of the starter-generator splines.
B.
Tools and Equipement (1) 10X Magnifying Glass (2) Solvent (MIL-PRF-680 or equivalent).
C.
Remove the left and right upper cowling doors. (1) Refer to Chapter 71, Engine Cowling and Nose Cap - Maintenance Practices.
D.
Complete a Cleaning/Inspection of the Starter-Generator Splines. (1) Remove the starter-generator. Refer to Chapter 80, Starter/Generator - Maintenance Practices. (2) Clean the starter-generator splines by wiping the splines with a damp cloth. Use solvent (Federal Specification MIL-PRF-680 or equivalent) if needed. (3) Examine for evidence of electrical arching damage (in the form of pitting) using a 10X magnifying glass. NOTE:
(4)
If there is any evidence of arching on the starter-generator drive splines, refer to Cessna SNL07-16 and P&WC S.I.L. NO. Gen-PT6-024 for additional information and inspection requirements.
Install the starter-generator. Refer to Chapter 80, Starter/Generator - Maintenance Practices.
E.
Restore Access (1) Install left and right upper cowling doors. End of task
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MODEL 208 MAINTENANCE MANUAL STARTER/GENERATOR BRUSHES - MAINTENANCE PRACTICES (Lear Siegler/Lucas Aerospace Model 23081-023 Only) 1.
General A.
Following procedures apply to only the Lear Siegler/Lucas Aerospace Model 23081-023 starter/generator.
WARNING: The model 23081-023A starter/generator is not Þeld serviceable. Parts and brushes are not interchangeable between the 23081-023 and 23081-023A. NOTE: 2.
3.
Model 23081-023 starter/generator brushes have a diagonal wear mark which allows identiÞcation of when brushes require replacement.
Description and Operation A.
Brushes carry current to starter/generator armature from external circuit (in START mode). They are made of a mixture of graphite and metallic powder. Brushes must be free to slide in their holders, so that they may follow small irregularities in curvature of commutator. In addition, brushes must be able to feed in as they wear. Proper pressure of brushes against commutator is maintained by brush holder springs. A low-resistance connection between brushes and brush holders is maintained by means of braided copper pigtails attached.
B.
Brushes require periodic inspection and replacement when wear indicator disappears. Each brush carries a diagonal wear mark which allows identiÞcation of when a brush requires replacement. Refer to Figure 201 for details.
Brush Removal/Installation A.
Remove Starter/Generator Brushes (Refer to Figure 201 ). NOTE: (1) (2) (3)
B.
If brushes are to be reused, mark each brush to allow reinstallation in brush holder from which it was removed.
Remove starter/generator. Refer to Starter/Generator - Maintenance Practices. Loosen screws and remove cooling fan/brush cover. Remove screws securing brush pigtails to brush holders, and, using a stiff wire hook, lift springs from brushes and remove brushes.
Install Starter/Generator Brushes (Refer to Figure 201 ). NOTE:
Ensure pigtails are separated as shown in Figure 201. Starter/generator brushes must be seated after installation.
CAUTION: Brushes must be installed in correct position with regard to armature rotation as shown in Figure 201. (1) (2) (3) 4.
Lift brush springs and insert brushes in brush holders; tighten brush pigtail retaining screws. Install cooling fan/brush cover and tighten retaining screws. Reinstall starter/generator. Refer to Starter/Generator - Maintenance Practices.
Brush Inspection/Check A.
Inspection/Checks (Refer to Figure 201). (1) Remove starter/generator. Refer to Starter/Generator - Maintenance Practices. (2) Remove cooling fan/brush cover. (3) Observe brush wear marks and compare with information shown in Figure 201. (4) Check brush holders for looseness, damage from arcing and broken brush holder springs.
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Starter/Generator Brush Installation Figure 201 (Sheet 1)
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5.
Starter/Generator Brush Seating and Run-ln A.
Brush Seating and Run-In Procedures. (1) Full brush seating is comprised of both coarse brush pre-seating by sanding combined with Þnal brush run-in by running the unit on a drive stand or operating the unit as a motor.
CAUTION: If functional faults are evident before line maintenance, or material defects are apparent during inspection, do not install new brushes. Replace the original brush inspection cover, and submit the unit to an FAA Authorized Accessory Overhaul Facility for repair and/or overhaul. CAUTION: Failure to fully pre-seat and run-in brushes on high starter current applications (starting current equal to or exceeding 300 percent of the machine’s DC current rating) may cause excessive sparking, burning and pitting of the commutator. NOTE: 6.
For low starting current applications, perform coarse brush pre-seating procedures. Final brush run-in, although not required, increases commutator and brush life.
Course Brush Pre-Seating A.
Pre-Seating Procedures. (1) Cut a strip of 180 grit (or Þner) sandpaper, slightly wider than brush contact area and one inch longer than the circumference of commutator.
WARNING: Methyl alcohol is both toxic and ßammable. Use in a well ventilated area free from sparks, ßame and hot surfaces. In case of eye contact, ßush with water and seek medical attention. In case of skin contact, wash with soap and water. (2) (3) (4)
Clean any four adjacent commutator bars with methyl alcohol. Tape leading edge of sandpaper to clean commutator bars with masking tape, so that taped end of sandpaper is in normal direction of armature rotation, and abrasive side of sandpaper faces away from commutator. Rotate armature in its normal direction of rotation so that sandpaper strip wraps around commutator. Trailing edge of the paper should cover masking tape.
CAUTION: Rapid release of the brush springs can damage the brushes. (5) (6) 7.
Remove brush and sandpaper residue from the unit using compressed air. Remove Þngerprints from the commutator surface using methyl alcohol. Lift brush springs and slide the brushes down the brush holder to contact commutator. Position brush leads over the brush spring clips.
Final Brush Seating A.
Final Brush Seating or Run-in. NOTE: (1)
The following instructions are for Þnal brush run-in by both drive stand operation and motoring methods. Because of the time factor, the drive stand method is preferred.
Final brush run-in by drive stand operation. (a) Use a variable speed drive stand capable of driving the unit at speeds between 5,500 and 12,000 RPM at 50 percent of maximum rated electrical load.
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MODEL 208 MAINTENANCE MANUAL (b)
Operate the starter/generator as a generator from 25 to 50 percent of maximum rated electrical load and at 75 percent of rated normal top operational speed for the following time periods:
BRUSH TYPE
TIME
Particle Flake
15 to 30 minutes
Cored
2 to 4 hours. (2)
Final brush run-in by motoring. (a) Obtain a power source with a 24.0 to 28.0 VDC, 50 ampere minimum power source. The following chart lists acceptable power sources and capacities.
SOURCE
CAPACITY REQUIREMENTS
Airplane Battery
24.0 to 28.0 VDC (50 ampere hour minimum)
Inverter
28.0 VDC, 50 ampere hour minimum output
Power Supply
28.0 VDC, 50 ampere hour minimum output (b)
Mount the unit securely in a V-block holder.
CAUTION: The switch must be open when connecting the power supply to the rheostat. CAUTION: The rheostat must read 10 ohms, +1 or -1 ohm, when set at maximum resistance setting. An inÞnite (open) reading can cause unit destruction due to excessive speeds. (c) (d) (e)
Hook-up electrical connections to starter/generator. Refer to Figure 202. Set the rheostat to its minimum resistance setting. Close the switch to energize the circuit. NOTE:
(f)
(g) (h)
If the unit rotates opposite its normal direction, check for improper internal and external lead connections.
Adjust the rheostat to set speed between 50 and 75 percent of the maximum rated operational speed. NOTE:
Final brush run-in by motoring of cored brushes takes approximately one hour for each 0.001 inch of commutator groove depth.
NOTE:
A properly seated and fully run-in portion of a brush has a smooth, semigloss surface.
Operate the unit as a motor until the brushes are Þne seated a minimum of 75 percent in the axial plane and 100 percent in the rotational plane. Refer to Figure 203. Disconnect the starter/generator from the motoring equipment.
WARNING: When using compressed air, regulate pressure to 40 PSI or less. Wear goggles or face shield to protect eyes. Use a dust Þlter respirator. Do not aim the air jet at exposed skin. (i) (j)
Remove carbon dust from the unit using dry Þltered compressed air. Install starter/generator on airplane. Refer to Starter/Generator - Maintenance Practices.
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Final Brush Run-in Set Up By Motoring Figure 202 (Sheet 1)
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Examples of Seated and Fully Run-in Brushes Figure 203 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL STARTER/GENERATOR COOLING AIR BLAST TUBE - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
This section provides description, operation and removal/installation information on the starter/generator cooling air blast tube.
Starter/generator incorporates a cooling air blast tube and hose assembly to augment cooling in ßight. Tube is mounted via brackets to upper right side of induction air plenum assembly on right side of engine. Flexible hoses clamped to forward and aft ends of blast tube direct cooling air from an inlet in right nose cap through blast tube and into inlet in starter/generator cooling fan/brush cover.
Blast Tube Removal/Installation A.
Remove Blast Tube (Refer to Figure 201). (1) Open upper right cowling door to gain access to starter/generator cooling air blast tube. (2) Loosen hose clamps (9) from forward and aft ends of cooling air blast tube (8) and remove forward and aft hoses (3) and (10). (3) Remove nuts (4) from forward and aft clamp bolts (2) and (7). (4) Remove blast tube (8) from engine.
B.
Install Blast Tube (Refer to Figure 201). (1) Position clamps (5) on blast tube (8) and align clamps with brackets (1) and (6). Install bolts (2) and (7) and nuts (4). (2) Install forward and aft blast hoses (3) ad (10) on blast tube (8) and tighten hose clamps (9). (3) Close upper right cowling door.
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Cooling Air Blast Tube Installation Figure 201 (Sheet 1)
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MODEL 208 MAINTENANCE MANUAL 300 AMP STARTER/GENERATOR COOLING AIR DUCT - MAINTENANCE PRACTICES 1.
General A.
2.
Description and Operation A.
3.
This section describes the operation and provides removal/installation information for the 300 amp starter/generator cooling air duct.
The 300 amp starter/generator incorporates a cooling air duct assembly to provide in-ßight cooling. The duct is mounted to the upper right side induction air plenum assembly located right of the engine. The duct is supported by brackets and clamps directly to the plenum and starter/generator cooling fan/brush cover.
Cooling Air Duct Removal/Installation A.
Remove air duct. Refer to Figure 201. (1) Open upper right cowling door to gain access to starter/generator cooling air duct. (2) Loosen clamps attaching duct to starter/generator inlet and air plenum. (3) Remove nuts securing clamps to forward and aft brackets. (4) Remove cooling air duct from starter/generator inlet and air plenum. Remove duct from airplane.
B.
Install air duct. Refer to Figure 201. (1) Position cooling air duct between starter/generator inlet and air plenum. Secure with clamps and nuts to forward and aft brackets. (2) Tighten clamps at attachment locations of duct to starter/generator inlet and air plenum. (3) Close upper right cowling door.
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Cooling Air Duct Installation Figure 201 (Sheet 1)
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