•
REVISION
MODEL
337, T337, F337 and FT337
SUPER SKYMASTER
SERIES
•
1965 THRU 1973
SERVICE MANUAL
CHANGE 2
5 JANUARY 2004
D2500C2-13
•
INSERT THE FOLLOWING CHANGED
PAGES INTO THE BASIC MANUAL
~~
•
Cessna
A Textron Company
Service Manual
1965
Thru
1973
•
MODEL 337,
T337, F337 and FT337
SUPER SKYMASTER SERIES
f)
Member of GAMA
FAA APPROVAL HAS BEEN OBTAINED ON TECHNICAL DATA IN THIS PUBLICATION THAT AFFECTS AIRPLANE TYPE DESIGN.
CHANGE 2 TO THE BASIC MANUAL INCORPORATES TEMPORARY REVISION 5 DATED 1 APRIL 1992,
TEMPORARY REVISION 6 DATED 1 JUNE 1992, AND TEMPORARY REVISION 7 DATED 17 MARCH 1995.
•
1 FEBRUARY 1973
COPYRIGHT © 2004
CESSNA AIRCRAFT COMPANY
WICHITA, KANSAS, USA
02500-2-13
Change 2
5 JANUARY 2004
CESSNA AIRCRAFT COMPANY
MODEL 337
•
SERVICE MANUAL
LIST OF EFFECTIVE PAGES
Dates of issue for original and changed pages are:
Original ............. 0 ...... ........ 1 February
Change ............. 1 ............. 1 October
Change ............ 2 ............. 5 January
1973
1990
2004
I
Total number of pages in this publication is 834, consisting of the following:
Page
No.
•
•
Change
Page
Change
Page
Change
No.
No.
No.
No.
No.
Title .................................... 2*
A ........................................ 2*
i thru vi ............................... 2*
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1-6 (Blank) ........................ 0
2-1 thru2-17 ..................... 0
2-18 .................................. 1
2-19 thru 2-24 ................... 0
2-25 thru 2-26 .................... 2"
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2-29 thru 2-32 .................... 2"
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14-3 thru 14-22 .................. 0
14-23 thru 14-28 ................ 2"
14-28A thru 14-28B ........... 2"
14-29 thru 14-32 ................ 0
15-1 thru 15-37 .................. 0
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NOTE:
Insert latest changed pages. Destroy superseded pages.
NOTE:
" Indicates pages changed, added, or deleted by the current change.
Change 2
Jun 5/2004
© Cessna Aircraft Company
I
A
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
•
TABLE OF CONTENTS
SECTION
•
•
PAGE
GENERAL ..•............•...•... ......................•.•.•.•.... ......
1-1
2
GROUND HANDLING, SERVICING, LUBRICATION, AND INSPECTION........
2-1
3
FUSELAG E ••..•.••...........••.........•.......•••....................
3.. 1
4
WINGS, BOOMS, AND EMPENNAGE......................................
4-1
5
LANDING GEAR AND HYDRAULIC SySTEM................................
5-1
6
AILERON CONTROL SYSTEM ••••••••••••••••••••••••••••••••••••••••••••
6-1
7
WING FLAP CONTROL SySTEM..........................................
7-1
8
ELEVATOR, ELEVATOR TRIM AND FLAP/ELEVATOR TRIM
INTERCONNECT SYSTEMS. • • •••• • • • • •• • • • •••• • • • • • •••••• • • • •• • ••• •• •• ••
8-1
9
RUDDER AND RUDDER TRIM CONTROL SySTEM.........................
9-1
10
ENGINES (NON-TURBOCHARGED) •••••••••••••••••••••••••••••••••••••••
10-1
10A ENGINES (TURBOCHARGED) ••••••••••••••••••••••••••••••••••••••••••••
10A-1
11
FUEL SYSTEM ••••••••••••••••••••••••••••••••••••••••••••••••••••••••••
11-1
12
PROPELLERS AND PROPELLER GOVERNORS............................
12-1
13
UTILITY SYSTEMS. • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • •
13-1
14
INSTRUMENTS AND INSTRUMENT SySTEMS.............................
14-1
15
ELECTRICAL SYSTEMS (THRU 1970 MODELS). •••• •• • • • • •• • • • • • • ••• • • • • • ••
15-1
15A ELECTRICAL SYSTEMS (BEGINNING WITH 1971 MODELS) •••••••••••••••••
15A-1
16
STRUCTURAL REPAIR...................................................
16-1
17
EXTERIOR PAINTING....................................................
17-1
18
WIRING DIAGRAMS •••••••••••••••••••••••••••••••••••••••••••••••••••••
18-1
Change 2
Jan 5/2004
© Cessna Aircraft Company
I
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
CROSS-REFERENCE LISTING OF POPULAR
NAME VS. MODEL NUMBERS AND SERIALS
All aircraft, regardless of manufacturer, are certified under model number designations. However popular
names are used for marketing purposes. To provide a consistent method of referring to the various aircraft,
model numbers will be used in this publication unless names are required to differentiate between versions of
the same basic model. The following table provides a cross-reference listing of popular name vs. model
numbers.
MODEL
POPULAR NAME
YEAR
SUPER SKYMASTER
1965
1966
337
337A
1966
1966
337A
337A
1967
337B
1967
TURBO-SYSTEM
SUPER SKYMASTER
I
MODEL
BEGINNING
SERIAL
ENDING
SERIAL
NUMBER
NUMBER
337-0002
337-0240
337-0307
337-0471
337-0526
337-0239
337-0305
337-0570
337-0755
337-0001
337-0470
1968
337B
337B
337B
337C
337-0756
337-0978
1969
337D
1970
1971
337E
337F
337F
337-0979
33701194
33701317
337-1193
33701316
33701398
33700306
1967
1967
T337B
T337B
T337B
337-0526
337-0570
1968
1969
1970
1971
T337C
T337D
T337E
T337F
337-0756
337-0979
33701194
33701317
337-0568
337-0755
337-0001
337-0978
337-1193
T337F
I
F337E
F33700001
F33700024
F337F
F33700025
F33700045
REIMS/CESSNA TURBO-SYSTEM
1970
F33700024
1971
FT337E
FT337F
F33700001
SUPER SKYMASTER
F33700025
F33700045
SKYMASTER
1972
337F
337F
33701399
33701450
33701448
33701462
337G
33701463
33701550
1972
F337F
F33700046
1973
F337G
F33700056
F33700055
F33700063
1972
1973
ii
© Cessna Aircraft Company
•
33701316
33701398
33700569
1970
REIMS/CESSNA
SKYMASTER
I
337-0469
337-0525
337-0568
1971
REIMS/CESSNA SUPER SKYMASTER
•
Jan 5/2004
Change 2
•
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
•
INTRODUCTION
1.
General
WARNING: ALL INSPECTION INTERVALS, REPLACEMENT TIME LIMITS, OVERHAUL TIME LIMITS,
METHODS OF INSPECTION, LIFE LIMITS, CYCLE LIMITS, ETC., RECOMMENDED BY
CESSNA ARE SOLELY BASED ON THE USE OF NEW, REMANUFACTURED, OR
OVERHAULED CESSNA-APPROVED PARTS. IF PARTS ARE DESIGNED,
MANUFACTURED, REMANUFACTURED, OVERHAULED, ANDIOR APPROVED BY
ENTITIES OTHER THAN CESSNA, THEN THE DATA IN CESSNA'S
MAINTENANCE/SERVICE MANUALS AND PARTS CATALOGS ARE NO LONGER
APPLICABLE AND THE PURCHASER IS WARNED NOT TO RELY ON SUCH DATA FOR
NON-CESSNA PARTS. ALL INSPECTION INTERVALS, REPLACEMENT TIME LIMITS,
OVERHAUL TIME LIMITS, METHOD OF INSPECTION, LIFE LIMITS, CYCLE LIMITS, ETC.,
FOR SUCH NON-CESSNA PARTS MUST BE OBTAINED FROM THE MANUFACTURER
ANDIOR SELLER OF SUCH NON-CESSNA PARTS.
•
A. The information in this publication is based on data available at the time of publication and is updated,
supplemented, and automatically amended by all information issued in Service Letters, Service
Information Letters, Service Bulletins, Service Newsletters, Supplier Service Notices, Publication
Changes, Revisions, Reissues and Temporary Revisions. Users are urged to keep abreast of the latest
amendments to this publication through information available at Cessna Authorized Service Stations or
through the Cessna Propeller Aircraft Product Support subscription services. Cessna Service Stations
also have been supplied with a group of supplier publications which provide disassembly, overhaul, and
parts breakdowns for some of the various supplier equipment items. Supplier's publications are updated,
supplemented, and specifically amended by supplier-issued revisions and service information which may
be reissued by Cessna. Supplier publications reissued by Cessna automatically amend this publication
and are communicated to the field through Cessna's Authorized Service Stations and/or through Cessna's
subscription services.
B.
Inspection, maintenance, and parts requirements for STC installations are not included in this manual.
When an STC installation is incorporated on the airplane, those portions of the airplane affected by the
installation must be inspected in accordance with the inspection program published by the owner of the
STC. STC installations can change systems interface, operating characteristics and component loads or
stresses on adjacent structures of the airplane. Cessna provided inspection criteria may not be valid for
airplanes with STC installations.
C. REVISIONS, REISSUES and TEMPORARY REVISIONS can be purchased from a Cessna Service
Station or directly from Cessna Aircraft Company at the following address:
Cessna Aircraft Company
Department 751 C
P.O. Box 7706
Wichita, Kansas 67277-7706
•
D.
Information in this Service Manual is applicable to all U.S. and Foreign Certified 337 Model series
airplanes within the following range of serial numbers; 337-0001 thru 33701550 and F33700001 thru
F33700063. Information unique to a particular country is identified in the chapter(s) affected.
E.
All supplemental service information concerning this manual is supplied to all appropriate Cessna Service
Stations so that they have the latest authoritative recommendations for servicing these Cessna airplanes.
Therefore, it is recommended that Cessna owners utilize the knowledge and experience of the Cessna
Service Organization.
Change 2
Jan 5/2004
© Cessna Aircraft Company
iii
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
2.
Cross-Reference Listing of Popular Name Versus Model Numbers and Serials.
A.
All airplanes, regardless of manufacturer, are certified under model number designations. However,
popular names are often used for marketing purposes. To provide a consistent method of referring to
these airplanes, the model number will be used in this publication unless the popular name is necessary to
differentiate between versions of the same basic model. The table on page ii provides a listing of popular
names, model number, and serial numbers.
•
B. The Cessna Super Skymaster Series (1965 thru 1973) Service Manual has been prepared to assist
maintenance personnel in servicing and maintaining Model 337 series airplanes. This manual provides
the necessary information required to enable the mechanic to service, inspect, troubleshoot, remove, and
replace components or repair systems.
NOTE:
This manual is not intended to cover model 337 series airplanes produced after 1973. For
manuals related to these airplanes, please refer to applicable listings in the Cessna Propeller
Aircraft Customer Care Supplies and Publications Catalog.
C. Technical Publications are also available for the various components and systems that are not covered in
this manual. These manuals must be utilized as required for maintenance of those components and
systems, and can be purchased from the manufacturer.
4.
Temporary Revisions.
A.
5.
Serialization
A.
6.
Additional information, which becomes available, can be provided by temporary revision. This service is
used to provide, without delay, new information, which will assist in maintaining safe flight/ground
operations. Temporary revisions are numbered consecutively. Temporary revisions are normally
incorporated into the maintenance manual at the next revision.
All airplanes are issued a serial number. This number is assigned as airplane construction begins and
remains with the airplane throughout its service life. This serial number appears on the airplane data plate.
Airplane serial numbers are used to identify changes within the text or within an illustration. The absence
of a serial number in text or illustration indicates the material is applicable to all airplanes.
•
Revision Filing Instructions
A.
Regular Revision
(1) Pages to be removed or inserted in the maintenance manual are determined by the effectivity page
(page A located at the front of this manual). Pages are listed in sequence by the two-element
number. The first number(s), which represent the manual section number, are followed by a dash and
then the page number for that section. When two pages display the same two-element number, the
page with the most recent Date of Page Issue must be inserted in the service manual. The date
column on the effectivity page must verify the active page.
B.
Temporary Revision
(1) File temporary revisions in the applicable section in accordance with filing instructions appearing on
the first page of the temporary revision.
(2) The rescission of a temporary revision is accomplished by incorporation into the maintenance manual
or by a superseding temporary revision.
iv
© Cessna Aircraft Company
Change 2
Jan 5/2004
•
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
•
7.
IDENTIFYING REVISED MATERIAL.
A.
Additions or revisions to text in an existing section will be identified by a revision bar in the outside margin
of the page and adjacent to the change.
B.
When technical changes cause unchanged text to appear on a different page(s), a revision bar will be
placed in the outside margin adjacent to the page number, providing no other revision bar appears on the
page. These pages will display the current revision date in the inside margin opposite of the page number.
C. When extensive technical changes are made to text in an existing section that requires extensive revision,
revision bars will appear the full length of text.
D. Revised and new illustrations. either a revision bar along the side of the page or a hand indicator directing
attention to the area will indicate new or revised information.
E.
8.
Warnings, Cautions and Notes
A.
•
9.
Changes to wiring diagrams are indicated by shaded areas.
Throughout the text in this manual, warnings, cautions and notes pertaining to the procedures being
accomplished are utilized. These adjuncts to the text are used to highlight or emphasize important points.
Warnings and Cautions precede the text they pertain to, and Notes follow the text they pertain to.
WARNING:
Calls attention to use of materials, processes, methods, procedures, or limits
which must be followed precisely to avoid injury or death to personnel.
CAUTION:
Calls attention to methods and procedures which must be followed to avoid damage to
the airplane or equipment.
NOTE:
Calls attention to methods that will make the job easier.
Propeller Aircraft Customer Care Supplies and Publications Catalog
A.
A Cessna Propeller Aircraft Customer Care Supplies and Publications Catalog is available from a Cessna
Service Station or directly from Cessna Aircraft Company. The address is:
Cessna Aircraft Company
Department 751 C
P.O. Box 7706
Wichita, Kansas 67277-7706
This 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 and aerofiche form.
10. Customer Comments on Manuals
A.
•
Cessna Aircraft Company has endeavored to furnish you with an accurate, useful and up-to-date manual.
This manual can be improved with your help. Please use the return card that is provided with your manual
to report any errors, discrepancies, and omissions in this manual as well as any general comments you
wish to make.
Change 2
Jan 5/2004
© Cessna Aircraft Company
v
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
•
•
THIS PAGE INTENTIONALLY LEFT BLANK
vi
© Cessna Aircraft Company
Change 2
Jan 5/2004
•
•
SECTION 1
GENERAL DESCRIPTION
TABLE OF CONTENTS
GENERAL DESCRIPTION.
Model 337-Series . .
Description . . . . .
Page
1-1
1-1
1-1
Aircraft Specifications
Stations· . .
Torque Values. . . .
1-1
1-1
'-1
•
1-1. GENERAL DESCRIPTION.
speed, full-feathering propeller. In addition, the
Model T337-Series aircraft engines are turbocharged.
1-2. MODEL 337-SERIES.
•
1-3. DESCRIPTION. Cessna Model 337-Series aircraft, descr!.bed in this manual, are twin-engine,
high-wing monoplanes of all-metal, semimonocoque
construction. The aircraft employs a fully-retractable tricycle landing gear with spring-steel main
gear struts. The steerable nose gear is an air/oil
filled oleo strut. Thru 1971, the landing gear is hydraulically-a~tuated. Beginning with 1972 models,
the landing gear is electrically-actuated. The wing
flaps are electrically actuated, and flight adjustable
trim is prOVided for the rudder and elevator systems.
Four-place seating is standard, but prOVisions are
made for the addition of optional seats. The engines
are placed in tandem on the fuselage centerline and
the empennage is mounted on twin tail booms. The
aircraft is powered by two six- cylinder, horizontallyopposed, air-cooled, fuel-injected Continental engines. Each engine turns an all-metal. constant-
1-4. AIRCRAFT SPECIFICATIONS. Leading particulars of these aircraft, with dimensions based on
gross weight, are listed in figure 1-1. If these
dimensions are to be used for constructing a hangar
or computing clearances, remember that such factors
as tire pressures, tire sizes and load distribution
may result in some dimensions that are considerably
different from those listed.
1-5. STATIONS. A station diagram is included in
figure 1- 2 to assist in locating equipment when a
written description is inadequate or impractical.
1-6. TORQUE VALUES. A chart of recommended
nut torques is provided in figure 1-3. These values
are recommended for all installation procedures
contained in this manual, except where other values
are stipulated. They are not to be used for checking
tightness of installed parts during ser..,ice.
1-1
•
MODEL 337 AND T337 SERIES
DESIGN GROSS WEIGHT
(Thru 337A) . . .
(337B and T337B)
(337C)
(T337C)
Take Off
Landing
(3370)
(T337D)
Take Off
Landing
(337E and On)
(T337E and On)
Take Off
Landing
............ .
FUEL CAPACITY (Total-Less Auxiliary Tanks)
Usable . . . . . . . . . . . .
(Thru 337F)
(Total-Including Auxiliary Tanks)
(Thru 337F)
Usable . . . . . . . . . . . .
(Total) . . . . . . . . . • . .
(337G)
Usable . . . . . . . . • . . .
(337G)
OIL CAPACITY (Total-Both Engines). . . . . . . . .
.
(With External Oil Filter and all Turbocharged Engines)
ENGINE MODEL
(337)
..................... .
(T337) . . . . . . . . . . . . . '.' . . . . . . .
PROPELLER (Constant-Speed, Full-Feathering, Both Engines) . .
PROPELLER (Constant-Speed, Full-Feathering, Forward Engine).
PROPELLER (Constant-Speed, Full-Feathering, Rear Engine)
MAIN WHEEL TIRES
Size (Standard) . . . . . . . . . . . . .
Pressure . . . . . . . . . . . . . . . .
Size (Optional: Beginning with 337E -Series)
Pressure . . .
NOSE WHEEL TIRE
Size • . . . . . . . . . . . . . . . . . .
Pressure . . . . . . . . . . . . . . . .
NOSE GEAR STRUT PRESSURE (Strut Extended)
WHEEL ALIGNMENT
Camber
....... .
Toe-In (Total-Both Wheels)
AILERON TRAVEL
Up . . . . . . .
Down . . . . . .
WING FLAP TRAVEL
Inboard Flaps . .
Outboard Flaps . . . . . . . . . . . . . . . . . . .
RUDDER TRAVEL (Perpendicular to Rudder Hinge Centerline)
Outboard . . . . . . . . . . . . . . .
Inboard . . . . . . . . . . . . . . . .
RUDDER TRAVEL (Parallel to Fin Water Line)
Outboard
. ...
• ...•.
Inboard . . . . . . . . . . . . . . . .
. ...
4200lbs
4400lbs
4500lbs
4500lbs
4400lbs
4500 lbs
4500lbs
4400 lbs
4630lbs
4630lbs
4400lbs
93 gal. (558 lbs)*
92 gal. (552 lbs)*
131 gal. (786Ibs)*
128 gal. (768 lbs) *
125 gal. (750 lbs)*
118 gal. (708 lbs)*
20 qt
22 qt
CONTINENTAL 10-360 Series
CONTINENTAL TSIO-360 Series
76" McCAULEY (Thru 337F)
78" McCAULEY (337G)
76" McCAULEY (337G)
6.00 x 6, 8-Ply Rating
55 psi **
18 x 5.5, 8-Ply Rating
64 psi**
70 psi
(Thru 337F) (337G)
15.0 x 6.00 x 6, 4-Ply Rating
42 psi **
35 psi
•
4 0 ± 1 0 30'
.06"
o to
0 to 25 + 1 0 _20
0 0 to 25 0 , + 1 ° _20
0
0
,
25° ± 2°
17°, + 0 0 _20
22° ± 2 0
15°, + 0° _2°
• FIGURED AT 6 POUNDS PER GALLON.
·*AT TIRE INSTALLATION, TO AVOID TIRE SLIPPAGE AND TO SET TIRE BEAD ON RIM,
OVERPRESSURE NOSE WHEEL TIRE TO 55 PSI, AND THEN REDUCE TIRE PRESSURE
TO 42 PSI. OVERPRESSURE STANDARD (6.00 x 6) MAIN WHEEL TIRES TO 70 PSI, AND
THEN REDUCE TIRE PRESSURE TO 55 PSI. OVERPRESSURE OPTIONAL (18 x 5.5)
MAIN WHEEL TIRES TO 80 PSI, AND THEN REDUCE TIRE PRESSURE TO 64 PSI.
Figure 1-1. Aircraft Spectflcations (Sheet 1 of 2)
1-2
•
•
ELEVATOR TRAVEL
Up (Thru 337C)
Up (337D and On)
Down (Thru 337C)
Down (337D and On) .
ELEVATOR TRIM TAB TRAVEL
Up (Thru 337B) . . . . . .
Up (337C) . . . . . . . . .
Up (337 D and On). • • • . .
Down, with Flaps Up (337). . . . . .
Down, with Flaps Up (337A thru 337C)
Down, with Flaps Up (337D and On). .
Down, with 2/3 Flaps (337 thru 337C) •
Down, with Full Flaps (337D and On) .
PRINCIPAL DIMENSIONS
Wing Span
(337 thru 337D) • . • . . . . .
(337E and On) . . • . . • . . .
(337E and On with Strobe Lights)
Tail Span (Overall)
Length
(337 -Series) .
(T337-Series) . . • . • . . . . . . . . . •
Fin Height (Maximum with Nose Gear Depressed)
Track Width . . . . . . . . . . . . . . . . .
BATTERY LOCATION . . . . . . . . . . . . . •
21 0 ± 30'
26 0 ± 1 0
15 0 ± 2 0
15 ° ± 1 °
15°±1°
20° ± 1 °
15°±1°
15° ± 1 °
10° ± 1 °
0° ± 1°
26°,+1°_2°
15°±1°
38'
38' 2"
38' 4"
10' 8-1/4"
29' 9"
29' 10"
9' 4"
8' 2"
Left side of front firewall
•
•
Figure 1-1. Aircraft Specifications (Sheet 2 of 2)
1-3
WING STATIONS
222.00
192.00
207.00
*
162.00
177.00
107. 60
135.60
79. 60
66.00 149.12t6. 3~30. 00
93.60
150.00
55. 50 42. 75
~
AIRPLANE
•
I 'I ( ( r 23. 00
Thru 1971 Models, the landing and taxi lights are located
in the wing leading edges. Beginning with 1972 Models,
the landing and taxi lights are installed in the lower nose
cap assembly.
0.00
BOOM STATIONS
(NOT THE SAME AS FUSELAGE STATIONS)
34.45
60.70
27.25
42 j 30
I
127'.00
19.85
I
I
5l. 50
110.50
138.75
168.00
I
I
I
83.25
I
I
96.75
124.50
•
I
153.25
STA 70.00 AT <t. OF BOOM CORRESPONDS
TO FUSELAGE STA. 193.90
FUSELAGE STATIONS
0.00
80.00
84.50
FIREWALL
65.00
120.00 135.45
I
- - , L . - - WL
55. 00
,--:.e---WL 49.12
114.18
----WLO.OO
*
**
~
FRONT SPAR AT ~ WING BOLT HOLE IS STA. 136.44
~
REAR SPAR AT ~ WING BOLT HOLE IS STA. 165.09
Figure 1-2.
1-4
Fuselage, Wing and Boom Stations
•
RECOMMENDED NUT TORQUES
•
NOTE
THE TORQUE VALUES STATED ARE POUND-INCHES. RELATED ONLY TO OIL-FREE CADMIUM PLATED THREADS.
FINE THREAD SERIES
,
TYPE OF NUT
TAP
SIZE
•
8-36
10-32
1/4-28
5/16-24
3/8-24
7/16-20
1/2-20
9/16-18
5/8-18
3/4-16
7/8-14
1-14
1-1/8-12
1-1/4-12
TENSION
SHEAR
TORQUE
TORQUE
STD
(NOTE 1)
ALT
(NOTE 2)
STD
(NOTE 3)
ALT
(NOTE 2)
12-15
20-25
50-70
100-140
160-190
450-500
480-690
800-1000
1100-1300
2300-2500
2500-3000
3700-5500
5000-7000
9000-11000
20-28
50-75
100-150
160-260
450-560
480-730
800-1070
1100-1600
2300-3350
2500-4650
3700-6650
5000-10000
9000-16700
7-9
12-15
30-40
60-85
95-110
270-300
290-410
480-600
660-780
1300-1500
1500-1800
2200-3300
3000-4200
5400-6600
12-19
30-48
60-106
95-170
270-390
290-500
480-750
660-1060
1300-2200
1500-2900
2200-4400
3000-6300
5400-10000
COARSE THREAD SERIES
8-32
10-24
1/4-20
5/16-18
3/8-16
7/16-14
1/2-13
9/16-12
5/8-11
3/4-10
7/8-9
1-8
1-1/8-8
1-1/4-8
(NOTE 4)
(NOTE 5)
12-15
20-25
40-50
80-90
160-185
235-255
400-480
500-700
700-900
1150-1600
2200-3000
3700-5000
5500-6500
6500-8000
7-9
12-15
25-30
48-55
95-100
140-155
240-290
300-420
420-540
700-950
1300-1800
2200-3000
3300-4000
4000-5000
NOTES
1. Covers AN310, AN315, AN345, AN362, AN363, AN366, MS20365, "1452", "EB", "UWN", "ZI200",
NAS679, MS21044, MS21042, MS21045 and other self-locking nuts.
2. When using AN310 or AN320 castellated nuts where alignment between bolt and cotter pin is not reached
using normal torque values, use alternate torque values or replace nut.
3. Covers AN316, AN320. AN7502 and MS20364.
4. Covers AN310, AN340, AN366, MS20365, and other self-locking anchor nuts.
5. Covers AN316, AN320 and MS20364.
•
The above values are recommended for all installation procedures contained in this book except where other
values are stipulated. They are not to be used for checking tightness of installed parts during service.
Figure 1-3. Torque Values
1-5/(1-6 blank)
•
SECTION 2
GROUND HANDLING, SERVICING. LUBHlCATlON, AND INSPECTION
TABLE OF CONTENTS
•
. GROUND HANDLING
Towing .
Hoisting .
Jacking .
Parking .
Tie-Down
Flyable Storage
Returning Aircraft To Service
Temporary Storage. . . . . .
Inspection During Storage . .
Returning Aircraft To Service
Indefinite Storage. . . • • . •
Inspection During Storage . •
Returning Aircraft To Service
Leveling • • •
SERVICING •
Fuel Tanks
Fuel Drains
Fuel Strainers
Engine Oil . . . . . . . . .
Engine Induction Air Filters .
Vacuum System Air Filters .
Battery . • • . . • • • . .
Tires • • . . . . • . • • .
Nose Gear Strut . . . . . .
Nose Gear Shimmy Dampener
Page
2-1
2-1
2-1
2-1
2-3
2-3
2-3
2-3
2-3
2-4
2-4
2-4
2-6
2-6
2-6
2-6
2-6
2-6
2-6
2-7
2-7
2-8
2-8
2-8
2-10
2-10
2-1. GROUND HANDLING.
2-2. TOWING. MOVing the airplane by hand is accomplished by using the wing struts or landing gear
struts as push pOints. A tow bar attached to the nose
gear is used for steering and maneuvering the airplane. The tow bar is provided as standard equipment and is stowed in the baggage compartment.
ICAUTION\
When towing the airplane, never turn the nose
wheel more than 39 degrees either side of
center or the nose gear will be damaged. Do
not push on control surfaces or empennage
surfaces. Depress airplane nose when tOWing.
•
2-3. HOISTING. The airplane may be lifted by
means of hoisting lugs which are provided as optional
equipment. Provisions for attaching the optional
hoisting rings to the front and rear carry-thru spars
are provided as standard equipment. If the optional
hoisting rings are used, a minimum cable length of
60 inches for each cable is required to prevent bending of the eyebolt-type hoisting rings. If desired, a
spreader jig may be fabricated to apply vertical force
Hydraulic Brake System
Hydraulic Reservoir • .
Hydraulic Pump Check .
Hydraulic Filter . . • .
HydraUlic Fluid Sampling
Oxygen System
Oxygen Face Masks. • •
CLEANING . . . • . • . •
Windshield and Windows
Plastic Trim. • . .
Aluminum Surfaces .
Painted Surfaces . .
Engine Compartment
Upholstery and Interior
Propellers. • • • . •
Wheels . . . . . . .
LUBRICATION . . . • •
Nose Gear Torque Links
Universal Joints . • . •
Downlock Pins and Overce!1ter Buttons
Nose Gear Cam Follows
Wheel Bearing Lubrication
Fuel Selector Valve Lubrication
Aileron Rod End Bearings .
Wing Flap Actuators
INSPECTION . . . . . . . .
2-10
2-10
2-10
2-10
2-10
2-11
2-11
2-11
2-11
2-11
2-11
2-11
2-11
2-11
2-12
2-12
2-12
2-12
2-12
2-12
2-12
2-12
2-12
2-12
2-12
2-22
\\- hen hLllStill~ the airplane, use a
hoist with a mimmum eapacity of three tons.
:,_, ','k ':YP,):"llii.
2-4. JACKING. Refer to figure 2-1 for jacking
procedures. Wing jack pOints and mounting screws
are stowed in the map compartment. The jack
pOints are to be installed just outboard of the wing
strut, in the bottom forward flange of the front
wing spar. Remove existing screws to install the
jack pOints and reinstall after jacking operation
has been completed.
I~~UTION\
When using the universal jack point, flexibility
of the gear strut will cause the main wheel to
slide inboard as the wheel is raised, tilting the
jack. The jack must then be lowered for a
second jacking operation. Jacking both wheels
simultaneously with universal jack points is
not recommended. Do not use brake casting
as a jacking point.
If the airplane is to be jacked with the rear engine
removed, the tail must be weighted to provide balance
while jacking. This weight is added by placing shot
bags on the horizontal stabilizer rear spar.
2-1
•
1968 MODEL TIE-DOWN RINGS
ARE RETRACTABLE
o
TAIL STAND
33" MINIMUM CLEARANCE
I REQUIRED FOR GEAR
f RETRACTION
NOTE
Wing jacks available from the Cessna Service Parts Center are
REGENT Model 4939-30 for use with the SE-576 wing stands.
Combination jacks are the REGENT Model 4939-70 for use without
wing stands. The 4939-70 jack (70-inch) may be converted to the
4939-30 jack (30-inch) by removing the leg extensions and replacing lower braces with shorter ones. The base of the adjustable
tail stand (SE-767) is to be filled with concrete for additional weight
as a safety factor. The SE-576 wing stand will also accommodate
the SANCOR Model 00226-150 jack. Other equivalent jacks, tail
stands, and adapter stands may be used.
UNIVERSAL JACK POINT
(PARTNO'l~
•
JACKING PROCEDURE:
1. Install wing jack pOints (Part No. 1400110-2, 2 reqd. ) just outboard of wing struts.
2.
Position wing jacks at wing jack points.
3.
Locate one or two people at the aft end of the tail booms to balance the airplane manually
as the wing jacks are raised. The airplane will become tail heavy as the wings are jacked.
4. Raise wing jacks evenly until desired height is reached.
5. Attach a weighted, adjustable tail stand to either boom tie-down ring.
6. Position nose jack at nose jack point and raise until airplane becomes steady.
7. Use the universal jack point to jack one wheel.
8. The nose may be raised either by jacking with the nose jack or by placing weight, such as
shot bags, along the stabilizer rear spar.
Figure 2-1. JaCking
2-2
•
•
2-5. PARKING. Parking precautions depend principally on local conditions. As a general precaution,
it is wise to set the parking brake or chock the wheels
and install the control lock. In severe weather, and
high wind conditions, tie down the airplane as outlined in paragraph 2-6 if a hangar is not available.
2-6. TIE-DOWN. When mooring the aircraft in the
open, head into the wind if possible. Secure control
surfaces with the internal control lock and set brakes.
I~~UTIONI
Do not set parking brakes during cold weather
when accumulated moisture may freeze the
brakes or when the brakes are overheated.
After completing the preceding, proceed to moor the
aircraft as follows:
a. Secure ropes, chains, or cables of 700 pounds or
more tensile strength to the wing tie -down fittings located at the upper end of each wing strut. Secure
opposite ends of ropes, chains, or cables to ground
anchors.
b. Secure ropes, chains, or cables of 700 pounds
or more tensile strength to the tie -down fitting on
each tailboom and fasten opposite end of ropes, chains,
or cables to a common ground anchor.
NOTE
•
In locations where heavy snow accumulations
occur, additional precautions should be taken
to support the tail section of the aircraft. Snow
accumulations on the horizontal stabilizer can
result in considerable weight on the tail, causing it to rotate downward, resulting in damage
to the ventral fins. Proper nose gear tie-down
and a simple tail support attached to one of the
tail boom tie-down fittings will protect against
such damage.
c. Secure the middle of a rope (do not use chain or
cable) to the nose gear trunnion (see figure 2-2). Pull
each end away at a 45 degree angle and secure to the
ground anchors.
d. These aircraft are equipped with a spring-loaded
steering bungee which affords protection against normal wind gusts. However, if extremely high wind
gusts are anttcipated, additional external locks may
be installed.
e. Install pitot tube cover.
f. On turbocharged aircr2ft, close rear cowl flaps.
NOTE
•
In areas subject to severe wind-driven rainstorms the turbocharged aircraft should De
hangared to reduce the possibility of water
getting into the rear engine induction system.
If hangar storage is not available, install a
cover with prominent red streamer on the
rear engine air inlet scoop.
2-7. FLYABLE STORAGE. Flyable storage is defined as a maximum of 30 days non-operational
storage and/or the first 25 hours of intermittent en-
gine operation.
NOTE
The aircraft is delivered from Cessna with
a Corrosion Preventive Aircraft Engine Oil
(Military Specification MIL-C -6529, Type II).
This engine oil is a blend of aviation grade
straight mineral oil and a corrosion preventive compound. This engine oil should be
used for the first 50 hours of engine operation. Refer to paragraph 2-21 for oil changes
during the first 50 hours of operation.
During the 30 day non-operational storage or the
first 25 hours of intermittent engine operation, every
seventh day the propellers shall be rotated through
five revolutions, without running the engines. If the
aircraft is stored outSide, tie-down in accordance
with paragraph 2-6. In addition, the pitot tube,
static air vents, air vents, openings in the engine
cowling, and other Similar openings shall have protective covers installed to prevent entry of foreign
material. After 30 days, the aircraft should be
flown for 30 minutes or ground run-up until oil has
reached operating temperature.
2-8. RETURNING AIRCRAFT TO SERVICE. After
flyable storage, returning the aircraft to service is
accomplished by performing a thorough pre-flight
inspection. At the end of the first 25 hours of engine
operation, drain engine oil, clean oil screens and
change external oil filter element. Service engines
with correct grade and quantity of engine oil. Refer
to figure 2 -5 and paragraph 2 -21 for correct grade
of engine oil.
2-9. TEMPORARY STORAGE. Temporary storage
is defined as aircraft in a non-operational status for
a maximum of 90 days. The aircraft 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, the 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 the 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:
a. Fill fuel tanks with correct amount and grade of
gasoline.
b. Clean and wax aircraft thoroughly.
c. Clean any oil or grease from tires and coat tires
with a tire preservative. Cover tires to protect
against grease and oil.
d. Either block up fuselage to relieve pressure on
tires or rotate wheels every 30 days to change supporting points and prevent flat spotting the tires.
e. Lubricate all airframe items and seal or cover
all openings which could allow moisture and/or dust
to enter.
2-3
NOTE
The aircraft battery serial number is recorded in the aircraft equipment list. To
assure accurate warranty records, the
battery should be reinstalled in the same
aircraft from which it was removed. H a
battery is returned to service in a different
aircraft, appropriate record changes must
be made and notification sent to the Cessna
Claims Department.
f. Remove battery and store in a cool dry place;
service the battery periodically and charge as required.
o. Attach a warning placard to the effect that the
propeller shall not be moved while the engine is in
storage.
2-10. INSPECTION DURING STORAGE.
a. Inspect airframe for corrosion at least once a
month and remove dust collections as freqliently as
possible. Clean and wax as required.
b. Inspect the interior of at least one cylinder
through the spark plug hole for corrosion at least
once a month.
•
NOTE
Do not move crankshaft when inspecting
interior of cylinder for corrosion.
NOTE
An engine treated in accordance with the following may be considered protected against
normal atmospheric corrosion for a period
not to exceed 90 days.
g. Disconnect spark plug leads and remove upper
and lower spark plugs from each cylinder.
NOTE
The preservative oil must be Lubricating
Oil-Contact and Volatile, Corrosion inhibited, Mll.-L-46002, Grade 1 or equivalent.
The following oils are approved for spraying
operations by Teledyne Continental Motors,
Nucle Oil 105-Daubert Chemical Co., 4700
So. Central Ave., Chicago, Winois; Petratect VA - Pennsylvania Refining Co., Butler,
Pennsylvania; Ferro-Gard 1009G-Ranco
Laboratories, Inc., 3617 Brownsville Rd.,
Pittsburgh, Pennsylvania.
h. Using a portable pressure sprayer, atomize
spray the preservative oil through the upper spark
plug hole of each cylinder with the piston in a down
position. Rotate crankshaft as each pair of cylinders
is sprayed.
i. After completing step "h, " rotate crankshaft so
that no piston is at a top pOSition. H the aircraft is
to be stored outSide, stop two- bladed propeller so
that blades are as near horizontal as possible to provide maximum clearance with passing aircraft.
j. Again, spray each cylinder without mOving the
crankshaft, to thoroughly cover all interior surfaces
of the cylinder above the piston.
k. Install spark plugs and connect spark plug leads.
1. Apply preservative oil to the engine interior by
spraying apprOximately two ounces of the preservative
oil through the oil filler tube.
m. Seal all engine openings exposed to the atmosphere, USing suitable plugs or non-hygroscopic tape.
Attach a red streamer at each point that a plug or
tape is installed.
n. If the aircraft is to be stored outside, perform
the procedures outlined in paragraph 2-6. In addition, the pitot tube, static source vents, air vents,
openings in the engine COWling and other similar
openings should have protecti.ve covers installed to
prevent entry of foreign material.
2-4
c. If at the end of the 90 day period, the aircraft
is to be continued in non-operational storage, again
perform procedures outlined in paragraph 2-9.
2-11. RETURNING AIRCRAFT TO SERVICE. After
temporary storage, use the following procedures to
return the aircraft to service.
a. Remove aircraft from blocks and check tires for
proper tire inflation. Check for proper nose gear
strut inflation.
b. Check battery and install.
C. Check that oil sump has proper grade aDd quantity of engine oil.
d. Service iilduction air filter and remove warning
placard.
e. Remove materials used to cover openings.
f. Remove, clean, and gap spark plugs.
g. While spark plugs are removed, rotate propellet several revolutions to clear excess rust preventive oil from cylinders.
h. Install spark plugs. Torque spark plugs to
330:30 lb-in and connect spark plug leads.
i. Check fuel strainer. Remove and clean filter
screen if necessary. Check fuel tanks and fuel lines
for moisture and sediment, drain enough fuel to
eliminate.
j. Perform a thorough pre-flight inspection, then
start and warm-up engine.
•
2-12. INDEFINITE STORAGE. Indefinite storage is
defined as aircraft in a non-operational status for an
indefjnite period of time. Engines treated in accordance with the follOwing may be considered protected
against normal atmospheric corrOSion, provided the
procedures outlined in paragraph 2-13 are performed
at the intervals specified.
a. Operate engine unW oil temperature reaches
normal operating range. Drain engine oil sump and
reinstall drain plug.
b. Fill oil sump to normal operating capacity with
corrOSion preventive mixture which has been thoroughly mixed and pre-heated to a minimum of 221°F
at the time it is added to the engine.
•
•
NOTE
/CAUTION\
Injecting corrosion-preventive mixture too
fast can cause a hydrostatic lock.
Corrosion preventive mixture consists of one
part compound MlL-C-6529, Type I, mixed
with three parts new lubricating oil of the
grade recommended for service. Continental
Motors Corporation recommends Cosmoline
No. 1223, supplied by E. F. Houghton & Co. ,
305 W. LeHigh Avenue, Philadelphia, Pa.
During all spraying operations, corrosion mixture is pre-heated to 221 ° to 250°F.
e. Do not rotate propeller after completing step
"d."
f. Remove all spark plugs and spray corrosionpreventive mixture, which has been pre-heated to
221 to 250°F, into all spark plug holes to thoroughly cover interior surfaces of cylinders.
g. Install lower spark plugs or install solid plugs,
and install dehydrator plugs in upper spark plug
holes. Be sure that dehydrator plugs are blue in
color when installed.
h. Cover spark plug lead terminals with shipping
plugs (AN4060-1) or other suitable covers.
i. With throttle in full open pOSition, place a bag
of desiccant in the carburetor intake and seal opening with moisture resistant paper and tape.
j. Place a bag of desiccant in the exhaust tail0
c. Immediately after filling the oil sump with corrosion preventative mixture, fiy the aircraft for a
period of time not exceed a maximum of 30 minutes.
d. With engine operating at 1200 to 1500 rpm and
induction air filter removed, spray corrosion preventive mixture into induction airbox, at the rate of
one- half gallon per minute, until heavy smoke
comes from exhaust stack, then increase the spray
until the engine is stopped.
...t:~)'
./~
!
CHAIN, CABLE, OR ROPE --__.
•
....:............
.....
.... ....
....,...
"
"'-"
-,/
......
....
,,'
..,
/.
':..
-",
.....
......
.......
. <....... .
.::".....
......~ ....." -',
....
RUDDER GUST
LOCK (BCTH
SIDES)
"'-"
.....
......
....
... ::., ...... .
....
'. /: //f8\tq) __ ~"
:"
.....
......
....
....
........
ROPE ONLY--...1
NOTE
•
Prior to the 1968 models, tie-down rings are
stowed in glove compartment. Beginning with
the 1968 models, the tie-down fittings are the
retractable type.
CONTROL LOCK
Figure 2-2.
Tie-Down Diagram
2-5
pipe(s) and seal openings with moisture resistant
tape.
k. Seal cold air inlet to the heater muff with moisture resistant tape.
1. Seal engine breather by inserting a protex plug
in the breather hose and clamping in place.
m. Seal all other engine openings exposed to atmosphere USing suitable plugs or non-hygroscope tape.
NOTE
Attach a red streamer to each place plugs or
tape is installed. Either attach red streamers
outside of the sealed area with tape or to the
inside of the sealed area with safety wire to
prevent wicking of moisture into the sealed
area.
n. Drain corrosion-preventive mixture from engine
sump and reinstall drain plug.
NOTE
The corrosion-preventive mixture is harmful
to paint and should be wiped from painted surfaces immediately.
o. Attach a warning placard on the throttle control
knob, to the effect that the engine contains no lubricating oil. Placard the propeller to the effect that it
should not be moved while the engine is in storage.
p. Prepare airframe for storage as outlined in
paragraph 2-9 thru step "f...
NOTE
As an alternate method of indefinite storage,
the aircraft may be serviced in accordance
with paragraph 2 -9 providing the aircraft is
run-up at maximum intervals of 60 days and
then reserviced per paragraph 2 -9.
2-13. INSPECTION DURING STORAGE. Aircraft
in indefinite storage shall be inspected as follows:
a. Inspect cylinder protex plugs each 7 days.
b. Change protex plugs if their color indicate~ an 1
unsafe condition.
c. If the dehydrator plugs have changed color in one
half of the cylinders, all desiccant material in the
engine shall be replaced with new material.
d. EVE'ry 6 months respray the cylinder interiors
with corrosion-preventive mixture.
NOTE
Before spraying, inspect the interior of one
cylinder for corrosion through the spark
plug hole and remove at least one rocker box
cover and inspect the valve mechanism.
2-14. RETURNING AmCRAFT TO SERVICE.
After indefinite storage, use the following procedure
to return the aircraft to service.
a. Remove aircraft from blocks and check tires for
correct inflation. Check for correct nose gear strut
inflation.
2-6
b. Check battery and install.
c. Remove all materials used to seal and cover
openings.
d. Remove warning placards posted at throttle and
propeller.
e. Remove and clean engine oil screen, then reinstall and safety. On aircraft that are equipped with
an external oil filter, install new filter element.
f. Remove oil sump drain plug and drain sump.
Install and safety drain plug.
•
NOTE
The corrosion-preventive mixture will mix
with the engine lubricating oil, so flushing
the oil system is not necessary. Draining
the oil sump will remove enough of the corrosion-preventive mixture.
g. Service and install the induction air filter.
h. Remove dehydrator plugs and spark plugs or
plugs installed in spark plug holes and rotate propeller by hand several revolutions to clear corrosionpreventive mixture from.cylinders.
i. Clean, gap, and install spark plugs. Torque
plugs to the value listed in Section 10.
j. Check fuel strainer. Remove and clean filter
screen. Check fuel tanks and fuel lines for moisture
and sediment, and drain enough fuel to eliminate.
k. Perform a thorough pre -flight inspection, then
start and warm-up engine.
1. Thoroughly clean aircraft and flight test aircraft.
2-15. LEVELING. Longitudinal leveling of the airplane is accomplished by backing out the two leveling
screws, located on the left Side of the airplane just
below the pilot's side window, and placing a level
across the screws. A level placed across the front
seat rails at corresponding pOints is used to level
the airplane late rally.
•
2-16. SERVICING.
2 -17. Servicing reqUirements are shown in the
Servicing Chart (figure 2-5). The follOWing paragraphs supplement this figure by adding details.
2-18. FUEL TANKS should be filled to capacity immediately after flight to retard moisture condensation.
The airplane may have an optional auxiliary fuel tank
installed in each wing between the tail boom and fuselage. The recommended fuel grade to be used is
listed in figure 2-5. Total fuel capacity of the standard and optional fuel tanks is given in the chart in
Section 1.
2-19. FUEL DRAINS are located at various pOints
in the fuel system to provide for drainage of water
and sediment. See Section 11.
2-20. FUEL STRAINERS. Each 100 hours, clean
the fuel strainers as outlined in Section 11. During
the 1967 model year, the strainer drain control was
removed from the instrument panel and relocated
adjacent to the engine oil dipstick. Access to the
strainer drain control is through the engine oil
•
•
:.
I
dipStick access door. Remove drain plugs and open
strainer drain at the intervals specified in figure 2-5
to drain water and sediment from the fuel system.
Also, during daily inspection of the fuel strainer, if
any water is found in the fuel strainer, there is a
possibility that wing tank sumps and fuel lines contain water. Therefore, all fuel drain plugs should be
removed and all water drained from the fuel system.
2-21. ENGINE OIL. Check engine lubricating oil
with the dipstick five to ten minutes after the engine
has been stopped. The aircraft should be in as near
a level position as possible when checking the engine
Oil, so that a true reading is obtained. Engine oil
should be drained while the engine is still hot, and
the nose of the aircraft should be raised slightly for
more positive draining of any sludge which may have
collected in the engine oil sump. Engine oil should
be changed every four months, even though less than
the speCified hours have accumUlated. Reduce these
intervals for prolonged operations in dusty areas, in
cold climates where sludging conditions exist, or
where short flights and long idle periods are encountered, which cause sludging conditions. Always
change oU, clean oil screens and clean and/or change
external filter element whenever oil on the dipstick
appears dirty. Detergent or ashless dispersant oil,
conforming to Continental Motors Specification No.
MHS-24A, shall be used in these engines. Multiviscosity oil may be used to extend the operating
temperature range, improve cold engine starting and
lubrication of the engine during the critical warmup period, thus permitting flight through wider ranges
of climate change without the necessity of changing
oil. The multi-viscosity grades are recommended
for aircraft engines subjected to wide variations in
ambient air temperatures when cold starting of the
engine must be accomplished at temperatures below
30°F.
a total of 50 hours have accumulated or
oil consumption has stabilized, then
change to detergent oil.
When changing engine oil, remove and clean oil
screens, or install a new filter element on aircraft
equipped with an external oil filter. An oil quickdrain valve may be installed. This valve provides a
quick and cleaner method of draining the engine oil.
This valve is installed in the oil drain port of the oil
sump. To drain the engine oil, proceed as follows:
a. Operate engine(s) until oil temperature is at
normal operating temperature.
b. (Front Engine) Remove cowling and open landing
gear doors.
c. In the nose landing gear door opening, remove
oil drain plug from engine sump and allow oil to drain
into a container. Reinstall and safety oil drain plug.
IWARNING'
Do not install quick-drain valve shown in figure 2-3 in the front engine. The valve will
interfere with nose landing gear retraction.
d. (Rear Engine.) Remove cowling side panels.
e. Attach a hose to the quick-drain valve in oil
sump, or place a flexible funnel down through small
spring-loaded door in bottom of cowling. Push up
on quick-drain valve until it locks open, and allow
oil to drain into a container.
f. After oil has drained, close quick-drain valve
as shown in figure 2-3. Remove hose or funnel.
g. On turbocharged engines, remove oil drain plug.
Reinstall and safety after draining oil.
h. Remove and clean oil screen or change external
oil filter element of each engine.
i. Service each engine with correct amount and
grade of engine oil.
NOTE
•
New or newly-overhauled engines should
be operated on aviation grade straight
mineral oil until the oil change. If a detergent or ashless dispersant oil is used
in a new or newly-overhauled engine,
high oil consumption might possibly be
experienced. The anti-friction additives
in detergent and dispersant oils will retard "break-in" of the piston, rings and
cylinder walls. This condition can be
avoided by the use of straight mineral oil
Beginning with Serial 337-0612 and all
T337, the aircraft are delivered from
Cessna with straight mineral oil (MIL-L6529, Type n, RUST BAN). If oil must
be added during the first 25 hours, use
only aviation grade straight mineral oil
(non-detergent) conforming to Specification No. MIL-L-6082. After the first 25
hours of operation, drain engine oil sump
and clean both the oil suction strainer and
oil pressure screen. If an external oil
filte~ is. installed •. change filter element
at thls hme. Refill sump with straight
mineral oil (non-detergent) and use until
Valve shown open. To close, twist
screwdriver until valve unlocks and
snaps down to closed position.
Figure 2-3. Quick-Drain Valve
2-22. ENGINE INDUCTION AIR FILTERS keep dust
and dirt from entering the induction system. The
value of maintaining the induction air filters in a good
clean condition can never be overstressed. More engine wear is caused through the use of dirty and/or
2-7
damaged air filters than is generally believed. The
frequency with which the filter should be removed
and cleaned will be .determined primarily by the airplane operating conditions. A good general rule, however, is to remove, clean, and inspect filters at least
every 50 hours of engine operating time and more
frequently if warranted by operating conditions. Some
operators prefer to hold a spare set of induction air
filters at their home base of operation so that a clean
set of filters are always readily available. Under
extremely dusty conditions, daily servicing of the
filters is recommended.
NOTE
Prior to airplane serial number 337-0634 a
permanent type filter element is used. This
permanent type filter has a wire mesh screen
around the inside and the outside of the filtering media. Beginning with airplane serial
number 337-0634 and all service parts, an
improved filter element is used. This improved filter has a perforated steel band
around the inside and the outside of the filtering media. The filters used with the
turbocharged engines are of a different shape,
but are serviced in the same manner as the
improved filter.
To service the inwction air filters, proceed as
follows:
a. Remove filter from airplane. For removal refer to Section 10 for the non-turbocharged engines or
Section lOA for turbocharged engines.
b. Clean filter by blowing with compressed air (not
over 100 psi) from direction opposite of normal air
flow. Normal air flow for the cylindrical filter is
from outside to inside. Arrows on filter case indicate direction of normal air flow on filters used with
turbocharged engines.
NOTE
Use care to prevent damage to filter element
when cleaning with compressed air. Never
use air pressure greater than 100 psi to clean
filter.
c. After cleaning as outlined in step "b, .. filter may
be washed, if necessary, with a mild household detergent and warm water solution. A cold water solution
may be used.
ICAUTION\
Do not use solvent or cleaning fluids to wash
either type filter. Use only a mild household
detergent and water solution when washing the
filters.
NOTE
The improved filter assembly may be cleaned
with compressed air a maximum of 30 times
or it may be washed a maximum of 20 times.
The filter should be replaced after 500 hours
of engine operation or one year, whichever
should occur first. However, the filter should
2-8
be replaced anytime it is damaged.
The permanent filter may be cleaned and reused
as long as it is not damaged. A damaged filter
may have the wire mesh screen broken on the
inside or the outside of the filter, or the filtering media may have sharp or broken edges.
However, any filter that appears doubtful
should be replaced.
•
d. After washing, rinse filter in clean water until
rinse water runs clear from filter. Allow water to
drain from filter and dry with compressed air (not
over 100 psi).
NOTE
The filtering panels of the filter may become
distorted when wet, but they will return to
their original shape when dry.
e. Be sure induction air box and air inlet wcts to
the engine are clean, inspect and replace filter if it
is damaged.
f. Install filters as outlined in Section 10 for nonturbocharged engines or Section lOA for turbocharged
engines.
2-23. VACUUM SYSTEM AIR FILTERS. On aircraft
equipped with a vacuum system, inspect the central
filter every 100 hours for damage and cleanliness.
Change central air filter element every 500 hours of
operating time and whenever suction gage reading
drops below 4.6 inches of mercury. Also, do not
operate the vacuum system with the fnter removed,
or a vacuum line disconnected as particles of dust or
other foreign matter may enter the system and damage the vacuum operated instruments. Change gyro
internal filters are overhauled. Beginning with the
the vacuum system. These instruments are not equipped with internal filters. The new instruments are
smaller with a beveled box type case. Also, these
gyro Instruments and related plumbing are used as
service parts.
•
2-24. BATTERY. Servicing involves adding distilled water to maintain the.electrolyte even with the
horizontal baffle plate at the bottom of filler holes,
checking the battery cable connections, and neutralizing and cleaning off any spilled electrolyte or
corrosion. Use bicarbonate of soda (baking soda)
and water to neutralize electrolyte or corrosion.
Follow with a thorough flushing with water. Brighten
cables and terminals with a wire brush, then coat
with petroleum jelly before connecting. The battery
box also should be checked and cleaned if any corrosion is noticed. Distilled water, not acid or "rejuvenators, " should be used to maintain electrolyte
level. Check the battery every 50 hours (or at least
every 30 days), oftener in hot weather. See Section
15 for detailed battery replacement and testing.
2-25. TIRES should be maintained at the air pressure specified in the chart of Section 1. When checking tire pressure, examine tire for wear, cuts,
bruises, and slippage.
•
•
FILLER VALVE
Remove valve core and attach
hose to filler valve
CCNTAINER
NOSE GEAR SHOCK STRUT
•
While extending and compressing strut,
keep end of hose below level of clean
hydraulic fluid.
Figure 2-4.
Filling Nose Gear Strut
SHOP NOTES:
•
2-9
NOTE
Recommended tire pressure should be maintained. Especially in cold weather, remember that any drop in temperature of the air
inside i:I. tire causes a corresponding drop in
pressure.
2-26. NOSE GEAR STRUT. The nose gear strut
requires periodic checking to ascertain that the strut
is filled with hydraulic fluid and is inflated to the
correct air pressure. When servicing the nose gear
strut proceed as follows:
a. Remove valve cap and reduce air pressure to
zero.
b. Remove valve core and attach hose and container
as shown in figure 2-4.
c. Lift nose of airplane, extend and compress strut
several times to expel any entrapped air, then lower
nose of airplane until strut is telescoped to its shortest length. Remove hose and container.
d. Install valve core and inflate strut to pressure
specified in Section 1.
NOTE
The nose landing gear shock strut will normally require only a minimum amount of
service. Maintain the strut extension pressure as shown in Section 1. Lubricate
landing gear as shown in figure 2-6. Check
the landing gear daily for general cleanliness, security of mounting, and for hydraulic fluid leakage. Keep machined surfaces
wiped free of dirt and dust, using a clean
lint-free cloth saturated with hydraulic flUid
(MIL-H-5606) or kerosene. All surfaces
Should be wiped free of excess hydrauliC
fluid or kerosene.
2-27. NOSE GEAR SHIMMY DAMPENER. The
shimmy dampener should be serviced at least every
100 hours. The dampener must be filled completely
with fluid, free of entrapped air, to serve its purpose.
To fill or add fluid to shimmy dampener while installed on airplane:
a. Remove filler plug from dampener.
b. Using a tow-bar, turn nose gear in the direction
that places thE' dampener piston at the end opposite
the filler plug.
c. Fil! with clean hydraUlic fluid.
d. Install and safety filler plug.
To fill shimmy dampener when it is removed from
airplane, proceed as follows:
a. Remove filler plug from dampener.
b. Submerge dampener in clean hydraulic fluid and
work dampener piston shaft in and out to remove any
entrapped air and ascertain complete filling of cylinder.
c. Reinstall plug before removing dampener from
hydraulic fluid.
NOTE
Keep shimmy dampener, especially the exposed portions of the dampener piston shaft
2-10
clean to prevent collection of dust and grIt
which could cut the seals in the dampener
barrel. Keep machined surfaces wiped free
of dirt and dust, using a clean lint-free cloth
saturated with hydraulic fluid (MIL-H-5606)
or kerosene. All surfaces should be wiped
free of excess hydraulic fluid or kerosene
•
2-28. HYDRAULIC BRAKE SYSTEMS should be
checked for the correct amount of fluid at least
every 100 hours. Add hydraulic fluid at the brake
master cylinders. Bleed the brake system of entrapped air whenever there is a spongy response
to the brake pedals.
2-29. HYDRAULIC RESERVOm. The reservoir
fluid level should be checked and replenished as necessary every 25 hours. Filling is accomplished by
using a pressure brake bleeder or Hydro Fill unit
attached to filler fitting on forward side of the firewall. Hydraulic fluid should be pumped into the
filler unit until fluid flows from the reservoir overboard vent line. The reservoir may also be filled
as outlined in paragraph 5-127 using the Hydro Test
Unit.
2-30. HYDRAULIC PUMP CHECK. The aircraft
may be equipped with the rear engine optional hydraulic system. Since either hydraulic pump will
operate the system, it is very difficult to determine
if one pump has failed. At each 100-hour inspection
a hydraulic pump check should be performed as
follows:
a. With front engine running, place master switch
to the OFF position.
b. Check that landing gear doors open.
c. Place master switch to ON position. Check
that landing gear doors close.
d. Start rear engine and shut down front engine.
e. Place master switch in the OFF position and
check that landing gear doors open.
f. Place master switch to ON position and check
that landing gear doors close.
•
2-31. HYDRAULIC FILTER. The screen in the hydraulic filter should be removed and cleaned with solvent (Federal SpeCification P-S-66l, or equivalent) at
the first 25 hours and the first 50 hours of operation,
thereafter at 100-hour inspections or whenever improper fluid circulation is suspected. Also, clean
rear filter when optional dual hydraulic system is
installed.
2-32. HYDRAULIC FLUID SAMPLING. At the
first 50 and first 100 hours, thereafter at each 500
hours or one year, whichever should occur first, a
sample of fluid should be taken and examined for
sediment and discoloration. This may be done as
follows:
a. Place master switch in OFF position.
b. With landing gear control handle in downneutral, actuate hydraulic hand pump to supply pressure to open landing gear doors.
c. Remove door open line from a door actuator
cylinder. Using the hydraulic hand pump, drain off
a small sample of hydraulic fluid into a non-metallic
container.
•
•
d. Reconnect door actuating cylinder line and inspect fluid coloration. If the fluid is clear and is not
appreciably darker in ~olor than new fluid, continue
to use the present fluid in the system.
e. If the fluid coloration is doubtful, insert a strip
of polished copper in the fluid. Keep the copper in
the fluid for six hours at a temperature of 70°F or
more. A slight darkening is permissible and there
should be no pitting or etching visible up to 20X
magnification.
2-33. OXYGEN SYSTEM.
IWARNING'
Do not rotate control lever to "ON" position
with outlet (low pressure) port(s) open to
atmosphere. Refer to Section 13.
2-34. OXYGEN FACE MASKS.
(Refer to Section 13. )
2-35. CLEANING.
2-36. Keeping the aircraft clean is important. Besides maintaining the trim appearance of the airplane,
cleaning reduces the possibility of corrosion and
makes inspection and maintenance easier.
•
2-37. WINDSHIELD AND WINDOWS should be cleaned carefully with plenty of fresh water and a mild
detergent, using the palm of the hand to feel and dislodge any caked dirt or mud. A sponge, soft cloth,
or chamois may be used, but only as a means of
carrying water to the plastic. Rinse thoroughly,
then dry with a clean moist chamois. Do not rub
the plastic with a dry cloth since this builds up an
electrostatic charge which attracts dust. Oil and
grease may be removed by rubbing lightly with a
soft cloth moistened with Stoddard solvent.
ICAUTION!
Do not use gasoline, alcohol, benzene, acetone,
carbon tetrachloride, fire extinguisher fluid,
de-icer fluid, lacquer thinner or glass window
cleaning spray. These solvents will soften
and craze the plastic.
Mter washing, the plastic windshield and windows
should be cleaned with an aircraft windshield cleaner.
Apply the cleaner with soft cloths, and rub with moderate pressure. Allow the cleaner to dry, then wipe
it off with soft flannel cloths. A thin, even coat of
wax, polished out by hand with clean soft flannel
cloths, will fill in minor scratches and help prevent
further scratching. Do not use a canvas cover on
the windshield or windows unless freezing rain or
sleet is antiCipated since the cover may scratch the
plastic surface.
•
2 -38. PLASTIC TRIM. The plastic trim instrument
panel, and control knobs need only to be wiped off
with a damp cloth. Oil and grease on the control
wheel and control knobs can be removed with a cloth
moistened with Stoddard solvent. Volatile solvents,
such as mentioned in paragraph 2-37, must never be
used since they soften and craze the plastic.
2-39. ALUMINUM SURFACES require a minimum of
care, but should never be neglected. The airplane
may be washed with clean water to remove dirt, and
with carbon tetrachloride or other 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 aircraft products.
2-40. PAINTED SURFACES. The painted exterior
surfaces of the aircraft, under normal conditions,
require a minimun of pOlishing or buffing. Approximately 15 days are required for acrylic or lacquer
paint to cure completely and approximately 90 days
are required for vinyl paint to cure completely; in
most cases, the curing period will have been completed prior to delivery of the airplane. In the event
that polishing or buffing is required within the curing
period, it is recommended that the work be done by
an experienced painter. 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 cause corrosion or make scratches should never be used. Remove stubborn oil and
grease with a cloth moistened with Stoddard solvent.
Mter the curing period, the airplane may be waxed
with a good automotive wax. A heavier coating of
wax on the leading edges of the wings and tail and
on the engine nose cap will reduce the abrasion
encountered in these areas.
2-41. ENGINE COMPARTMENT. Cleaning is essential
to minimize any danger of fire, and for proper inspection of components. The engine and engine compartment may be washed down with a suitable solvent,
and then dried thoroughly. Refer to Section 10.
2 -42. UPHOLSTERY AND INTERIOR cleaning prolong the life of upholstery fabrics and interior trim.
To clean the interior, proceed as follows:
a. Empty all ash trays.
b. Brush or vacuum clean the carpeting and upholstery to remove dirt.
c. Wipe leather and plastic surfaces with a damp
cloth.
d. Soiled upholstery fabrics and carpeting may be
cleaned with a foam-type detergent, used in accordance with the manufacturer I s instruclions.
e. 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 in the fabric to be
cleaned. Never saturate the fabric with a volatile
solvent; it may damage the padding and backing
materials.
f. Scrape sticky materials with a dull knife, then
spot-clean the area.
2-11
2-43. PROPELLERS should be wiped off occasionally with an oily cloth to clean off grass and bug
stains. In salt water areas this will assist in corrosion-proofing the propeller.
2-44. WHEELS should be washed periodically and
examined for corrosion, chipped paint, and cracks
or dents in the wheel castings. Sand smooth,
prime, and repaint minor defects.
2-45.
LUBRICATION.
2 -46. LUBRICA TION requirements are shown on the
Lubrication Chart (figure 2-6). Before adding grease
to grease fittings, wipe off all dirt. Lubricate until
new grease appears around parts being lubricated,
and wipe off excess grease. The following paragraphs supplement this figure by adding details.
2-47. NOSE GEAR TORQUE LINKS. Lubricp.te
torque links every 50 hours. When operating in
dusty conditions, more frequent lubrication is
recommended.
2 -48. UNIVERSAL JOINTS. It is important that
all pivot pOints and sliding surfaces of the universal
joints be lubricated. Lubricate with SAE 90 gear oil
at installation and at each lOa-hour inspection. Apply gear oil to each pivot point and sliding surface of
the universal joint so that the oil will work between
the mOVing surfaces.
2-49. DOWNLOCK PINS AND OVERCENTER BUTTONS. At each laO-hour inspection, clean with solvent and inspect for sharp edges the downlock pins,
over center buttons, and main landing gear struts
where they contact the pins and buttons. Smooth all
sharp edges. Do not paint the "tracks" on the struts
made by the pins and buttons. Lubricate down lock
pins, overcenter buttons, and strut with general
purpose grease. Also, clean and lubricate the cam
surface of the downlock switch bracket.
2-5Q. NOSE GEAR CAM FOLLOWERS. At the first
500-hour inspection, remove plugs in stud of cam
followers and lubricate with general purpose grease.
Lubricate cam followers at each 500-hour inspection,
using automotive type rubber tipped grease gun when
lubricating cam followers. There is no need to reinstall plugs in cam follower studs.
•
2-51. WHEEL BEARING LUBRICATION. It is
recommended that nose and main wheel bearings be
cleaned and repacked at the first 100-hour inspection
and at each 500-hour inspection thereafter. If more
than the usual number of take-off and lanc:!.:.ngs are
made, extensive taxiing is required, or the airplane
is operated in dusty areas or under seacoast conditions, it is recommended that cleaning and lubrication of wheel bearings be accomplished at each 100hour inspection.
2-52.
FUEL SELECTOR VALVE LUBRICATION.
It is now recommended that the fuel selector valve
detents and valve shaft be lubricated at each 100hour inspection. Apply lubrication to each detent of
the valve and to the valve shaft where it protrudes
from the valve cover boss.
2-53. AILERON ROD END BEARING. The actuating rod attach point is exposed to the weather through
a small opening in the upper leading edge of the aileron. Therefore, periodic inspection and lubrication is required to prevent corrosion of the bearing
in the rod end. At each lOa-hour inspection, disconnect the control rods at the aileron and inspect
each rod end ball for corrosion. If no corrosion is
found, wipe the surface of the rod end balls with
general purpose oil and rotate the ball freely to distribute the oil over its entire surface and connect
the control rods. If corrosion is detected during
inspection, replace the rod end.
•
2-54. WING FLAP ACTUATORS. On aircraft prior
to 337-0240, clean screw jack threads of the wing
flap actuator with solvent and brUSh, and lubricate
screw jack threads as speCified in figure 2-6. Beginning with Serial 337-0240, the wing flap actuator
jack screw threads require no lubrication.
•
2-12
•
.......
'
.......
'
. . ::3t~:.
.
.......
.....................
............
......
..•.........
~ ~,.~: <: :f~:,: ~:;·: :~: .
.........
0
•
"
*RECOMMENDED FUEL:
AVIATION GRADE---I00/130 MINIMUM GRADE
*100/130 low lead aviation fuel with a lead content limited to 2 cc per gallon is also approved.
HYDRAULIC FLUID
SPEC. NO. MIL-H-5606
OXYGEN:
SPEC. NO. MIL-O-27210
_RECOMMENDED ENGINE OIL:
AVIATION GRADE---SAE 30 OR SAE IOW30 BELOW 40°F.
SAE 50 ABOVE 40°F.
•
• MULTI-VISCOSITY OIL WITH A RANGE OF SAE 10W30 IS RECOMMENDED FOR IMPROVED
STARTING AND TURBOCHARGER CONTROLLER OPERATION IN COLD WEATHER. DETERGENT OR DISPERSANT OIL, CONFORMING TO CONTINENTAL MOTORS SPECIFICATION
MHS-24A, MUST BE USED.
Figure 2-5. Servicing Chart (Sheet 1 of 4)
2-13
o
DAILY
1
FUEL TANKS:
Fill after each flight.
for details.
2
FUEL TANK SUMP DRAINS:
Drain water and sediment before first flight of day and after each refueling.
to paragraph 2-19 for details.
Keep full to .·etard condensation. Refer to paragraph 2-18
OXYGEN CYLINDER (OPTIONAL MODEL 337) (STANDARD MODEL T337):
Check for anticipated requirements before each flight. Refer to Section 13 for
details.
4
FUEL STRAINERS:
Drain water and sediment before first flight of day.
5
OIL DIPSTICK:
Check on preflight. Add oil as necessary.
6
OIL FILLER CAP:
Whenever oil is added, check that oil filler cap is tight and oil filler door is
secure.
7
14
8 9
50HOURS
INDUCTION Am FILTERS:
Service every 50 hours; oftener under dusty conditions. Refer to paragraph 2-22
for details.
•
BATTERY:
Check electrolyte level every 50 hours (or at least every 30 days), oftener in hot
weather. Refer to paragraph 2-24 for details.
ENGINE OIL SYSTEM:
Change engine oil and external filter element every 50 hours. Without external
filter, change oil and clean oil screen EVERY 25 HOURS. Reduce these intervals
under severe operating conditions. Refer to paragraph 2-21 for details.
16
HYDRAULIC FILTER:
See under 100 hours.
15
HYDRAULIC FLUID CONTAMINATION CHECK:
See under 500 hours.
o
10
Refer to paragraph 2-21 for details.
PITOT AND STATIC PORTS:
Check for obstruction before first flight of the day.
o
100HOURS
VACUUM RELIEF VALVE FILTER:
Check air inlet filter for cleanliness. Remove, flush with solvent, and dry with
compressed air. Replace air filter at each engine overhaul.
Figure 2-5. ServiCing Chart (Sheet 2 of 4)
2-14
Refer
3
11 12
•
•
•
0
100 HOURS (Cont)
17
BRAKE MASTER CYLINDERS:
Check fluid level and refill as required with hydraulic fluid.
20
SHIMMY DAMPENER:
Check fluid level and refill as required with hydraulic fluid.
paragraph 2 -27 for details.
4
1&
FUEL STRAINERS:
Remove bowl and filter screen and clean every 100 hours.
2-20for details.
Refer to
Refer to paragraph
HYDRAULIC FILTER:
Remove and clean filter screen at first 25 and first 50 hours of operation: thereafter, at each 100-hour inspection. Refer to paragraph 2-31 for details.
OSOOHOURS
•
15
HYDRAULIC FLUID CONTAMINATION CHECK:
At the first 50 and first 100 hours, thereafter at each 500 hours or one year,
whichever occurs first, make a hydraulic fluid sampling test as outlined in
paragraph 2-32.
21
VACUUM SYSTEM AIR FILTERS:
Replace central air filter every 500 hours. Replace gyro instrument air filters
at instrument overhaul. Refer to paragraph 2-23 for details.
~
18
TIRES:
Maintain proper tire inflation as listed in chart in Section 1. Also refer to
paragraph 2-25.
19
NOSE GEAR SHOCK STRUT:
Keep strut filled and inflated to correct pressure.
details.
15
•
AS REQUIRED
Refer to paragraph 2-26 for
HYDRAULIC FLUID RESERVOIR AND FILLER:
Check fluid level at least every 25 hours through sight gage in reservoir and
fill as required. Refer to paragraph 2-29 for details .
Figure 2-5. Servicing Chart (Sheet 3 of 4)
2-15
D
13
•
AS REQUIRED (Cont)
GROUND SERVICE RECEPTACLE (PRIOR TO 1967 MODELS) (OPT):
Connect to 24-volt, DC, negative-ground power unit for cold weather starting and
lengthy ground maintainance of the electrical system. Master switch should be turned
on before connecting a generator type external power source; it should be turned off
before connecting a battery type external power source. Refer to Section 10.
ICAUTION!
Be certain that the polarity of any external power source or batteries
is correct (positive to positive and negative to negative). A polarity
reversal will result in immediate damage to semiconductors in the
airplane's electrOnic equipment.
13 GROUND SERVICE
RECEPTACLE (1967 MODELS AND ON) (OPT):
Connect to 24-volt, DC, negative-ground power unit for cold weather starting and
lengthy ground maintenance of the airplane's electrical equipment with the exception
of electronic equipment. Master switch should be turned on before connecting a
generator type or battery type external power source. Refer to Section 10.
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.
FUSES:
Replace as required with the following fuses:
PROTECTS
LOCATION
Clock
Upper left forward firewall.
S-1091-2
Front Cowl Flaps
At cowl flap motor.
AGC-2
Rear Cowl Flaps
At cowl flap motor.
AGC-3
Cigarette Lighter
Forward side of instrument
panel just left of center.
SPE-6
Alternators
(Auxiliary Field
Circuit)
Upper left forward firewall.
S-1091-5
Figure 2-5. Servicing Chart (Sheet 4 of 4)
2-16
NUMBER
•
•
• o~
METHOD OF APPLICATION
FREQUENCY (HOURS)
<9> (t-)
,...
HAND
GREASE
GUN
~
OIL
CAN
WHERE NO INTERVAL IS SPECIFIED,
LUBRICATE AS REQUIRED AND
WHEN ASSEMBLED OR INSTALLED.
SYRINGE
(FOR POWDERED
GRAPffiTE)
NOTE
The military specifications listed below are not mandatory,
but are intended as guides in choosing satisfactory materials.
~cts of most reputable manufacturers meet or exceed
these specifications.
LUBRICANTS
POWDERED GRAPHITE
MIL-G-81322AGENERAL PURPOSE GREASE
GH
MlL-G-23827 AIRCRAFT AND INSTRUMENT GREASE
GL - MlL-G-21164 HIGH AND LOW TEMPERATURE GREASE
OG- MlL-L-7870 GENERAL PURPOSE OIL
PL - VV-P-236
PETROLATUM
6 0 - MlL-L-2105B MULTI PURPOSE GEAR OIL GRADE 90
,G -
caR -
•
REFER TO SHEET 4
CAM FOLLOWERS
AUiO REFER TO
PARAGRAPH 2-50
NEEDLE BEARING
THRUST BEARING
AUiO REFER TO
PARAGRAPH 2-47
ALSO REFER TO PARAGRAPH 2-51-"""'---
•
Figure 2 -6. Lubrication (Sheet 1 of 5)
2-17
CONTROL COLUMN
NEEDLE BEARING
ROLLERS
t
~.
~f~
.
t
*
GR
THRUST BEARINGS
~
•
. . NEEDLE BEARINCS
GI
!
FLAP BELLCRANK
NEEDLE BEARINGS
• I~
NEEDLE BEARING
GI~O"'':b y
e
~~""':
~/~~
•..:'"
,-
.... , !)-_.
~
__ •
•
FLAP BELLCRANK NEEDLE BEARINGS
THRU SERIAL 337 -0239
A~O
REFER TO
PARAGRAPH 2-53
•
Beginning with serial 337-0240, no lubrication is required on the wing flap actuator screw jack threads. Also, refer to
paragraph 2 -54.
COLLAR
SCREW
HOUSING
AILERON BELLCRANK
NEEDLE BEARINGS
COLLAR
*
PL
BATTERY TERMINALS
ELEVATOR TRIM
TAB ACTUATOR
Figure 2 -6.
2-18
Change 1
Lubrication (Sheet 2 of 5)
•
RUDDER BARS AND PEDALS
BEARING BLOCK
HALVES
OG
OG
ALL LINKAGE
POINT PIVOTS
GR
PARKING BRAKE
CABLE AND CONDUIT
*
GH
•
FUEL SELECTOR
VALVE CONTROLS
OG
PARKING BRAKE
HANDLE SHAFT
REFER TO
PAAAGii(
THRU 33701316
AND F33700024
r&
GH
BEGINNING WITH
33701317 AND
F33700025
LANDING GEAR
UNIVERSAL JOINTS
•
ALL PIANO IDNGES
ALSO REFER TO
PARAGRAPH 2-52 FUEL SELECTOR VALVES
Figure 2-1.
Lubrication (Sheet 3 of 5)
2-19
SPRA Y BOTH SIDES OF SHADED AREAS WITH
ELECTROFILM LUBRI-BOND "A, .. wmCH IS
AVAIlABLE IN AEROSOL SPRAY CANS, OR
AN EQUIVALENT LUBRICANT.
•
MAIN GEAR THRUST BEARINGS
NOSE GEAR OOWNLOCK
•
ALSO REFER TO
PARAGRAPH 2-43
ALSO REFER TO
PARAGRAPH 2-43
OOWNLOCK PIN
CAM SURFACE
OVERCENTER BUTTON
ALSO REFER TO
PARAGRAPH 2-43
Figure 2-1. Lubrication (Sheet 4 of 5)
2-20
•
•
~
PL
CONTROL QUADRANT LEVERS
•
•
GH
PROPELLER SYNCHRONIZER CONTROL
FOUL-WEATHER AND
CABIN DOOR WINDOW
INSERT GROOVES
~
GH
NOTE
Sealed bearings require no lubrication.
McCauley propellers are lubricated at overhaul and require no other lubrication.
Do not lubricate roller chains or cables except under seacoast conditions. Wipe with a clean,
dry cloth.
Lubricate unsealed pulley bearings, rod ends, Olite bearings, pivot and hinge points, and any
other friction point obviously needing lubrication, with general purpose oil every 1000 hours or
oftener if required.
Paraffin wax rubbed on seat rails will ease sliding the seats fore and aft.
Lubricate door latching mechanism with MIL-G-81322A, applied sparingly, to friction
points, every 1000 hours or oftener if binding occurs. No lubrication is recommended
on the rotary clutch.
•
Lubricate quadrant controls with petrolatum on levers only within a one-inch radius from
pivot hole .
Figure 2-1.
Lubrication (Sheet 5 of 5)
2-21
INSPECTION
To avoid repetition throughout the inspection, general pOints to be checked are given below. In the inspection,
only the items to be checked are listed; details as to how to check, or what to check for, are excluded. The
inspection covers several different models. Some items apply only to specific models, and some items are
optional equipment that may not be found on a particular airplane. Check FAA Airworthiness Directives and
Cessna Service Letters for compliance at the time specified by them. Federal Aviation Regulations require
that all civil aircraft have a periodic (annual) inspection as prescribed by the administrator, and performed by a
person designated by the administrator. The Cessna Aircraft Company recommends a 100-hour periodic inspection for the aircraft.
'
•
CHECK AS APPLICABLE:
MOVABLE PARTS for: lubrication, servlcmg, security of attachment, binding, excessive wear,
safetying, proper operation, proper adjustment, correct travel, cracked fittings, security of
hinges, defective bearings, cleanliness, corrosion, deformation, sealing, and tensions.
FLUID LINES AND HOSES for: leaks, cracks, dents, kinks, chafing, proper radius, security,
corrosion, deterioration, obstructions, and foreign matter.
METAL PARTS for: security of attachment, cracks, metal distortion, broken spotwelds, corrosion,
condition of paint, and any other apparent damage.
WIRING for: security, chafing, burning, defective insulation, loose or broken terminals, heat deterioration, and corroded terminals.
BOLTS IN CRITICAL AREAS for: correct torque in accordance with the torque values given in the
chart in Section 1, when installed or when visual inspection indicates the need for a torque check.
FILTERS, SCREENS. AND FLUIDS for: cleanliness, contamination and/or replacement at specified
intervals.
AffiPLANE FILE.
Miscellaneous data, information, and licenses are a part of the airplane file. Check that the following documents are up-to-date and in accordance with current Federal Aviation Regulations. Most of the items listed
are required by the United States Federal Aviation Regulations. Since the regulations of other nations may
require other dC'cuments and data, owners of exported aircraft should check with their own aviation officials
to determine their individual requirements.
To be displayed in the aircraft at all times:
1. Aircraft Airworthiness Certificate (FAA Form 8100-2).
2. Aircraft Registration Certificate (FAA Form 8050-3).
3. Aircraft Radio Station License, if transmitter installed (FCC Form 556).
To be carried in aircraft at all times:
1. Weight and Balance, and asso,ciated papers (Latest copy of the Repair and Alteration
Form, FAA Form 337, if applicable).
2. Aircraft Equipment List.
To be made available upon request:
1. Aircraft Log Book and Engine Log Books.
•
ENGINE RUN-UP.
Before beginning the step-by-step inspection, start, run up, and shut down the engine in accordance with
instructions in the Owner's Manual. Ouring the run-up, observe the following, making note of any discrepancies or abnormalities:
l. Engine temperatures and pressures.
2. Static rpm.
3. Magneto drop (See Owner's Manual).
4. Engine response to changes in power.
5. Any unusual engine noises.
6. Propeller response (See Owner's Manual).
7. Fuel tank selector valve; operate engine on each tank position and off position l()n~
enough to make sure the valve functions properly.
8. Idling speed and mixture; proper idle cut-off.
9. Alternator and ammeter.
10. Suction Gage.
ll.
Fuel flow indicator.
12. Optional hydraulic pump (see paragraph 2-30).
2-22
•
•
SCOPE AND PREPARATION.
If the engine is NOT equipped with an external oil filter, change engine oil and clean the oil screens EVERY 25
HOURS of engine operation.
The 50-hour inspection includes a visual check of the engine, pr(ljJeller, and aircraft exterior for any apparent
damage or defects; an oil change and filter element change on aircraft equipped with an external oil filter; and
accomplishment of lubrication and servicing requirements. Remove propeller spinner and engine cowling, and
replace after the inspection has been completed.
The 100-hour (or annual) inspection includes everything in the 50-hour inspection. Also loosen or remove all
fuselage, wing, boom, empennage, and upholstery inspection doors, plates, and fairings as necessary to perform a thorough, searching inspection of the airplane. On those aircraft with inspection plates on the tunnel
cover, it is not necessary to remove the tunnel cover during inspection, remove only the inspection plates on
the tunnel cover. Replace after the inspection has been completed.
NOTE
Numbers appearing in the "AS SPECIFIED" column refer
to the data listed at the end of the inspection chart.
•
AS SPECIFIED
EACH 100 HOURS
PROPELLER.
EACH 50 HOURS
1. Spinner and spinner bulkhead--------------------------------------------------------
•
Blades-------------------------------------~--------------------------------------
•
3. Hub-------------------------------------------------------------------------------
•
4. Mounting nuts--- ------ ------------------------------------ -------- -------- - -- --- ---
•
5. Governor and control---------------------------------------------------------------
•
2.
6. Unfeathering accumulator----------------------------------- -------- --- --- -- - - ---- -7.
Synchronizing system -------------------------------------------------------------
•
1
•
8. Anti-Ice electrical wiring ----------------------------------------------------------
•
9. Anti-Ice brushes, slip ring, and boots -----------------------------------------------
•
ENGINE COMPARTMENT.
Check for evidence of oil, hydraulic fluid, and fuel leaks, then clean entire engine and compartment,
if needed, prior to inspection.
•
1.
Engine oil, screen, filler cap, dipstick, drain plug, and external filter element ----------
2. Oil cooler-------------------------------------------------------------- __________ _
•
2
•
2-23
AS SPECIFIED
EACH 100 HOURS
EACH 50 HOURS
3. Induction air filters (Also see paragraph 2-22) ---------------------------------------4.
Induction airbox, air valves, doors, and controls -------------------------------------
5. Cold and hot air hoses -------------------------------------------------------------6.
Engine baffles - - --- --- -- - -- - - -- - -- - - - -- - -- - -- ---- -- - --- - - -- - - - - - - - - - -- - - --- - - -- ----
7.
Cylinders, rocker box covers, and push rod housings -----------:-----------------------
8.
Crankcase, oil pan, accessory section. and front crankshaft seal ----------------------
9.
Metal lines and nuid hoses ----------------------------------------------------------
10. Intake and exhaust systems (Also refer to Section 10)
11.
I~nition
---------------------------------
harness - - ---- -- ---- -- -- -- - --------- -- ---- --------- --- ----- -- - - ---- --- - ----
•
•
•
•
•
•
•
•
•
12. Spark plugs and compression check ------------------------- .. -----------------------13.
Crankcase breather lines -- -- ------------ ---------- ---- ---------- ---- - - --------- - ---
14.
Electrical wiring - ------------ ----------------------------- -- -------- - ----- -- ------
15.
Vacuum pump, oil separator, and relief valve ----------------------------------------
16.
Vacuum relief valve filter ----------------------------------------------------------
17.
Engine controls and linkage ---------------------------------------------------------
18.
Engine shock mounts, engine mount structure, and ground straps -----------------------
19.
(Exhaust type heaters) Cabin heater valves, doors, and controls -----------------------
20. Starter, solenoid, and electrical connections -----------------------------------------
•
•
•
•
•
•
•
13
14
•
•
23.
Alternator brushes, brush leads, and slip ring ---------------------------------------
24.
Voltage regulator mounting and electrical leads --------------------------------------
25.
Magnetos (externally) and electrical connections --------------------------------------
26.
Magneto breaker compartment (Also refer to Section 10)
27.
Magneto timing to engine -----------------------------------------------------------
28.
Fuel injection fuel-air control unit, fuel pump, fuel manifold valve, fuel lines, and
nozzles ---------------------------------------------------------------------------
29.
Firewall - ---- -- ------ -- --- - -- -- - --- ---- ---------- -- --'-- ---- - --- ------ ----- -- - - - ---
30.
Engine cowling - - -- -- ----- - ------ ----- ----------------- ------- ----- --- -------------
31. Cowl flaps controls and motors -----------------------------------------------------32.
2-24
Hydraulic pump (5) -------- ------ --- --- --------------------------------------------
3
15
•
4
21. Starter brushes, brush leads, and commutator --------------------------------------22. Alternator, and electrical connections -----------------------------------------------
•
•
5
•
•
16
• 16
•
•
•
•
•
•
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
•
AS SPECIFIED
EACH 100 HOURS
EACH 50 HOURS
33. Turbocharger ..................................................................................... ........................................
•
34. Turbocharger pressurized vent lines to fuel pump, discharge nozzles,
and fuel flow gage......................................................................................................................
•
35. Turbocharger mounting brackets ..............................................................................................
36. Waste gate, actuator, and linkage, and controllers ...................................................................
•
37. All oil lines to turbocharger, waste gate, and controllers ...........................................................
•
•
38. For airplanes equipped with an internal combustion heater: Check ventilating and
combustion air inlets, exhaust outlet, fuel and drain lines, electrical connections,
combustion air blower, and air tube connections ......................................................................
39. Turbocharger oil line check valves ............................................................................................
7
21
23
•
40. Engine fuel injection nozzle removal and cleaning ....................................................................
FUEL SYSTEM
1 . Fuel strainer, drain valves and controls .....................................................................................
2. Fuel strainer screens and bowls ............................................................................................... .
3. Electric fuel pumps and electric connections ........................................................................... .
•
•
•
4. Fuel tanks, fuel sump tanks, fuel lines, drains, filler caps, and placards .................................. .
•
•
5. Drain fuel and check tank interior, attachment, and outlet screens .......................................... .
6. Fuel vents and vent valves ........................................................................................................
7. Fuel selector valve and placards ............................................................................................. .
8. Fuel quantity gages and transmitter units ................................................................................ .
6
•
•
•
•
•
9. Vapor return lines and check valves ......................................................................................... .
10. Engine Primer ............................................................................................................................
•
11. Turbocharger vent system .........................................................................................................
12. Perform a fuel quantity indicating system operational test.
Refer to Section 14 for detailed accomplishment instructions .................................................. .
22
AIRFRAME
1. Aircraft exterior ....................... .............................................................................................. .....
•
2. Aircraft structure ........................... ............................................................................................
3.
4.
5.
6.
7.
•
•
Windows, windshield, and doors ..............................................................................................
•
Seat stops, seat rails, upholstery, structure, and seat mounting ..............................................
•
Seat belts and attaching brackets .............................................................................................
•
Control column bearings, sprockets, pulleys, cables, chains, and turnbuckles.........................
•
Control lock, control wheel, and control column mechanism ....................................................
•
8. Instruments and markings .........................................................................................................
•
Change 2
Jan 5/2004
© Cessna Aircraft Company
2-25
I
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
AS SPECIFIED
EACH 100 HOURS
EACH 50 HOURS
•
9. Gyro filter replacement ..............................................................................................................
10. Vacuum system central air filter ................................................................................................
11. Magnetic compass compensation ............................................................................................ .
12. Instrument wiring and plumbing ................................................................................................
13. Instrument panel, shock mounts, ground straps, cover, decals, and labeling .......................... .
14. Defrosting, heating, ventilating systems, and controls ............................................................. .
15. Cabin upholstery, trim, sun visors, and ashtrays ...................................................................... .
16. Area beneath floor, lines, hoses, wires, and control cables ..................................................... .
17. Electrical horns, lights, switches, circuit breakers, fuses, and spare fuses .............................. .
18. Exterior lights .............................................................................................................................
19. Pitot and static systems .............................................................................................................
20. Stall warning sensing unit, and pitot and stall warning heaters ................................................ .
21. Electronic equipment and controls ............................................................................................
22. Antennas ....................................................................................................................................
23. Battery, battery box, and battery cables ................................................................................... .
24. Battery electrolyte level (Also see paragraph 2-24) .................................................................. .
25. Oxygen system (Also see Section 13) ...................................................................................... .
26. Oxygen supply, masks, and hoses ........................................................................................... .
27. De-Ice system plumbing ............................................................................................................
28. De-Ice system components .......................................................................................................
I
29. De-Ice system boots ..................................................................................................................
30. Wings - front spar cap, rear spar cap, and front spar web ....................................................... .
CONTROL SYSTEMS
In addition to the items listed below, always check for correct direction of movement,
correct travel and correct cable tension.
2-26
1. Cables, terminals, pulleys, pulley brackets, cable guards, turnbuckles, and fairleads ..............
•
2. Chains, terminals, sprockets, and chain guards........................................................................
3. Trim control wheels, indicators, actuator, and bungee ........................... ...................................
•
4. Travel stops .......................................................... ........................ ............................ .................
•
5. All decals and labeling...............................................................................................................
•
6. Elevator downspring system....................... .................................................... ...........................
•
© Cessna Aircraft Company
•
Change 2
Jan 5/2004
•
•
AS SPECIFIED
EACH 100 HOURS
EACH 50 HOURS
7. Flap rollers and tracks, flap electrical indicating system, flap mechanical indicating
system, flap controls, flap electric motor brake and transmission, and flap/elevator
trim intercoMect system -------------------------------------------------- - -----
8. Rudder pedal assemblies and linkage ------------.. ---------------------------------9. Skin and structure of control surfaces and trim tabs ---------------------------------
10. Balance weight attachment -------------------------------------------------------11. Elevator trim tab actuator lubrication and tab free-play inspection --------------------12. Trim tabtnspection ______________________________________________________________ _
13. Trim tab control system __________________________________________________________ _
•
•
•
•
· 18171
•
•
LANDING GEAR.
•
1. Brake fluid, lines and hoses, linings, discs, brake assemblies, and master
cylinders -----------------------------------------------------------------------.
•
2. Main gear wheels, wheel bearings, spring struts, and tires --------------------------
•
3.
Nose gear strut and shimmy dampener servicing -------------------------------- ___ .
4.
Nose gear wheel, wheel bearings, strut, steering system, shimmy dampener, tire,
and torque links - - ---------------------------------------------------------- -- __ _
5.
Parking brake system -----------------------------------------------------------.
•
•
•
LANDING GEAR RETRACTION SYSTEM.
NOTE
When performing inspection of the landing gear
retraction system, a hydraulic power source
can be used. Refer to Section 5 for test stand
operation procedures.
1.
Operate the landing gear through five fault-free cycles, noting cycling time. Refer to
5-----------------------------------------------------------------------
•
Check landing gear doors for at least 1/2 inch clearance with any part of landing gear
during operation, and for proper fit when closed -----------------------------------.
•
3.
Check down position of main gear struts. Refer to Section 5
----------------------.
•
4.
Check main gear dawnlock engagement. Refer to Section 5
------------------------
•
5.
Check overcenter adjustments of retracted main gear dawnlock. Refer to Section
5
---------------------------------------------------------------------------
•
Check operation of downlock cam. Refer to Section 5 -----------------------_______
•
Section
2.
•
6.
Chauge 1
2-27
AS SPECIFIED
EACH 100 HOURS
EACH 50 HOURS
7 Check main gear uplock hook operation.
(Refer to Section 5.) __________________________
•
8· Check that main gear snubbing action occurs. (Refer to Section 5.) _____________________
•
9 Check adjustment and operation of main gear up and down indicator switches, nose
gear up and down indicator switches, and nose gear safety switch. (Refer to Section 5.) Also check indicator lights for proper operation. ______________________________ _
10 Check nose gear downlock adjustments.
(Refer to Section 5. ) __________________________ _
Check nose gear uplock operation. (Refer to Section 5.) ______________________________ _
11
12 Check adjustment of landing gear handle up-down switch. (Refer to Section 5.) ----------13 Check operation of landing gear handle lockout solenoid. (Refer to Section 5. )------- ____ _
14 Check all hydraulic system components for security, hydraulic leaks, and any
apparent damage to components or mounting structure. --------- _______________________ _
15 Check universal jOints for cracks and excessive wear. -------------------------- ______ _
16
Check gear and door linkage for security, wear of pivot points and bearings, and for
distortion or other damage ------------------------ _______________________________ _
17 Check main gear strut-to-saddle attachment ---------------__________________________ _
18 Check torque of adapter-to-pivot shaft attaching bolts, and resafety ---------- __________ _
19
Check condition of all springs ---------------------- ________________________________ _
20
Clean hydraulic filter (Refer to Section 2. ) __________________________________________ _
21
Clean small in-line filters at each end of restrictor check valve between main
gear actuator and main gear downlock cylinders. Also, clean in-line filter in
nose gear up line on forward side of front firewalL ___________________________________ _
22
•
•
•
•
•
•
•
•
•
•
24 Check roller clearances on steering cam
(Refer to Section 5.) _________________________
•
11
&
Hydraulic fluid contamination check (Refer to Section 2. ) _____________________________ _
23 Check security and operation of emergency hand pump _______________________________ _
10
•
12
•
•
NOTE
A high-time inspection is merely a 100-hour inspection with the addition of an engine overhaul. Continental Motors Corporation, Inc. recommends overhaul at 1500 hours for the 10-360 Series engines,
and overhaul at 1400 hours for the TSIO-360 Series engines. At time of engine overhaUl, engine
accessories, turbochargers, controllers, waste gate valves, and waste-gate actuators should be overhauled. Engine propellers and governors should be overhauled at 1200 hours of engine operating
time. Refer to Section 12 for specific information.
1 First 25 hours; each 100-hour inspection thereafter.
2 First 25 hours, refill with straight grade mineral oil (non-detergent) and use until a total of
50 hours have accumulated or oil consumption has stabilized, then change to detergent oil.
Thereafter, change oil each 25 hours if the engine is NOT equipped with an external filter.
2-28
•
CESSNA AIRCRAFT COMPANY
MODEL 337
•
SERVICE MANUAL
3.
At each instrument overhaul, replace filter.
4.
Each 200 hours for Delco Remy or each 1500 hours for Prestolite.
5.
Each 500 hours.
6.
Each 1000 hours, or to coincide with engine overhauls.
7.
It is recommended that the internal combustion heater be removed from the aircraft for a complete
inspection and necessary overhaul operations at the expiration of 500 hours of operation or after each
heating season, whichever occurs first (refer to Cessna Multi-engine Service Information Letter ME82-17, or
latest revision).
8.
Replace central filter each 500 hours; gyro filters at instrument overhaul. See paragraph 2-23.
9.
Anticipated requirements before each oxygen flight.
10. At first 1OO-hour inspection; at next 1OO-hour inspection after new shear washers installed.
11. At first 25 hours and first 50 hours of operations; at each 1OO-hour inspection thereafter.
12. First 50 and first 100 hours, thereafter at each 500 hours or one year, whichever comes first.
13. Replace fluid hoses at engine overhaul or after 5 years, whichever comes first.
14. General inspection every 50 hours. Refer to Section 10 and 1OA for 100 hour inspection.
15. Each 50 hours for general condition and freedom of movement. These controls are not repairable. Replace
at each engine major overhaul.
16. Check timing each 200 hours; check breaker compartment each 500 hours, unless timing is off.
•
17. Check that snap rings are properly located between spacers and actuator mounting damp. Check that
mounting damp bolts are torqued to 20-25 Ib-in. Apply white lacquer torque putty to bolt for future
inspections. Inspect guard block for condition and attachment.
18. Inspect trim tab hinge for evidence of damage. Inspect hinge pin for proper installation and safety. Inspect
push-pull rod and actuator rod end bearing for evidence of binding and damage. Inspect push-pull rod
attach-bolt at the actuator and trim tab horn for proper safetying, nut with cotter pin.
19.
Inspect system for operation and tab for freedom of movement. Check tab travel, and adjust if required,
refer to Section 1 of this manual.
20. Accomplish in accordance with Service Letter ME78-2 and any supplements or changes thereto.
21.
Replace turbocharger oil line check valves every 1000 hours of operation (Refer to Cessna MUlti-engine
Service Bulletin MEB92-4 Revision 2, or latest revision).
22.
Fuel quantity indicating system operational test is required every 12 months. Refer to Section 14 for detailed
accomplishment instructions.
23. At the first 1OO-hour inspection on new, rebuilt or overhauled engines remove and clean the fuel injection
nozzles. After the initial inspection has been accomplished, the fuel nozzles must be deaned at 300-hour
intervals or more frequently if fuel stains are noted.
•
Change 2
Jan 5/2004
© Cessna Aircraft Company
2-29
I
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
2-55.
COMPONENT TIME LIMITS
1. General
A.
Most components listed throughout Section 2 must be inspected as detailed elsewhere in this
section and repaired, overhauled or replaced as required. Some components, however, have
a time or life limit, and must be overhauled or replaced on or before the specified time limit.
NOTE:
Overhaul - Item can be overhauled as defined in FAR 43.2 or it can be replaced.
NOTE:
Replacement - Item must be replaced with a new item or a serviceable item that is
within its service life and time limits or has been rebuilt as defined in FAR 43.2.
•
B. This section provides a list of items that must be overhauled or replaced at specific time limits.
Table 1 lists those items that Cessna has mandated must be overhauled or replaced at
specific time limits. Table 2 lists component time limits that have been established by a
supplier to Cessna for the supplier's product.
C.
2.
In addition to these time limits, the components listed herein are also inspected at regular time
intervals set forth in the Inspection Charts, and can require overhaul/replacement before the
time limit is reached based on service usage and inspection results.
Cessna-Established Replacement Time Limits.
A.
The following component time limits have been established by Cessna Aircraft Company.
•
Table 1: Cessna-Established Replacement Time Limits
REPLACEMENT
TIME
OVERHAUL
Restraint Assembly, Pilot, Copilot,
and Passenger Seats
10 years
NO
Trim Tab Actuator
1,000 hours or 3 years,
whichever occurs first
YES
Vacuum System Filter
500 hours
NO
Vacuum System Hoses
10 years
NO
Pitot and Static System Hoses
10 years
NO
Vacuum Relief/Regulator Valve Filter
(If Installed)
500 hours
NO
Engine Compartment Flexible Fluid
Carrying Teflon Hoses (CessnaInstalled) Except Drain Hoses
(Drain hoses are replaced
on condition)
10 years or engine overhaul,
whichever occurs first
(Note 1)
NO
Engine Air Filter
500 hours or 36 months,
whichever occurs first
(Note 9)
NO
COMPONENT
2-30
© Cessna Aircraft Company
Change 2
Jan 5/2004
•
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
•
•
3.
COMPONENT
REPLACEMENT
TIME
Engine Compartment Flexible Fluid
Carrying Rubber Hoses (CessnaInstalled) Except Drain Hoses
(Drain hoses are replaced
on condition)
5 years or engine overhaul,
whichever occurs first
(Note 1)
NO
Engine Mixture, Throttle, and
Propeller Controls
At engine TBO
NO
Check Valve (Turbocharger
Oil Line Check Valve)
Every 1,000 hours of
operation
(Note 10)
NO
Oxygen Bottle - Light Weight Steel
(lCC-3HT, DOT-3HT)
Every 24 years or 4380 cycles,
whichever occurs first
NO
Oxygen Bottle - Composite
(DOT-E8162)
Every 15 years
NO
Engine Driven Dry Vacuum Pump
Drive Coupling
(Not lubricated with engine oil)
6 years or at vacuum
pump replacement,
whichever occurs first
NO
Engine Driven Dry Vacuum Pump
(Not lubricated with engine oil)
500 hours
(Note 11)
NO
Standby Dry Vacuum Pump
500 hours or 10 years,
whichever occurs first
(Note 11)
NO
OVERHAUL
Supplier-Established Replacement Time Limits
A.
The following component time limits have been established by specific suppliers and are
reproduced as follows:
Table 2: Supplier-Established Replacement Time Limits
COMPONENT
REPLACEMENT
TIME
ELT Battery
(Note 3)
NO
Vacuum Manifold
(Note 4)
NO
Magnetos
(Note 5)
YES
Engine
(Note 6)
YES
Engine Flexible Hoses
(Note 2)
NO
Auxiliary Electric Fuel Pump
(Note 7)
YES
Propeller
(Note 8)
YES
OVERHAUL
(TCM-Installed)
•
Change 2
Jan 5/2004
© Cessna Aircraft Company
2-31
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
NOTES:
Note 1: This life limit is intended to not allow flexible fluid-carrying Teflon or rubber hoses in a deteriorated
or damaged condition to remain in service.
Replace engine compartment flexible Teflon
(AE3663819BXXXX series hose) fluid-carrying hoses (Cessna-installed only) every ten years or at
engine overhaul, whichever occurs first. Replace engine compartment flexible rubber fluid-carrying
hoses (Cessna-installed only) every five years or at engine overhaul, whichever occurs first (this
does not include drain hoses). Hoses that are beyond these limits and are otherwise in a
serviceable condition, must be placed on order immediately and then be replaced within 120 days
after receiving the new hose from Cessna.
•
Note 2: For TCM engines, refer to Teledyne Continental Service Bulletin SB97-6, or latest revision.
Note 3: Refer to FAR 91.207 for battery replacement time limits.
Note 4: Refer to Airborne Air & Fuel Product Reference Memo No. 39, or latest revision, for replacement
time limits.
Note 5: For airplanes equipped with Slick magnetos, refer to Slick Service Bulletin SB2-80C or latest
revision for time limits.
For airplanes equipped with TCM/Bendix magnetos, refer to Teledyne Continental Motors Service
Bulletin No. 643, or latest revision, for time limits.
Note 6:
Refer to Teledyne Continental Service Information Letter SIL98-9, or latest revision, for time limits.
Note 7:
Refer to Cessna Service Bulletin MEB94-3 Revision 2/Dukes Inc. Service Bulletin NO. 0003, or
latest revision.
Note 8:
Refer to the applicable McCauley Service Bulletins and Overhaul Manual for replacement and
overhaul information.
Note 9: The air filter can be cleaned, refer to Section 2 of this service manual and for airplanes equipped
Refer to Donaldson Aircraft Filters Service
with an air filter manufactured by Donaldson.
Instructions P46-9075 for detailed servicing instructions. The address for Donaldson Aircraft Filters
is:
•
Customer Service
115 E. Steels Corners RD
Stow OH. 44224
CAUTION: DO NOT OVER-SERVICE THE AIR FILTEA. OVER-SERVICING INCREASES THE
RISK OF DAMAGE TO THE AIR FILTER FROM EXCESSIVE HANDLING. A
DAMAGEDIWORN AIR FILTER MAY EXPOSE THE ENGINE TO UNFILTERED AIR
AND RESULT IN DAMAGE/EXCESSIVE WEAR TO THE ENGINE.
Note 10: Replace the turbocharger oil line check valve every 1,000 hours of operation (Refer to Cessna
Service Bulletin MEB92-4 Revision 2, or latest revision).
Note 11: Replace engine driven dry vacuum pump not equipped with a wear indicator every 500 hours of
operation, or replace according to the vacuum pump manufacturer's recommended inspection and
replacement interval, whichever occurs first.
Replace stand-by vacuum pump not equipped with a wear indicator every 500 hours of operation or
10 years, whichever occurs first, or replace according to the vacuum pump manufacturer's
recommended inspection and replacement interval, whichever occurs first.
For a vacuum pump equipped with a wear indicator, replace pump according to the vacuum pump
manufacturer's recommended inspection and replacement intervals.
2-32
© Cessna Aircraft Company
Change 2
Jan 5/2004
•
.0
•
•
SECTION 3
FUSELAGE
TABLE OF CONTENTS
Page
3-1
FUSELAGE . . . . .
3-1
Windshield and Window s
3-1
Description
3-1
Cleaning
3-1
Waxing .
3-2
Repairs.
3-2
Scratches
3-6
Cracks .
3-6
Windshield
3-6
Description
.
Removal and Installation (thru aircraft
Serials 33701462 and F33700055).. 3-6
Removal and Installation (Beginning with
aircraft Serials 33701463 and
3-6
F33700056
... .
3-7
Windows . . . . . . . . . . .
3-7
Foul- Weather Window
3-7
Description . . . . . .
3-7
Removal and Installation
3-7
Emergency Window .
3-7
Description . .
3-7
Installation
3-7
Fixed Cabin Windows
3-7
Description . .
3-7
Removal and Installation
Movable Window (Thru Aircraft serials
33701462 and F33700055
3-7
Description . . . . . . . . . . 3-7
Cabin Door . . . . . . . . . . . . . . 3-7
Description . . . . . . . . . . . . 3-7
Removal and Installation (Thru Aircraft
Serials 33701462 and F33700055 . . 3-7
Removal and Installation (Beginning with
Aircraft Serials 33701463 and F33700056 . . . . . . . . . . . . . . 3-7
Cabin Door Latch (Thru Aircraft Serials
33701462 and F33700055) . .
3-16
Description . . . . . . . . . . 3-16
Adjustment . . . . . . . . . . 3-16
Indexing Cabin Door Handle . . . 3-16
Cabin Door Latch (Beginning With Aircraft
Serials 33701463 and F33700056
3-16
Description . . . . . . .
3-16
Adjustment . . . . . . .
3-16
Indexing Cabin Door Handle
3-16
Baggage Door . . . . . . .
3-16
Description . . . . . .
3-16
Removal and Installation
3-16
3-16
Seats . . . . . . .
3-16
Individual Seats
3-16
Description
3-16
Removal and Installation
3-18
Bench Seats . . . . . . . .
3-18
Description . . . . . .
3-18
Removal and Installation
3-18
Power Seat . . . . . . . .
3-18
Description . . . . . .
3-18
Removal and Installation
3-18
Quick AttaChing 5th and 6th Seats
3-18
Description . . . . . .
Removal and Installation
3-18
Seat Repair . . . .
3-18
Cabin Upholstery. . . .
3-18
Description . . . .
3-18
Materials and Tools
3-18
Soundproofing .
3-18
Description . .
3-18
Cabin Headliner . .
3-33
Description . .
3-33
Removal and Installation (Thru
Aircraft Serials 33701462
and F33700055). . . . . . . . 3-33
Removal and Installation (Beginning
with Aircraft Serials 33701463
and F33700056) .
3-33
Upholstery Side Panels
3-33
Windlace
3-33
Description
3-33
Carpeting . . .
3-33
Description
3-33
Safety Belts . .
3-36
Description
3-36
Shoulder Harness
3-36
Description .
3-36
Cargo Tie-Downs
3-36
Description .
3-36
Outside Step . . .
3-36
Description .
3-36
Removal and Installation
3-36
Maintenance
3-36
Cargo Pack . .
3-36
Description
3-36
Removal
3-36
Installation
3-37
Rigging Front Cowl Flaps with Cargo
Pack (Non-Turbocharged) . . . . .
3-37
3-l. FUSELAGE
Windows and skin laps.
3-2. WlNDOWS AND WINDSHIELD.
3-4. CLEANING (Refer to Section 2).
3-3. DESCRIPTION. The windshield and windows
are Single-piece acrylic plastic panels held by formed retainers secured to the fuselage with screws and
nuts. Both Windshield and Window s are sealed, on
installation with 3C-200 sealant, Churchill Chemical
Corporation. 579.6 Sealer, Prestite Engineering
Company, may be used to seal creVices, voids around
3-5. WAXING. Waxing will fill in minor scratches
in clear plastic and help protect the surface from
further abrasion. Use a good grade of commercial
wax applied in a thin, even coat. Bring wax to a high
polish by rubbing lightly with a clean, dry flannel
cloth.
3-1
3-6. REPAIRS. Damaged window panels and windshield may be removed and replaced if the damage
is extensive. However, certain repairs as precribed in the following paragraphs can be made successfully without removing the damaged part from
the aircraft. Three types of temporary repairs for
cracked plastic are possible. No repairs of any kind
are recommended on highly- stressed or compound
curves where the repair would be likely to affect the
pilot's field of vision. Curved areas are more diffibult to repair than flat areas and any repaired area
is both structurally and optically inferior to the original surface.
sively finer grade of abrasives until the scratches
disappear.
c. When the scratches have been removed, wash
the area thoroughly with clean water to remove all
gritty particles. The entire sanded area will be
clouded with minute scratches which must be removed to restore transparency.
d. Apply fresh tallow or buffing compound to a
motor-driven buffing wheel. Hold the wheel against
the plastic surface, moving it constantly over the
damaged area until cloudy appearance disappears.
A 2000-foot-per-minute surface speed is recommended to prevent overheating and distortion.
3-7. SCRATCHES. Scratches on clear plastic surfaces can be removed by hand- sanding operations
followed by buffing and poliShing, if steps below are
followed carefully.
a. Wrap a piece of No. 320 (or finer) sandpaper
or abrasive cloth around a rubber pad or wood
block. Rub the surface around the scratch with a
circular motion, keeping the abrasive constantly
wet with clean water to prevent scratching the surface further. Use minimum pressure and cover an
area large enough to prevent the formation of ''bull'seyes" or other optical distortions.
b. Continue the sanding operation, using progres-
NOTE
•
Polishing can be accomplished by hand but it
will require a considerably longer period of
time to attain the same result as a buffing
wheel.
e. When buffing is finished, wash the area thoroughly and dry with a soft flannel cloth. Allow the surface
to cool and inspect the area to determine if full transparency has been restored. Then apply a thin coat of
hard wax and polish the surface lightly with a clean
flannel cloth.
•
WOOD REINFORCEMENT
8
ALWAYS DRILL END OF CRACK CUSHION OF
RUBBER
TO RELIEVE STRAIN
OR FABRIC
SOFT WIRE
LACING
CEMENTED
FABRIC PATCH
TEMPORARY
REPAIR
OF CRACKS
SANDDIG REPAIR
Figure 3-1. Repair of Windows and Windshield
3-2
•
8
•
NOTE
579.6 sealer is used to seal
crevices and voids around
windows, skin laps and fasteners on cabin top down to
floorboard on each Side of
cabin to prevent leaks.
Detail
B
DetanA
REFER TO FIGURE 3-3
JO
•
o
17
7
J4
Detail
Detail
E
Detail
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
C
Trim
Windshield
Center Strip
EC-1202 Tape
Air Temp Probe
579. 6 presstite Sealer
Screw
Cabin Top Skin
Rubber Moulding
Figure 3-2.
10.
11.
12.
13.
14.
15.
16.
17.
18.
Inner Window
Window
Felt Seal
Cabin Skin
579.6 Presstite Sealer
Catch
Latch Handle
EC -801 Sealer
Rubber Seal
0
THRU AIRCRAFT SERIALS
33701462 AND F33700055
Cabin Window Retainers and Seal (Sheet 1 of 2)
3-3
•
1
Detail
A
II
II
II
•
LI
DetailC
16
Detail
F
I
DetailD
Detail
E
BEGINNING WITH AmCRAFT SERIALS
33701463 AND F33700056
1.
2.
3.
4.
5.
6.
7.
8.
Cabin Skin
Doubler
Clip
Headliner
Retainer Strip
Window Trim
Rubber Moulding
Inner Window
Figure 3-2.
3-4
9.
10.
11.
12.
13.
14.
15.
16.
Nut
Screw
Window
3C -200 Sealant
Retainer
Clip Catch
579.6 Presstite Sealer
Trim Panel
Cabin Window Retainers and Seal (Sheet 2 of 2)
17.
18.
19.
20.
21.
22.
23.
24.
Cover
Release Handle
Cotter Pin
Pin
Windshield
Pilots Window
Latch Handle
Catch
•
•
DOUBLER
FLANGE
---BRACKET
A
COWL DECK
SKIN
~--RECEPTACLE
TYPICAL 4 PLACES
Detail A
•
SCUPPER DRAIN
""'--:JIIJ.....:::il--COTTER PIN
'----CLAMP
•
ROUTES INTO CENTER
FUSELAGE BAY FORWARD
OF FIREWALL•
Figure 3-3. Scupper Drain
3-5
NOTE
Rubbing the plastic surface with a dry cloth
will build up an electrostatic charge which
attracts dirt particles and may eventually
cause scratching of the surface. After the
wax has hardened, diSSipate this charge by
rubbing the surface with a slightly damp
chamois. This will also remove the dust
particles which have collected while the
wax is hardening.
f. Minute hairline scratches can often be removed
by rubbing with commercial automobUe body cleaner
or fine-grade rubbing compound. Apply with a soft,
clean, dry cloth or imitation chamois.
3-8. CRACKS. (Refer to figure 3-1.)
a. When a crack appears in a panel, drill a hole
at the end of the crack to prevent further spreading.
The hole should be approximately 1/8 inch in diameter, depending on the length of the crack and
thickness of the material.
b. Temporary repairs to fiat surfaces can be
effected by placing a thin strip of wood over each
side of the surface and then inserting small bolts
through the wood and plastic. A cushion of sheet
rubber or airplane fabric should be placed between
the wood and plastic on both sides.
c. A temporary repair can be made on a curved
surface by placing fabric patches over the affected
areas. Secure the patches with airplane dope, Specification No. MIL-D-5549; or lacquer, Specification
No. MIL-L-7178. Lacquer thinner. Specification
No. MIL-T-6094 can also be used to secure the
patch.
d. A temporary repair can be made by drilling
small holes along both sides of the crack 1/4 to 1/8
inch apart and lacing the edges together with a soft
wire. Fine-strand antenna wire makes a good temporary lacing material. This type of repair is used
as a temporary measure only, and as soon as facUities are available the panel should be repaired.
3-9. WINDSHIELD.
3-10. DESCRIPTION. The windshield is a singlepiece formed, acrylic plastic panel. A center strip
supports the center portion of the Windshield. Thru
aircraft serials 33701462 and F33700055 the windshield was set in felt sealing strips and held in place
by formed retainer strips riveted to the fuselage.
Beginning with aircraft serials 33701463 and F33700056 the felt sealing strips are no longer used, also
the retainer strips are held in place by screws and
nuts for easy removal. Refer to figure 3-2 for sealing.
3-11. HEMOVAL AND INSTALLATION. (THRU
AIRCRAFT SERIAL 33701462 AND F33700055).
Refer to figure 3-2, sheet 1 of 2.
3-6
a. Remove the screws, trim, and attaching parts
at the center strip and top retainer.
b. Drill out rivets securing the retainer strips at
the front of the Windshield, taking care to protect all
items under the instrument deck which may interfere
with safe removal of rivets.
c. Ease Windshield straight forward, out of side
retainer strips.
d. Clean all retainer strips and channels, using
P-S-661 solvent or equivalent.
e. Inspect all retainers and strips for damage, and
repair or replace as necessary.
f. Carefully inspect the felt seal around perimeter
of new windshield for looseness or damage.
g. Apply sealer around perimeter of fuselage opening so that as windshield is pressed into place, a good
seal is formed.
h. Work the windshield straight back into the side
retainers.
i. Install top retainer strip using screws and nuts.
j. Install lower retainer strips around the front
exterior of the windshield. Apply additional sealer
as needed to ensure a positive seal.
•
NOTE
Screws and self-locking nuts may be used to
replace the rivets securing the lower retainer
strips to the cowl deck. These screws should
be at least No.6, and tightened sufficiently to
ensure a good seal.
k. Apply a new sealing tape to centerstrlp and install.
1. Remove excess sealer with a stick of very soft
wood or rolled paper, then wipe clean using P-S-661
solvent or equivalent.
•
3-12. REMOVAL AND INSTALLATION. (BEGINNING WITH AIRCRAFT SERIALS 33701463 AND
F33700056). Refer to figure 3-2, sheet 2 of 2.
a. Remove sun visors and upper Windshield mouling.
b. Remove screws securing upper inside retainer.
c. Remove screws securing outside center strip.
d. Remove screws securing lower outside retainer.
e. Ease windshield forward, at the bottom, out of
the side retainer strips and from under the cabin top
skin.
f. Clean all retainer strips and channels, using a
putty knife.
g. Inspect all retainers for damage and repair or
replace as necessary.
h. Reverse the preceding steps for installation.
i. When installing a new windshield check fit and
carefully file or grind away excess plastic.
j. After installation remove excess sealer from
inside and outside of windshield.
•
•
3-13. WINDOWS.
3-14. FOUL-WEATHER WINDOW.
3-15. DESCRIPTION. Refer to figure 3-4.
3-16. REMOVAL AND INSTALLATION. (Refer to
figure 3-4).
Remove screws (13) from hinges and remove window.
Remove latch handle by removing screw (15).
3-17. EMERliENCY WINDOW.
3-18. DESCRIPTION. Refer to figure 3-5. The
jettisonable emergency window assembly is held in
place by tabs inserted in slots at the top of the window opening, angles at the sides, and drilled studs
keyed with pull-away pins at the bottom.
•
3-19. INSTALLATION. (Refer to figure 3-5).
a. Clean away old sealing compoWld and position
window in opening. Trim metal frame to conform
to fuselage cutout.
b. Apply sealer aroWld fuselage cutout where window contacts lip.
c. Bond 1/4" rOWld rubber seal to window using
EC-880 adhesive (Minnesota Mining and Mfg. Co.).
Break the gloss on the seal with sandpaper, to ensure positive bond.
d. Insert tabs along top of window and press window into place. Bend side tabs aroWld angles, and
insert pull-away pins into drilled studs at bottom of
window.
e. Check emergency release mechanism and reinstall placarded glass.
f. Caulk carefully aroWld window frame from outside using Presstite compound as illustrated. Wipe
away excess.
'
g. Repaint window frame, if necessary, to conform
to aircraft paint scheme.
3-20. FIXED CABIN WINDOWS.
3-21. DESCRIPTION. Thru aircraft serials 33701462 and F33700055 the center and rear windows
are set in rubber channel and sealed with Presstite
compound. They are held in place by retainer strips
secured with rivets and screws. Beginning with aircraft serials 33701463 and F33700056 the center and
rear Windows are sealed with 3C-200 sealant and
held in place by retainers secured to the fuselage
with screws and nuts.
•
3-22. REMOVAL AND INSTALLATION. (Refer to
figure 3-2).
a. Remove screws as necessary and remove decorative trim around windows.
b. Straighten interior window clips and remove interior windows.
c. Remove screws, nuts and/or rivets as necessary
to remove retainers. Then remove window.
d. THRU AIRCRAFT SERIALS 33701462 AND F33700055, apply new rubber seals, and press firmly into
sealing compoWld.
e. BEGINNING WITH AIRCRAFT SERIALS 33701463 AND F33700056, remove old sealer with a
putty knife. Apply new sealer and install window.
f. Reinstall retainers, interior Windows and decorative trim.
g. Seal any gaps aroWld outside of windows with
Presstite compoWld.
3-23. MOVABLE WINDOW. (THRU AIRCRAFT
SERIALS 33701462 AND F33700055). Refer to figure
3-6.
3-24. DESCRIPTION. The movable window, hinged
at the top, is installed in the door. The window assembly, that is the clear plastiC and frame unit, may
be replaced by removing the hinge pins and disconnecting the Window stop. To remove the frame from
the plastic panel, drill out the blind rivets at the
frame splice. When replacing the plastiC panel in a
frame, make sure that the sealing strip and an adequate coating of Presstite No. 579.6 sealing compOWld is used aroWld all edges of the plastic panel.
3-25. CABIN DOOR.
3-26. DESCRIPTION. (Refer to figure 3-6.) THRU
AIRCRAFT SERIALS 33701462 AND F33700055, the
cabin door is installed on the right hand side of the
fuselage, hinged at the forward side and incorporates
a window which may be opened for ventilation while
the aircraft is on the groWld. BEGINNING WITH
AIRCRAFT SERIALS 33701463 AND F33700056, the
cabin door consists of two sections. The upper portion lifts out and up and is held in the open position by
a over-center arm arrangement. The lower portion
folds down and acts as an entrance step. Each section of the door has its own door handle and latching
mechanism. The upper portion may be opened for
ventilation while the aircraft is on the ground.
3-27. REMOVAL AND INSTALLATION. (THRU
AIRCRAFT SERIALS 33701462 AND F33700055).
(Refer to figure 3-6).
a. Remove cotter pin and pin from the door extension arm on the forward door post.
b. Remove interior trim panel forward of the door
post for access to the hing pins.
c. Remove cotter pins from the hinge pins, remove
hinge pins and door.
d. For reinstallation reverse the preceding steps.
3-28. REMOVAL AND INSTALLATION (BEGINNING
WITH AIRCRAFT SERIALS 33701463 AND F33700056).
(Refer to figure 3-6).
a. Remove bolt securing door extension arm to upper
door.
b. Remove screws securing upper hinge and remove
upper door.
c. Remove screws securing lower door stop chains.
d. Remove screws securing lower hinge and remove
lower door.
e. To reinstall reverse the preceding steps.
f. When fitting a new door, some trimming of the
door skin and some reforming with a soft mallet may
be necessary to achieve a good fit.
3-7
•
14
---./'
18
19
APPLY LOCTITE SEALANT,
GRADE C, TO THREADS AT
INSTALLA TION
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Tab
Inner Retainer
Windshield
Center Strip
Support
Outer Retainer
Striker
Seal
Window
Insert
11.
12.
13.
14.
15.
16.
17.
18.
19.
20~
Handle
Roll Pin
Sleeve Nut
Hinge
Bushing
Retainer
Bracket
Trim
Rubber Washer
Temperature Gage
7
10
I
8
9
11
Figure 3-4.
3-8
Windshield and Foul-Weather Window (Sheet 1 of 2)
•
•
NOTE
Refer to figure 3 -2 for
windshield sealing .
•
7
•
1.
2.
3.
4.
5.
6.
7.
Inner Retainer
Nut
Screw
Windshield
Inner Center Strip
Outer Center Strip
Outer Retainer
Figure 3-4.
Windshield and Foul-Weather Window (Sheet 2 of 2)
3-9
2
•
LUBRICATE ALL PINS
WITH PARAFFIN BEFORE
INSTALLING WINOOW
3
•
12
THRU AIRCRAFT SERAII1)
33701462 AND F33700055
BEGINNING WITH AIRCRAFT
SERAII1) 33701463 AND F33700056
Figure 3-5.
3-10
1.
2.
3.
4.
5.
6.
7.
S.
9.
10.
Nuts
Slots
Screws
Pull Pins
Tabs
Bend-Over Tabs
Drilled Pins
Pin
Release Handle
Cotter Pin
11. Spacer
12.
13.
14.
15.
16.
17.
IS.
19.
20.
Emergency Window Installation
Cover
Rubber Moulding
Inner Window
Outer Window
Felt Seal
Window Structure
Rubbl!r Seal
579.6 Presstite Sealer
3C -200 Sealant
•
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Washer
Hinge Pin
Window Arm
Window
Window Hinge
Striker
Nutplate (For Armrest)
Door Stop Spring
Door Arm
Cotter Pin
3_-t-~'\
REFER TO FIGURE 3-7
•
9
Detail
•
Lubricate sliding surfaces of the door stop
mechanism with grease, Specification MILG-21164, applied sparingly if binding occurs.
FiglJre 3-6.
C~.bin
A
Gri nd or file cam surfaces of door stop brackets
to effect a door closing force of 10 - 2 + 4 pounds
measured perpendicular to the door at the trailing
edge in the latch area .
Door Installation (Sheet 1 of 3)
3-11
•
I:
2
•
8
~.~
1.
2.
3.
4.
5.
6.
7.
8.
9.
Nut
Washer
Retainer
Window
Hinge
Hinge Pin
Door Assembly
Pawl
Decal (Hi-Visibility Orange)
Figure 3-6.
3-12
f If
/
~~­
/
Detail
A
Cabin Door Installation (Sheet 2 of 3)
•
::;;'-2 l
i.
'7\ .~.
I
5/~'
\
\
\
\.
\
\
I
•
.
J4
15
13
9
•
11
1.
2.
3.
4.
5.
6.
7.
12
Screw
Washer
Adaptor
Shim
Chain
Spri
Hingr:; (Door Chain)
8. Nut
Figure 3-6.
9. Spacer
~~. Claw Latch
12.• Actuat
Springor Assembly
13. Knob
14.
15 Scr
D ew (Special)
• oor
16. Link
C2.btn Door Installation (Sh eet 3 of 3)
3-13
•
• Apply Loctite, Grade B
upon installation.
• • 337-0116 THRU 337-0239
18
23
.BEGINNING WITH 337-1134
AND F33700001
20
24
• • 337-0001 THRU 337-0115 AND
337-0240 THRU 337-1133
17
"
THRU AmCRAFT
SERIAL 337-0755
1. Top Bolt Guide
2. Bolt
3. Side Bolt Guide
4. Base Bolt Guide
5. Latch Base Plate
6. Side Bolt Guide Assembly
7. Bellcrank
B. Roll Pin
9. Bracket
10. Door Handle Spring
11. Escutcheon
12. Inside Door Handle (Typical)
13. Retaining Clip
14. Placard
15. Washer
16. Bracket
17. Bracket
lB. Shaft Assembly
19. Bolt Push Rod
20. OutSide Handle
21. Pull Bar
22. Rotary Clutch
23. Guide
24. Cover
25. Spring
26. Setscrew
27. Thumb Button
Figure 3-7.
3-14
*337-0116 THRU 337-0239
**337-0001 THRU 337-1133
BEGINNING WITH AIRCRAFT
SERIAL 337-0756 AND F33700001
Cabin Door Latch (Sheet lof 2)
•
•
•
BEGINNING
33701463 ANDWITH
F33700056
23
•
27
25.
26.
27
28..
Lock Link
Act uating Link
Bracket A ssembly
Shaft
29.
30 Turnb uckle
114 .
15
3
1. Sc.....
•
2.
3
••
5.
6.
4
Latch A
T
ssembly
Bracket
orque Tube
PIvot A
PIn
ssembly
16:
7.
8•
9.
10.
12.
Rod
Pu sh Rod
Spacer
Knob
PIate
17.
18.
19 •
20.
21.
234.
=
•
Gear
Ha ellSupport
e
31'• Shaft
Rod End
Door
Ha ndle Outside
32.
A
:!.
(Support)
Washer
Cam
BuShi.;sembIY
• C...er P'
35. Nut
••
.
ye-Bolt
36 E
37. Bearl
::.
40.
Lock
41 . Striker
42.
Support
Torsl::~
Gear
Roll PI.
Shim
~I:I.~ S:p:n.:~ ~-: :~2~2.~lB~e~a~n.~~::::~ ~~';
_____________________
2
ng Su
. Pivot
Bearing
Pin pport
______
FIgure 3-7.
Unk
43. Nut-Bolt
Cabin Door Lat ch (Sheet 2 of 2)
3-15
3-29. CABIN DOOR LATCH. (THRU AIRCRAFT
SERIALS 33701462 AND F33700055).
3-30. DESCRIPTION. The cabin door latch is a
push-pull bolt type, utllizing a rotary clutch for
positive bolt alignment. As the door is closed, teeth
on the underside of the bolt engage the gear teeth on
the clutch, aligning the bolt With the slot. The inside
handle is rotated into the "LOCK" position and the
door is drawn in snugly as the bevel on the bolt slides
into the slot. Beginning with 337-0526 the door latch
is equipped with a bolt lockout. This lockout holds
the bolt retracted unill the door is closed, preventing
damage caused by closing the door with the bolt in
locked position.
3-31. ADJUSTMENT. Vertical adjustment of the
rotary clutch is afforded by slotted holes in the
cover plate. This adjustment ensures sufficient
gear-to-bolt engagement. Refer to Section 2 for
lubrication requirements.
3-32. INDEXING CABIN DOOR HANDLE. (Refer to
figure 3-7.) When the inside handle is removed, it
must be positioned in relation to rotary latch operation upon installation. The following procedure may
be used for indexing the inside cabin door handle:
a. Temporarily install handle (12) apprOximately
vertical.
b. Move handle (12) back and forth until handle
centers in spring-loaded position.
c. Without rotating shaft assembly, remove and
reposition handle to vertical position, if necessary.
d. Mark this pOsition and install escutcheon (11)
80 CLOSE index aligns with mark.
e. Press escutcheon against panel to seat prongs.
f. Install handle to align with CLOSE mark on
escutcheon.
g. Ensure bolt (2) clears doorpost and teeth engage
clutch gear when handle is in CLOSE pOSition.
d. With the door closed and handle (14) in the lock
position observe position of lock link (25) in relation
to actuating link (26) through observation hole on the
rear latch assembly.
e. Bushing should be bottomed out in the radius in
lock link (25) on the latch.
f. If bushing is not bottomed out in the lock link (25),
open door and adjust rod end (30) 1/2 turn at a time
until bushing bottoms out.
•
NOTE
If tolerance cannot be obtained, check for
ware of shaft (16), bushing (22), lock link
(24), actuating link (26), bearing or rod
assembly (3). Replace worn parts.
g. Repeat steps c, d, e, and f for forward latch assembly.
h. After forward latch assembly is adjusted recheck
rear assembly.
i. Remove door handle and reinstall upholstery
panel, trim and arm rest, then reinstall handle.
3-36. INDEXING CABIN DOOR HANDLE. With outside handle flush with door skin, install inside handle,
horizontal, pointing forward with latches fully overcenter.
3-37. BAGGAGE DOOR. (THRU AIRCRAFT SERIALS
33701462 AND F33700055).
3-38. DESCRIPTION. Refer to figure 3-8.
3-39. REMOVAL AND INSTALLATION.
a. Remove screw from door stop arm.
b. Remove bolts securing the door to the hinges and
remove door.
c. To install reverse the preceding steps.
•
NOTE
3-33. CABIN DOOR LATCH. (BEGINNING WITH
AIRCRAFT SERIALS 33701463 AND F33700056).
3-34. DESCRIPTION. The cabin door latch on the
upper section of the door consists of two latch assemblies actuated by push-pull rods connected to a
90 gear assembly, which is rotated by the door
handle. The lower section of the door has two claw
latches which are installed on a spring loaded rod,
rotated by the release knob.
0
3-35. ADJUSTMENT. (Refer to figure 3-7.) Adjustment of the lower door latch is accomplished by the
poSitioning of the e.ccrentic spacers under the latch
plate on the door post, and the number of shims installed under the plate. Adjust the upper door latch
as follows:
a. Remove arm rest, door handle, loosen trim and
remove upholstery panel.
b. Reinstall handle (14) temporarily.
c. With the door closed and the handle in the lock
position check that door skin is flush with cabin skin
and that lock link (25) is snug with striker pin.
3-16
On the bonded baggage door forming of the
flange to align the door with cabin skin
is not recommended as forming of the
flange could cause damage to the bonded
area.
3-40. SEATS.
3-41. INDIVIDUAL SEATS.
a. RECLINING BACK.
b. VERTICAL ADJUSTABLE.
c. ARTICULATING RECLINE/VERTICAL ADJUST.
3-42. DESCRIPTION. These seats are manually
operated throughout their full range of operation.
Seat stops are provided to limit fore-and-aft travel.
3-43. REMOVAL AND INSTALLATION. (Refer to
figure 3-9).
a. Remove seat stops from seat rails.
b. Disengage the seat adjustment pins.
c. Slide seat fore-and-aft to disengage sea:
from rails.
•
•
NOTE
A bonded baggage door is
installed beginning with
33701317 and F33700025.
THRU AmCRAFT
SERIAL 337-0755
3
2
•
t-
*THRU AmCRAFT SERIAL 33701316
AND F33700024
NOTE
Forming of the flange on the
bonded baggage C:loor is not
permissible as forming of
the flange could cause damage
to the bonded area.
15
1. Baggage Door
13
2
•
BEGINNING WITH AmCRAFT SERIAL
337-0756 AND F33700001
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
Outside Handle
Lock
Seal
Swing Stop
Striker Plate
Shim
Hinge Bracket
Pin
Hinge
Cotter Pin
Inside Handle
Pan
Latch Assembly
Cam
Figure 3- 8. Baggage Door
3-17
d. Lift seat out of aircraft.
e. To reinstall, reverse the preceding steps. Ensure all seat stops are installed.
IWARNINGt
It is extremely important that seat stops
are installed. Acceleration and deceleration could possibly permit seat to become
disengaged from the seat rails and create
a hazardous situation, especially during
take-off and landing.
3-44. BENCH SEATS
a. DOUBLE- WIDTH BOTTOM/INDIVIDUAL
RECLINING BACKS.
b. DOUBLE-WIDTH BOTTOM/INDIVIDUAL
RECLINING BACKS, FOLD UP.
3-45. DESCRIPTION.
a. The double width bottom/indiVidual reclining
back seat used thru aircraft serials 33701462 and
F33700055 is secured to the floorboard with four
studs and locks. The seat has no fore-and-aft
adjustment.
b. The double width bottom/individual reclining
back, fold up seat is installed on two seat tracks,
running from the left hand side of the cabin to approximately th~ center. A latch mechanism on the
right hand side of the cabin holds the seat in place.
With the seat backs folded forward, the seat may be
folded up to the left hand side of the cabin for access
to the baggage area.
3-46. REMOVAL AND INSTALLATION. (Refer to
figure 3-9).
a. DOUBLE WIDTH BOTTOM/INDIVIDUAL RECLINING BACK. Removal is accomplished by releasing the four lock clips from the studs and lifting
the seat from the studs.
b. DOUBLE WIDTH BOTTOM/INDIVIDUAL RECLINING BACK FOLD UP SEAT.
1. Fold seat backs forward against seat bottom.
2 . Release seat latch and fold seat up against
the cabin wall.
3. Remove seat stops and pivot rod.
4. Slide seat inboard out of the seat tracks and
remove from the aircraft.
5. Reverse the preceding steps for reinstallation. Be sure to install seat stops on seat tracks.
3-47. POWER SEAT.
3-48. DESCRIPTION. An electric motor, geared to
a screwjack actuator, operates the mechanism which
raises and lowers the seat.
3-49. REMOVAL AND INSTALLATION. (Refer to
figure 3-9).
After disconnecting the electrical leads at the quick
disconnects on the floorboard remove seat in accordance with paragraph 3-43. When installing the seat,
3-18
the wires may be reversed without affecting seat operation. Limit switches are not needed, as actuator,
free-Wheels, at each end of travel.
3-50. QUICK ATTACHING FIFTH AND SIXTH
SEATS.
•
3-51. DESCRIPTDN. (Refer to figure 3-9). This
seat consists of a seat bottom and a seat back which
are installed separate. The seat bottom and back
are held in place with Velcro fasteners. The Velcro
(Hook) is attached to the seat bottom frame and on
the back of the seat back. The Velcro (Pile) is sewn
to the carpet and on the aft cabin wall upholstery. A
strap sewn to the seat back with a snap on the end
connects the seat back to the seat bottom. This seat
has no adjustment.
3-52. REMOVAL AND INSTALLATION. Removal
and installation of this seat is accomplished by lifting
up on the seat bottom and pulling forward on the seat
back. To reinstall set seat bottom in place and place
seat back in place so Velcro Hook and Pile match and
press.
3-53. SEAT REPAIR. Replacement of defective
parts is recomme..Jed in repair of seat mechanism.
The square-tube aluminum framework is 6061 aluminum, heat-treated to a T-6 condition. Except in an
area of stress concentration (close to a hinge or
bearing point), a crack may be heli-arc welded.
Torch welds are not feasible because excessive heat
destroys heat-treatment of the structure. Refer to
figure 3-10 for replacement of seat back cams on
reclining seat backs.
3-54. CABIN UPHOLSTERY.
•
3-55. DESCRIPTION. Due to the wIde selection of
fabrics, styles and colors, it is impossible to depict
each particular type of upholstery. The following
paragraphs describe general procedures which serve
as a guide in removal and replacement of upholstery.
Major work, if possible, should be done by an experienced trim mechanic. If the work must be done by a
mechanic unfamiliar with upholstery practices, the
mechanic should take detailed notes during removal
of each item to facilitate replacement.
3-56. UPHOLSTERY MATERIALS AND TOOLS will
vary with each job. Scissors for trimming and a
dull-bladed putty knife for wedging materials beneath
the retainer strips are the only tools required for
most upholstery work. Industrial rubber cement Is
used to hold soundproofing mats and fabric edges in
place. Refer to Section 16 for thermo-plastic repairs.
3-57. SOUNDPROOFING.
3-58. DESCRIPTION. 337-Series aircraft are insulated With spun glass, mat-type insulation. Two
types of vibration dampening materials are used .
•
•
NOTE
Seat back cam stops are not used
on the copilot's seat since the seat
back reclines fully, resting on a
support bracket on the next seat
aft.
REC LINING BACK
7
PILOT AND COPILOT
•
15
•
1. Rod
2. Screw
3. Washer
4. Spacer
5. Bushing
6. Nut
7. Seat Back
8. Cam
9. Pawl
10. Fore-and-Aft Adjustment Handle
11. Housing
12. Spring
13. Fore-and-Aft Adjustment Pin
14. Roller
15. Recline Handle
Figure 3-9.
12
Seat Installation (Sheet 1 of 13)
3-19
•
NOTE
Seat back cam stops are not
used on tbe copilot's seat
since the seat back reclines
fully, resting on a support
bracket on tbe next seat aft.
Detail
A
•
VERTICAL ADJUST
1. Screw
2. Washer
3. Spacer
4. Bushing
5. Nut
6. Seat Back
7. Recline Handle
8. Pawl
9. Spring
10. Torque Tube
11. Channel
12. Roller
13. Housing
14. Fore-and-Aft Adjustment Pin
15. Bearing Block
16. Vertical Adjustment Handle
17. Fore-and-Aft Adjustment Handle
18. Bellcrank
19. Roll Pin
Figure 3-9.
3-20
Lr.!"""--12
Seat Installation (Sheet 2 of 13)
•
•
1. Recline Handle
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
Spring
Seat Bottom
Recline Actuator
stop
Seat Back
Roll Pin
Channel
Pin
Spacer
Motor and Transmission
Fore-and-Aft Adjustment Pin
Cotter Pin
Fore-and-Aft Adjustment Handle
Vertical Adjustment Switch
Bellcrank
•
8
9
12
•
11
POWER SEAT
Figure 3-9.
Seat Installation (Sheet 3 of 13)
3-21
•
ARTICULA TING REC LINE/
VERTICAL ADJUST
PILOT AND COPILOT SEAT
3
NOTE
Nut on adjustment screw (2) is
rotated 180 0 BEGINNING WITH
AmCRAFT SERIAL 33701399
AND F33700046.
•
Detail
A
9
Detail
1.
2.
3.
4.
5.
6.
7.
8.
9.
·bz
B
3
Articulating Adjustment Handle
Adjustment Screw
Bellcrank
Adjustment Nut
Seat Back
Seat Bottom
Channel
Torque Tube
Seat Structure
10.
11.
12.
13.
14.
.15.
16.
17.
Housing
Bushing
Washer
Roller
Vertical Adjustment Handle
Bearing Block
Fore-and-Aft Adjustment Pin
Fore-and-Aft Adjustment Handle
Figure 3-9.
3-22
Seat Installation (Sheet 4 of 13)
10
•
NOTE
•
Nut on adjustment screw (2) is
rotated 180 0 BEGINNING WITH
AIRCRAFT SERAILS 33701399
AND F33700046.
5
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
Adjustment Pin
Spring
Fore/Aft Adjustment Handle
Seat Bottom
Articulating Adjustment Handle
Bellcrank
Adjustment Screw
Seat Back
Trim Bracket
Spacer
Bracket
Spacer
Seat Structure
Bushing
Roller
, ,
10
I,
,I
I.
;..I
,J
T!~
15
13
14
DetaHA
•
Figure 3-9.
Seat Installation (Sheet 5 of 13)
3-23
.1
ARTICULATING RECLINE/
VERTICAL ADJUST
PILOT AND COPILOT SEAT
•
Detail
A
A
Detail
B
1. Articulating Adjustment Handle
2.
3.
4.
5.
6.
7.
8.
9.
10.
Bearing Block
Adjustment Screw
Bellcrank
Torque Tube
Nut (Screw Assembly)
Seat Back
Seat Bottom
Channel
Seat Structure
11.
12.
13.
14.
15.
16.
Figure 3-9.
3-24
Housing
Roller
Vertical Adjustment Handle
Fore-and-Aft Adjustment Handle
Spring
Adjustment Pin
Seat Installation (Sheet 6 of 13)
•
•
ARTICULA TING REC LINE
COPILOT SEAT
8
•
Detail
A
A
~
•
1.
2.
3.
4.
5.
6.
7.
8.
Articulating Adjustment Handle
Bearing Block
Adjustment Screw
Bellcrank
Torque Tube
Nut (Screw Assembly)
Seat Back
Seat Bottom
.~
"\ f~
12 13
12
9.
10.
11.
12.
13.
14.
15.
Roll Pin
Torque Tube
Bushing
Housing
Roller
Adjustment Pin
Fore-and-Aft Adjustment Handle
Figure 3-9.
Seat Installation (Sheet 7 of 13)
3-25
RECLINING BACK
CENTER SEAT
9
~~-Il
•
23
21
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
Rod
Screw
Washer
Spacer
Bushing
Washer
Washer
Nut
Seat Back
Cam
Pawl
Pin
Shim
~
/'
20
14. Roller
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
Housing
Bracket
Spring
Fore -and -Aft Handle
Cotter Pin
Adjustment Pin
Seat Stop
Retainer
Support
Cover
Recline Handle
Figure 3-9.
3-26
15
14
Seat Installation (Sheet 8 of 13)
•
INFINITE ADJUSTABLE
SEAT BACK
CENTER SEAT
•
•
7
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
Infinite Adjul,;tment Handle
Adjustment Screw
Bellcrank
Adjustment Nut
Seat Back
Seat Bottom
Seat Frame
Spacer
Bushing
Roller
Washer
Housing
Adjustment Pin
Spring
Fore-and-Aft Adjustment Handle
Figure 3-9.
15
J3~~
12'1~T
10
8
9
Seat Installati.op (Sheet 9 of 13)
3-27
•
9
•
14
J
A
9
DOUBLE-WIDTH BOTTOM
INDIVIDUAL RECLINING BACKS
CENTER SEAT
Detail
A
1.
2.
3.
4.
Bolt
Washer
Spacer
Bushing
5. Nut
6. Seat Back
7. Cam Stop
Figure 3-9.
3-28
8. Pawl
9. Recline Handle
10.
1l.
12.
13.
14.
Stop
Nutplate
Stud
Lock
Seat Frame
Seat Installation (Sheet 10 of 13)
•
•
1. Seat Bottom (Upholstery Removed)
Guide Pin
Stud
Lock Nut
Bushing
Bolt
Washer
Nut
Seat Back
10. Spacer
11. Spring
12. Bellcrank
13. Nut (Screw Assembly)
14. Pivot Rod Assembly
15. Seat Track
16. Seat Frame
17. Recline Handle
18. Control
19. Seat Back Release Handle
20. Knob
21. Screw Assembly
2.
3.
4.
5.
6.
7.
8.
9.
9
•
DOUBLE-WIDTH BOTTOM
INDIVIDUAL RECLINING BACKS
FOLD-UP CENTER SEAT
16
1fT
675
15
•
Figure 3-9.
Seat Installation (Sheet 11 of 13)
3-29
•
5
9
•
13
RECLINING BACK
5TH AND 6TH SEAT
1.
2.
3.
4.
5.
6.
7.
Recline Handle
Seat Bottom
Screw
Washer
Spacer
Seat Back
Bushing
Figure 3-9.
3-30
8. Nut
9. Cam
10. Pawl
11. Adjustment Pin
12. Spring
_13, _l:_o_re_-_and-Aft Handle
14. Roller
Seat Installation (Sheet 12 of 13)
•
•
2
5
:i
~
.~.
Detail
•
,
~
A
Rotated 90%
5TH AND 6TH SEAT
1.
2.
3.
4.
5.
6.
•
Seat Bottom
Seat Back
Seat Frame
Velcro Fastener (Hook)
Aft Cabin Wall
Velcro Fastener (Pile)
3
Detail
B
Rotated 90%
Figure 3-9.
Seat Installation (Sheet 13 of 13)
3-31
Q)
CD
CLEVIS BOLT (REF)
SEAT BACK (RE F)
•
2.50" R. (CONSTANT AT EACH NOTCH)
I
//0
REPLACEMENT CAM:
CDPAWL (REF)
1414230-1 (SINGLE
ADJUSTABLE SEAT)
1414230-2 (FULL
WIDTH REAR SEAT)
1414111-5 (VERTICALLY
ADJUSTABLE SEAT)
•
MENT PROCEDURE:
a. Remove seat from aircraft.
b. Remove plastic upholstery panels from aft side of seat back, loosen upholstery retaining
rings and upholstery material as required to expose the rivets retaining the old cam assembly.
c. Drill out existing rivets and insert new cam assembly (2). Position seat back so that pawl (3)
engages first cam slot as shown.
d. Position the cam so each slot bottom aligns with the 2.50" radius as shown.
e. Clamp securely in this position and check travel of cam. Pawl must contact bottom of each cam
slot. Using existing holes in seat frame, drill through new cam and secure with MS20470AD6
rivets.
f.
Reinstall upholstery, upholstery panels
Figure 3-10.
3-32
am
seat.
Reclining Seat Cam Replacement
•
•
The Model 337 uses a brush-on type compound on the
inner surfaces of the baggage and cabin area. The
Model 337A and on use, in addition to the brush-on
type, a sheet type that is held in place with an epoxy
adhesive. Cabin upholstery and carpeting also assist
in reducing noise level.
3-59. CABIN HEADLINER.
3-60. DESCRIPTION. Thru aircraft serials 33701462 and F33700055 the headliner is made of cloth
with channel strips running lengthwise of the headliner. Suspension Wires are cormected to the channel
strips and to the cabin top. Beginning with aircraft
serials 33701463 and F33700056 the headliner is a
four piece moulded headliner with the overhead console installed between the left and right sections.
3-61. REMOVAL AND INSTALLATION. (THRU
AIRCRAFT SERIALS 33701462 AND F33700055.)
(Refer to figure 3-11).
a. Remove overhead console as follow s:
1. Remove fuel selector handles.
2. If an oxygen system is installed, remove oxygen control handle knob.
3. Remove the four console attaching" screws.
4. Pull aft end of console down and securely
hold oxygen pressure gage While unscrewing bezel
from gage.
•
•
I~AUTIONl
Use care in removing pressure gage to avoid
damaging pressure line.
5. Detach console light electrical wires at
quick-discormects, and remove console.
b. If an oxygen system is installed, discormect outlets by lifting caps and using a spanner wrench to unscrew the cap assemblies.
c. Remove dome light lens assemblies by pulling
straight out. They are retained by plug button type
prongs.
d •. Remove fresh air outlets. Refer to Section 13.
e. Remove sunvisors and coat hanger hook.
f. Remove all visible retainers securing headliner.
g. Work headliner free from metal tabs securing
fabric.
h. With zippers open, begin detaching suspension
wires from charmel at front of headliner. Continue
working aft until headliner is free.
L If charmels are to be removed from headliner,
tag for proper installation sequence.
j. Prior to installation, be sure items concealed
by headliner are secure. Use wide cloth tape to
secure loose wires to fuselage. Check all openings
in wing root and seal if necessary. Straighten any
tabs distorted during removal.
k. If soundproofing panels were loosened or removed, cement in place.
1. Insert charmels into new headliner.
m. Attach charmels at rear of cabin with screws
and, working forward, cormect wire hangers suspending headliner from cabin top.
n. Secure forward ends of charmels to bulkhead.
o. With the zippers closed, work around the edges,
securing the headliner with pointed tabs and cement.
Maintain contour without distorting charmels. The
longitudinal beads should be checked periodically for
straightness.
p. Replace trim, console, fresh air ducts and all
other items that were removed. Observe the
"CAUTION" in this paragraph.
3-62. REMOVAL AND INSTALLATION. (BEGINNING WITH AIRCRAFT SERIALS 33701463 AND
33700056). (Refer to figure 3-11).
a. If an oxygen system is installed, remove oxygen
control handle knob.
b. Discormect outlets by lifting caps and using a
sparmer wrench to unscrew the cap assemblies if
installed.
c. Unscrew oxygen gage face if installed.
d. Unscrew fresh air outlets and light assemblies
and remove.
e. Remove cabin flood light lens.
f. Remove sun visors.
g. Remove screws securing console and remove
console.
h. Remove shoulder harnesses.
i. Remove any screws securing headliner under
console.
j. Lift headliner out of the retainer along the outboard edge and remove headliner.
k. To install reverse the preceding steps .
3-63. UPHOLSTERY SIDE PANELS. Removal of
side panels is accomplished by removing the seats
for access, then removing all attaching parts.
Screws secure side panels to the fuselage. The door
panel is attached with automotive type upholstery clips.
A dull putty knife makes an excellent tool to pry these
clips loose. Do not over-tighten sheet metal screws.
Larger screws may be used in enlarged holes if the
added length causes no interference with fuel lines,
electrical wiring or any other components.
3-64. WINDLACE.
3-65. DESCRIPTION. The windlace is primarily a
decorative trim around the door opening on the aircraft thru 1972 models. On 1971 and 1972 models
the windlace is installed on the baggage door opening
and on the lower half of the cabin door opening.
Sheet metal screws or rubber cement may be used to
secure the windlace.
3-66. CARPETING.
3-67. DESCRIPTION. Cabin area and baggage compartment carpeting is held in place by rubber cement,
small sheet metal screws, scuff plates and retaining
strips through 1970 model aircraft. Beginning with
1971 model aircraft Velcro fastening strips are also
used to secure carpeting in the tunnel areas and access plate locations for quick-removal and inspection .
When fitting a new carpet, use the old one as a pattern
for trimming and marking screw holes.
3-33
•
•
BEGINNING WITH AIRCRAFT SERIAL
33701463 AND F33700056
1.
2.
3.
4.
5.
Channels
Suspension Wires
Zippers
Headliner
Retainer Strip
Figure 3-11.
3-34
Cabin Headliner
•
•
_BEGINNING WITH AmCRAFT
SERIALS 33701484 AND F33700064
5
9
3
2
Detail
1
FOURIO'~\~
A
'\~
~I
Detail B
BEGINNING WITH 33701317
AND F33700025
C
Detail
BEGINNING WITH 33701463
AND F33700056
....... -.~•......" ...••..
17
&1
PLAC~
~4
•
...
F
Detail
THRU 33701316
AND F33700024
13
6
13
F
Detail
E
Detail
1. Shoulder Harness
•
2.
3.
4.
5.
Clip
Cover
Screw
Spacer
6. Bolt
7.
8.
9.
10.
11.
Figure 3-12.
Trim Panel
Spacer Assembly
Firewall
Washer
Nut
12.
13.
14.
15.
16.
17.
"j
1t
D
Latch Assembly
Seat Belt
Nut Plate
Bracket
Hook
Spring
Seat Belt and Shoulder Harness Installation
3-35
•
1.
2.
3.
4.
Figure 3-13.
3-6B. SAFETY BELTS.
5. Eye Bolt
6. Floorboard
7. Nut Plate
B. Latch Assembly
Seat Rail
Clamp Half
Washer
Bolt
Cargo Tie Downs
(Refer to figure 3-12).
3-69. DESCRIPTION. Safety belts are installed for
each seat. Replace belts if frayed or cut, if latches
are defective, or stitching is broken. Replace worn
or defective attaching parts.
3-70. SHOULDER HARNESS. (Refer to figure 3-12).
3-71. DESCRIPTION. Individual shoulder harnesses
may be installed for each seat. Component parts are
replaced as outlined in paragraph 3-69.
NOTE
When reinstalling the steps assembly it is
necessary to substitute blind rivets for
standard rivets.
3-77. MAINTENANCE is limited to keeping mounting screws secure and replacing safety pad (2). The
pad may be replaced with coarse grade, fabric backed,
waterproof emery cloth. Cut cloth to size and secure
with waterproof epoxy adhesive.
3-7B.
CARGO PACK.
3-72. CARGO TIE-DOWNS. (Refer to figure 3-13).
3-73. DESCRIPTION. Provisions for cargo tiedowns vary with each seating arrangement. Several
different fittings are mounted into existing nutplates
on the cabin floor. A cargo net is available for use
with the four and five-place seating arrangements.
A hat shelf may be installed thru 1972 models. The
shelf folds up and locks against the rear cabin wall
when not in use.
3-74. OUTSIDE STEP. (Refer to figure 3-14).
3-75. DESCRIPTION. An outside step is available
as optional equipment thru 337-0239 and standard
equipment beginning with 337-0240 thru 33701462
and F33700001 thru F33700055. Beginning With aircraft serials 33701463 and F33700056 the lower portion of the cabin door acts as entrance step.
3-76. REMOVAL AND INSTALLATION. (Refer to
figure 3-14).
a. Drill out rivets securing plate (4) to fuselage
skin.
b. Remove screws securing support structure (3)
to seat rails (5).
c. Reverse the preceding steps for reinstallation.
3-36
•
3-79. DESCRIPTION. The cargo pack is constructed
of glass fiber with a corrugated aluminum floorboard.
It is secured to the bottom of the fuselage with screws
and Rivnuts. A hinged door on the left side of the
pack provides access for loading cargo.
REMOVAL. (Refer to figure 3-15).
a. Remove cabin step from step support arm.
3-BO.
NOTE
When removing cabin step, carefully peel back
safety pad on the step to expose the two screw
heads. Use a small amount of Methyl Ethyl
Ketone applied to the underside of pad to assist
in peeling it back. Protect the adhesive backing of the pad from dirt and other foreign material. When installing the pad, reactivate the
adhesive backing by wiping lightly with a cloth
dampened with Methyl Ethyl Ketone. However,
if appearances is not objectionable, two small
holes may be cut in safety pad on the step to
expose the two screw heads. Replace safety
pad if it is badly worn.
b. Remove screws securing fairing and seal around
step support arm.
•
•
............
........
';".-- ---:')'------
...",
. ::::.......
,
: : :":.: ":~" ..........
:......,.'.'.'.'
""-".-..
\\
"
~
.
. ...... '. :~..
~'" .I.......•:.?):. : : : . .
"-'"
.
.......
:.:....
'\..
..........
.•....••.•.
........
...\.....
..........
...<:......:.
.........~~~~:::.:.....:..........:.:::::::::::>
..•..•
.. . :::.~
....
/
•...:.<:•....
...........~.. . .
..............
".
"<.,/
..../, .. ,•...\ .•..
, .....,
..••.
........ .
.....
/
....:.
'.
14
;:-1'.;.. .,(;;.. "--..~
«:)
....
1. Step
Detail
•
•
2.
3.
4.
5.
A
Safety Pad
Support
Plate
Seat Rail
THRU AIRCRAFT SERAIL 33701462 AND F33700055
Figure 3-14. Outside Step
c. Position a support under the pack and remove
all screwS attaching the pack to the aircraft.
d. Clean sealing compound from aircraft fuselage
with Stoddard solvent.
e. If cargo pack is not to be reinstalled, remove
cowl flap push-pull rod extensions and rerig cowl
flaps in accordance with Section 10.
f. Reinstall button head screw in Rivnuts.
3-81. INSTALLATION. Prior to positioning the
cargo pack under the aircraft, remove old sealant
from pack and inspect all Rivnuts in the bottom of
the fuselage for obstructions. Apply 576.1 Permagum (Presstite Engineering Company) or equivalent
sealant around perimeter of pack where it will contact the aircraft fuselage to seal the pack against
entry of moisture.
a. Remove step as outlined in paragraph 3-80.
b. Move the pack into position under the aircraft.
Raise the aft end of the pack threading step support
through opening in right side of pack and place support under pack.
c. Raise the forward end of the pack and align the
two forward holes in the pack rim with the two front
Rivnuts. Install two screws to support the forward
end of the pack.
NOTE
Install lock washers and flat washers under
the heads of all pack attaching screws.
d. Raise the aft end of the pack and install two
attaching screws.
e. Check pack for proper alignment, then install
and tighten all pack attaching screws.
f. Position the rubber seal and fairing around step
support tube. Check alignment and proper fit of
fairing, then install fairing retaining screws.
g. Install step. See note in paragraph 3-80.
h. Install front engine cowl flap actuator rods in
accordance with paragraph 3-82 on the non-turbocharged aircraft only.
3-82. RIGGING FRONT COWL FLAPS WITH CARGO
PACK (NON-TURBOCHARGED).
NOTE
When the cargo pack is installed on the nonturbocharged aircraft, . the standard front
engine cowl flap rods are replaced with
longer rods for additional engine cooling.
a. Remove front cowl flap lower push-pull rods by
disconnecting at torque tube arm and at cowl flap.
b. Connect longer push-pull rods to the torque tube
arms.
c. With master switch on and lower end of pushpull rod disconnected, place left cowl flap lever in
"CLOSED" poSition. Allow cowl flap motor to operate to the closed position and turn master switch off.
3-37
......
Detail
•
..................
....:::.
"
A
'0_.
THRU 33701462 AND F33700055
......•.
,
,
"
,.
,..-q'.. )
.-'.:!
( .........,.
....
~
...
'\
(~r~"··--~·
'.
NOTE
When the cargo pack is installed,
standard front engine cowl flap
rods are replaced with longer
rods except on turbocharged
aircraft.
1..-_--6
7
1.
2.
3.
4.
5.
6.
7.
8.
9.
Outside Step
Step Support Tube
Cargo Pack Structure
Escutcheon
Seal
Seal
Quick-release Fastener
Access Door
Door Lock
Detail
•
B
.THRU 33701462 AND F33700055
Figure 3-15.
Cargo Pack Installation
d. Connect push-pull rod to cowl flaps, but do not
secure at this time.
e. Measure the distance from trailing edge of cowl
flaps to cowl skin. The cowl flaps should be open
1. 75 inches with the front cowl flap indicator in the
closed position. This opening is measured at the cowl
flap trailing edge, perpendicular to the cowl fla~ contour. Be sure that control rod ends have sufficIent
thread engagement, then tighten rod end jam nuts.
f. Operate cowl flaps several times to check cowl
3-38
flap operation.
NOTE
Refer to Section 10 for rigging of the complete cowl flap system. For rigging of the
front cowl flaps on the turbocharged aircraft with or without cargo pack, refer to
Section lOA.
•
•
SECTION 4
WINGS, BOOMS, AND EMPENNAGE
TABLE OF CONTENTS
Wings . • • • . . . .
Description • . • •
Removal at.d Installation
Wing Struts • . • • • • •
Description • • • • . •
Removal and Installation
Booms • . • . • • • • . •
Description • • • • . . •
Removal and Installation • • • •
Empennage
Description • . • • • • •
Page
4-1
4-1
4-1
4-1
• 4-1
4-1
4-2
4-2
4-2
4-2
4-2
4-2. DESCRlPl'ION. The wings are all-metal externallift strut braced wing panels of semi-monocoque
construction. Wing structure consists of a forward
and rear spar, ribs for attachment of the skin and a
integral boom support structure.
4-3. REMOVAL AND INSTALLATION.
a. Remove the wing to fuselage, strut and aft boom
fairings for the wing being removed.
b. Drain all fuel from aircraft.
c. Remove empennage as a unit from tail booms .
. (See paragraph 4-12.)
d. Remove tall boom from wing being removed.
(See paragraph 4-9.)
e. Disconnect and remove parts of fuel selector
valve control as necessary to remove Wing. Refer
to Section 11 for details of selector valve control.
f. Disconnect aileron carry-thru cable at turnbuckle above cabin headliner and pull cable into wing
root area.
g. Disconnect applicable flap cables from actuator
above cabin headliner rear access opening. Remove
cable guards and pulleys as necessary to pull cables
into wing root area.
NOTE
It is recommended to secure flap in streamlined position with tape during wing removal
to prevent damage, since flap wlll swing freely.
h. Disconnect flexible hoses, plumbing, electrical
wiring, and any other items that would interfere with
wing removal, at or near the wing root area.
1. Support opposlte wing as a safety precaution.
Support wing being removed.
•
4-2
4-2
4-2
4-2
4-2
4-2
4-2
4-2
4-2
4-2
wing stand.
4-1. WINGS.
•
Removal and Installation
Vertical Fin • • • . • . •
Description . . . • . •
Removal and Installation
Horizontal Stabilizer . . .
Descritpion • • • • . • •
Removal and Installation
Mooring Rings . . • • . •
Description . . • • • •
Removal and Installation
If cables routing through wing strut were not
pulled from wing during boom removal, do
so before detaching strut. Refer to paragraph
4-6 and figure 4-2 for wing strut removal.
j. Detach wing strut from wing being removed.
k. Detach wing from fuselage and place on padded
NOTE
Figure 16-5 illustrates wing and fuselage
support stands which can be manufactured
locally of any suitable wood.
1. Reverse the preceding steps to install the Wing.
Rig fuel selector valve control and flight control
systems in accordance with procedures outlined in
the applicable Sections of this Service Manual.
NOTE
Torque wing-to-boom and boom-to-empennage
attaching screws to values shown in figure 4-3.
m. Refuel aircraft and check for leaks. Check operation of all systems and equipment that may have
been affected by wing removal.
4-4. WING STRUTS.
4-5. DESCRIPTION. Each wing has a single lift
strut which transmits a part of the wing load to the
lower portion of the fuselage. The strut consists of
an aluminum "I" shaped extrusion with forged fittings
at each end for attachment to the wing and lower fuselage. Cable guides are attached to the front and rear
of the strut for control cable routing. Each strut
assembly is covered with a elliptical shaped fairing
and a cup fairing at each end.
4-6. REMOVAL AND INSTALLATION.
a. Remove screws from lower strut cup fairing.
b. Pull upper strut cup fairing from recess in
boom support.
c. When removing left hand strut disconnect pitot
line, also pitot heater wiring if installed.
d. Remove screws from strut fairing and remove
fairing.
e. Disconnect control cables at turnbuckles and
pull cables out of cable guides.
f. Place support under wing and remove upper and
lower attaching bolts. Then remove strut.
g. To install reverse this procedure. Rig in accordance with Sections 6, 7, 8 and 9.
4-1
4-7. BOOMS.
4-8. DESCRIPTION. Tall booms are elliptical in
section and constructed of formed bulkheads, extruded stringers and stressed skins. Cables, electrical wiring, and plumbing for various equipment is
routed through the boom structure. (See figure 4-3.)
4-9. REMOVAL AND INSTALLATION.
a. Remove empennage from booms as a unit. (See
paragraph 4-12.)
b. If right boom is being removed, disconnect flap/
elevator trim interconnect at trim cable and at clamps
inside the boom. Also disconnect the rudder cable.
c. If left boom is being removed, disconnect elevator and rudder cables and any other items that would
interfere with boom removal.
d. Remove aft boom-,to-wing fairings.
e. Support boom and remove attaching screws. Pull
boom aft and work cables and electrical wiring out of
the boom.
f. Reverse this procedure to install booms. Refer
to figure 4-3 for torque values for boom attachment
screws. Rig control systems as necessary. Refer
to Sections 8 and 9.
4-10. EMPENNAGE.
4-11. DESCRIPTION. The empennage is of conventional aluminum all metal design consisting of a horizontal stabilizer, elevator, dual ventral fin and dual
fin and rudder.
4-12. REMOVAL AND INSTALLATION.
NOTE
The empennage should always be removed
from the tail booms instead of removing
the booms with the empennage attached to
them, because of the possibility of twisting or otherwise distorting the stabilizer.
a. Remove stabilizer fairings, also fin, boom and
stabilizer cover plates as needed for access to control cables.
b. Disconnect elevator trim tab cables at turnbuckles in the right boom.
c. Release tension on rudder cables at the turnbuckle located in the stabilizer, then disconnect
rudder cables at the bellcranks inside each end of
the stabilizer.
d. Remove cable guards and pulleys as necessary
to pull cables forward of boom-to-empennage junction.
e. Unhook elevator downspring in left vertical fin
(thru Model 337-0755). Release tension on the elevator cables at the turnbuckles located in the left hand
wing strut. Then disconnect the cables from the bellcrank located in the left fin.
f. Remove cable guards and pulleys as necessary to
pull the cables forward of the boom-to-empennage
junction.
g. Disconnect all electrical wires routed through
boom to empennage.
h. Check for and disconnect any other items that
would interfere with empennage removal.
i. Support empennage, remove attaChing screws,
4-2
and pull empennage aft to remove.
j. Reverse this procedure to install the empennage.
Torque boom-to-empennage attaching screws to
values shown in figure 4-3.
k. Rig control surfaces as necessary. Refer to
Sections 8 and 9.
1. Check operation of flashing beacon and tail
navigation lights.
•
4-13. VERTICAL FIN.
4-14. DESCRIPTION. The fins are of all-metal
construction consisting of a forward and rear spar
with ribs for attachment of the skin and rudder attachment brackets. The left fin houses the elevator
bellcrank. A elevator balance weight is located in
each fin. (See figure 4-4.)
4-15. REMOVAL AND INSTALLATION.
a. Remove the empennage in accordance with paragraph 4-12.
b. Remove the rudder in accordance with Section 9.
c. Remove the elevator in accordance With Section
8.
d. If right fin is being removed, pull elevator trim
cables aft into area between fin and stabilizer.
e. Remove pulleys and cable guards as necessary.
f. Check for and disconnect any other items that
would interfere with fin removal.
g. Support fin and remove forward and rear spar
attaching bolts, then pull the fin outboard to remove.
h. Reverse this procedure to install the fins.
i. Rig control systems as necessary. (Refer to
Sections 8 and 9. )
4-16. HORIZONTAL STABILIZER.
•
4-17 . DESCRIPTION. The horizontal stabilizer is
of all-metal construction, consisting of a forward
and rear spar with ribs for the attachment of the skin.
The elevator trim tab actuator and both rudder bellcranks are located inside the stabilizer. (Refer to
figure 4-5.)
4-18. REMOVAL AND INSTALLATION.
a. Remove the empennage in accordance with paragraph 4-12.
b. Remove the elevator and rudder in accordance
with Sections 8 and 9.
c. Remove the vertical fin in accordance with paragraph 4-15.
d. To install the stabilizer reverse this procedure
and rig in accordance With Sections 8 and 9.
4-19. MOORING RINGS.
4-20. DESCRIPTION. Thru aircraft serial 337-0755
eye-bolt type mooring rings were installed in both
wings and booms. Beginning with aircraft serial
337-0756, retractable mooring rings are installed in
the wings and booms.
4-21. REMOVAL AND INSTALLATION. Refer to
figure 4-6 for removal and installation.
•
•
2
REFER TO SECTION 6
•
12
* 1968 MODEL 337C-SERIES & ON
NOTE
Recessed areas around rivet heads on
the leading edge of the wing are filled
with Bonttte No. BTO-20, or equivalent
compound.
REAR SPAR ATTACHMENT
FRONT SPAR ATTACHMENT
•
1.
2.
3.
4.
5.
6.
7.
8.
Wing Tip
Navigation Light
Fuel Filler Access Door
Fuel Transmitter Access Door
Landing and Taxi Lights
Stall Strip
Main Fuel Tank Cover
Auxiliary Fuel Tank Cover
9.
10.
11.
12.
13.
14.
15.
16.
Bolt
Washer
Nut
Inboard Flap
Boom Support Structure
Outboard Flap
Aileron Trim Tab
Aileron
Figure 4-1. Wing
4-3
THRU AmCRAFT SERIALS 33701462 AND F33700055
VIEW LOOKING OOWN
RIGHT STRUT
HOLE FOR
RUDDER CABLE
•
HOLE FOR
ELEVATOR
TAB UP
CABLE
1.
2.
3.
4.
5.
6.
7.
8.
Upper Fairing
Fairing
Nut
Washer
Bolt
Strut
Fairlead
Lower Fairing
•
LE FT WING STRUT
VIEW LOOKING DOWN
LEFT STRUT
HOLE FOR
RUDDER CABLE
NOTE
Before installing a strut, slide upper
and lower fairings on strut. However,
the fairings may be removed and re- .
placed without removing the strut, since
they are split, then riveted to a doubler
and sealed with Presstite No. 579.6
sealer.
Figure 4-2.
4-4
Wing Strut (Sheet 1 0(2)
•
•
VIEW LOOKING DOWN
mGHTSTRUT
HOLE FOR
ELEVATOR
TAB UP
CABLE
HOLE FOR
RUDDER CABLE
NOTE
Before installing Ii strut, slide upper
and lower fairing on strut. However,
the fairing may be removed and replaced without removing the strut, since
they are split, then riveted to a doubler
and sealed with Presstite No. 579. 6
sealer.
HOLE FOR
ELEVATOR
TAB DOWN
CABLE
BEGINNING WITH AIRCRAFT SERIALS
33701463 AND F33700056
LEFT WING STRUT
8
•
VIEW LOOKING DOWN
LEFT STRUT
•
1.
2.
3.
4.
5.
6.
7.
8.
Upper Fairing
Fairing
Nut
Washer
Bolt
Strut
Fairlead
Lower Fairing
HOLE FOR
ELEVATOR
UP CABLE
Fig'.l!"'" 4-2.
Wing Strllt (Sheet 2 of 2)
4-5
B
•
INBOARD SIDE OF RIGHT BOOM
..
"
..
. . ...
2
.. .
.
'
•
3
Detail
B
7
NOTE
Detail
1.
2.
3.
4.
5.
6.
7.
*
A
Tail Boom
Cover Plate
Nutplate
Screw
Empennage
Wing Boom Support
Fairing
Figure 4 -3.
4-6
Tail Booms
Torque Boom attaching screws
to (90 TO 95 LB-IN).
•
•
4
OUTBOARD SIDE OF LEFT FIN
Detail
Detail
•
A
B
8
Detail
1.
2.
3.
4.
5.
6.
Upper Tip
Fin
Lower Tip
Upper Bearing
Middle Bearing
Lower Bearing
C
DUMMY NA VIGA TION LIGHT (LEFT TIP)
NAVIGATION LIGHT (RIGHT TIP ONLY)
Figure 4-4.
Vertical Fin
4-7
•
•
c
Detail
Detail
1.
2.
3.
4.
5.
6.
7.
8.
9.
C
Nut Plate
Fin Structure
Stabilizer
Bolt
Washer
Elevator Hinge
Bearing
Nut
Fitting (Fin)
Figure 4-5 .
4-8
B
Horlzontal
.
Stabilizer
•
•
BEGINNING WITH 1968 MODEL 337C-SERIES
I
LEFT HAND WING
LEFT BOOM
9
10
4
~
_ _ _- 2
2
THRU 33701194 AND F33700009
BEGINNING WITH 33701195, F33700010
AND ALL SERVICE PARTS
2
•
1. Doubler
2. Mooring Ring
3. Lower Spar Cap
4. Spring
5. Spacer
6.
7.
8.
9.
10.
Inner Bracket
Outer Bracket
Web Stiffener
Bracket
Cable Guide
4
Figure 4-6. Retractable Mooring Rings
SHOP NOTES:
•
4-9/(4-10 blank)
SECTION 5
•
LANDING GEAR, BRAKES AND HYDRAULIC SYSTEM
TABLE OF CONTENTS
•
•
LANDING GEAR SYSTEM (Thru 33701398
and F33700045) . . .
Description . . .
Operation . . . . .
Main Gear System .
Description . .
Trouble Shooting
Strut Removal .
Strut Installation
Main Gear Actuator.
DeSCription . .
Removal
Disassembl y. .
Inspection of Parts
..
••
Parts Repair/Replacement
Assembly.
Installation . . . . . . .
Linkage . . . . . . . . . . .
Description . . . . . . .
Universal Joint and Adapter
Removal . . . . . . .
Installation of Removed Universal
Joints and Adapters . . . .
Installation of New Universal
Joints and Adapters . . . .
Saddle and Pivot Shaft Removal.
Saddle and Pivot Shaft Installation
Uplock Mechanism . . . . . . . .
Description . . . . . . . . .
Operation . . . . . . . . . .
Removal of Uplock Mechanism
and Release Actuator . . . .
Release Actuator Disassembly .
Inspection of Parts . . . . . .
Assembly . . . • . . . . . .
Installation of Uplock Mechanism
and Release Actuator
.
Downlock Mechanism . . . . . .
Description . . . . . . . .
Operation . . . . . . . • .
Removal of Downlock Mechanism
and Downlock Actuator . . . .
Disassembly, Inspection of Parts and
assembly of Downlock Actuator. .
Installation of Downlock Mechanism
and Downlock Actuator
. . .
Main Gear Door System. . . . . . .
Description . . . . . . . . . .
Operation . . . . . . . . . . .
Removal of Main Gear Wheel Doors
and Actuators . . . . . . . .
Disassembly of Main Gear Wheel
Door Actuator (Thru 33701426
and F33700035) .
Inspection of Parts
Assembly . . . . . . . . . .
Page
5-5
5-5
5-5
5-6
5-6
5-6
5-7
5-7
5-9
5-9
5-9
5-9
5-11
5-14
5-14
5-14
5-15
5-15
5-15
5-15
5-15
5-16
5-16
5-16
5-16
5-19
5-19
5-20
5-20
5-20
5-20
5-20
5-22
5-22
5-22
5-22
5-22
5-22
5-22
5 22
5-22
5-23
5-24
5-24
Disassembly of Main Gear Wheel
Door Actuator (Beginning with
5-25
33701427 and F33700036) . .
5-25
Inspection of Parts . . . . . .
5-25
Assembly . . . . . . . . . •
Installation of Main Gear Wheel
5-25
Doors and Actuator . . • • .
Removal of Main Gear Strut Doors
and Actuator. . . . . . . . . . 5- 26
Disassembly. Inspection and Assembly
of Strut Door Actuator. . . .
5-27
Installation of Main Gear Strut
Doors and Actuator. .
5-27
Main Gear Wheels and Axles
5-27
Description . . . . . .
5-27
Operation . . . . . . . .
5-27
5-27
Main Gear Wheel Removal. . .
5-27
Main Gear Wheel Disassembly.
Main Gear Wheel Inspection and
Repair . . . . . . • . . .
5-27
Main Gear Wheel Assembly . .
5-29
Main Gear Wheel Installation
5-29
Main Gear Wheel Axle Removal
5-29
Main Gear Wheel Axle
Installation . . . . . .
5- 29
Main Gear Wheel Alignment
5- 29
Wheel Balancing
5-29
Brake System . . .
5-32
Description . .
5-32
Operation . . .
5-32
Trouble Shooting
......
5- 32
Brake Master Cylinder Removal . . 5-33
Brake Master Cylinder Disassembly 5-33
Inspection of Parts . . . . . . . . 5-33
Brake Master Cylinder Assembly . . 5-33
Brake Master Cylinder Installation. 5-33
Bleeding Brake System . .
5- 35
Wheel Brake Removal. . .
5-35
5-35
Wheel Brake Disassembly.
Wheel Brake Installation .
5-35
Brake Lining Replacement.
5-35
Parking Brake System . . . . .
5- 36
Description (Prior to 337-0240)
5-36
Operation . . . . . . . . . .
5-36
Removal . . . . . . . . . .
5-36
Installation . . . . . . . . .
5- 36
Description (337-0240 thru 3370931) . .
5-36
Operation .
. .. .
5-36
Removal . . . . . . . . . . .
5-36
Installation . . . . . . . . .
5-36
Rigging . . . . . . . . . . .
5-36
Description (Beginning with 3370932) .
5-36
Operation
5-36
Removal
5-36
5-1
•
Installation
Rigging
Nose Gear System
Description .
Operation
.
Trouble Shooting . .
.
Removal of Shock Strut and
Trunnion
..
Removal and Installation of
Trunnion
..
Removal and Disassembly of
Lower Strut
..
Removal and Installation
Locking Collar.
.
Assembly and Installation of
Lower Strut
Nose Gear Shimmy Dampener
Description
Operation
Removal
.
Disassembly .
Inspection of Parts
Assembly
Installation
Torque Links.
Description
Removal
Installation
Nose Gear Uplock Mechanism
Description
Operation
Removal.
Disassembly, Inspection and
Assembly of Nose Gear
Uplock Actuator
Installation
Nose Gear Downlock Mechanism
Description
Operation
Removal.
Disassembly of Nose Gear Actuator
(Thru 33701426 and F33700035).
Inspection of Parts .
Assembly
Disassembly of Nose Gear Actuator
(Beginning with 33701427 and
and F33700036) .
Inspection of Parts .
Assembly
Installation
Nose Gear Door System .
Description
Operation
.
Removal of Aft Nose Gear Door
Installation of Aft Nose Gear Door
Removal of Forward Doors and
5-2
5-36
5-39
5-39
5-39
5-39
5-39
5-41
5-41
5-41
5-41
5-43
5-44
5-44
5-44
5-44
5-44
5-44
5-44
5-44
5-44
5-44
5-44
5-44
5-46
5-46
5-46
5-46
5-46
5-47
5-48
5-48
5-48
5-48
5-48
5-49
5-50
5-50
5-50
5-50
5-50
5-52
5-52
5-52
5-52
5-52
Actuator.
..
•
..
Disassembly, Inspection of Parts
and Assembly of Nose Gear Door
Actuator
.. •
Installation of Forward Doors and
Actuator
Nose Wheel Steering System .
Description
Operation
Trouble Shooting .
Removal of Nose Wheel Steering Cam.
Installation of Nose Wheel Steering
Cam
Nose Gear Wheel .
Description
.
Operation
.
Wheel Removal .
Wheel Disassembly.
Inspection .
Assembly of Wheel .
Installation
.
Landing Gear Hydraulic Power.
Description
•
Operation
Hydraulic Tools and Equipment
Hydro Test Unit.
•
Operation .
Flow Regulation
Connecting Test Unit to Aircraft
Disconnecting Test Stand from
Aircraft.
Bleeding Aircraft HydrauliC
System
Use of Test Stand to Leak Test
Hydraulic System and
Components
Cycling Landing Gear .
Checking Landing Gear Cycle
Time
Hydro Fill Unit
.
Installation of HydrauliC Fittings.
HydrauliC System Components .
General Description
Hydraulic Component Repair.
Repair Versus Replacement.
Repair Parts and Equipment .
Equipment and Tools
Hand Tools.
Compressed Air
Engine-Driven Hydraulic Pump
Description
Operation
Removal
Trouble Shooting
Disassembly.
5-52
5-52
5-52
5-52
5-52
5-52
5-52
5-52
5-55
5-55
5-55
5-55
5-55
5-55
5-55
5-55
5-55
5-57
5-57
5-57
5-57
5-57
5-58
5-58
5-58
•
5-59
5-59
5-59
5-59
5-62
5-62
5-62
5-62
5-62
5-63
5-63
5-63
5-63
5-63
5-63
5-63
5-63
5-63
5-63
5-64
5-64
•
•
•
Inspection of Pump
Assembly
.,
Installation
Hydraulic Fluid Filter
Description .
Operation
.
.
Removal
•..
Disassembly (Thru 33701339 and
F337000023)
..
Inspection of Parts . .
Assembly (Thru 33701330 and
F33700023)
. • . .
Disassembly (33701331 and
F33700024 thru 33701462
and F33700055). • . . •
Inspection of Parts.
.
Assembly (33701331 and F33700024 thru 33701462 and
F33700055)
Assembly
Installation
•
Hydraulic Power Pack
Description
Operation
Removal.
Trouble Shooting
Disassembly.
Manifold Disassembly
Disassembly of Components
Secondary Relief Valve
(Prior to 1968 Models)
Primary Relief Valve.
Priority Valve .
System Inlet Check Valve
Standpipe and Filter
Door Vent Valve
.
Landing Gear Handle-Release
Mechanism
..
Assembly of Power Pack
Door Vent Valve
Standpipe and Filter
System Inlet Check Valve
Priority Valve .
Primary Relief Valve .
Secondary Relief Valve (Prior
to 1968 Models)
Assembly of Manifold. •
Landing Gear Handle-Release
Mechanism
..
Installation of Manifold .
Bench-Testing the Power Pack.
Pressure Adjustments
Test Equipment
Connecting Test Unit
Handle-Release Mechanism
5-64
5-65
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-68
5-71
5-72
5-76
5-76
5-76
5-76
5-76
5-76
5-76
5-76
5-76
5-77
5-77
5-77
5-77
5-77
5-77
5-77
5-79
5-79
5-79
5-80
5-80
5-80
5-81
5-81
Secondary Relief Valve (Prior
to 1968 Models)
•
Primary Relief Valve .
Priority Valve .
Door Vent Valve .
Reservoir Leakage Test.
Installation of Power Pack.
Field- Testing the Power Pack
(Installed in Aircraft). .
Primary and Secondary Relief
Adjustment
.
Adjustment of Priority Valve.
Handle- Release Adjustment
Checking Handle-Release to
Neutral .
.
Checking Time-Delay Valve
Checking Priority Valve.
Checking Primary (System)
Relief Valve .
Checking Secondary (Hand
Pump) Relief Valve (Prior
to 1968 Models)
Checking for Suction Air
Leakage..
. .
Bleeding Time - Delay Valve.
Emergency Hand Pump
Description
Removal.
Disassembly.
Trouble Shooting
Inspection of Parts
Assembly
Installation .
Bleeding .
Door Close Lock Valve
Description
Removal.
Disassemble.
Inspection of Parts
Assembly
Installation
Landing Gear Electrical Circuits.
Description
.
Switch Adjustm ent
.•
Rigging of Main Landing Gear
Rigging of Adjusting Support .
Rigging of Downlock Mechanism
Rigging of Uplock Mechanism
Rigging of Up Indicator Switches
Rigging of Down Indicator
Switches
Rigging of Doors.
•
5-82
5-82
5-82
5-83
5-83
5-84
5-84
5-84
5-84
5-84
5-85
5-85
5-85
5-86
5-86
5-86
5-86
5-86
5-86
5-87
5-87
5-88
5-88
5-88
5-88
5-88
5-88
5-88
5-88
5-90
5-90
5-90
5-90
5-90
5-90
5-92
5-92
5-92
5-93
5-94
5-99
5-99
5-99
5-3
•
Adjustment of Snubber Valve.
5-99
Rigging of Nose Gear.
5-99
Rigging of Downlock Mechanism
5-99
Rigging of Uplock Mechanism
5-99
Rigging of Down Indicator Switch. 5-99
Rigging of Up Indicator Switch
.5-101
Rigging of Safety Switch.
.5-lOl
Rigging of Doors .
.5-101
Rigging of Power Pack Switch and
Lockout Solenoid .
.5-101
Rigging of Up-Down Switch
.5-101
Rigging of Gear Handle Lockout .5-101
Hydraulic System Schematics
.5-lO2
LANDING GEAR SYSTEM (Beginning with
33701399 and F33700046
5-145
Description
5-145
Operation
5-145
Main Gear System
5-145
Description
5-145
5-146
Trouble Shooting
Strut Removal and Installation
5-148
5-148
Main Landing Gear Actuator .
5-148
Description
Removal.
5-148
5-149
Disassembly .
5-149
Inspection of Parts .
5-150
Parts Repair/Replacement
5-150
Assembly
5-150
Installation
5-150
Linkage .
5-150
Description
5-150
Saddle and Pivot Shaft Removal.
Saddle and Pivot Shaft Installation 5-150
5-151
Uplock System .
5-151
Downlock System .
5-151
Gear Door System
5-151
Wheel Door Actuator
5-151
Wheels and Tires
Wheel and Axle Removal and
5-152
Installation
5-152
Wheel Alignment .
5-152
Wheel Balancing
5-152
Brake System
5-152
Parking Brake System
5-152
Nose Gear System
5-152
Nose Gear Assembly
5-152
Nose Gear Strut
5-152
Shimmy Dampener
5-152
Torque Links
5-152
Uplock Mechanism
5-152
Downlock Mechanism
5-152
Nose Gear Actuator
Removal and Installation of Nose
5-4
Gear Uplock and Release Act,1ator 5-152
Disassembly, Inspection and
Assembly
.
5-152
Nose Gear Door System .
.
5-152
Nose Wheel Door Removal and
Installation
5-152
Nose Wheel Steering System.
5-152
Nose Gear Wheel .
5-152
Landing Gear Hydraulic Power
5-152
Hydraulic Tools and Equipment
5-152
Hydraulic Power System Components.
5-152
General Description
5-152
Hydraulic Component Repair.
5-152
Repair Versus Replacement .
5-152
Repair Parts and Equipment .
5-152
Equipment and Tools
5-152
Hand Tools.
.
5-153
Compressed Air
5-153
Power Pack
5-153
Description
5-153
Removal.
5-153
Disassembly .
5-154
Inspection .
5-154
Assembly
5-154
Installation
5-156
Pressure Switch
5-156
Adjustment
5-156
Gear Manifold Assembly
5-156
Disassembly.
5-156
Inspection .
5-156
Assembly
5-156
Door Manifold Assembly
5-158
Disassembly .
5-158
Inspection .
5-158
Assembly
5-158
Emergency Hand Pump
5-158
Description
5-158
Removal
5-158
Disassembly .
5-158
Inspection .
5-159
Assembly
5-159
Installation
5-160
Rigging Main Landing Gear
5-160
Rigging Adjusting Support
5-160
Rigging Downlock Mechanism
5-161
5-162
Rigging Uplock Mechanism
Rigging Up Indicator Switches
5-166
5-166
Rigging Down Indicator Switches
Rigging Doors
5-166
Adjustment of Snubber Valves
5-166
Rigging of Nose Gear.
5-166
Hydraulic and Electrical System Schematics
.
.
•
•
NOTE
•
This Section is divided into two parts. Part 1 covers the landing
gear system for aircraft through Serial No. 33701398 and F33700045. Part 2 covers the landing gear system for aircraft beginning with Serial No. 33701399 and F33700046. Part 1 contains information which is also applicable to aircraft described
in Part 2. To avoid repetition of information, the reader is referred back to this information in Part 1. A separate set of hydraulic schematic diagrams is provided for aircraft described
in each Part of this Section. These diagrams may be found at
the end of each part of this Section.
PART 1
(THRU SERIALS 33701398 AND F33700045)
5-1. LANDING GEAR SYSTEM.
•
5-2. DESCRIPTION. A hydraulically-operated, tricycle retractable landing gear is employed on the aircraft. The hydraulic power system includes equipment required to provide a flow of pressurized hydraulic fluid to the landing gear system. Main components of the hydraulic system include the enginedriven hydraulic pump, located on the right rear
accessory pad of the front engine; the hydraulic filter,
located in the pump pressure line, at the forward side
of the front firewall; the hydraulic power pack, located in the cabin on the aft left side of the front firewall, behind the instrument panel; and the emergency
hand pump, mounted on a support beneath the floorboard, immediately in front of the pilot and copilot
seats, on the aircraft centerline.
NOTE
5-3. OPERATION.
NOTE
Refer to the hydraulic schematic diagrams to
trace the flow of hydraulic fluid as outlined
in the following steps.
.'
poSition selected by the landing gear control lever.
e. When the gear has moved to the full-up or fulldown pOsition, the uplock or downlock switches are
actuated, causing the solenoid-operated door control
valve to move to the door-closed position. Then the
fluid flows through the valve to close the doors.
f. After the doors are closed, pressure builds up
in the system until the 3 to 9-second time-delay
valve, operated by pressure from the door-close
line, opens and permits fluid to flow to the handlerelease valve, returning the handle to neutral.
g. As the gear control handle returns to neutral,
it moves the gear selector spool in the power pack,
which again permits fluid to circulate freely through
the pump, into the power pack manifold, and back to
the reservoir.
a. Filtered hydrauliC fluid from the engine-driven
hydraulic pump enters the power pack, where a passage connects to the primary relief valve. With
landing gear control handle in either up - neutral or
down - neutral, fluid circulates back through the
pump (unloaded).
b.When the control lever is moved out of neutral,
fluid flows through a check valve to the solenoidoperated door control valve and to the gear priority
valve.
c. Fluid flows through this door control valve
(which is in the door-open position when the handle
is moved out of neutral) and opens the doors. The
gear priority valve remains closed while the door
system is being operated, because the door system
operates at less pressure than is required to open
the priority valve.
d. After the doors are open, pressure builds up
until the gear priority valve opens and permits fluid
first to unlock, then to move the landing gear to
either the up or down pOsition, depending on the
Prior to 1968 models, a secondary relief
valve, which also serves as the emergency
hand pump relief valve, opens at a higher
pressure than the primary relief valve.
Beginning with 1968 modelS, the secondary
relief valve is deleted from the hydrauliC
system. This also includes relocation of
the primary relief valve, in the hydraulic
circuit, to a position downstream of the
engine-driven hydrauliC pump check valve.
This prevents loading of the engine-driven
pump when the emergency hand pump is
operated. Delete references to the secondary relief valves for 1968 modelS.
h. When extending the landing gear with the emergency hand pump, fluid flows directly to the solenoidoperated door control valve and to the gear priority
valve, where it first opens the doors, then extends
the gear through the same passages and lines used by
the regular system. A check valve prevents fluid
from entering the inlet passage from the enginedriven hydrualic pump.
i. In case of an electrical failure, the door control
valve will move to the door-open position, and remain in this position.
j. A door vent valve in the power pack, relieves
any pressure from thermal expanSion in the door
5-5
system, to keep the doors closed, while the aircraft
is parked.
NOTE
This valve is not installed on some early
power packs. However, replacement
power packs (either new or remanufactured) have the valve installed.
5-4. MAIN GEAR SYSTEM.
aft and inboard to stow the wheels in the lower fuselage, beneath the rear horizontal firewall. The firewall has raised contours just above the stowed position of the wheels. Each strut is attached to a saddle, which is rotated by a universal joint. The two
universal joints are operated by one main landing
gear rotary actuator, which is a double-acting cylinder, powering a rack and pinion gear. A downlock
and double-acting downlock release cyUnder is provided for each main gear, and a single-acting uplock
release cylinder operates both main gear uplocks.
•
5-5. DESCRIPTION. Main landing gear struts rotate
5-6. TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
REMEDY
Reservoir fluid level low.
Refill reservoir.
Engine-driven pump failure
or internal leakage.
Repair or replace engine pump.
Air leakage in engine pump
suction line.
Repair or replace suction lines or
fittings.
Fluid leak in door or gear line.
Tighten or replace lines.
Defective piston seal in door
or gear cylinder.
Repair or replace defective parts.
Excessive internal Power
Pack leakage.
Remove and repair or replace
Power Pack.
Broken or distorted universal
joint.
Replace universal jOint and adapter
as an assembly.
Sheared tapered pins or bolts
at actuator shaft, universal
joint, or adapter.
Replace defective parts.
GEAR OPERATES, BUT
DOORS WILL NOT OPEN.
Solenoid valve jammed or stuck
in door-closed position.
Repair or replace solenoid valve.
Repair any damage to doors or door
operating linkage.
GEAR UNLOCKS BEFORE
DOORS ARE FULL-OPEN.
Priority valve setting low.
Adjust valve setting.
Priority valve leaking or stuck
open.
Remove Power Pack and repair
or replace valve.
Adapter-to-saddle pivot shaft
not tight, permitting shear movement between adapter and saddle
shaft.
Remove bolts and shear washer.
Clean any metal from serrations
of pivot shaft and adapter. Install
a new shear washer, and reinstall
bolts and safety. Refer to paragraph 5-46 for indexing of bolts
in slotted holes during assembly.
LANDING GEAR OPERATION
EXTREMELY SLOW.
ONE MAIN GEAR WILL
NOT RETRACT OR EXTEND.
ONE MAIN GEAR LAGS
BEHIND DURING RETRACTION.
5-6
•
•
•
5-6. TROUBLE SHOOTING (Cont) .
TROUBLE
PROBABLE CAUSE
UNEVEN OR EXCESSIVE
TIRE WEAR.
AIRCRAFT LEANS
TO ONE SIDE.
•
Dragging brake.
Jack wheel and check brake.
Wheel bearings not adjusted
properly.
Tighten axle nut properly.
Defective actuators.
Repair or replace actuators.
Incorrect tire inflation.
Inflate to correct pressure.
Wheels out of alignment.
Align wheels.
Wheels out of balance.
Balance wheels.
Sprung landing gear spring.
Replace spring.
Bent axle.
Replace axle.
Incorrect tire inflation.
Inflate to correct pressure.
Landing gear attaching
parts not tight.
Tighten loose parts; replace
defective parts.
Sprung landing gear spring.
Replace spring.
Bent axle.
Replace axle.
Different quantity of fuel
Refuel airplane.
""
in wing tanks.
Structural damage to landing
gear bulkhead components.
ONE OR MORE UPLOCKS OR
Incorrect rigging.
DOWN LOCKS DO NOT OPERATE
5-7. MAIN GEAR STRUT REMOVAL. (Refer to
figure 5-1.)
a. Remove bench-type rear seat or individual center seats.
b. Remove carpeting and access covers from area
of landing gear bulkhead.
c. Jack the aircraft in accordance with procedures
outlined in Section 2.
Replace damaged parts.
Rig per applicable paragraph.
•
(20).
h. Remove bolts securing axle (15) and brake torque
plate to strut, noting numbers of and marking pOSition
of wheel alignment shims (14), so that shims may be
installed in exactly the same pOSition.
i. With master switch OFF, place landing gear
handle up, and operate emergency hand pump until
main gear downlock re leases.
j. Disconnect brake hose from swivel fitting at
block near saddle (2); cap openings.
k. Remove inboard bolt (4) and barrel nut (1) securing strut to saddle (2).
1. Remove bolts (5) securing clamp (6) and strut to
saddle.
m. Carefully work strut out through door openings,
leaving brake line (9) attached to strut.
n. Remove brake line (9) from clips (10) on strut.
e. Remove bolts securing back plates to brake cylinder and remove back plates.
f. Remove cotter pin (22) and axle nut (19); remove
wheel from axle.
g. Disconnect brake hose (13) from brake assembly
(16) and plug or cap openings.
5-8. STRUT INSTALLATION. (Refer to figure 5-1.)
a. Install brake line (9) in Clips (10) on strut.
b. Carefully work strut through door opening into
position on saddle.
c. Install inboard bolt (4), barrel nut (1), bolts (5)
NOTE
If a new strut is to be installed, complete
steps "d" thru ''h'', and step "n".
d. Remove hub cap retainer screws (21) and hub cap
•
REMEDY
5-7
TIGHTEN ONLY
FINGER-TIGHT
BEFORE DRILLING
•
151
MAKE SURE STRUT IS
FORWARD AS FAR AS
IT WILL GO BEFORE
DRILLING
Detail
(RH Installation shown)
Detail
A
DetailB
D
I
•
1. Barrel Nut
2. Saddle
3. Strut
4. Hex-Head Bolt
5. Internal Wrenching Bolt
6. Strut Clamp
7. Deleted
S. Pivot Shaft
9. Brake Line
10. Clip
11. Hose
12. Union
13. Brake Hose
14. Wheel Alignment Shim
15. Axle
15A. Fitting
16. Brake
NOTE
17. Bolt
IS. Wheel
It is necessary to drill a hole in the downlock
19. Axle Nut
stop of new main gear struts prior to initial
20. Hub Cap
installation. Refer to paragraph 5-S for in21. Screw
structions to be used in conjunction with de22. Cotter Pin
23. High-Strength Bolt taU "D" of this figure.
24. Bushing
Figure 5-1. Main Landing Gear
5-S
DetailC
20
•
•
•
and clamps (6) securing strut to saddle (2).
NOTE
It is necessary to drill a hole in the downlock
stop of new main gear struts prior to initial
installation. Refer to detail "D" for instructions
to be used in conjunction with this paragraph.
d. When installing a new strut, complete steps "a"
and ''b'', and install inboard bolt (4), barrel nut (1),
aft bolt (5) and clamp (6) securing strut to saddle (2).
Tighten aft bolt (5) only finger-tight (tightening bolt
too tight will raise forward end of clamp). Do not
drill hole in downlock stop until after completion of
step ''n''.
e. Connect brake line to swivel fitting at back near
saddle.
f. Inspect axle for straightness and for damage to
threads; if damaged or bent, install new part.
g. Insert mounting bolts (17 and 23) through torque
plate, axle and alignment shims. Position shims according to reference marks made at time of disassembly.
h. Position axle assembly to strut, install nuts and
tighten.
i. Slide wheel on axle, using care to prevent damage
to threaded surface of axle •
j. Install axle nut (19) on axle and tighten until a
slight bearing drag is obvious when wheel is rotated.
k. Loosen axle nut only enough to aUgn with nearest
cotter pin hole and install cotter pin (22).
1. Install back plates and cylinder bolts •
m. Install hub cap (20) and retainer screws (21).
n. Connect brake hose (13) to brake assembly (16).
o. If a new strut (3) is being installed, move strut
at wheel, aft as far as it will go; this will move upper
inboard end of strut forward.
p. Make sure upper inboard end of strut is forward
as far as it will go, and checking from underneath,
line up hole in forward arm of saddle (2) with tab on
downlock stop.
q. Using a size ''V'' (.377) drill, Une drill up through
hole in saddle arm, through downlock stop tab.
r. Install forward bolt (5), and tighten both bolts (5),
securing clamp (6) to strut (3) and saddle (2).
s. Bleed brakes in accordance with instructions outlined in paragraph 5-73.
t. Check rigging of main landing gear in accordance
with paragraph 5-264.
u. Remove aircraft from jacks and check wheel alighment in accordance with figure 5-12.
v. Install upholstery and access panels.
w. Install rear seat.
5-9. MAIN LANDING GEAR ACTUATOR. (Refer to
figure 5-3.)
•
5-10. DESCRIPTION. The main gear actuator is a
double-acting, cylinder-type actuator, powering a
rack and pinion gear. The actuator is located between,
and connected to the main landing gear strut saddles
by means of universal joints and adapters. When the
gear control handle is moved to the gear -up pOSition,
hydraulic pressure is routed to the main gear actuator,
moving the pinion gear and rack, causing the main
gear struts to rotate aft and inboard into the stowed
position. Moving the gear control handle to the geardown position, reverses the movement of the rack and
pinion, rotating the main gear struts forward and down
ward into the gear -down position.
5-11. REMOVAL. (Refer to figure 5-2.)
a. Remove rear or center seats.
b. Remove carpeting and access covers from area
of actuator.
c. Jack aircraft in accordance with instructions
outlined in Section 2.
d. With master switch OFF, place landing gear control handle in the gear -up position and use emergency
hand pump to rotate main landing gear as necessary
for access and clearance.
e. Mark all parts in their correct relationship to
each other, before removal.
f. Remove tapered pins (13) securing universal
jOints (26) to actuator shaft (15).
g. Remove tapered pins (13) securing both universal
joints to adapters (14).
h. Remove bolts attaching both adapters (14) to
saddle pivot shafts (10) and slide both adapters inboard
on universal joints as far as possible.
i. Remove shear washers (12) between adapters and
pivot shafts.
j. Disconnect and cap or plug all hydraulic lines at
the actuator (18), without disturbing the fittings installed
in the actuator.
k. Remove horizontal angle (19) above center part of
actuator.
1. Remove bolts securing vertical angles (17 and 20)
to structure under center of actuator. The vertical
angles (17 and 20) may be left attached to the actuator.
m. Remove four mounting bolts at forward end of
actuator.
n. Remove bolts securing vertical angles (22 and 24)
to structure under forward end of actuator.
o. Remove bolts attaching horizontal angle (21) above
forward end of actuator at each side, and lift horizontal angle (21), (with vertical angles 22 and 24 attached),
upward to remove.
p. Work adapter end of universal joint up, then
slide inboard end of universal joint outboard until
clear of actuator shaft, and remove universal joint
with adapters attached.
q. Lift forward end of actuator just clear of structure under the actuator, slide actuator forward until
aft end can be lifted free, then work actuator from
aircraft.
5-12. DISASSEMBLY.
(Refer to figure 5-3.)
NOTE
Leading particulars of the actuator are as
follows:
Cylinder Bore Diameter
Piston Diameter . . .
Piston Rod Diameter. .
Cylinder Stroke . • . . .
Shaft Rotation (Loaded).
Shaft Rotation (Unloaded) .
2.996 in.
2.992 in.
1. 222 in.
3.38 in.
180 D (Min)
187 D (Max)
a. Remove screw (18), then remove end gland (4)
and metering pin (1) by unscrewing end gland from
5-9
•
NOTE
Lubricate NTA -4860 thrust bearing (7)
with MIL-G-2U64 grease on installation.
(Service each 500 hours thereafter. )
•
21
NOTE
If the installation of bearing (11) is not a ltght
press fit (tight enough to hold the bearing in
Washers (29) and spacers (9) are used
as required to eliminate end play from
pivot shaft.
position and prevent rotation in support (27),
prime bearing and joining surface of support
with Grade "1'" Primer and seal with Retaining Compound 75 (Lodite Corporation).
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Outboard Support
Saddle
Bolt
Strut
Strut Clamp
Internal Wrenching Bolt
Thrust Bearing
Thrust Bearing Race
Spacer
Pivot Shaft
11.
12.
13.
14.
15.
16.
17.
18.
19.
Bearing
Shear Washer
Tapered Pin
Adapter
Actuator Shaft
Lower Center Horizontal Angle
Right Center Vertical Angle
Main Landing Gear Actuator
Upper Center Horizontal Angle
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
Left Center Vertical Angle
Upper Forward Horizontal Angle
Left Forward Vertical Angle
Lower Forward Horizontal Angle
Right Forward Vertical Angle
Tapered Pin Washer
Universal Joint
Inboard Support
Barrel Nut
Washer
Figure 5 -2. Main Gear Actuator and Linkage Installation
5-10
•
•
•
14
9. Support
10. Sector
11. Bearing
12. Bushing
13. Thrust Washer
14. Cap
15. Roll Pin
16. Piston
1. Metering Pin
2. Locknut
3. Plug
4. End Gland
5. O-Rillg
6. Back-Up Ring
7. Cylinder Body
8. Cap Plug
17. Snap Ring
lB.
19.
20.
21.
22.
23.
24.
Screw
Back-Up Ring
O-Ring
Back-Up Ring
O-Ring
Back-Up Ring
O-Ring
Figure 5-3. Main Gear Actuator
cylinder body (7).
b. Remove cap plug (8) and, using a phenolic block,
drive piston (16) from cylinder body (7). Use care
when removing piston to prevent damage.
c. Cut safety wire and remove pins (15) from cy- -Under body.
d. Remove cap (14) from cylinder body. This removes bearing (11), bushing (12) and thrust washer
(13).
•
NOTE
Unless defective, do not remove bearings
(11) and thrust washers (13) from cap or
cylinder body.
e. Remove sector (10) from cylinder body.
f. Remove support (9), using a phenolic block and
tapping from cylinder body. Tap out from smaller
end of support.
g. Remove bushings (12) from bearings (11) in cap
and cylinder body.
h. Remove snap ring (17) and metering pin (1) from
end gland (4).
i. Remove and discard all O-rings and back-up
rings.
5-13. INSPECTION OF PARTS. Perform the following inspections to ascertain that all parts are in serviceable condition.
a. Thoroughly clean all parts in ~t:\lvp.nt (Federal
5-11
•
0--
END OF ACTUATOR
END OF ACTUATOR
•
1.
Measure distances "A" and "B" when the actuator piston is bottomed in UP and DOWN positions.
2.
Subtract "A" from "B" to establish actual travel of rack.
3.
Subtract 2.815" (travel needed to operate landing gear) from this actual travel, to establish
unused travel.
4.
Subtract one-half of this unused travel from "B" to establish the distance from the end of the
actuator to the rack. This is the DOWN RIGGING POSITION of the actuator.
NOTE
Accomplishing the procedure listed above divides the unused travel equally at
each end of the actuator, and establishes a DOWN RIGGING POSITION for any
particular actuator.
Figure 5-4. Main Gear Actuator Down Rigging Position
5-12
•
•
-
-.750"
ADAPTER
-------TtT-
-- ------ -t----1,1
I
J
•
-
.310"
1. Before drilling and reaming these tapered pin holes, be sure alignment is as follows:
a. Main gear actuator must be in the rigging position specified in figure 5-4.
b. Inboard end of universal joint must be attached securely to the actuator shaft, with
tapered pins tightened.
c. Landing gear must be down and locked.
d. Adapter must be installed, with shear washer in place, and attaching bolts must be
in center of slotted holes in adapter.
2. Locate the inboard tapered pin hole as shown, drill and ream, and install tapered pin with
speclal washer and nut.
NOTE
start with a No. 21 drill, then use a 7/32-inch drill, then a 1/4-inch
straight reamer. After a smooth 1/4-inch hole has been obtained, use
a B and S No. 2 Taper Reamer (or equivalent), removing only enough
material to permit the smaller end of the tapered pin to be flush; it must
not protrude more than 1/16 inch. Install the special tapered pin washer
with its flat side against.the nut.
3. After installing the inboard tapered pin, rotate landing gear and repeat step "2" for the
outboard pin.
NOTE
•
The tapered pin holes may be drilled and reamed with parts installed
in the airplane, or an initial hole may be located and drilled, and parts
may be removed and reassembled for bench-drilling and reaming. It
is not critical that the tapered pin holes be exactly 90° to each other,
nor is it critical that they be exactly perpendicular to and through the
centerline of the parts. A tolerance of :1:5° is permissible .
Figure 5-5. Tapered Holes for Universal Joint Replacement
5-13
SpecUicatlon P-S-66., or equivalent).
b. Inspect all threaded surfaces for cleanliness and
freedom from cracks and wear.
c. Inspect cap (14), bushings (12), sector (10), support (9), piston (16) and cyUnder body (7) for cracks,
scratches, scoring, wear or surface irregularities
which may affect their function or the overall operation of the actuator.
d. Inspect bearings (11) for roller operation and for
scores, scratches and Brlnnel marks.
5-14. REPLACEMENT/REPAm OF PARTS.
a. Repair of small parts of the actuator is impractical. Replace all defective parts with serviceable
parts. Minor scratches or scores may be removed
by polishing with abrasive crocus cloth (Federal
Specification P-C -458), providing their removal does
not affect the operation of the actuator.
b. Install new a-rings and back-up rings during
assembly.
5-15. ASSEMBLY. (Refer to figure 5-3.)
NOTE
Lubricate sector aDd piston rack gears with
MIL-G-23827 lubricant. Apply lubricant
sparingly. Over-greasing may cause ccmtamination of the hydraullc cylinder with grease,
which may work past back-Up ring (6) and
a-ring (5).
•
i. Install cap (14), USing attaching roll pins (15).
Safety Wire roll pins.
j. Install new back-up rings (19) and a-ring (20)
in bore of end gland (4), and install back-up ring
(21) and a-ring (22) in groove on end gland (4).
k. Install metering pin (1) in end gland (4), and install snap ring (17) on metering pin.
1. Install end gland and metering pin assembly in
cylinder, and tighten until end gland is tight in cylinder. Install, tighten and safe.ty Allen screw (18).
m. Install end cap (8) at end of actuator assembly.
n. Adjustment of metering pin, causing a snubbing
action in the actuator, is accomplished as outlined
in paragraph
NOTE
Use MIL-G-21164 lubricant on support (9),
bearings (11) and sector (10) when installing
parts in cylinder.
a. If bearings (11) are being replaced, insert thrust
washer (13) in cylinder body and press bearing (11) in
until seated against thrust washer and retaining base
in cyUnder body. Install thrust washer and bearing
in cap (14).
b. Lubricate bearing and insert bushing (12) in
bearing in cylinder body.
c. Lubricate and install back-up ring (6) and a-ring
(5) in groove of cylinder bore.
d. Install support (9) in cylinder body, tapping it in
until seated against retaining base in cylinder.
NOTE
Ensure that cutout in support (9) will
align with piston when piston is installed.
e. Lubricate and install back-up rings (23) and O-ring
(24) in groove on piston (16).
f. Slide piston (16) into cylinder body so that flat
portion of piston rack aligns with cutout in support
(9). Push piston to bottom of cylinder body bore.
Use care to prevent damage to back-up and O-rings
in cylinder bore and on piston.
NOTE
Be sure that gear teeth on sector rotate
in the correct direction so that piston
can be extended.
g. Place sector (10) in cylinder, aligning first tooth
on sector with first tooth on piston rack with piston
retracted.
h. Lubricate bearing (11) and insert bushing (12) in
cap (14).
5-14
5-16. INSTALLATION. (Refer to figure 5-2.)
a. Work actuator (18), With center vertical angles
(17 and 20) attached, into pOSition, and lUt aft end to
clear structure. Sllde aft until forward end of actuator will clear lower structure.
b. If adapters (14) were removed from universal
joints (26), place in position on universal jOints, but
do not install tapered plns (13).
c. Work adapters into position, shUting actuator
from side to side as necessary for clearance.
d. Ensure that all parts are aligned as marked
during removal, then install tapered pins and tighten.
e. Position upper forward horizontal angle (21),
with forward vertical angles (22 and 24) attached, and
install mounting bolts.
f. Install bolts securing forward vertical angles
(22 and 24) to structure under forward end of actuator.
g. Install four mounting bolts through upper forward
horizontal angle (21) to actuator.
h. Install outer vertical angles (17 and 20) to structure. If center vertical angles were not attached to
actuator, attach with four bolts (refer to paragraph
•
5-11).
i. Install upper center horizontal angles (19) to
structure.
j. Connect all hydraulic lines to actuator.
k. Slide shear washers (12) into posttton between
adapters and pivot shafts. Since the pivot shaft and
adapters are both serrated, the shear washers may be
reused once by turning them so that the serration
marks are 90 0 to their original position.
1. With landing gear down and locked, and main
gear actuator in down poSition, install bolts securing
adapters to pivot shafts (10). Torque bolts to 250 Ibin.
!CAUTION\
Use only a phenoliC hammer or equivalent
when seating adapter and pivot shaft into
shear washer. Serious damage to parts
may otherwise result.
•
•
m. Seat the serrations of the adapter and pivot
shaft into the shear washer by hammering on adapter.
Retorque and safety wire bolts in pairs.
n. Bleed hydraulic system in accordance with paragraph 5-163.
o. Rig landing gear in accordance with paragraph
5-264.
p. Check wheel alignment in accordance with figure
q. Remove aircraft from jacks.
r. Install access panels and upholstery.
s. Install seats.
5-17. MAIN LANDING GEAR LINKAGE. (Refer to
figure 5-2. )
5-18. DESCRIPTION. The main landing gear linkage
consists of two pivot shafts, two universal joints, two
adapters, two saddles, two strut clamps with bearings,
pins, bushings and attaChing parts. The linkage provides the connection between the main landing gear
actuator and the main landing gear struts. The landing
gear struts are clamped in saddles which are rotated
by pivot shafts connected to the main gear actuator
shaft, which rotates the main gear struts into the retracted or extended position.
5-19. REMOVAL OF UNIVERSAL JOINTS AND
ADAPTERS. (Refer to figure 5-2. )
a. Remove rear or center seats.
b. Remove carpeting and access covers from landing
gear bulkhead •
c. Jack aircraft in accordance with procedures outlined in Section 2.
NOTE
With master switch OFF, use emergency hand
pump to rotate main landing gear as necessary
for access and clearance when removing bolts
and tapered pins. Mark all parts in their
correct relationship to each other before removal.
d. Remove tapered pins securing universal joint
to adapter shaft.
e. Remove tapered pins securing both universal
joints to adapters.
f. Remove bolts attaChing both adapters to saddle
pivot shafts, and slide both adapters inboard on universal joints as far as they will go. Remove shear
washers between adapters and pivot shafts.
NOTE
It is necessary to sUde both adapters inboard,
regardless of which universal joint is being
removed, so that the main gear actuator
may be lihifted laterally.
•
5-20. INSTALLATION OF REMOVED UNIVERSAL
JOINTS AND ADAPTERS. (Refer to figure 5-2.)
NOTE
5-12.
•
j. Shift main gear actuator in the opposite direction
and remove the other universal joint .
k. After universal jOints have been removed, adapters may be removed.
g. Disconnect, plug or cap all hydrauliC lines at the
main gear actuator.
h. Remove actuator supporting structure as necessary to allow actuator to be shifted laterally.
i. Work adapter end of universal joint up, then
slide inboard end of universal joint outboard until it
clears actuator shaft; remove universal joint.
The following procedure is to be used when
the same parts are being reinstalled.
a. If adapters were removed from universal joints,
place adapters in position on universal jOints, but do
not install tapered pins.
b. Work adapters and universal jOints into position,
shifting main gear actuator from side to side as necessary for clearance.
c. Ensure that all parts are aligned as marked during removal, then install and tighten tapered pins.
d. Install all parts securing main gear actuator,
and connect hydraulic lines.
e. Slide new shear washers into position between
adapters and pivot shafts. Since the shaft and adapter
are both serrated, the shear washers may be reused
once by turning them so the serration marks are 90°
to their original position.
f. With landing gear down and locked, and main
gear operated to the rigging position (refer to paragraph 5-264), install bolts securing adapters to pivot
shafts. Tighten bolts to 250 Ib-in. Using an E-6
rivet gun with suitable flat rivet set (or hammer and
rod), seat serrations of the adapter and pivot shaft
into shear washer. Retorque and safety bolts in
pairs.
g. Operate landing gear through several cycles tobleed any air from the system, checking for proper
operation.
h. Remove aircraft from jacks and install all parts
removed for access.
5-21. INSTALLATION OF NEW UNIVERSAL JOINTS
AND ADAPTERS. (Refer to figure 5-2.)
NOTE
The following procedure is to be used when
new parts are to be installed.
a. Position adapters (14) on undrilled end of universal joints (26) and work into position, shifting main
gear actuator (18) from side to side as necessary for
clearance.
b. Align tapered pin holes in inboard end of universal joints with corresponding holes in main gear actuator shaft. Install tapered pins (13) and tighten.
c. Install actuator support angles and bolts (refer
to paragraph 5-16). Connect hydraulic lines.
d. Ensure that main gear actuator remains in the
down position.
.e. Manually move landing gear to down and locked
position. Maintain this position .
f. Slide shear washers (12) into position between
adapters (14) and pivot shafts (10). Since the pivot
shaft and adapters are both serrated, the shear
washers may be reused once by turning them so that
5-15
the serration marks are 90 u to their original position.
g. Install bolts securing adapters to pivot shafts
and torque to 250 Ib-in.
[cAurloNl
Use only a phenolic hammer or equivalent
when seating adapter and pivot shaft into
shear washer. Serious damage to parts
may otherwise result.
h. Seat serrations of adapter and pivot shaft into
shear washer by hammering on adapter. Retorque
and safety bolts in pairs.
i. Bleed hydraulic system in accordance with
paragraph 5-163.
NOTE
The tapered pin holes may be drilled and
reamed with parts installed in the aircraft,
or an initial hole may be located and
drilled and parts then removed and reassembled for drilling and reaming on
bench. It is not critical that the tapered
pin holes be exactly 90° to each other, nor
is it critical that they be exactly perpendicular to and through the centerline of
the parts. A tolerance of .! 5° is permissible.
j. Locate, drill and ream tapered holes through
adapters and universal joints as follows:
NOTE
Maintain dimensions called out in Figure 5-5.
1. Ensure that alignment is as follows
before drilling and reaming.
(a) Main gear actuator must be in the
down pOSition.
(b) Landing gear must be down and
locked.
(c) Inboard end of universal joint must
be attached securely to actuator shaft with nuts on
tapered pins tightened.
(d) Adapter must be installed with shear
washer in place; attaching bolts must be in center of
slotted holes in adapter.
NOTE
Start with a No. 21 drill, then use a 7/32inch drill, then a 1/4 -inch straight reamer.
After a smooth 1/4-inch hole has been obobtained, use a B & S No. 2 taper reamer
(or equivalent), removing only enough material to permit the smaller end of the tapered
pin to be flush. The smaller end of the tapered
pin must not protrude more than 1/16-in. Install the special tapered pin washer with flat
side against the nut.
2. Locate inboard tapered hole as shown in
figure 5-4. Drill and ream hole. Install tapered pin
(13) with special washer (25) and nut (refer to figure
5-2. )
5-16
3. After instalUng inboard tapered pin, rotate landing gear (or remove necessary parts) and repeat step "2" for the outboard pin.
k. Bleed hydraulic system in accordance with figure
5-163.
1. Lubricate universal jOints in accordance with
Section 2.
m. Rig landing gear in accordance with paragraph 5264.
n. Check wheel alignment as shown in figure 5-12.
o. Remove aircraft from jacks.
p. Install access panels, upholstery and seats.
•
5-22. REMOVAL OF MAIN GEAR SADDLE AND
PIVOT SHAFT. (Refer to figure 5-2.)
a. Remove main gear strut In accordance with paragraph 5-7.
b. Remove universal joint and adapter in accordance
with paragraph 5 -19.
c. Remove bolts (3) attaching saddle (2) to pivot
shaft (10).
d. Pull pivot shaft inboard until clear of outboard
bearing support.
e. Allow saddle (2), thrust bearing (7), bearing race
(8) and spacers (9) to slide outboard as pivot shaft is
pulled Inboard.
NOTE
Note number of and thickness of spacers
(9) between thrust bearing race and inboard bearing support.
f. When shaft is clear of outboard bearing support
(1), lift outboard end of shaft and slide saddle off shaft.
Remove remaining bearing parts from shaft.
g. Move main gear actuator (18) as required for
clearance and pull pivot shaft inboard to remove from
aircraft.
•
5-23. INSTALLATION OF MAIN GEAR SADDLE AND
PIVOT SHAFT. (Refer to figure 5-2.)
a. Position pivot shaft (10) through inboard forging
(27). Slide spacers (9), thrust bearing race (8), thrust
bearing (7) and saddle (2) onto shaft (10).
b. During installation, lubricate thrust bearing and
race as specified in Section 2 and figure 5-2.
NOTE
Spacers (9) are used as required to remove
end play from pivot shaft without caUSing it
to bind.
c. Position outboard end of pivot shaft ill bearing in
outboard support forging (1). Check end play of shaft
and adjust with shims (29) as necessary.
d. Install bolts (3) securing saddle (2) to pivot shaft
(10).
e. Install universal joint (26) and adapter (14) in
accordance with paragraph 5 -20.
f. Install main gear strut (4) In accordance with
paragraph 5-8.
5-24. UPLOCK MECHANISM. (Refer to figure 5-6.)
5-25. DESCRIPTION. The uplock is a hook (or pawl)
•
•
12
.050" TO . 100"
Detail
A
3
b
I
•
A
/'
,
--~--@-
10
PRIOR TO 1971 MODEL
10
SERIAL NO. 337-0114
THRU 337 -0905
•
1.
2.
3.
4.
5.
Uplock Release Actuator
End Fitting
Bushing
Clevis
Hook Pivot Bolt
Detail
B
6.
7.
8.
9.
10.
SERIAL NO. 337-0906
AND ON
Slotted Hole
Bushing
Main Gear Strut
Up lock Hook
Bracket
11.
12.
13.
14.
15.
Uplock Switch
Spring-Loaded Push-Pull Rod
Block and Washer
Bellcrank
Plate Assembly
Figure 5 -6. Main Gear Uplock Installation (Sheet 1 of 2)
5-17
•
7
Detail
A
~
5
'''-......
4
"
I
I
_12
•
1971 MODELS AND ON
Uplock Release Actuator
End Fitting
Bushing
Clevis
Hook Pivot Bolt
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
Hanger
Shim
Uplock Stop
Uplock Hook Support
Spring-Loaded Push-Pull Rod
Main Gear Strut
Bushing
Uplock Hook
Block and Washer
Bellcrank
Uplock Switch
Bracket
Figure 5-6. Main Gear Uplock Installation (Sheet 2 of 2)
5-18
•
•
t=:
*
Used ONLY on main landing gear uplock actuator.
All other parts used on main landing gear downlock actuators.
~1
cr
:
I
10
•
1. End Fitting
2.
3.
4.
5.
6.
7.
8.
9.
10.
Nut
Back-Up Ring
O-Ring
Fitting
O-Ring
Spring
Ball
Ball
Back-Up Ring
11.
12.
13.
14.
15.
16.
17.
18.
19.
O-Ring
Barrel and Valve Body
Piston and Rod
Back-Up Ring
O-Ring
O-Ring
Spring
Spring
End Plug
*
Figure 5-7. Lock and UnlOCk Actuator Assembly
which is spring-loaded to the loclted pOSition and hydraulically operated to the unlocked position. The installation consists of one hydraulic uplock release
actuator, two clevises, two uplock hooks, two bellcranks, two &{Iring-loaded push-pull rods, two uplock
swl tches and attaching parts.
•
5-26. OPERATION. The uplock hook is moved into
the locked position when the main gear strut strikes
the upper part of the hook, causing the hook to rotate
to the locked position by cam action. The springloaded push-pull rod maintains the locked position
until the cam action is reversed by actuation of the
uplock release actuator, which is linked to the pushpull rod by a clevis and bellcrank. The uplock indi-
cator switches are actuated when the gear is in the
up and locked pOSition.
5-27. REMOVAL OF MAIN GEAR UPLQCK MECHANISM AND RELEASE ACTUATOR. (Refer to figure
5-6. )
a. Remove rear or center seats.
b. Remove upholstery and access panels from area
of main landing gear bulkhead.
c. Jack aircraft in accordance with procedures outlined in Section 2.
d. Remove end fitting (2) from actuator shaft (1) and
bellcrank (14).
e. Disconnect push-pull rod (12) from uplock hook
(9) and bellcrank (14).
5-19
f. Remove clevis (4) from push-pull rod (12) by
loosening locknut after noting distance from outboard
end of clevis to mounting bracket. Remove push -pull
rod.
g. Disconnect electrical connection to uplock switch
Specification P-C -458), providing their removal doe.
not affect the operation of the unit.
5-30. ASSEMBLY. (Refer to figure 5-7.)
NOTE
(11).
h. Remove uplock hook pivot bolt (5) and remove
hook from aircraft.
1. Uplock switch, bracket and hook may be disassembled after removal from aircraft.
j. Disconnect hydraulic lines at actuator 0) and
plug or cap openings.
k. Remove actuator mounting bolts and actuator.
5-2S. DISASSEMBLY OF UPLOCK RELEASE ACTUATOR. (Refer to figure 5-7.)
NOTE
Leading particulars of the actuator are
as follows:
Cylinder Bore Diameter .
Piston Diameter • . . . .
Piston Rod Diameter. . .
Stroke (except maingear
downlock)
(Total at 1. 0 GPM) • . .
Stroke (maingear downlock
total travel) . . .
Stroke (to unseat valvel . .
0.750 +.002, -.000 in.
0.748 +.000, -.002 in.
0.343 +.001, -.002 in.
O. Sl2 in. (max)
O. S4 :t .04 in.
0.719 :t . 031 in.
Install all new O-rings and back-up
rings during lock cylinder assembly.
a. Install new O-rings (16 and 15) and back-up ring
(14) in grooves on piston and rod 03).
b. Install new O-ring (11) and back-up ring (10) In
groove of barrel and valve body (12).
c. Slide piston and rod (13) into barrel and valve
body (12). Use care to prevent damage to O-rings
and back-up rings.
d. Insert springs (17 and 18). Install and safety
end plug (19) or end fitting (I) to barrel and valve
body (12).
e. Insert balls (S and 9) and spring (7) in barrel and
valve body (12).
f. Install new O-ring (6) on fitting (5). Install and
tighten fitting (5).
5-31. INSTALLATION OF MAIN GEAR UPLOCK
MECHANISM AND RELEASE ACTUATOR. (Refer to
figure 5-6. )
a. Position actuator (1) to align with holes in mounting bracket. Install bolts, washers and nuts; tighten.
NOTE
a. Remove fitting (5), spring (7) and balls (S and 9).
b. Cut safety wire and unscrew end plug (19) from
barrel and valve body (12).
c. If end fitting (1) is installed, loosen nut (2) and
remove end fitting from barrel and valve body.
d. Remove springs (17 and IS) and piston and rod
(13) from barrel and valve body.
e. Remove and discard all O-rings and back-up
rings.
5-29. INSPECTION OF PARTS. Make the following
inspections to determine that all parts are in a serviceable condition.
a. Inspect all threaded surfaces for cleanliness,
cracks and excessive wear.
b. Inspect spring (17) for breaks and distortion.
The free length of the spring must be 2.95± .09 inches
and compress to 1. 969 inches under a 22.5 ± 2.2 lb.
load.
c. Inspect spring (18) for breaks and distortion. The
free length of the spring must be 2. 9S:t .09 inches and
compress to 1. 969 inches under a 10. 6 :t 1. 1 lb. load.
d. Inspect spring (7) for breaks and distortion. The
free length of the spring must be .446:t .015 inches
and compress to .359 inches under a .IS ± .02 lb. load.
e. Inspect plug (19) or fitting (I), piston and rod (13),
barrel and valve body (12), balls and ball seats for
cracks, chips, scratches, scoring, wear or surface
irregularities which may affect their function or the
overall function of the unit.
f. Repair of most parts of the lock cylinder is impractical. Replace defective parts With serviceable
parts.
g. Minor scratches and scores may be removed by
polishing with fine abrasive crocus cloth (Federal
5-20
•
Actuator position, when installed, must be in
same relation to attaching components as is shown
in figure 5-6. The longer part containing the
check valve toward the right side of the aircraft.
b. Connect and tighten hydraulic lines at actuator.
c. Assemble uplock hook (9), switch bracket (10) and
switch (11), leaving screws through switch and slotted
holes in bracket, loose for adjustment.
d. Position uplock hook (9) and install hook-pivot
bolt (5), washers, bushing and nut in slotted holes (6)
in support structure (DO NOT TIGHTEN.)
e. Position spring-loaded push-pull rod (12) through
hole in bracket. Install clevis (4) and lock nut. (Maintain measurement distance noted in step "f" of paragraph 5 -27 . )
f. Connect push-pull rod (12) to uplock hook (9) and
bellcrank (14).
g. Install end fittings (2) to shaft of act.Jator (I) and
bellcrank (14).
h. Connect electrical connection at uplock switch
(11).
i. Bleed aircraft hydraulic system in accordance
with paragraph 5-163.
j. Rig main gear uplock and uplock switch in accordance with paragraphs 5-267 and 5-268.
k. Torque uplock hook pivot bolt (5) nut to 90-100
lb. in.
1. Rig landing gear in accordance with paragraph
5-264.
m. Install access panels, upholstery and seats.
5-32. OOWNLOCK MECHANISM.
5-S. )
(Refer to figure
•
•
•
24
20
•
I
ii~~~~
1.J-
•
14
Due to a buildup of production tolerances,
it may be necessary to place a 10caUyfabricated shim or spacer between downlock switch (16) and bracket (14) to ensure
adequate contact between switch actuator
and downlock stop bracket on strut.
•
1.
2.
3.
4.
5.
6.
7.
B.
9.
10.
Barrel Nut
Shim
Eyebolt
Nut
Washer
Bolt
Star Washer
Adjusting Support
Main Gear Strut
Adjusting Wedge
1l.
12.
13.
14.
15.
16.
17.
lB.
19.
21
Downlock Pin
Leaf Spring
Downlock
Switch Bracket
Stop Bolt
Switch
Actuator Pin
Roll Pin
Overcenter Arm
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
Downlock Pivot Bolt
Overcenter Release Bolt
Bumper
Overcenter Spring
Outboard Support
Actuator
End Fitting
Clevis
Shim
Clamp
Figure 5-B. Main Gear Downlock Installation
5-21
5-33. DESCRIPTION. The installation consists of
an overcenter arm, a hydraulic downlock actuator,
a downlock assembly containlng an adjustable downlock pin, an adjusting support, a downlock switch and
attaching parts.
5-34. OPERATION. The hydraulically-operated
downlocks (pawls) contain adjustable downlock pins
which wedge under the forward edge of the struts
to lock the landing gear in the down position. The
downlocks are moved out of the way by the downlock
actuators before gear retraction.
5-35. REMOVAL OF MAIN GEAR DOWNLOCK MECHANISM AND OOWNLOCK ACTUATOR. (Refer to figure 5-S.)
a. Remove rear or center seats.
b. Remove upholstery and access panels in area of
landing gear bulkhead.
c. Jack aircraft in accordance with procedures outlined in Section 2.
d. With master switch OFF, place landing gear
control handle in the gear -up position and operate
emergency hand pump until main gear downlocks release.
e. Release hydraulic pressure and pull downlocks
(13) aft for access.
f. Remove actuator pin (17) securing downlock (13)
to arm of actuator (25).
g. Disconnect hydraulic lines from downlock actuator (25) and cap or plug openings.
h. Remove screw securing switch (16) to bracket
(14); remove bracket from downlock (13).
i. Remove clevis bolt and bushing from actuator
end fitting (26).
j. Remove actuator mounting bolts and remove
actuator from aircraft.
k. Disconnect overcenter spring (23) from overcenter arm (19).
1. Remove downlock pivot bolt (20) and remove
downlock assembly (13) from aircraft.
m. Remove bolts securing adjusting support (S)
to outboard support (24). Remove fore - and - aft
adjusting bolt (6). Remove adjusting support assembly from aircraft.
n. Remove eyebolt (3) and overcenter spring (23)
from aircraft.
o. Remove cleviS (27), shims (2S) and clamp (29)
from supporting structure.
NOTE
Parts removed as assemblies may be disassembled after removal from aircraft.
5-36. DISASSEMBLY, INSPECTION OF PARTS AND
ASSEMBLY OF MAIN GEAR DOWNLOCK ACTUATOR.
Main gear uplock and downlock actuators are identical
except for end fittings. Refer to figure 5-7 and paragraphs 5-2S thru 5-30 for procedures for disassembly,
inspection and assembly of main gear downlock actuators.
5-37. INSTALLATION OF MAINGEAR DOWNLOCK
MECHANISM AND DOWNLOCK ACTUATOR. (Refer
to figure 5-S.)
a. Install eyebolt (3), with overcenter spring (23)
5-22
attached. Install washer and nut and tighten.
b. Install clevis (27), shims (28) and clamp (29)
to supporting structure and tighten.
c. Position adjusting support (S) and install attaching bolts loosely.
d. Install fore - and - aft adjusting bolt (6) With
nuts and washers, but do not tighten.
e. Assemble downlock (13), overcenter arm (19),
bumper (22), overcenter release bolt (21), downlock
pin (11), leaf spring (12), sWitch bracket (14) and stop
bolts (15), loosely.
f. Posmon downlock assembly into adjusting support and install downlock pivot bolt (20), bushing,
washer and nut, but do not tighten.
g. Connect overcenter spring (23), to overcenter
arm (19).
.
h. Position actuator (25) into supporting structure,
install mounting bolts, and tighten.
•
I~AUTION\
Applying too much torque to mounting screws
in downlock switch may crack switch case.
1. Install switch (16) to bracket (14).
j. Connect and tighten hydraulic lines.
k. Position actuator end fitting (26) into clevis (2n
Install and tighten cleviS bolt, bushing, washers and
nut.
1. Bleed hydrauUc system in accordance with paragraph 5-163.
m. Rig downlock mechanism in accordance with
paragraph 5-266.
n. Install access panels, upholstery and seats.
5-3S. MAIN GEAR DOOR SYSTEM.
5 -39. DESCRIPTION. The main gear door system
consists of two main wheel well doors, two main gear
door actuators, two strut doors with one actuator and
torque tube, linkage and attaching .parts. The main
gear doors open for extension or retraction of the
landing gear and close again after the cycle has been
completed.
5-40. OPERATION. Each main gear wheel door is
operated by a double-acting hydraulic cyUnder. Both
strut doors are linked through a torque tube to one
double -acting hydraulic cylinder.
5-41. REMOVAL OF MAINGEAR WHEEL DOORS AND
AC TUATORS. (Refer to figure 5 -9. )
a. Remove rear or center seats.
b. Remove upholstery and access panels from area
of landing gear bulkhead.
c. Jack aircraft in accordance with procedures outlined in Section 2.
d. Release hydraulic pressure.
e. Remove bolts securing actuator rod ends to doors.
f. Disconnect hydraulic Unes from actuators and cap
or plug openings.
g. Remove bolt securing actuator to support and remove actuator from aircraft.
h. Support door and remove bolts through forward
and aft hinge brackets. Remove door from aircraft.
1. Mark hinges and brackets on door before disassembly to provide for realignment and location for
reassembly.
•
•
•
....
..
4
....
::..........
.........- .,-::
:::....
::.'-
.....
(:.
...... : ...
....~.•.~:~ ...... -..--.
...........
.......
....
.......
"-:"
6
10
Detail
A
•
13--------'\
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Actuator Support
Actuator
Rod End
Right Arm Assembly
Torque Tube
Support
Actuator Arm Assembly
Bearing Block
Bearing
Left Arm Assembly
11.
12.
13.
14.
15.
16.
17.
Bushing
Spacer
Aft Hinge
Push - Pull Rod
Forward Hinge
Bushing
Spacer
Figure 5-9. Main Gear Doors Installation (Sheet 1 of 2)
•
5-42. DISASSEMBLY OF MAIN GEAR WHEEL DOOR
ACTUATOR. (Thru Serials 33701426 and F33700035)
(Refer to figure 5-10, sheet 1.)
a. Unlock cylinder by applying hydraulic pressure
to port in clevis end (22) of actuator.
b. Loosen locknut (2) and remove rod end (1)
from piston rod. Remove locknut from piston.
c. Remove safety wire from knurled nuts (13) and
loosen knurled nuts.
d. Remove gland end (5) from barrel (17), using a
strap wrench on barrel.
e. Remove clevis end (22) from barrel, then push
piston (7) from barrel. Use care when pushing
piston from barrel, to prevent loss of balls (12).
f. Remove spacer (6) from barrel. Spacer (6) is
used only in the main landing gear wheel door actuator cylinders.
g. Remove O-ring (4) and back-up ring (3) from
gland end (5).
h. Apply a sharp blast of air to hydraulic port of
clevis end (22) to remove plunger (18), washer (11),
and race (10). Remove spring (21) from clevis end.
i. Remove and discard O-rings and back-up rings
from barrel, piston, and plunger.
5-23
•
".
.....
•••••••• 0: ••••••••••
..........
.....
. . - ~."
1.
2.
3.
4.
5.
6.
7.
5
Support
Forward Hinge Bracket
Aft Hinge Bracket
Aft Hinge
Left Door
Forward Hinge
Door Actuator
Figure 5-9.
5
Detail
(Refer to figure 5-10, stteet 1.)
NOTE
Install new O-rings and back-up rings
during cylinder assembly.
a. Install O-ring (19) and back-up ring (20) in
groove on plunger (18).
b. Insert spring (21) and plunger (18) into clevis
5-24
•
Main Gear Doors Installation (Sheet 2 of 2)
5-43. INSPECTION OF PARTS. Make the follOwing
inspections to ascertain that all parts are in a serviceable condition.
a. Inspect all threaded surfaces for cleanliness
and for freeness from cracks and excessive wear.
b. Inspect spring (21) for evidence of breaks and
distortion. The free length of the spring must be
1. 055 inches and compress to .875 inch under a 35
± 3. 5 pound load.
c. Inspect gland end (5), spacer (6), piston (7),
barrel (17), plunger (18), and clevis end (22) for
cracks, chips, scratches, scoring, wear or surface
irregularities which may affect their function or the
overall function of the door actuator cylinder.
5-44. ASSEMBLY.
A
end (22). Install washer (11) and race (10) over end
of plunger (18).
c. With knurled nuts (13) on barrel (17), install
O-rings (14) and back-up rings (15) in grooves on
barrel.
d. Install O-ring (9) and back-up rings (8) in groove
on piston (7) and install balls (12) in holes of piston.
e. Insert piston into barrel. Be sure that all six
balls are in place in piston as piston is inserted in
barrel.
f. Screw barrel (17) into cleviS end (22). Tighten
barrel down snugly against race, then tighten
knurled nut.
g. Insert space (6) in barrel (17). Spa.cer (6) is
used only in the main landing gear wheel door actuators.
h. Install O-ring (4) and back-up ring (3) in bore
groove of gland end (5), lubricate piston rod and
slide gland end over rod. Tighten gland end on
barrel, aligning hydraulic port fittings of the gland
end with the port fitting in the clevis end.
i. Tighten knurled nuts (13) to a torque value of
130 ± 10 lb. in. Install lockwire on both knurled nuts.
j. Install locknut (2) and rod end (1).
•
•
11
10
1
*6
5
2
34
13
•
1.
2.
3.
4.
5.
6.
7.
Rod End
Nut
Back- Up Ring
O-Ring
Gland End
Spacer
Piston
B.
Back- Up Ring
9. O-Ring
10. Race
11. Washer
12. Balls
13. Nut
14. O-Ring
15. Back- Up Ring
lB.
19.
20.
21.
22.
Name Plate
Barrel
Plunger
O-Ring
Back-Up Ring
Spring
Clevis End
Figure 5-10. Landing Gear Doo!" Actuator (Sheet 1 of 2)
5-45. DISASSEMBLY OF MAIN GEAR WHEEL DOOR
ACTUATOR. (Beginning with Serials 33701427 and
F33700036) (Refer to figure 5-10, sheet 2.)
a. Loosen check nut (2) and remove rod end (1) and
check nut from piston (7).
b. Remove retaining ring (3) from cylinder (9) .
c. Remove retainer (4), packing (5) and gland (6),
then remove piston (7).
d. Remove retainers (4) and packing (5) from piston
(7).
5-46. INSPECTION OF PARTS.
a. Inspect all threaded surfaces for cleanliness,
cracks, and excessive wear.
b. Inspect gland (6), piston (7) and cylinder (9) for
cracks, chips, scoring, wear or surface irregularities which might affect their function or the overall
function of the actuator.
NOTE
•
16.
17.
Repair of most parts of the actuator is
impractical. Replace defective parts
with serviceable parts. Minor scratches
may be removed by polishing with fine
abrasive crocus cloth (Federal Specification P-C -458), providing their removal
does not affect operation of the actuator.
5-47.
ASSEMBLY.
(Refer to figure 5-10, sheet 2.)
NOTE
Install all new packing and back-up rings
during assembly. Lubricate all packing
and back-up rings with Petrolatum or
MIL-4-5606 hydraulic fluid during assembly.
a. Install retainers (4) and packing (5) in grooves
of piston (7).
b. Insert piston assembly into cylinder (9).
c. Install packing (5) on gland (6); install on rod of
piston (7).
d. Install packing (5), retainer (4) and retaining
ring (3).
e. Install check nut (2) and rod end (1).
5-48. INSTALLATION OF MAIN GEAR WHEEL
DOORS AND ACTUATOR. (Refer to figure 5-9.)
a. Inspect door assembly, hinges, hinge brackets,
actuator end fittings, actuator, hydraulic lines and
attaching parts for distortion, cracks and damage
before installation.
b. Check hinge bushings, bearings and actuator
end fittings for lubrication prior to installation.
(Refer to Section 2. )
c. Assemble door, door hinge brackets and hinges
before installation, using reference marks made in
step "i" of paragraph 5-41. Do not tighten bolts.
d. POSition door hinges into wheel well hinge
5-25
•
4
4
NOTE
Lubricate packings before assembh
with Petrolatum or MIL-H-5606 .
hydraulic fluid.
1. Rod End
2. Check Nut
3. Retaining Ring
4. Retainer
5. Packing
6. Gland
7. Piston
8. Bearing
9. Cylinder
•
Figure 5-10. Landing Gear Door Actuator (Sheet 2 of 2)
brackets and install bolts, washers and nuts; tighten.
e. Tighten bolts in door hinges and brackets.
f. Manually open and close doors several times
checking for binding distortion and flair-in to surrounding structure and opposite door.
NOTE
When installing new doors, trimming and
hand-forming at the edges may be necessary
to achieve a good fit and to permit actuators
to lock. The doors must clear the gear by
at least l/2-inch during retraction.
g. Position inboard end of actuator into support.
Install bolt, washer and nut; tighten.
h. Position actuator rod end into door bracket.
Install bolt, washer and nut; tighten.
i. Connect hydraulic lines to actuator.
j. Bleed aircraft hydraulic system in accordance
with paragraph 5-163.
k. Rig the doors in accordance with paragraph 5-270.
1. Install access panels and upholstery.
m. Install rear or center seat.
5-49. REMOVAL OF MAIN GEAR STRUT DOORS
AND ACTUATOR. (Refer to figure 5-9.)
5-26
NOTE
Steps "a" thru "g" are for removal
of the actuator only.
a. Remove rear or center seat.
b. Remove upholstery and access panels from area
of landing gear bulkhead.
c. Jack aircraft in accordance with procedures outlined in Section 2.
d. Release hydraulic pressure.
e. Remove bolt securing actuator rod end to actuator arm.
f. Disconnect hydraulic lines from actuator and
plug or cap openings.
g. Remove bolt securing actuator to actuator support, and remove actuator from aircraft.
h. Before removal of doors, mark all parts on
door, linkage and attaching parts, to provide for alignment and location during reinstallation.
i. Remove bolts securing push-pull rods to right
and left arm assemblies.
j. Support door and remove bolts securing door
hinges to support structure, and remove door from
aircraft.
k. Remove bolts through actuator arm, torque tube
and shaft of left arm assembly.
1. Remove bolts through torque tube and shaft of
•
•
right arm assembly.
m. Remove bolts securing left bearing block to
structure; slide actuator arm inboard to clear left
arm assembly shaft.
n. Slide torque tube off shaft of right arm assembly
and remove torque tube, actuator arm and left bearing block from aircraft.
o. Remove four bolts securing bearing block to
structure; slide left arm assembly, with bearing block
attached, outboard, and remove from aircraft.
5-50. DISASSEMBLY, INSPECTION AND ASSEMBLY
OF MAIN GEAR STRUT DOOR ACTUATOR.
NOTE
Refer to paragraphs 5-42 thru 5-47 and
figure 5-10, sheets 1 and 2.
5-51. INSTALLATION OF MAIN GEAR STRUT DOORS
AND ACTUATOR. (Refer to figure 5 -9. )
NOTE
Steps "m" thru ''q'' are for installation
of the actuator only.
•
•
a. Inspect door assembly, hinge brackets, actuator
end fittings, actuator, hydraulic lines and attaching
parts for distortion, cracks and damage, before installation.
b. Check hinge bushings, bearings and actuator
end fittings for lubrication prior to installation.
(Refer to Section 2. )
c. Assemble bearing block and left arm assembly,
pOSition into structure and install four bolts using
reference marks made in step "h" of paragraph 5-49.
d. Assemble torque arm, left bearing block, actuator ar~ and right arm assembly with installing bolts.
e. Shde torque tube and actuator arm onto shaft of
left arm assembly; align holes and install bolts, using
reference marks.
f. Align holes in left bearing block to structure,
and install bolts loosely.
g. Align holes in torque tube and right arm assembly, and install bolts.
h. Tighten bolts in left bearing block and bearing
block-supporting left arm assembly to structure,
and tighten all loose bolts.
i. Assemble door, hinges, brackets and push-pull
rod. Use reference marks made in step "h" of paragraph 5-49; do not tighten bolts.
j. Position door hinges into support structure; install bolts and tighten.
k. Install bolts securing push -pull rods to right and
left arm assemblies, and tighten all loose bolts.
1. Manually open and close doors several times
checking for binding, distortion and flair-in to su-lrounding structure and opposite door.
m. Posit~on actuator into actuator support, install
bolt, and tIghten.
n. POSition actuator rod end into actuator arm, install bolt, and tighten.
o. Connect hydraulic lines to actuator.
p. Rig doors in accordance with paragraph 5 -270.
q. Install access panels, upholstery and seats.
5-52. MAIN LANDING GEAR WHEELS AND AXLES.
(Refer to figure 5 -11. )
5-53. DESCRIPTION. Each main gear wheel assembly consists of two wheel halves, two tapered roller
bearing assemblies, one tube, one tire, one steel
brake disc and attaching parts for each of the two
maingear wheel assemblies. Each main gear axle
assembly consists of one axle, one axle nut, wheel
alignment shims as required, and axle mounting
bolts and nuts.
5-54. OPERATION. The main gear wheels are freerolling on independent axles until the hydraulic brake
system is actuated.
5-55. REMOVAL OF MAIN GEAR WHEELS. (Refer
to figure 5-1. )
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. Remove hub cap retainer screws and hub cap.
c. Remove bolts securing brake back plates to brake
cylinder and remove back plates.
d. Remove cotter pin and axle nut.
e. Remove wheel from axle, using care not to damage axle threads.
5-56. DISASSEMBLY OF MAIN GEAR WHEELS.
(Refer to figure 5-11.)
!WARNING'
Injury can result if tire is not completely
deflated before attempting to separate wheel
halves.
a. Deflate tire completely and brake loose tire
beads from wheel flanges; use care to prevent damage
to wheel flanges.
b. Remove wheel thru-bolts and separate wheel
halves.
c. Remove tire, tube and brake disc.
d. If bearing cups are to be replaced, proceed as
follows:
NOTE
Bearing cups are a press-fit and should
be removed only if replacement is necessary.
1.
Heat wheel half in boiling water for 15
minutes.
2. Press out bearing cup and press in new
cup while wheel is still hot.
5-57. INSPECTION AND REPAIR OF MAIN GEAR
WHEELS.
a. Clean metal parts and grease felts in solvent
and dry thoroughly.
b. Inspect wheel halves for cracks; replace if damaged. Sand out nicks, gouges and corroded areas.
Where protective coating has been removed, clean
thoroughly, prime and repaint with aluminum lacquer.
c. If excessively warped or scored, or worn to a
thickness of O. 430-inch for the standard 6. OOX6, 8Ply wheel and brake assembly, or O. 325-inch for
5-27
•
~
/
13
•
NOTE
Some wheel brakes have
''kidney-shaped'' washer
installed under the head
of bolts (14).
1.
2.
3.
4.
5.
6.
7.
8.
9.
Snap Ring
Grease Seal Ring
Grease Seal Felt
Grease Seal Ring
Bearing Cone
Outboard Wheel Half
Nut
Washer
Tire
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
Tube
Inboard Wheel Half
Bearing Cup
Brake Disc
Bolt
Pressure Plate
Anchor Bolt
Brake Line Fitting
Washer
Nut
Figure 5-11. Main Wheel and Brake
5-28
20.
21.
22.
23.
24.
25.
26.
27.
28.
Bolt
Washer
Bleeder Valve
Brake Cylinder
Piston and O-Ring
Brake Lining
Torque Plate
Brake Lining
Back Plate
•
•
the optional 18.00 x 5.5, 8-Ply wheel and brake
assemb~y. brake disc should be replaced with a new
part. Sand smooth small nicks and scratches.
d. Replace damaged or discolored bearing cups and
cones. Mter cleaning, repack bearing cones with
clean wheel bearing grease before installation. (Refer
to Section 2 for grease type. )
5-58. ASSEMBLY OF MAIN GEAR WHEELS. (Refer
to figure 5-11.)
a. Insert tire in tube. Position outboard wheel half
in tire, aligning valve stem with hole in wheel half,
and align Slippage marks on tire and wheel.
b. Insert wheel thru-bolt through brake disc. Position disc in inboard wheel half, using thru-bolts as a
guide. Ensure disc is seated.
c. Place wheel halves together. Ensure tube is not
pinched, and secure with thru-bolts, washers and
nuts. Torque nuts to valve marked on wheel. Uneven
or improper torque may cause bolt failure with resultant wheel failure.
•
5-59. INSTALLATION OF MAIN GEAR WHEELS.
(Refer to figure 5-1.)
a. Slide wheel assembly on axle, using care to prevent damage to threaded surface of axle.
b. Screw axle nut onto axle and tighten until a
slight bearing drag is obvious when the wheel is rotated.
c. Loosen axle nut only enough to align to the nearest cotter pin hole and install cotter pin.
d. Install shim, brake back plate and cylinder bolts .
Safety wire bolt beads.
e. Install hub cap and retainer screws.
f. Remove aircraft from jacks.
5-60. REMOVAL OF MAIN GEAR WHEEL AXLES.
(Refer to figure 5 -1. )
a. Remove main gear wheel in accordance with
procedures outlined in paragraph 5-55.
b. Remove bolts securing axle, bushings and brake
torque plate to strut.
NOTE
Note number and position of wheel alignment
shims. Mark shims and axle so they may be
reinstalled in exactly the same position.
5-61. INSTALLATION OF MAIN GEAR WHEEL
AXLES. (Refer to figure 5-1.)
NOTE
Inspect axle for straightness and damage
to threads; replace if damaged or bent.
•
a. Insert mounting bolts through brake torque
plate, bushings, axle and alignment shims. Position
shims according to reference marks made at time of
disassembly.
b. Position axle assembly to strut. Install nuts
and tighten .
c. Install main gear wheel in accordance with paragraph 5-59.
5-62. MAIN GEAR WHEEL ALIGNMENT. (Refer to
figure 5-12. )
a. Alignment of maingear wheels is of primary importance in that misalignment adversely affects landing and take-off, roll characteristics, tire wear and
steering of the aircraft during ground operations.
Severe misalignment can cause malfunction and failure of some of the major components of the landing
gear system.
b. Alignment should be checked with landing gear
rigged correctly. (Refer to paragraph .) Removal
and installation of major main gear system components, evidence of uneven or excessive tire wear or
obvious damage to the system require a wheel alignment check, and correction, if necessary.
c. Alignment tolerances are set with the cabin and
fuel tanks empty, and will give apprOximately zero
toe-in and zero camber at normal gross weight.
d. If aircraft is normally operated at less than
gross weight, and abnormal tire wear access, realign
wheels to attain ideal setting for the load conditions
under which the aircraft normally operates.
e. Always use the least number of shims possiblp.
to obtain zero toe-in and zero camber at normal
operating load conditions.
l. To check wheel alignment tolerances, proceed
as follows:
1. Check to see that fuel tanks are empty
and aircraft is on a level surface.
2. Two aluminum plates, l/B-inch thick and
apprOximately IB-inches square, with grease applied
to their contacting sides, placed under each main gear
wheel, will allow the wheels to move free of friction,
between the tire and ground surface.
3. After placing greased plates under main
gear wheels, rock aircraft wings to allow wheels to
normalize.
4. Place a straight edge, long enough to extend between and approximately 12 -inches outboard of
each wheel, in front of the main wheels, and touching
the front-center of the tires.
5. Ensure that straightedge is level and
blocked up to just below wheel axle nut.
6. Place a carpenter's square against
straightedge and let it touch the wheel just below the
axle nut. Measure toe-in at edges of wheel flange.
The differences in measurements at both wheels is
the toe-in for one wheel (half of total toe-in). Toe-in
(total of both wheels) valves are contained in figures
1-1 and 5-11.
7. Place a protractor level vertically against
the outboard flanges of the wheel. If the top of the
wheel inclines inboard, a negative camber will result.
If the top of the wheel inclines outboard, a positive
camber reading will result. Positi ve camber should
be obtained. Camber valves are contained in figures
1-1 and 5-11.
B. Refer to paragraphs 5-60 and 5-61 for
procedures for removal and installation of axles and
shims.
5-63. WHEEL BALANCING. Since uneven tire wear
is usually the cause of wheel unbalance, replacing the
tire will probably correct this condition. Tire and
tube manufacturing tolerances permit a specified amount of static unbalance. The lightweight point of the
tire is marked with a red dot on the tire Sidewall, and
5-29
REFER TO FIGURE 1-1 FOR
CAMBER AND TOE-IN REQUIREMENTS.
•
PLACE CARPENTER'S ~UARE
AGAINST STRAIGHTEDGE AND
LET IT TOUCH WHEEL JUST
BELOW AXLE
ALUMINUM PLATES, APPROXIMATELY
18" ~UARE, PLACED UNDER WHEELS - - -.........\
GREASE BETWEEN PLATES
NOTE
. Rock wheels before
checking wheel alignment.
BLOCK STRAIGHTEDGE AGAINST
TIRES JUST BELOW AXLE HEIGHT
FRONT VIEW OF CAMBER CHECK
TOP VIEW OF TOE-IN CHECK
•
Measure camber by reading protractor level
held vertically against outboard flanges of
wheel.
Measure toe-in at edges of wheel flange. Difference in measurements is toe -in for one wheel.
(half of total toe-in. )
POSITIVE CAMBE"]
CARPENTER'S SQUARE
,
,,
I
I
....
FORWARD
INBOARD.
NOTE
Setting toe-in and camber within these tolerances while the cabin and fuel tanks are empty will give
approximately zero toe-in and zero camber at gross weight. Therefore, if normal operation is at
less than gross weight and abnormal tire wear occurs, realign the wheels to attain the ideal setting
for the load conditions. Refer to sheet 2 of this figure for shims availability and their usage. Always use the least number of shims possible to obtain the desired result.
Figure 5-12. Main Wheel Alignment (Sheet 1 of 2)
5-30
•
•
SHIM
PART
NO.
POSITION OF
THICKEST CORNER
OR EDGE OF SIDM
0541157-1
0541157-2
1241061-1
0441139-5
0441139-6
•
0541111-2
. CORRECTION IMPOSED ON WHEEL
TOE-IN
TOE-OUT
POS. CAMBER
NEG. CAMBER
AFT
.063"
----
0°4'
----
FWD
----
.063"
UP
DOWN
• 008"
----
UP& FWD
UP& AFT
DOWN & FWD
DOWN & AFT
UP& FWD
UP& AFT
DOWN & FWD
DOWN & AFT
UP& FWD
UP& AFT
DOWN & FWD
DOWN & AFT
UP& FWD
UP& AFT
DOWN & FWD
DOWN & AFT
----
.008"
----
0°28'
---.006"
----
0°28'
.006"
2°44'
2°46'
----
----
2°46'
2°44'
0°10'
0°25'
----
-------
0°25'
0°10'
.253"
0°21'
0°51'
-------
.235"
-------
0°51'
0°21'
1 °10'
1°51'
----
----
.125"
.117"
----
----
.117"
.125"
----
----
----
----
.235"
----
----
.253"
----
.375"
.323"
----
.375"
----
----
----
----
---.028"
----
.028"
0°4'
----
.323"
----
----
----
----
1°51'
1°10'
,
.
1241061-1
044113~-6
.0441139-5
0:>41157-2
0541157-1
0541111-2
1241061-1
0441139-6
0441139-5
0541157-2
0541157-1
0541111-2.
SHIM NO.
COLUMN 1
•
0
0
0
0
0
0
0
0
0
1
1
0
0
0
1
1
1
0
1
1
2
2
0
0
1
2
2
2
0
0
0
0
0
0
0
0
Max. number of
shims to be used
with shims in
column 1.
COLUMN 2
Figure 5-12. Main Wheel Alignment (Sheet 2 of 2)
5-31
5-64. BRAKE SYSTEM. (Refer to figures 5-11and
5-14. )
ed to the wheel cylinder assembly, run through the
torque plate, and allow the cylinder to move laterally
to compensate for lining wear. Brake linings are
bonded to the back plate and pressure plate with rivets.
The fixed shoes are on one side of the brake diSC,
while the movable piston and pressure plate are exactly opposite on the other side. The master cylinders
are connected to the rudder pedals, with plumbing
routed down the main gear struts to the wheel cylinders.
5-65. DESCRIPTION. The brake system is manually
actuated and hydraulically-operated. The wheel-mounted brake disc is straddled by a double hydraulic
piston assembly which mounts to a torque plate, anchored to the axle-attaching bolts. Two pins, moont-
5-66. OPERATION. The master cylinders are operated by pressing the toe portion of either the pilot or
copilot rudder pedals. The brakes are individually
actuated, and may be used to steer the aircraft while
taxiing.
the heavyweight point of the tube is marked with a contrasting color line (usually near the valve stem). When
installing a new tire and/or tube, place these marks
adjacent to each other. H a wheel becomes unbalanced
during service, it may be statically rebalanced.
Wheel balancing equipment is available from the Cessna Service Parts Center.
•
5 -67 . TROUBLE SHOOTING.
TROUBLE
DRAGGING BRAKES.
BRAKES FAIL TO OPERATE.
5-32
PROBABLE CAUSE
REMEDY
Brake pedal or linkage binding.
Lubricate pivot points; repair or
replace defective parts.
Weak or broken piston return
spring in master cylinder.
Repair or replace master cylinder.
Parking brake control improperly
adjusted.
Adjust properly.
Parking brake check valves
not releaSing (337A & on and
T337 Series).
Replace defective valves.
Insufficient clearance between
Lock-O-Seal and piston in
master cylinder.
Adjust per figure 5-13.
Restriction in hydraulic lines or
restricted passages in compensating sleeve in master cylinder.
Clean out restrictions. Flush
brake system with denatured alcohoL.
Repair or replace master cylinder.
Warped or badly scored brake
disc.
Replace brake disc and linings.
Damage or accumulated dirt
restricting free movement of
wheel brake parts.
Clean and repair or replace brake
parts as necessary.
Insufficient fluid in master cylinder or air trapped in brake
system.
Fill and bleed brakes.
Worn or damaged O-ring seal
in master cylinder or wheel
brake cylinder.
Replace O-rings.
Worn or damaged Lock-O-Seal
in master cylinder.
Replace Lock-O-Seal.
Too much clearance between
Lock-Q-Seal and piston in
master cylinder.
Adjust per figure 5-13.
•
•
•
•
5-67. TROUBLE Sl-JOOTING (Cont) .
TROUBLE
BRAKES FAIL TO OPERATE
(Cont).
PROBABLE CAUSE
Brakes too hot from -extensive
use.
Check that pistons are free after
overheating brakes.
Internally swollen hoses and/or
swollen O-rings due to use of
wrong kind of hydraulic fluid
in brake system.
Replace hoses and O-rings. Flush
system with denatured alcohol.
Fill and bleed with proper fluid.
Pressure leak in brake system.
Tighten loose connections; repair
or replace defective parts.
Brake linings worn out.
Replace brake linings.
Oil, grease, or other foreign
material on brake linings, or
new linings just installed.
Clean linings with carbon tetrachloride, then taxi the aircraft
slowly, applying the brakes several
times to condition the linings. New
linings must also be conditioned.
5-68. REMOVAL OF BRAKE MASTER CYLINDERS.
(Refer to figure 5-14.)
a. Drain hydraulic fluid from cylinder before removal.
b. Disconnect hydraulic lines and plug or cap opel'ings.
c. Remove pin securing clevis end of piston rod to
support bracket on right-hand cylinder and/or actuator arm on left-hand cylinder.
d. Remove pins securing cylinder to mounting bracket on left cylinder and/or actuating arm on right
cylinder.
e. Remove cylinder from aircraft.
5-69, DISASSEMBLY OF BRAKE MASTER CYLIN DER. (Refer to figure 5-13.)
a. Remove setscrew (11) securing cylinder cover
(10) into cylinder body (13).
b. Unscrew cylinder cover (10) and remove piston
assembly (9) and cover from cylinder, using care to
prevent damage to internal surfaces and parts.
c. Remove piston return spring (17) from cylinder.
d. Remove nut (1), piston spring (2), piston (3),
lock-O-seal (4) and compensating sleeve (5) from
piston rod (9), using care not to damage lock-O-seal.
e. Remove and discard O-ring (15) from piston (3).
f. Remove jam nut (8), clevis (7) and cover (10)
from piston rod (9).
g. Remove filler plug (6) from cover (10), and check that vent hole in plug is not restricted.
•
REMEDY
5-70. INSPECTION OF BRAKE MASTER CYLINDER.
a. Inspect threaded surfaces for damage, cracks
and excessive wear.
b. Inspect passages in compensating sleeve (5) for
restrictions. Inspect internal cylinder walls, piston
rod (9) and piston (3) for wear, scoring, scratches
or surface irregularities which may affect their function or the overall operation of the master cylinder.
c. Inspect springs for breaks or distortion and dimensions as follows:
Piston return spring (free length) 2-3/8 to 2-5/8 in.
Piston spring (free length)
.375 to .385 in.
5-71. ASSEMBLY OF BRAKE MASTER CYLINDER.
(Refer to figure 5-13.)
NOTE
Replace defective parts and O-rings prior
to asse mbly. Use clean hydraulic fluid
as a lubricant during asse mbly.
a. Install jam nut (8) and clevis (7) onto piston rod
(9) and insert into cover (10).
b. Install filler plug (6) in cover (10), and tighten.
c. Assemble piston rod (9), compensating sleeve
(5), lock-O-seal (4), piston (3), piston spring (2)
and nut (1), maintaining 0.040 :I: 0.005-inch spacing
between lock-O-seal and piston. (Refer to cutaway
section in figure 5-13.)
d. Install piston return spring (17) toward piston
(3), insert into cylinder, using care to prevent damage and to ensure that piston return spring is seated
into bottom of cylinder.
e. Screw cover into cylinder snugly and tighten
setscrew (11).
5-72. INSTALLATION OF BRAKE MASTER CYLINDER. (Refer to figure 5-14.)
a. Position master cylinder cleviS into support
bracket for right -hand cylinder and/or into actuating
arm for left-hand cylinder. Install pins, washers
and cotter pins.
b. Position lower end of master cylinder into actuating arm on right-hand cylinder and/or into mounting
bracket for left-hand cylinder. Install pins.
c. Connect and tighten hydraulic lines.
d. Remove filler plug and fill reservoir with clean
hydraulic fluid.
e. Bleed brake system in accordance with paragraph
5-73.
5-33
•
7
•
7
&
3
•
9
t---2
10
11
I
~
I
12
1
17
13
5
14
4
15
3
2
1
1&
11
•
0.040 ± 0.005 INCH
DO NOT DAMAGE
LOCK-O-SEAL
ADJUSTMENT OF PISTON
1. Nut
2. Piston Spring
3. Piston
4. Lock-O-Seal
5. Compensating Sleeve
6. Filler Plug
7. Clevis
B. Jam Nut
9. Piston Rod
10. Cover
11. Setscrew
12. Cover Boss
Figure 5-13. Brake Master Cylinder
5-34
13.
14.
15.
16.
17.
lB.
Body
Reservoir
O-Ring
Cylinder Chamber
Piston Return Spring
Screw
•
•
5-73. BLEEDING BRAKE SYSTEM. (Refer to figures
5-11 and 5-14.)
Ensure parking brake is OFF before
bleeding brake system.
a. Connect a clean hydraulic pressure source to
one wheel cylinder bleeder valve.
b. Remove filler plug in master cylinder on same
side as wheel cylinder, and install a suitable fitting,
with flexible hose attached, into filler hole.
c. Immerse the free end of the flexible hose in a
container, with enough clean hydraulic fluid to cover
end of hose.
d. Loosen the wheel cylinder bleeder valve, and unscrew approximately one turn.
e. As fluid is pumped into the system, observe the
immersed end of the hose at the master cylinder for
evidence of air being forced from the system.
f. When air bubbling has ceased, tighten the bleeder valve.
g. Remove the hydraulic pressure source and install
bleeder valve cover.
h. Remove the fitting, with flexible hose attached,
from master cylinder, and install filler plug.
i. Repeat the preceding procedures for the oppOSite
wheel cylinder and brake master cylinder.
•
5-74. REMOVAL OF WHEEL BRAKES. (Refer to
figure 5-1. )
a. Drain hydraulic fluid and disconnect brake hose
from wheel cylinder assembly.
b. Remove main gear wheel in accordance with paragraph 5-55.
c. Slide brake cylinder out of torque plate, and remove pressure plate from anchor bolts.
e. Inspect cylinder bore for scoring or surface defects. Replace if defective.
f. Inspect anchor bolts for nicks and gouges. Minor nicks and gouges may be sanded smooth to prevent
binding with pressure plate or torque plate.
g. If anchor bolts are to be replaced, they should
be pressed out, and new bolts installed by tapping
them in with a non-metallic hammer.
h. If excessively warped or scored, or worn to a
thickness of O. 430-inch for the standard 6. 00 x 6, 8Ply wheel and brake assembly, or O. 325-inch for
the optional 18. 00 x 5.5, 8-Ply wheel and brake as sembly, brake disc should be replaced with a new
part. Sand smooth small nicks and scratches.
5-76. ASSEMBLY OF WHEEL BRAKES. (Refer to
figure 5-11.)
a. Lubricate all internal wheel brake cylinder
parts with clean hydraulic fluid.
b. Install O-rings, and install pistons in cylinder.
c. Place pressure plate on anchor bolts.
d. Assemble brake disc to wheel. (Refer to paragraph 5-58.)
5-77. INSTALLATION OF WHEEL BRAKES. (Refer
to figure 5-1.)
a. If torque plate was removed, reinstall torque
plate, bushings and axle in accordance with paragraph
5-61.
b. Position wheel brake cylinder anchor bolts through torque plate.
c. Install wheel to axle in accordance with paragraph 5-59 .
d. Connect brake hose to wheel cylinder fitting.
e. Bleed brakes in accordance with paragraph 5-73.
5-78. BRAKE LINING REPLACEMENT.
figure 5-11.)
NOTE
To remove torque plate, it is necessary
to remove axle assembly. (Refer to paragraph 5-60.)
5-75. DISASSEMBLY OF WHEEL BRAKES. (Refer
to figure 5-11.)
a. Disassemble main gear wheel to remove brake
disc. (Refer to paragraph 5-56.)
(WARNING'
When using carbon tetrachloride, work in a
well ventilated area, and wear rubber gloves.
b. Clean all metal parts with carbon tetrachloride
and dry thoroughly.
c. Remove and discard all O-rings. Install new 0rings during assembly.
(Refer to
NOTE
It is not necessary to remove wheels to
reline brakes.
a. Remove bolts (20), washers (21) and back plates
(28).
b. Pull brake cylinder from torque plate (26) and
slide pressure plate (15) off anchor bolts (16).
c. Place back plate (28) on a table with lining side
down flat. Center a 9/64-inch (or slightly smaller)
punch in rolled rivet, and hit the punch crisply with
a hammer. Punch out all the ri vets securing the
linings to the back plates (28) and pressure plate (15)
in the same ma-nner.
NOTE
A rivet setting kit, Part No. R561, is available
from the Cessna Service Parts Center. This
Kit consists of a small anvil and punch.
NOTE
•
Brake linings should be replaced when th'~'-· are
worn to a minimum thickness of 3/32-in(,,,
d. Check brake linings for damage and maximum
permissible wear.
d. Clamp the flat sides of the anvil in a vise.
e. Align new lining (27) on back plate (28), and place
brake rivet in hole with rivet head in lining. Place
ri vet head against anvil.
f. Center the rivet setting punch on the lips of the
rivet. While holding down firmly against the lining,
5-35
hit the punch with a hammer to set the rivet. Repeat blows on the punch until lining is firmly against
back plate. Realign the lining on the back plate and
install remaining rivets.
g. Install a new lining on the other back plate and
pressure plate (15) in the same manner.
h. Position pressure plate (I5) on anchor bolts (16),
and place cylinder (23) in position so anchor bolts
slide into torque plate (26).
i. Install back plates (28) with bolts (20) and washers
(21). Safety the bolts.
5-79. PARKING BRAKE SYSTEM. (Refer to figure
5-14. )
5-80. DESCRIPTION. (Prior to 337-0240) The
parking brake system utilizes a handle and ratchet
mechanism, connected by cables to linkage at the
brake master cylinders.
5 -81. OPERA TION. Turning and pulling out on the
handle depresses both master cylinder piston rods.
The ratchet locks the handle In this pOSition until the
handle is turned and released.
5-82. REMOVAL. (Refer to figure 5-14.)
a. Remove cotter pin and pin (7); remove control
(4).
b. Remove boits, spacers (5), washers and nuts
from clamp (11) and tab on housing (9).
c. Turn handle (12) to clear catch (10), and remove
housing (9) and tube (8) from aircraft.
d. Remove pin and cotter pin attaching control cable
end (4) from forward end of bellcrank (2).
e. Remove bolt, nut, washers and spacer (1) attaching bellcrank (2) to mounting brackets attached to
channel.
f. Remove pin and cotter pin attaching clamp (3) to
bellcrank (2); remove bellcrank (2).
g. Loosen nuts securing parking brake control (4) to
angle on forward end of channel; remove cable (4)
from slots in angles.
h. Remove clamps (14) at master cylinders and disconnect control at master cylinders.
5-83. INSTALLATION. (Refer to figure 5-14.)
a. Install control cable (4) at master cylinders and
install clamps (14) loosely.
b. Route control cable (4) through slots in angle on
forward end of channel; attach clamp (3) to lower tabs
of bellcrank (2) with pin and cotter nino
. c. Install bolt, spacer (1), bellcrank (2), washers
and nut to mounting brackets attached to channel.
d. Attach control cable end (4) to forward end of
bellcrank (2).
e. Install handle (12), housing (9) and tube (8), and
install bolts, spacers (5), washers and nuts to clamp
(11) and tab on housing (9).
f. Install control (4) in slot of tube (8), and install
pin (7) and cotter pin.
g. Turn and pull handle (12) to engage parking brake.
Shift cable housings in clamps at master cylinders
and adjust nuts at slots in angle on channel for cable
adjustment.
h. Tighten all attaching hardware and clamps. Ensure that catch (10) engages in slot in housing (9).
5-36
5-84. DESCRIPTION. (337-0240 thru 337-0931.)
The parking brake system consists of two parking
brake valves, a single control cable, attaching parts
and connecting lines, hose and linkage.
5-85. OPERATION. pulling out on the handle engages both valves, each of which is connected to a
brake master cylinder.
•
5-86. REMOVAL. (Refer to figure 5-14.)
a. Remove access panel at forward left-hand side
of cabin, under instrument panel.
b. Drain hydraulic brake fluid.
c. Disconnect hydraulic lines and hose from valves
and cap or plug openings.
d. Disconnect control cable from clamps on valves.
e. Remove bolts securing valves to structure and
remove valves from aircraft.
5-87. INSTALLATION. (Refer to figure 5-14.)
a. POSition valves to structure; install mounting
bolts and tighten.
b. Connect hydraulic lines and hose to valves and
tighten.
c. Connect control cable to valve lever arms.
d. Fill brake systems with clean hydraulic fluid
and bleed systems in accordance with paragraph 5-73.
e. Rig parking brake control in accordance with
paragraph 5 -88.
5-88. RIGGING PARKING BRAKE. (Refer to figure
5-14. )
a. Push parking brake control full IN. Then, pull
OUT l/4-inchfor cushion, and lock in this position.
b. Connect control cable to aft valve lever, while
lever is full aft, with approximately l/2-inch of housing protruding through the clamp.
c. Attach control cable to forward valve lever,
while lever is full forward.
d. Check that arm on valve has full travel for OFF
and ON posi tion.
•
5-89. DESCRIPTION. (Beginning with 337-0932)
The parking brake consists of a parking brake valve
a control cable, attaching parts and connecting lines,
hose and linkage.
5-90. OPERATION. The parking brake c.ontrol cable
actuates the parking brake valve. When the control
is full IN, the valve must release pressure. When the
control is pulled OUT, the parking brake valve must
trap hydraulic pressure in its corresponding brake
system as the brake pedals are operated. The parking brake valve utilizes a spring attached to the valve
lever arm to ensure unlocked brakes.
5-91. REMOVAL. (Refer to figure 5 -14. )
a. Remove access panel at forward left-hand side
of cabin, under instrument panel.
b. Drain hydraulic fluid.
c. Disconnect hydraulic brake lines and hose from
valve and cap or plug openings.
d. Remove bolts securing valve to structure and
remove valve from aircraft.
5-92. INSTALLATION. (Refer to figure 5-14. )
•
.. ,
.~
DetailB
A
B
O:'-,D ,\
D
\
,
/
'.
;o~,
/~/,' 1
~, "-!
. ~"k:
~
'0
.~
'r."'fO
,~ r
•
~
13
12/
5
9
10
.~1l1
• ""l.
i
'I
"-<>,
1
LJ,·
5
~ ,
4
6
11
•
19
1. Spacer
2. Bell c rank
3. Clamp
4. Parking Brake Control
5. Spacer
6. Pulley
7. Pin
B. Handle Tube
•
9. Housing
Pilot's rudder pedal sup10. Catch
ports are welded to rudder
11. Clamp
bars at 337-0112 and on,
12. Handle
and all service parts _ _...JI
13. Roll Pin
14. Clamp
15. Pivot Pin
16. Spring
17. Left Brake Link
lB. Left Master Cylinder
19. Right Brake Link
20. Right Master Cylinder
PRIOR TO MODEL 337 A
Detail
A
Figure 5-14. Brake System (Sheet 1 of 2)
5-37
•
.....
........
....
I
roY'
I ~
2
TO RIGHT
MASTER CYLINDER
Details
A
•
TO LEFT MASTER
CYLINDER
7
MODEL 337A AND T337 SERIES
THRU SERIAL NO. 337-0931
TO RIGHT
MASTER CYLINDER
NOTE
SERIAL NO. 337-0932
ANn ON
1.
2.
3.
4.
5.
Parking Brake 'Control
Aft Valve
Aft Valve Lever
Forward Valve Lever
Forward Valve
6.
7.
8.
9.
Parking Brake Valve
Bracket
Valve Lever
Valve Lever Stop
Prior to serial number 337-0347 the
parking brake control was attached to
the forward lever (4) with a clevis pin.
Beginning with serial number 337-0347
the control is attached to the lever with
a bolt, spacers, washers, and a selflocking nut
Figure 5-14. Brake System (Sheet 2 of 2)
5-38
•
a. Position valve to structure; install and tighten
bolts.
b. Connect hydraulic lines and hose to valve and
tighten.
c. Connect control cable to valve lever arm.
d. Fill brake systems with clean hydraulic fluid
and bleed systems in accordance with paragraph 5-73.
e. Rig parking brake control in accordance with
paragraph 5 -93.
5-95. DESCRIPTION. The nose gear consists of a
pneudraulic shock strut assembly, mounted in a trunnion which pivots in heavy-duty needle bearings, a
steering collar, shiming dampener, up lock and downlock mechanisms, steering cam and lock, nose wheel,
tire, tube, hub caps, bearings, seals, and a doubleacting hydraulic actuator for extension and retraction.
A separate, single-acting hydraulic actuator unlocks
the uplock hook.
5-93. RIGGING PARKING BRAKE. (Refer to figure
5-14. )
a. Push parking brake control full IN. Then pull
OUT 1/4-inch for cushion, and lock in this position.
b. Connect control cable to valve lever with lever
against stop.
c. Check that arm on valve has full travel for OFF
and ON positions. Shift control housing in clamps
as required to obtain correct travel.
5-96. OPERA TION. When the gear control 'handle is
moved into the gear-up pOSition, the nose gear retracts
forward and upward to its stowed position beneath the
front engine. The steering collar at the top of the
strut contains rollers which engage tracks to cause
the nose gear to rotate 90° during retraction, so that
the nose wheel lies flat while in the retracted position.
The nose gear actuator contains the nose gear actuator
at the aft end. Initial movement of the actuator disengages the downlock before retraction begins. The nose
gear uplock hook is released by the uplock actuator
before gear extension begins.
5-94. NOSE GEAR SYSTEM. (Refer to figure 5-15.)
5-97. TROUBLE SHOOTING.
TROUBLE
•
REMEDY
UNEVEN OR EXCESSIVE
TIRE WEAR.
Loose torque links.
Add shim washers and replace
parts as necessary .
HYDRAULIC FLUID
LEAKAGE FROM NOSE STRUT.
Defective strut seals and/or
defects in lower strut.
Replace defective seals; stone out
small defects in lower strut. Replace lower strut if badly scored
or damaged.
NOSE STRUT WILL NOT
HOLD AIR PRESSURE.
Defective air filler valve
or valve not tight.
Check gasket and tighten loose valve.
Replace defective valve.
Defective O-ring at top of
strut.
Replace O-ring.
Result of fluid leakage at
bottom of strut.
Replace defective seals; stone out
small defects in lower strut. Replace lower strut if badly scored
or damaged.
Nose strut attachment loose.
Secure attaching parts.
Shimmy dampener lacks fluid.
Service shimmy dampener.
Defective shimmy dampener.
Repair or replace dampener.
Loose or worn steering components.
Tighten loose parts; replace if
defective.
Loose torque links.
Add shim washers and replace
parts as necessary.
Loose wheel bearings .
Replace bearings if defective;
tighten axle nut properly.
Nose wheel out of balance.
Balance nose wheel.
NOSE WHEEL SHIMMY.
•
PROBABLE CAUSE
5-39
SERIAL 337-0501 & ON
5
SERIAL 337-0393 THRU 337-0500 &
ALL PRIOR SERIALS USING SK337-4
•
9
•
+.06"
7.12" -.00"
Section
A-A
THRU SERIAL 337-0392
NOTE
When installing new upper torque
link, remove material from lug
on torque link as required to obtain specified dimension at full
extension of strut.
1.
2.
3.
4.
5.
6.
Nose Gear Actuator
Downlock Mechanism
Roller
Steering Collar
Roller
Uplock Roller
7.
8.
9.
10.
11.
12.
Trunnion Assembly
Needle Bearing
Inner Race
Lower Strut
Fork
Wheel
Figure 5-15. Nose Gear
5-40
13.
14.
15.
16.
17.
18.
Lower Torque Link
Upper Torque Link
Safety Switch
Cover
Welded Roller Support
Clamped Roller Support
•
•
5-98. REMOVAL OF SHOCK STRUT AND TRUNNION
ASSEMBLY.
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. With master switch OFF, place gear control
handle in the gear-up position, and use emergency
hand pump to open nose gear wheel doors and to unlock downlock mechanism.
c. Remove floor covering on each side of tunnel
at firewall in cabin for access to trunnion pivot bolts.
d. Tag and disconnect leads to squat switch on lower
torque link, and remove wiring clamps along routing.
e. Remove bolts securing aft nose gear door links
to trunnion arms.
f. Remove nose gear wheel in accordance with
paragraph 5 -149.
IWARNING'
Do not unscrew air filler valve core unless
strut is completely deflated. Loosening the
filler valve or valve core while the strut is
pressurized can result in injury and will strip
the last few threads of the valve or valve core.
•
g. Deflate shock strut completely in accordance
with procedures outlined in Section 2.
h. Remove bolt securing nose gear actuator and
downlock mechanism to top of nose gear, and remove downlock mechanism from aircraft.
1. Remove trunnion pivot bolts through access
holes in rudder cable pulley brackets on each side
of tunnel at firewall in cabin area.
j. Work nose gear forward evenly, tapping with
a non-metallic mallet if necessary, and remove nose
gear from aircraft.
5-99. REMOVAL AND INSTALLATION OF TRUNNION. (Refer to figure 5-16.)
NOTE
After the nose gear has been removed,
. remove the trunnion as follows:
a. Deflate strut if it has not already been deflated.
(Refer to warning in paragraph 5-98. )
b. Remove bolt at top of strut.
NOTE
Since the upper bolt also secures the orifice
piston assembly inside the strut, use a 5/16inch diameter guide pin, 2-1/4 inches in length,
to drive out the bolt. Center the guide pin and
leave it in place to retain the orifice piston
assembly.
•
c. Remove steering collar and washers from top
of strut.
d. Pull upper strut down, out of trunnion.
e. Thrust bearing, at lower end of trunnion, may be
removed, if desired. Clean with solvent and lubricate
with MIL-G-81322A grease before installation.
f. Reverse the preceding steps to install trunnion.
NOTE
Service shock strut before installation.
5-100. REMOVAL AND DISASSEMBLY OF LOWER
STRUT. (Refer to figure 5-16.)
NOTE
This procedure may be used to separate
the upper and lower struts, leaving the
upper strut and trunnion installed in the
aircraft. Most shock strut seals and
parts subject to wear, may be replaced,
without nose gear removal and complete
disassembly.
a. Jack nose wheel a sufficient distance to permit
lower strut to be pulled froni upper strut. (Refer to
Section 2. )
b. Deflate strut completely in accordance with procedures outlined in Section 2. (Refer to warning in
paragraph 5 -98. )
c. Disconnect upper torque link from lower torque
link, noting positions of washers and spacer.
d. Disconnect leads from safety switch.
e. Remove lock ring from groove inside lower end
of upper strut. A small access hole is provided at
the lock ring groove to facilitate removal of lock ring.
NOTE
Hydraulic fluid will drain as lower strut
is pulled from upper strut.
f. Use a straight, sharp pull to separate the upper
and lower struts. Invert lower strut and drain remaining fluid.
g. Remove lock ring (24) and bearing (25) from top
end of lower strut.
h. Slide packing support ring (28), scraper ring (31)
retaining ring (32), and lock ring (33) from lower strut
noting relative position and top side of each ring; wire
together if desired .
i. Remove O-ring (27) from outer groove in packing support ring (28), Remove back-up ring and 0rings from inner groove in packing support ring.
j. Remove bolt, washer and nut attaching fork to
lower strut, and pull base plug (38) and assembled
parts out of lower strut. Remove O-rings and
metering pin from base plug.
NOTE
Nose gear fork and lower strut area press-fit
drilled on assembly. Separation of these
parts is not recommended, except for replacement of parts.
5-101. REMOVAL AND INSTALLATION OF LOCKING
COLLAR. (Refer to figure 5-16.) After removal of
lower strut, remove locking collar and related parts
at lower end of upper strut as follows:
a. Remove bolt securing upper torque link, cover
and electrical clamp for safety switch leads.
b. Remove upper torque link, noting position of
spacers and washers. Pull cover forward to remove.
5-41
/~:
3
,.,
4
14
/
15
5
/
A
•
.
1&
27
9
10
Section
A
11
I
1.
•
.cil ,
~~y
I /"'"
Roller
2. Bolt
,/'
3. Roller
4. Steering Collar
5. Nut
6. Washer
7. Bearing
B. Trunnion
9. Uplock Roller
10. Bolt
11. Bearing
12. Inner Race
13. Bearing
14. Filler Valve
15. O-Ring
16. Orifice Piston Support
17. Upper Strut
lB. Race
19. Thrust Bearing
20. RaCf;!
21. Locking Collar
22. Spring
23. Retaining Ring
24. Lock Ring
25. Bearing
......--39
NOTE
Two races (20) are used thru serial no. 337-0229,
and on -0240, -0241, and -0242. Due to a change
in the strut, one race (20) is used on all other
serials.
Figure 5-16. Nose Gear Shock Strut
5-42
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.
37.
38.
39.
40.
41.
Lower Strut
O-Ring
Packing Support Ring
O-Ring
Back-Up Ring
Scraper Ring
Retaining Ring
Lock Ring
Fork
Placard
Metering Pin
O-Ring
Base Plug
Nut
O-Ring
Bolt
•
2
13
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
Retaining Ring
O-Ring
Bearing Head
Bushing
Steering Cam
Plug
Lock-O-Seal
Barrel
Shimmy Dampener Support
Shaft
Piston
Roll Pin
Back-Up Ring
9
Figure 5-17. Shimmy Dampener
c.
d.
e.
f.
Remove retaining ring below locking collar.
Disconnect centering springs.
Slide collar down to remove.
Reverse prececing steps to install locking collar.
5-102. ASSEMBLY AND INSTALLATION OF LOWER
STRUT. (Refer to figure 5-16.)
a. Thoroughly clean all parts in solvent and examine them carefully. Replace all worn or defective
parts, and all rubber or plastic seals and rings with
new parts.
NOTE
•
Packing support rings with different width
inner grooves and various seals have been
used in the strut. On packing support rings
with the Wide groove, install a contoured
rubber back-up ring above and below the
O-ring. If strut is equipped with a packing
support ring having the narrow groove, install
one contoured rubber back-up ring belm'! the
O-ring. If any struts are found with Teflon
or leather back-up rings installed in the packing
support ring inner groove, replace with the
contoured back-up rings above and below the
O-ring.
b. Assemble and install the lower strut by reversing the procedures outlined in paragraph 5-100. Note
that bearing (25) must be installed with beveled edge
up (next to lock ring).
c. Used sparingly, Dow Corning DC-4 compound
is recommended for O-ring lubrication. All other
internal parts should be liberally coated with hydraulic fluid during assembly.
d. Sharp metal edges should be smoothed with #400
emery paper, then cleaned. Tape or other coverings
should be used to protect seals where possible. Remove after seals are past edges .
e. Cleanliness and proper lubrication, along with
careful workmanship, are important during shock
strut assembly.
f. When installing lock ring (33), l'osition the lock
5-43
ring so that one of its ends covers the small access
hole in the lock ring groove.
g. Temporary bolts or pins of correct diameter
and length are useful tools for holding parts in correct relationship during assembly and installation.
h. Service shock strut in accordance with procedures
outlined in Section 2 after installation.
d. Install O-rings (2) on bearing beads (3) and slide
into barrel.
e. Install outer retaining rings (1) into barrel and
check dampener piston and rod for binding by pushing
rod for full travel in both directions.
I. Fill and service dampener as outlined in Section
2; install lock-Q-seal (7) and plug (6).
5-103. NOSE GEAR SHIMMY DAMPENER. (Refer to
figure 5-17. )
5-110. INSTALLATION.
DESCRIPTION. The shimmy dampener, a
self -contained hydraulic cylinder, is attached to a
shimmy dampener support on top of the nose gear
tunnel, immediately forward of the firewall in the
engine compartment, and to the steering cam,
mounted to the steering cam support, also atop the
tunnel in the engine compartment.
5-104.
5-105. OPERATION.
When the steering system reacts too rapidly, the shimmy dampener maintains
pressure against the steering cam by means of a
piston which permits a restricted flow of hydraulic
fluid from either end of the cylinder to the other,
through an orifice in the piston.
REMOVAL (Refer to figure 5-17.)
a. Remove bolt attaching barrel (8) to steering cam
5-106.
•
(Refer to figure 5-17.)
a. Position rod end of shaft (10) into support bracket
(9). Install bolt, washers, nut and cotter pin.
b. Position mounting lug into steering cam (5) bracket.
c. Check for clearance between cylinder and structure, while turning nose gear wheel from side to side.
5-111. TORQUE LINKS.
(Refer to figure 5-18.)
5 -112. DESCRIPTION. The torque links align the
lower strut with the nose gear steering system, but
permit shock strut action.
5-113.
REMOVAL.
IWARNING'
ALWAYS DEFLATE NOSE GEAR SHOCK STRUT
BEFORE DISCONNECTING TORQUE LINKS.
(5).
b. Remove bolt attaching shaft (10) to shimmy dampener support (9).
c. Remove dampener from aircraft.
DISASSEMBLY. (Refer to figure 5-17.)
a. Push clevis end of shaft (10) to limit of travel
toward cylinder.
b. Remove plug (6) and lock-O-seal (7), using care
not to damage lock-O-seal. Drain hydraulic fluid from
barrel.
c. Remove retainer rings (1), O-rings (2) and bearing beads (3) from barrel ends.
d. Slide piston assembly from barrel.
e. Remove roll pin (12) from piston (11), and slide
piston from shaft.
5-107.
INSPECTION OF PARTS.
a. Clean metal parts with solvent, and dry thoroughly.
b. Inspect parts for cracks, excessive wear, scoring or surface defects which may affect their function
or the overall operation of the dampener.
c. Replace defective parts with new parts.
5-108.
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. Deflate shock strut completely as outlined in
Section 2.
c. Remove upper bolt from upper torque link.
d. Remove lower bolt from upper torque link, and
carefully remove torque link from strut.
e. Remove lower bolt from lower torque link, and
carefully remove torque link from strut.
•
INSTALLATION.
a. Hold lower torque link in place on fork and determine clearance between torque link and lug on
fork, using feeler gages.
b. If clearance exceeds 0.013 inch, install 0.010
inch shims as required.
c. Check torque link for freedom of movement
after torquing.
5-114.
NOTE
Torque upper and lower torque link attach
bolts to 33 -38 lb in, and install cotter pin on
outside of nut.
NOTE
Install new 0 -rings and lubricate internal
parts liberally with clean hydraulic fluid
during assembly.
5-109. ASSEMBLY.
(Refer to figure 5-17.)
a. POSition piston (11) on shaft (10) and install roll
pin (12).
b. Install O-ring (2) and back-up ring (13) on piston
(11) and slide piston and shaft into barrel (8). Use
care to prevent damage to O-ring.
c. Install inner retaining rings (1) in both ends of
barrel (8).
5-44
d. Install upper torque link and check for freedom
of movement after torqUing.
e. Connect upper and lower torque links and tighten
nut finger-tight. (Do not torque. )
I. If difficulty is encountered in mating the torque
links, remove the lower torque link and shift shim or
shims, if installed, as required to the opposite side.
g. If step "f" does not correct the problem, check
the lugs on the barrel and fork for misalignment.
h. Rig squat switch in accordance with paragraph
5-277.
•
I.
'
•
•
•
1.
2.
3.
4.
5.
Upper Torque Link
Grease Fitting
Spacer
Cover
Nose Gear Strut
6.
7.
8.
9.
10.
Bushing
Switch Actuator
Shim
Lower Torque Link
Safety Switch
11.
12.
13.
14.
15.
Nut
Tab Washer
Switch Bracket
Lockwasher
Nut
Figure 5-18, Torque Links
5-45
•
~~)+-
OPERATED BY UPLOCK
ROLLER ON NOSE GEAR
2
1
Detail
A
\ez:"'-.J.Hr-t--- 4
1.
2.
3.
4.
5.
6.
7.
•
Actuator Support Bracket
Uplock Hook
Uplock Actuator
Up Indicator Switch Bracket
Needle Bearing
Bearing Race
Up Indicator Switch
Figure 5-19. Nose Gear Uplock Installation
Fore and aft adjustment is provided by slotted holes
in the actuator mounting bracket.
NOTE
Grease fittings and torque link bushings should
not be removed except for replacement of parts.
Excessively worn parts should be replaced.
5-115. NOSE GEAR UPLOCK MECHANISM.
to figure 5-19. )
(Refer
5-116. DESCRIPTION. The nose gear uplock mechanism is a hydraulically-unlocked hook that is springloaded to the locked pOSition. The installation consists of one single -acting hydraulic actuator, one hook
assembly, one indicator switch and attaching parts.
5-117. OPERATION. The uplock hook engages a
roller on the upper left side of the nose gear strut.
5-46
5-118. REMOVAL. (Refer to figure 5-19.)
a. Remove pin securing uplock arm to actuator (3)
and disconnect leads to SWitch.
b. Remove bolt securing uplock hook (2) to structure, and remove hook from aircraft.
c. Disconnect hydraulic lines from act.mtor and
cap or plug openings.
d. Mark location of bolts securing actuator to
slotted holes in support (1). Remove bolts and actuator from aircraft.
e. Indicator switch (7) and bearings (5) may be
disassembled after removal from aircraft.
5-119. DISASSEMBLY, INSPECTION AND ASSEMBLY
•
•
NOTE
Refer to Section 2 for
lubrication requirements.
•
l.
2.
3.
4.
Nose Gear Actuator
Packing
Back- Up Ring
Nut
5.
6.
7.
8.
Hook
Thin Washer
Nose Gear Trunnion
Bolt
9. Thick Washer
10. Rod End Assembly
11. Nose Gear Down
Indicator Switch
Figure 5-20. Nose Gear Downlock Installation
OF NOSEGEAR UPLOCK ACTUATOR. Refer to paragraphs 5-27 thru 5-30 and figure 5-7.
•
5-120. INSTALLATION. (Refer to figure 5-19.)
a. Position actuator (3) to support. Locate in
slotted holes, aligning the marks made during removal.
b. Connect hydraulic lines to actuator.
c. Assemble needle bearing (5) and race (6) into
uplock hook assembly and lubricate bearings in accor-
dance with procedures outlined in Section 2.
d. Position uplock hook assembly to mounting holes,
and install bolt securely.
e. Install indicator switch (7) to bracket (4), and
connect leads.
f. Install pin securing actuator to uplock hook arm.
g. Rig nose gear uplock and bleed hydraulic system
in accordance with applicable paragraphs.
5-47
19
11
•
14
10
~.
2
1. Thin Washer
2. Hook
3. Spring Guide
4. Spring
5. Thick Washer
6. Shield
7. Rod End
8. Crossbar
9. Hook
10. Bolt
11. Locknut
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
Back-Up Ring
O-Ring
Pin
Roll Pin
Bearing End
Setscrew
O-Ring
Plston
Back-Up Ring
O-Ring
Balls
Bushing
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
Head
Spring
Back- Up Ring
O-Ring
Plunger
Washer
Race
Locknut
Barrel
Name Plate
Locknut
•
Figure 5-2L Nose Gear Actuator (Sheet 1 of 2)
5-121. DOWNLOCK MECHANISM.
5 -20. )
(Refer to figure
5-122. DESCRIPTION. The nose gear downlock is
a hook at the piston rod end of the nose gear actuator.
The installation consists of the hook assembly, indicator switch, lock pins and attaching parts to the nose
gear actuator and strut.
5-123. OPERATION. The hook, at the piston rod end
of the nose gear actuator, contains an internal lock
to hold mechanism over center. Adjustment is provided by the rod end of the actuator piston rod.
5-124. REMOVAL. (Refer to figure 5-20.)
a. Jack aircraft in accordance with procedures
outlined in Section 2.
b. Remove bolt securing actuator (1) and downlock
mechanism to top of trunnion (7) and remove down5-48
lock mechanism from aircraft.
c. Disconnect hydraulic lines from actuator and
cap or plug openings.
d. Remove bolt securing actuator to structure;
remove actuator from aircraft.
5-125. DISASSEMBLY OF NOSE GEAR ACTUATOR.
(Thru 33701426 and F33700035)(Refer to figure 5-21,
sheet 1.)
a. Unlock cylinder by applying hydraulic pressure
to port in head (24).
b. Loosen locknut (11) at end of piston rod and un screw parts (l thru 10) as an assembly from piston rod.
c. Mark barrel (32) and head (24) so that same end
of barrel may be reinstalled in head (24) when reassembling actuator. Remove safety wire from locknuts (31 and 34).
d. Remove setscrew (17) in bearing end (16) and
loosen locknut (34). While using a strap wrench on
•
•
NOTE
14
Before assembly, lubricate a-Rings
and Back- Up Rings with Petrolatum
or MIL-H-5606 hydraulic fluid.
11
•
17
21 20
18
1.
2.
3.
4.
5.
6.
7.
8.
9.
Bolt
Thin Washer
Hook
Crossbar
Rod End
Nut
Back-Up Ring
Packing
Pin
10. Roll Pin
ll. Bearing End
12. Packing
13. Piston
14. Back-Up Rings
15. Packing
16. Cylinder
17. Locknut
18.
19.
20.
21.
22.
23.
24.
25.
26.
Nut
Thin Washer
Thick Washer
Thin Washer
Hook
Thick Washer
Spring Guide
Spring
Thick Washer
Figure 5-21. Nose Gear Actuator (Sheet 2 of 2)
•
barrel (32), remove bearing end (16) from barrel.
e. Pull piston (19) from barreL using care to prevent loss of balls (22) as piston is removed from
barrel.
f. Remove setscrew (17) from head (24) and loosen
locknut (31). USing a strap wrench on barrel (32),
remove head (24) from barrel.
g. Remove O-ring (18) from head (24), and remove
plunger (28) and parts (25 thru 30), by applying a
sharp blast of air in the vent hole located in head (24).
h. Remove all a-rings and back-up rings.
i. Disassemble hook assembly.
5-126. INSPECTION OF PARTS.
Make the following
inspections to determine that all parts are in a serviceable condition.
a. Inspect all threaded surfaces for cleanliness
and for cracks and excessive wear.
b. Inspect spring (4) for breaks and distortion.
The free length of the spring must be 2.460 ± .080
inches, and compress to 2.00 inches under a 19. 5 ±
1. 95 pound load.
c. Inspect spring (25) for breaks and distortion.
The free length of the spring must be 1. 055 inches,
and compress to .875 inch under a 35.0 ± 3.5 pound
load.
d. Inspect hooks (2 and 9), spring guide (3), bearing
end (16), piston and stop assembly (19), barrel (32).
5-49
head (24) and bushing (23) for cracks, chips, scratches
scoring, wear or surface irregularities which may
affect their function or the overall operation of the
actuator.
e. Repair of most parts of the actuator assembly
is impractical. Replace defective parts with serviceable parts. Minor scratches and scores may be
removed by polishing with fine abrasive crocus cloth
(Federal Specification P-C -458), providing their
removal does not affect the operation of the unit.
NOTE
Install all new O-rings and back-up
rings during assembly of the actuator.
5-127. ASSEMBLY.
a. Install O-rhlg (27) and back-up ring (26) in groove
on plunger (28).
b. Insert spring (25) and plunger (28) into head (24).
Install stop washer (29) and race (30) over end of
plunger (28) and install O-ring (18) in groove in head
(24).
c. With locknut (31) in barrel, screw barrel (32)
into head (24) until tapped hole in head is aligned
with hole in barrel.
NOTE
Ensure that marked end of barrel is
installed in head (24). Barrel should
tighten against race to prevent any
movement between stop washer and race.
d. Install and tighten setscrew (17) in head (24).
Tighten locknut (31).
e. Install O-ring (21) and back-up rings (20) in
groove on piston; install balls (22) into holes of piston.
l. Insert piston into barrel. Ensure that all six
balls are in place in piston.
g. Install O-rings (18 and 13) and back-up ring (12)
into grooves in bearing end (16).
h. With locknut (34) on barrel, screw bearing end
(I6) on barrel until tapped hole in bearing end (16) is
aligned with hole in barrel (32). Install and tighten
setscrew in bearing end (16). Tighten locknut (34).
NOTE
Centerline of hook pins and centerline of
bushing hole must align within . 005 inch
with cylinder locked at a length of 13.580
± .031 inches from centerline of hook pins
to centerline of bushing (23) in head (24).
i. Install locknut (11) on end of piston. Assemble
and install hook assembly on piston.
NOTE
When assembling hook assembly, lubricate
as specified in Section 2.
5-128. DISASSEMBLY OF NOSE GEAR ACTUATOR.
(Beginning with 33701427 and F33700036)(Refer to
figure 5 -21, sheet 2. )
a. Unlock cylinder by applying hydraulic pressure
5-50
to port in cylinder (16).
b. Loosen nut (6) at end of piston rod. Unscrew
parts (1,2,3,4,5,26,24,23,22,21,20, 19, and 18) as
an assembly from piston rod. Remove nut (6) from
piston rod.
c. Remove safety wire from locknut (17); loosen
locknut (17), using spanner wrench, if necessary, and
unscrew cylinder (16) from bearing end (11).
d. Pull piston (13) from cylinder (16).
e. Remove packing (12) from bearing end (11).
l. Remove back-up rings and packings.
g. Disassemble hook assembly, noting relative
arrangement of parts for reassembly.
5-129. INSPECTION OF PARTS. Make the following
inspections to determine that all parts are in a serviceable condition.
a. Inspect all threaded surfaces for cleanliness and
cracks or excessive wear.
b. Inspect spring (25) for breaks and distortion.
The free length of the spring must be 2. 406 ± • 080
inches, and compress to 2.00 inches under a 19.8
± 2.0 pound load.
c. Inspect hooks (3 and 22), spring guide (24), bearing end (11), piston (13), cylinder (16) and bushing in
end of cylinder for cracks, scratches, scoring, wear
of surface irre!tularities which might affect their funetion or the overall operation of the actuator.
d. Do not remove pins (9) unless they are damaged
and should be repaired.
e. Repair of most parts of the actuator assembly
is impractical. Replace defective parts. Minor
scratches and scores may be removed by polishing
with fine abrasive crocus cloth (Federal Specification
P-C -458), prOViding their removal does not affect
the operation of the unit.
5-130. ASSEMBLY.
•
•
NOTE
Install all new packings and back-up
rings during assembly. Before assembly,
lubricate O-rings and back-up rings with
Petrolatum or hydraulic fluid.
a. Install back-up rings (14) and packing (15) in
grooves of piston (13).
b. Insert piston into cYl1Dder (16).
c. Install locknut (17) over threads of cylinder (16),
and screw cylinder into bearing end (11).
d. Install packing (8), back-up ring (7) and nut (6)
on threads of piston (13).
e. Tighten and safety locknut (17).
l. Assemble and install hook assembly on piston
(13).
NOTE
When assembling hook assembly, lubricate
as specified in Section 2.
5-131. INSTALLATION. (Refer to figure 5-20.)
a. Position nose -gear -actuator into support and
•
I.
,
Detail
A
19
11
•
6
21
Detail
*Flanges of bearings must
face inboard.
.
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
Aft Nose Gear Door
Hinge
Nose G~ar
Male Rod End
Female Rod End
Hinge
Actuator Mounting Bracket
Hinge
Bushing
10.
11.
12.
13.
14.
15.
16.
17.
18.
Actuator
Rod End
Bearing
Bearing Block
Bearing Lock Plate
Right Tube and Bellcrank
Left Tube and Bellcrank
Bearing Lock Plate
Bearing Block
19.
20.
21.
22.
23.
24.
25.
26.
27.
B
Bearing
Rod End
Bellcrank
Bushing
Spacer
Eyebolt
Bearing
Hinge
Right Nose Gear Door
Figure 5 -22. Nose Gear Door Mechanism
5-51
install bolt and nut securing forward end of actuator
to structure and tighten. Install cotter pin.
b. Position hook end of actuator to top of nose gear
trunnion; install bolt and tighten.
c. Connect hydraulic lines to actuator.
d. Bleed hydraulic system in accordance with paragraph 5-163.
e. Rig nose gear actuator in accordance with paragraph 5-274.
5-132. NOSE GEAR DOOR SYSTEM.
5-133. DESCRIPTION. (Refer to figure 5-22.) The
nosegear door system consists of a right and left forward door, an aft gear door, one double-acting hydraulic actuator, linked through a torque tube, to the forward doors, hydraulic connections and attached parts.
The aft door is connected by adjustable links to the
nose gear.
5-134. OPERATION. The aft nose gear door, linked
mechanically to the nose gear, opens as the nose extends and closes as the nose gear retracts. The forward nose gear doors open for extension and retraction
of the landing gear and close again, after the cycle is
completed, through movement of the hydraulic actuator.
5-135. REMOVAL OF AFT NOSE GEAR DOOR. (Refer to figure 5-22.)
a. Remove bolts securing links to aft nose gear door.
b·. Remove hinge pin from hinge, and remove door
from aircraft.
5-136. INSTALLATION OF AFT NOSE GEAR DOOR.
(Refer to figure 5-22. )
a. Position door hinge half into hinge on structure,
and install hinge pin.
b. Position door links into door brackets; install
bolts and attaching hardware and tighten.
NOTE
When installing new doors, trimming and
hand-forming at the edges may be necessary
to achieve a good fit and permit actuators to
lock. The doors must clear the gear by at
least 1/2-inch during retraction.
5-137. REMOVAL OF FORWARD DOORS AND ACTUATOR. (Refer to figure 5-22.)
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. With master switch OFF, place gear control
handle to gear-up pOSition, and operate emergency
hand pump until doors are open.
c. Release hydraulic pressure and remove pin securing actuator rod end (11) to right tube bellcrank (15).
d. Disconnect hydraulic lines from actuator and cap
or plug openings.
e. Remove bolt securing actuator to bracket and
remove actuator from aircraft.
f. Remove pins securing bellcrank rod ends to right
and left tube bellcranks.
g. Support door and remove hinge pivot bolts securing hinges to brackets and remove door from aircraft.
h. Remove bolts securing right tube bellcrank (I5)
5-52
to left tube bellcrank (16) and telescope together to
slide ends from bearing blocks (13), and remove
right and left bellcranks from aircraft.
1. Remove bolts securing bearing blocks to structure,
and remove bearing blocks from aircraft, noting position of bearing blocks to structure.
j. Inspect parts for damage, cracks and excessive
wear. Replace faulty parts.
•
5-138. DISASSEMBLY, INSPECTION OF PARTS AND
ASSEMBLY OF NOSE GEAR DOOR ACTUATOR. Refer
to paragraphs 5-42 thru 5-47 and figure 5-10.
5-139. INSTALLATION OF FORWARD DOORS AND
ACTUATOR. (Refer to figure 5-22.)
a. Position bearing blocks (13) to structure, in same
pOSition as noted during removal.
b. Lubricate parts during assembly and installation
as specified in Section 2.
c. Assemble right and left tube and bellcrank assemblies loosely and telescope together. Position
ends into bearing blocks and align holes in bellcrank
tubes. Install and tighten bolts.
d. Assemble door hinges to door. Position hinges
into brackets, install and tighten hinge pivot bolts.
e. Position bellcrank rod ends (20) to right and
left tube bellcranks. Install pivot pins and cotter
pins.
f. Manually move door to closed position and check
for binding in hinges and linkage.
g. If necessary, hand form and trim doors to fit •
h. Position actuator clevis end into bracket, install
and tighten bolt.
i. Connect hydraulic lines to actuator.
j. Position actuator rod end (ll) into right tube bellcrank (5) and install pivot pin, washer and cotter pin.
k. Bleed hydraulic system in accordance with paragraph 5-163.
1. Rig nose gear doors per paragraph 5-278.
m. Remove aircraft from jacks.
•
5-140. NOSE WHEEL STEERING SYSTEM.
'5-14l. DESCRIPTION. The system consists of a
steering cam and lock assembly, a push-pull rod,
bellcrank, linkage and attaching parts.
5-142. OPERATION. Steering is accomplished by
use of the rudder pedals. A spring-loaded pungee is
connected between the rudder arm and steering cam
by-a push -pull rod and bellcrank. The steering cam
turns the nose gear on the ground, but is locked in
neutral as the gear retracts. The bungec then acts
as a rudder trim bungee. The nose wheel is steerable
up to approximately 15° each side (f. neutral, after
which the brakes may be used for a maximum deflection of about 39° each side of neutral.
5-143. TROUBLE SHOOTING. (Refer to Section 9. )
5-144. REMOVAL OF NOSE WHEEL STEERING CAM.
(Refer to figure 5-23.)
a. Jack aircraft in accordance with procedures outlined in Section
b. With master switch OFF, place gear control
handle in gear-up position and operate emergency
hand pump until nose gear is retracted enough to
gain access to steering cam bolt.
•
I.
~~
6
I
I
/
5
/
.
•
2
I~AurloNl
16
17
1. Steering Cam Lock
2. Steering Cam
3. Spring
4. Spacer
5. Steering Cam Support
6. Rudder Bar
7. Spacer
8. Bungee
9. Rod End
10. Bellcrank
•
~
10
Figure 5-23.
Shims (19) should not be allowed
to increase nose gear actuator
locking or unlocking pressures.
11. Spacer
Bearing
Push-Pull Rod
Retainer
Boot
Clamp
Rod End
Bearing
Shim
Bumper
12.
13.
14.
15.
16.
17.
18.
19.
20.
Nose Wheel Steering System (Sheet 1 of 2)
5-53
NOTE
Remainder of nose gear steering system
is unchanged from that shown in sheet 1
of this figure.
•
337 -0501 & ON
AND SERVICE
PARTS
A
•
A
. 030"
± .010"
.020"
± .010"
A-A
CLAMPED ROLLER SUPPORT
Proper clearance between steering cam and
roller may be attained by adjusting roller
support, which is clamped to the nose gear,
up or down as required. It may be necessary
to file the steering cam when installing a new
cam.
WELDED ROLLER SUPPORT
Clearance between steering cam and roller is
non- adjustable on the welded type roller support. However, it may be necessary to file
the steering cam when installing a new cam.
When tightening clamp bolts, also be sure
to keep roller aligned with centerline of lock.
Figure 5-23. Nose Wheel Steering System (Sheet 2 of 2)
5-54
•
•
c. Disconnect door actuator rod end from right
tube bellcrank. (Refer to figure 5-22.)
d. Remove bolt securing shimmy dampener to
steering cam. (Refer to figure 5 -17. )
e. Remove bolt securing push-pull rod to steering
cam. (Refer to figure 5 -23. )
f. Remove bolt securing steering cam lock (1) and
spring (3) to steering cam support (5). Remove cam
lock and spring from aircraft.
g. Remove bolt securing steering cam (2) to cam
support (5) and remove cam from aircraft.
h. Push-pull rod (13), bellcrank (10) and bungee (8)
may be removed by removal of bolts at attach points.
i. Inspect all removed parts for damage and excessive wear.
loose from wheel flanges. Use care to prevent damage to wheel flanges .
b. Remove wheel thru-bolts (10) and separate
wheel halves (6 and 9. )
c. Remove tire (7) and tube (8).
5-151. INSPECTION OF NOSE GEAR WHEEL.
a. Clean all metal parts and grease seal felts with
solvent.
b. Replace damaged or discolored bearing cups (11)
and cones (5) (refer to figure 5 -24. )
c. After cleaning, repack bearing covers and cups
with wheel bearing grease before installation. (Refer
to Section 2. )
NOTE
•
5-145. INSTALLATION OF NOSE WHEEL STEERING
CAM. (Refer to figure 5-23.)
a. Position steering cam (2) to cam support (5). Install and tighten bolt.
b. Position cam lock (1), spring (3) and spacers (4)
to cam support (5). Install and tighten bolt.
c. Position shimmy dampener to steering cam bracket and install bolt. (Refer to figure 5 -17 . )
d. Position push-pull rod end (17) into steering cam
bracket and install bolt.
e. Installation of push-pull rod (13), bellcrank (10)
and bungee (8) may be accomplished by installing
bolts at attach points.
f. Position door actuator rod end to right tube
bellcrank and install pin. (Refer to figure 5-22. )
g. Rig nose wheel steering as outlined in Section 9 .
h. Remove aircraft from jacks.
5-146. NOSE GEAR WHEEL.
(Refer to figure 5-24.)
5-147. DESCRIPTION. The nose gear wheel assembly
consists of two cast wheel halves, two tapered roller
bearing assemblies, one tire, one tube, two hub caps
grease seals and attaching parts. The wheel is mounted to the fork of the nose gear strut on an axle.
5-148. OPERATION. The nose gear wheel is freerolling on an independent axle and is used to taxi the
aircraft during ground operations.
5-149. REMOVAL OF NOSE GEAR WHEEL. (Refer
to figure 5-24. )
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. Remove axle bolt (21).
c. Insert a long punch through one axle ferrule (16)
to tap out ferrule at opposite side of fork.
d. Remove both ferrules and pull wheel from fork.
e. Remove spacers (19) and axle tube (20) before
disassembling wheel.
5-150. DISASSEMBLY OF NOSE GEAR WHEEL.
fer to figure 5-24.)
Ir-=-W~A"""R=""'N~IN-=-G~a
•
Inju!:'y can result if tire is not completely
deflated before attempting to separate wheel
halves.
a. Deflate tire completely and break tire beads
(Re-
Bearing cups are a press -fit and should be
removed only if replacement is necessary.
d. If bearing cups are to be replaced, proceed as
follows:
1. Heat wheel half in boiling water for 15
minutes.
2. Press out bearing cup and press in new
cup while wheel is still hot.
e. Replace cracked wheel halves. Minor nicks,
scratches or scores may be sanded smooth.
f. Where protective finish has been removed,
clean, prime and repaint with aluminum lacquer.
g. Inspect tire and tube for damage; replace if damaged.
5-152. ASSEMBLY OF NOSE GEAR WHEEL. (Refer
to figure 5-24.)
a. Insert tire in tube. Position wheel half with hole
for valve stem in tire. Align valve stem with hole in
wheel half and carefully work valve stem through hole.
Align tire and tube balance marks per paragraph 5-58.
b. Place wheel halves together, ensuring tube is
not pinched.
NOTE
Uneven or improper torque of wheel thru-bolt
nuts can cause bolt failure with resultant wheel
failure.
c. Secure wheel halves with wheel thru-bolts and
torque to valve marked on wheel.
d. Install grease seals, bearing cones, snap rings
and hub caps.
e. Install tire to set tire beads, then adjust to
pressure specified in figure 1-1.
5-153. INSTALLATION OF NOSE GEAR WHEEL
(Refer to figure 5 -24. )
.
a. Assemble spacers (19) and axle tube (20) into
wheel.
b. Position wheel in fork (13) anti install ferrules
(16) into fork. Tap with non-metallic hammer until
seated.
c. Install axle bolt (21) and tighten until a slight
bearing drag is obvious, then back off axle nut (17)
to align nearest cotter pin hole and install cotter
pin (18).
d. Remove aircraft from jacks.
5-55
•
2
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
~2.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
Snap Ring
Grease Seal Ring
Grease Seal Felt
Grease Seal Ring
Bearing Cone
Wheel HaU
Tire
Tube
Wheel HaU
Thru-Bolt
Bearing Cup
Bolt
Fork
Washer
Nut
Bucket
Nut
Cotter Pin
Spacer
Axle Tube
Axle Bolt
Washer
Nut
Hub Cap
21
16-----
17
19
* serials
Bushings not used on later
or service parts,
which have smaller holes.
Figure 5-24. Nose Wheel
5-56
•
PRESS.
lACK
FUJW
VALVE
RET.
uNE
O'FLOW
DIVIDER VALVE
RETURN
UNE
RESERvom
•
PRESSURE
FILTER
PUMP
REGULATOR
Figure 5-25. Simplified Schematic of Hydro Test
5-154. LANDING GEAR HYDRAULIC POWER.
5-155. DESCRIPTION. (Refer to paragraph 5-2.)
5-156. OPERATION.
•
(Refer to paragraph 5-3.)
A hydraulic test unit may be assembled locally, if
desired. Specifications for a test unit are listed in
the following chart.
1.
Flow
1. 25 gpm
5-157. HYDRAULIC TOOLS AND EQUIPMENT.
2. Reservoir
I gallon
5-158. HYDRO TEST UNIT. A special portable hydraulic servicing unit is available from the Cessna
Service Parts Center. The Hydro Test unit combines a motor-driven pump, pressure jack, pressure
gage, reservoir and controls into a compact unit.
The Hydro Test, or its equivalent, is indispensible
for servicing, testing and rigging of the landing gear
system.
3. Check valve
Aft of pump in
pressure line
4.
3 gpm, 10 micro in
pressure line after
pump and before
relief valve.
When using the Hydro Test, make sure
personnel are in the clear before cycling
the landing gear. Apply hydraulic pressure
carefully; gear and door operations are
rapid when hydraulic flow is set near the
full capacity of the Hydro Test Unit. .
Filter
5. Relief Valve
Pressure line after
filter and discharging to reservoir.
6. Relief Valve
Setting
1700.0 crack to 15000'
psi (min) reseat.
7.
2000 psi dual on
pressure line and
snubbed.
Pressure Gage
5-57
PRESSURE GAGE
PUMP MOTOR SWITCHES
PRESSUREJACK:-~~____----~--~~~--FLOW VALVE
T\
•
VENT HOSE
SUCTION HOSE
BYPASS VALVE
FLOW INDICA TOR
PRESSURE HOSE
Figure 5-26. Hydro Test Unit
8. Temperature Gage
50 0 to 200 0 at pump
outlet.
9. Suction Hose and
Lines
-8 (1/2 inch tube
size)(min)
10. Pressure Hose and
Line
-4 (1/4 inch tube
size)(min)
d. Cap all hoses and stow on rack when not in use.
e. Avoid contamination of test stand fluid by checking condition of fluid in aircraft system before connecting test stand.
f. Before disconnecting test stand, check that aircraft reservoir is full; fluid may Siphon from aircraft
reservoir to test stand if idle for a period of time.
•
NOTE
ll.
Power Input
3 hp (desired) 2 hp
(min)
I~AUTIONl
Means should be provided to keep connections
to aircraft system clean and free of foreign
material at all times.
5-159. OPERATION.
a. Always open bypass valve before starting test
stand motor. This will permit motor to start under
a no-load condition and will contribute to the service
life of the test stand unit.
b. Operation of the test stand with bypass and lockout valves closed at the same time should not be continued for more than one minute.
c. Avoid continuous operation of the test stand under
high-pressure - low flow condition; this will cause
rapid heating of the fluid supply. When pressure is
no longer needed, open bypass valve to relieve pressure.
5-58
The Hydro Test unit is a precision test
instrument as well as a hydraulic power
source. The retention of its accuracy and
the length of its service life depends on good
care and proper operation.
5-160. FLOW REGULATION. The following procedure is used to adjust the test unit flow to any valve
desired for a specific operation, with the test unit
connected to the aircraft hydraullcsystem and the
aircraft on jacks.
a. Open bypass valve and lockout valve.
b. Start test unit pump motor.
c. Close bypass valve.
d. Open flow valve, then slowly close it until indicator in flow gage sight glass aligns with mark indicating desired flow. To read flow indicator, match
line on widest part of indicator with fixed line on
external part of gage.
5-161. CONNECTING TEST UNIT TO AIRCRAFT.
(Refer to figure 5-25.)
•
•
a. Remove front engine cowling for access.
b. Disconnect hydraulic pump suction hose from
firewall fitting, connect test stand suction hose to
fitting and cap disconnected pump hose.
c. Disconnect hydraulic pump pressure hose from
fitting in filter at firewall, connect test stand pressure
hose to the fitting and cap disconnected pump pressure
hose.
d. Connect test stand vent hose to aircraft reservoir
-venlline protruding below lower edge of firewall, using care to ensure that Une is Wiped clean and is free
of any foreign material. If line is dirty internally,
remove, clean and reinstall.
d. Connect test stand electrical cable to appropriate
power source.
5-162. DISCONNECTING TEST STAND FROM AmCRAFT. (Refer to figure 5-25.)
a. Check that landing gear is down and locked and
gear doors are closed.
b. With bypass valve closed and lockout valve open,
operate test stand until aircraft hydrauUc reservoir
is full; then open bypass valve and stop test stand
pump motor.
c. Disconnect all test stand hoses from aircraft
immediately, beginning with the suction hose. If the
suction hose is left connected, fluid will siphon from
the aircraft reservoir to the test stand reservoir.
d. Connect all aircraft hydrauliC lines and install
engine cowling.
•
5-163. BLEEDING AmCRAFT HYDRAULIC SYSTEM.
NOTE
There is only one reason for having to bleed
the hydraulic system: the entrance of considerable air into the hydraulic system. The
most probable means of air entering the system
are: permitting reservoir fluid level to become
too low; air leaks in the engine - driven pump
or pump suction line and poor maintenance
procedures when connecting fluid lines or replacing components.
a. Jack aircraft as outlined in Section 2.
b. Connect test stand in accordance with paragraph
5-161.
c. Use test stand to operate landing gear through
five complete cycles.
d. Use only clean filtered hydraulic fluid (MIL-H5606) to fill hydraulic system and test stand.
e. Hydraulic fluid preservative (MIL-H-6083) may
be used for flushing and storage of hydrauliC components.
5-164. USE OF TEST STAND TO LEAK TEST HYDRAULIC SYSTEM AND COMPONENTS. (Refer to
figure 5-25. )
a. Jack aircraft in accordance with procedures outlined in Section 2.
•
[CAUTIONI
When testing any actuator by applying pressure
to one port of the cylinder, always have the
opposite port open to atmospheric pressure,
otherwise excessive pressure may be built up
due to the differential in piston areas. The
rod side of the piston has less area than the
head side. All lines, fittings, actuators and
any other parts subjected to hydraulic dead-end
pressure in excess of 2275 psi for any length of
time shall be considered faulty due to over
streSSing and shall be replaced.
b. Connect test stand pressure hose to system or
component to be tested. Use suitable fittings to make
connection (refer to paragraph 5-161). The power
pack must be bypassed.
c. Set flow valve for minimum flow.
d. Set locknut valve cracked open.
e. Set bypass valve open.
f. Set pressure jack out approximately 1-1/2 inches.
g. Start test stand pump motor.
h. Slowly close bypass valve until pressure reaches
1950 psi.
i. Close lockout valve to trap fluid, then stop test
stand pump motor immediately.
j. Screw pressure jack in, increasing pressure to
2200 psi, and hold 5 minutes.
k. Check for leaks while system or component is
under pressure.
1. After completion of tests, open test stand lockout valve to relieve pressure and discmnect test unit
from system or component (refer to paragraph 5-162).
m. Remove aircraft from jacks.
5-165. CYCLING LANDING GEAR.
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. Connect test stand as outlined in paragraph 5-161.
c. Set test stand flow valve closed, lockout valve
open and bypass valve open.
d. Start test unit pump motor.
e. Slowly close bypass valve completely.
f. Observe fluid flowing through test unit sight
gage. When all air bubbles have diSSipated, operations may be continued.
g. Use landing gear control handle in aircraft to
operate the gear through cycles.
NOTE
Gear cycling can be prolonged by slowly
opening the test unit bypass valve part
way. This will bleed off part of the pump
flow.
h. After tests are completed, open test stand bypass valve and stop test stand motor.
1. Disconnect test stand in accordance with paragraph 5-162.
j. Remove aircraft from jacks.
5-59
2
1
POSITION O-RING
INSTALL NUT
•
COVER THREADS WITH A PLASTIC THIMBLE
OR TAPE,APPLY PETROLATUM TO O-RING,
THEN ROLL IT UP INTO POSITION AGAINST
NUT. REMOVE THIMBLE OR TAPE AFTER
O-RING IS IN POSITION.
THESE THREADS MUST Nor PROTRUDE
BELOW NUT. POSITION NUT EXACTLY
AT TOP OF NON-THREADED AREA.
3
4
INSTALL ELBOW IN THREADS UNTIL
O-RING CONTACTS CHAMFER,
AND NUT CONTACTS FACE OF BOSS
•
ROTATE NUT AND FITTING TOGETHER TO
RETAIN THE ORIGINAL POSITION OF THE
l'iUT ON THE FITTING.
HOLD NUT STATIONARY, TURN FITTING
TO DESIRED POSITION.
5
J I.
~INSTAll O.RINGS CAREFULLY. MOST HYDRAULIC LEAKS
ARE CAUSED BY CARELESS INSTAllATION.
Figure 5 -27. Installation of Hydraulic Fittings (Sheet 1 of 2)
5-60
•
•
1
2
INSTALL NUT
POSITION BACK-UP
RING & O-RING
APPLY PETROLATUM TO BACK-UP RING
AND 0- RING, THEN WORK THEM UP INTO
POSITION AGAINST NUT. TURN NUT DOWN
UNTIL O-RING IS PUSHED DOWN FIRMLY
AGAINST LOWER THREADS.
POSITION NUT WITH RECESS DOWN.
3
4
INSTALL ELBOW IN THREADS UNTIL
O-RING CONTACTS FACE OF BOSS
WITH NUT HELD, TURN FITTING IN
1~
TURNS
1-1/2 TURNS PLUS A
HYDRAULIC LINE.
ROTATE NUT AND FITTING TOGETHER TO
RETAIN THE ORIGINAL POSITION OF THE
NUT ON THE FITTING.
ATTACH LINE TO FITTING.
5
TIGHTEN NUT UNTIL IT CONTACTS BOSS
•
INSTALL O-RINGS CAREFULLY. MOST HYDRAULIC LEAKS
ARE CAUSED BY CARELESS INSTALLATION .
Figure 5-27. Installation of Hydraulic Fittings (Sheet 2 of 2)
5-61
5-166. CHECKING LANDING GEAR CYCLE TIME.
NOTE
When the hydraulic system is suspected of
malfunction because the landing gear
cycle time is slow, it could be caused by
low fluid in the power pack reservoir,
causing the hydraulic system to be full
of air. This procedure Will purge air
from the system and fill the reservoir.
a. Cycle landing gear through two complete cycles
in accordance with paragraph 5-165.
b. With landing gear extended, place gear handle in
full-up pOSition and record time required for gear to
retract and doors to close. Time should not exceed
ll. 5 seconds (plus 4 seconds, minus 2 seconds) plus
the time required for the time delay valve to operate
(3 to 9 secoOC:s at room temperature; colder t'!mperatures will cause a longer delay).
c. With landing gear retracted, place gear handle
in full-down pOSition and record time required for
gear to extend and doors to close. Time should not
exceed 10. 5 seconds (plus 4 seconds, - 2 seconds)
Olus the time required for the time delay valve to
operate (3 to 9 seconds at room temperature; colder
temperatures will cause a longer delay).
NOTE
If time is within limits when operated by
a test stand, but exceeds limits when
operated by the engine - driven hydraulic
pump, there is internal leakage in the
ITEM
pump. Repair or replace pump. If time
exceeds limits when operated either by
test stand or hydraulic pump, internal
leakage is in the hydraulic system. Check
actuators for internal leakage; refer to paragraph 5-164, and repair or replace actuators
as required. If actuators are not defective,
power pack internal leakage is indicated.
Repair or replace power pack.
•
5-167. HYDRO FILL UNIT. A portable special
filler can (Part No. SE350), with a manually-operated pump, is available from the Cessna Service Parts
Center. In addition to prOviding a handy means of
filling hydraulic reservoirs, the unit may be used to
bleed brake systems.
5-168. INSTALLATION OF HYDRAULIC FITTINGS.
(Refer to figure 5-27.) Most hydraulic leaks are
caused by careless installation of O-rings and
fittings. The figure illustrates correct methods of
installing hydraulic fittings and may be used as a
guide during removal and installation of hydraulic
system components.
5-169. HYDRAULIC SYSTEM COMPONENTS.
5-170. GENERAL DESCRIPTION. The hydraulic
power system includes equipment required to provide
a flow of pressurized hydraulic fluid to the retractable
landing gear system. Main components of the hydraulic system are listed in the folloWing chart. A detailed
description and removal, disassembly, assembly and
installation procedure for each component is included,
beginning with paragraph 5-177.
PURPOSE
LOCATION AND ACCESS
ENGINE -DRIVEN
HYDRAULIC PUMP
To provide a flow of pressurized
hydraulic fluid to the system.
of starter. Remove upper cowling.
HYDRAULIC FILTER
To filter fluid from the pump before entering remainder of system.
Upper left side of front firewall. Re
move upper engine cowling.
HYDRAULIC POWER
PACK
(1) To "load" the engine-driven
pump when landing gear handle is
moved out of neutral.
Aft left side of front firewall, behind
instrument panel.
•
Front engine accessory section, aft
(2) To provide a reservoir of
hydraulic fluid.
(3) To afford control of gear and
door systems through use of
valves and appropriate passages.
EMERGENCY HAND
PUMP
5-62
To provide emergency hydraulic
pressure through use of hand
pump.
Floorboard, just forward of front
seats. Remove cover.
•
•
DOOR CLOSE LOCK
VALVE
(BEGINNING WITH
33701427 & F33700052)
To hold wheel door actuators in
the closed position by pressure
trapped in the door close line.
5-171. HYDRAULIC COMPONENT REPAIR. Since
emphasis here is on repair and not overhaul of the
basic components of the hydraulic system, it is unlikely that the mechanic will go through all of the procedures outlined. Instead, he will repair the particular item which is causing the difficulty.
NOTE
To isolate the item causing the malfunction,
refer to the trouble shooting charts in paragraph
5-6, and if pOSSible, check with test stand.
5-172. REPAIR VERSUS REPLACEMENT. Often
the moderate trade -in price for a factory - rebuilt
component is less than the accumulated cost of
labor, parts, and (often time-consuming) trial and
error adjustment. Repair or replacement of a component will depend on the time, equipment and skilled
labor that is locally available.
•
LOCATION AND ACCESS
PURPOSE
ITEM
5-173. REPAIR PARTS AND EQUIPMENT. Repair
parts may be ordered from the applicable Parts Catalog. Test equipment may be ordered from the Special
Tools and Support Equipment Catalog. Both publications are available from the Cessna Service Parts
Center.
5-174. EQUIPMENT AND TOOLS.
5-175. HAND TOOLS. The following hand tools are
necessary for repair work on the power pack and
other hydraulic components.
Snap Ring Pliers
Strap Wrench (for removing door solenoid
and various cylinder barrels of the hydraulic actuators)
Pin Punches
Duck-bill Pliers
Box and Open-end Wrenches
Needle-nose Pliers
In left-hand engine comparbnent;
remove left engine cowl.
pair, are various 1/4" ~luminum rods, ground to a
gradual taper, and hooks, formed from brass welding
rod, to extricate small plungers from hydrauliC ports.
The hook, formed on brass welding rod, must not be
over 1/16 - inch in length, so as not to scratch or score
the bore. Various sizes of Allen wrenches may be
welded .or brazed to "T" handles for use when removing, installing or adjusting the various internal wrenching plUJs or valves.
5-176. COMPRESSED Am. The easiest way to remove some hydraulic parts in inaccessible galleries
of the power pack is a quick blast of compressed air
from behind. Parts can be blown out in seconds,
which would otherwise take endless "fishing" operations to extricate. An air hose and nozzle are common-sense tools.
5-177. ENGINE-DRIVEN HYDRAULIC PUMP.
fer to figure 5-28.)
(Re-
5-178. DESCRIPTION. The engine-driven hydraulic
pump is a gear-type pump, and is mounted on the
right rear accessory pad of the front engine.
5-179. OPERATION. The pump is driven at approximately 1-1/3 times crankshaft speed and supplies a
controlled flow of hydraulic fluid to the power pack
and hydraulic systems when the gear control handle
is moved from neutral position. When gear control
handle is in neutral, fluid circulates freely through
the pump into the power pack and back to the reservoir. Pump flow is controlled to approximately one
gallon -per -minute.
5-180. REMOVAL.
a. Remove upper right cowling from front engine.
b. Disconnect hydraulic lines from pump and cap
or plug openings.
c. Remove mounting nuts and remove pump from
aircraft.
d. Remove and discard mounting gasket.
Locally fabricated items, handy for power pack re-
SHOP NOTES:
•
5-63
•
5-181. TROUBLE SHOOTING.
PROBABLE CAUSE
TROUBLE
REMEDY
HAND PUMP DOES NOT BUILD
UP PRESSURE, BUT ENGINE
PUMP OPERATES GEAR
PROPERLY.
Faulty hand pump plunger check
valve or O-ring.
Remove and inspect hand pump
plunger; replace pa I ts as needed.
Faulty system inlet check valve
or hand pump inlet. check valve.
Remove Power Pack and repair
or replace check valves.
ENGINE PUMP WILL NOT
OPERA TE GEAR BUT
EMERGENCY HAND PUMP
WILL OPERA TE GEAR.
Fluid level low in reservoir.
Refill reservoir.
Engine pump or pump line
failure.
Repair or replace pump nr broken
pump line. Refill reservu,:·.
Faulty primary relief valve.
Remove Power Pack, repair or
replace primary relief val \', .
ENGINE PUMP OR EMERGENCY PUMP WILL NOT
BUILD PRESSURE IN
SYSTEM.
No fluid in reservoir.
Refill reservoir.
Broken gear or door line.
Repair or replace hydraulic line.
Door solenoid valve jammed or
sticking at mid-travel.
Repair solenoid valve.
Faulty secondary relief valve.
Remove Power Pack, repair or
replace secondary relief valve.
5-182. DISASSEMBLY. (Refer to paragraph 5-28.)
a. Plug all ports and clean outside of pump with
solvent.
b. Position pump, shaft down, in a vise and tighten
vise on pump mounting flange just enough to retain
pump in the vise. Index mark pump housing (3) and
front plate (12) to ensure correct reassembly.
(~AUTIONI
Do not pry sections apart with a screw driver
or other instrument, as scratches caused by
the tool, will prevent sealing of mating surfaces when reassembled.
c. Remove cap screws and washers (I and 2), and
lift off rear housing (3), by rocking from side-to-side
and sliding off gear shafts and dowel pins (13).
NOTE
In case of sticking, tap sides lightly with a
hard, non-metallic hammer. Further
disassembly of pump housing is not
necessary.
ITEM
GEARS AND SHAFTS
5-64
.-
d. Remove idler gear assembly (16).
e. Remove snap ring (4) from drive shaft, and exercise care not to scratch bearing surface of drive
shaft.
f. Remove gear (5) and Key (6) from drive shaft
(11).
g. Remove remaining snap ring (4) from drive
shaft (11).
h. Remove drive shaft (ll) from front housing (12)
by pulling it out of housing by splined end.
1. Remove diaphragm (15) from front plate (12)
by prying with a sharp tool.
j. Remove phenolic back-up gasket (7) and protector gasket (14) from front plate (12).
k. Remove diaphragm seal (8) from front plate (12~
1. Remove snap ring (10) and drive shaft seal (9)
from bore in front plate (12).
•
5-183. INSPECTION OF PUMP. Clean all metal
parts with cleaning solvent and dry with filtered
compressed air. Prior to assembly, inspect all
parts as follows.
INSPECTION
Inspect drive gear shaft for
broken splines.
REPAIR
Replace shaft if damaged.
•
GEARS AND SHAFTS (Cant)
FRONT PLATE ASSEMBLY
•
REPAm
INSPECTION
ITEM
REAR HOUSING
Inspect both the drive gear and
idler gear shaft at bearing points
and shaft seal area~ for rough
surfaces and excessive wear. If
shafts measure less than . 4360 in
bearing area, they should be replaced.
Replace drive gear shaft.
Inspect gear face for scoring and
excessive wear. If gear width is
below .1950, drive gear or idler
gear should be replaced.
Replace drive gear.
Visually inspect snap rings on
idler gear shaft. They should
be in grooves.
Replace if necessary.
Visually inspect edges of gear
teeth to see if they are too sharp.
Break sharp edge with emery cloth.
Visually insper.t he:l:·;·,_., !'.JI"
scratrhes or sCllrills. \oleasure
J. n. of hcaril1!-:"s. 11 1. V. measures
mOIl' th.-In .4400. front plate
ciho~lci be replaced.
R'C'place front plate assembly.
(Bearings are not available as
separate items. )
Visually insped bearin~s for
proper positioning. Bearings
should be flush with islands in
groove pattern. Splits in bearings should be in line with dowel
pin holes and in position closest to
the respective dowel pin hole.
Replace front plate assembly if bearings are out of position. (Bearings
are not available as separate items. )
Visually inspect inside gear
pockets for excessive scoring or
wear. Also measure I. D. and
depth of gear pockets. I. D.
should not exceed 1.691 and
depth should not exceed. 1972.
If badly scored or wear exceeds di-
Replace idler gear shaft.
Replace idler gear.
mens ions given, replace rear housing
assembly.
Visually inspect bearings for
scratches or scoring. I. D.
should not exceed. 4400.
If I. D. of bearing exceeds dimensions
given, replace rear housing assembly.
Visually inspect bearings for
proper pOSitioning. Splits in
bearings should be in line with
dowel pins and in position closest
to the respective dowel pin.
If bearings are out of pOSition,
5-184. ASSEMBLY.
replace rear housing. (Bearings
are not available as separate itemS. )
is available from the Cessna Service Parts
Center.
NOTE
.:
Diaphragm (15), phenolic gasket (7), protector
gasket (14). diaphragm seal (8), drive gear snap
rings (4), shaft seal (9), snap ring (10), copper·
crush washer (2) and Key (6) should be replaced
with new parts when reassembling hydraulic
pump. A Major Seal Repair Kit (part No. 2024077). conSisting of the parts listed in this note,
a. Install new shaft seal (9) in front plate, with flat
metal side of seal in front plate and the tapered internal part of seal toward pump shaft splines.
NOTE
Press shaft seal just deep enough to allow
snap ring (16) to be installpd in groove.
5-65
DOUBLE UP SEAL
SINGLE UP SEAL
,---=. . INSTALL "OPEN" END
INSTALL "CLOSED"
END TOWARD PUMP
SHAFT SPUNES
Section
•
TOWARD PUMP SHAFT
SPLINES
A-A
(USED ON EARLY
SERIAL NO. PUMPS)
(USED ON ALL LATER SERIAL NO.
PUMPS AND ALL SERVICE PARTS)
11
~
1
Cap Screw
Copper Crush Gasket
Rear Housing Assembly
Snap Ring
5. Gear
6. Key
1.
2.
3.
4.
O-RING AND PLUG
INST ALLED HERE
17
4
16
7.
8.
9.
10.
11.
Phenolic Back-Up Gasket
Diaphragm Seal
Shaft Seal
Snap Ring
Drive Shaft
4
15
14
DRAIN LINE FITTING
INSTALLED HERE
12.
13.
14.
15.
16.
17.
•
Front Plate Assembly
Dowel Pin
Protector Gasket
Diaphragm
Idler Gear
Idler Gear Shaft
Figure 5-28. Hydraulic Pump Assembly
b. Install snap ring (10) in groove in front plate (12)
with sharp edge of snap ring toward shaft splines.
c. Place diaphragm seal (8) on front plate (12), with
flat side of seal down (cup side of seal up). Using a
dull pointed tool, work diaphragm seal to bottom of
grooves in front plate. Ensure that seal is all the
way down tn grooves of front plate.
d. Press protector gasket (14) and phenolic back-up
gasket (7) tnto cup of diaphragm seal.
e. Place diaphragm (15) on top of phenolic back-up
gasket with bronze face of diaphragm up, next to the
gears. The two small depressions on the bronze face
must match the two depressed areas in the rear housing.
5-66
NOTE
Protector gasket (14), phenolic back-up
gasket (7) and diaphragm (15) must fit
inside cup of diaphragm seal (8).
f. Coat drive shaft (11) with grease to prevent damage to seal (9) as drive shaft is installed.
g. Work drive shaft (ll) through shaft seal (9) and
into position.
h. Install snap ring (4) in groove on shaft next to
diaphragm.
i. Place Key (6) in slot in drive shaft and inatall
gear (5) over Key in shaft.
•
•
NOTE
16
15
18
19
With installation of
rear engine optional
hydraulic system,
some line routing is
changed in the area
of the Power Pack.
13---1
12
11~O----24
29
•
•
33701331 THRU 33701462 AND
F33700024 THRU F33700055
Detail
A
3
~
31
FIREWALL CONNECTIONS
NOTE
A face-spanner wrench, (Tool
#418, available from Armstrong
Bros., 5200-5300 W. Armstrong
Ave., Chicago, Ill.) or equivalent, may be used to remove end
gland from hydraulic filter, seriallized 33701331 thru 33701462
and F33700024 thru F33700055.
1. Screw
2. Lockwasher
3. End Fitting
4. Filter Disc
5. Snap Ring
6. Spider
7. Ball
8. Spring
9. O-Ring
10. Body
11. Pump Drain Line
12. Hydraulic Filter
13. Hydraulic Pump
14. Pump Pressure Line
15. Pump Suction Line
FILL
SUCTION
0
o
PRESSURE
o
LANDING
GEAR
DOWN
o
LEFT SIDE OF FIREWALL (LOOKING AFT)
16. Mounting Bolt
LANDING 17. Washer
GEAR UP 18. Mounting Bracket
o
19. Filler Elbow
20. Power Pack
21. Mounting Bracket
22. Washer
DOOR
23.
Mounting Bolt
CLOSE
24. Hand Pump Suction Line
o
25. Hand Pump Pressure Line
26. Door Close Line
DOOR
27. Door Open Line
OPEN
28. Gear Up Line
o
29. Gear Down Line
30. Overboard Vent Line
31. Cover
Figure 5-29. Hydraulic Power System
5-67
j. Install snap ring (4) In groove of shaft (11) next
to gear (5).
k. Install idler gear assembly (16).
1. Slide rear housing assembly (3) over gear shafts
until dowel pins 03) are engaged.
m. Install cap screws (1) with copper crush washer
(2) on the 1-3/4 inch long screw which passes through
the suction port of the pump. Tighten cap screws
evenly to torque valve of 7 -10 lb-ft.
n. Rotate pump shaft by hand. Pump will have a
small amount of drag, but should turn freely after a
short peri od of use.
5-185. INSTALLATION. (Refer to figure 5-29.)
a. Install a new mounting gasket on accessory pad.
b. Grease pump splines lightly with all purpose
grease, and slide pump into pOSition, rotating pump
splines as necessary for smooth meshing of splines.
c. Install mounting nuts and tighten. Connect hydraulic lines to pump.
d. To prevent initial dry-running of pump, loosen
suction hose fitting at pump inlet and disconnect
power pack reservoir drain line from firewall fitting.
e. Connect suitable filler unit to reservoir filler
elbow; hold finger over open end of reservoir drain
line fitting at firewall and fill reservoir until fluid
is forced from loosened end of suction hose.
L Tighten suction hose fitting, connect reservoir
drain line and disconnect filler unit.
g. Install engine cowling.
5-186. HYDRAULIC FLUID FILTER.
figure 5-29.)
(Refer to
5-187. DESCRIPTION. The hydraulic fluid filter
consists of a filter body, spring, check-ball. filter
diSC, inlet and outlet fittings for hydraulic lines and
attaching parts.
5-188. OPERATION. The filter is located in the
pump pressure line at the forward side of the front
firewall. and filters the hydraulic fluid from the
pump before it enters the power pack. The filter
contains a bypass valve which will open and supply
the system with fluid if the filter disc should become
clogged.
5-189. REMOVAL.
a. Remove cowling from front engine.
b. Disconnect hydraulic lines at filter and cap or
plug openings.
c. Remove filter from aircraft.
5-190. DISASSEMBLY. (Thru 33701330 and F33700023).
a. Cut safety wire and remove screws securing end
fitting (3) to filter body (10).
b. Remove filter disc (4) from end fitting, using
care to prevent damage to parts.
c. Remove snap ring (5), spider (6), check-ball
(7) and spring (8) from filter body.
d. Remove and discard O-ring (9).
5-191. INSPECTION OF PARTS.
a. Clean all metal parts with solvent (Federal Specification P-S-661).
b. Inspect all parts for damage; replace faulty parts.
c. Use care to keep dirt or foreign material from
5-68
parts after cleaning or during assembly.
5-192. ASSEMBLY. (Thru 33701330 and F33700023).'
a. Insert spring (8), check-ball (7) and spider (6)
into filter body (10) and install snap ring (5).
b. Lubricate new O-ring (7) with hydraulic fluid
and install on filter body (10).
c. Install disc (4) in end fitting (3), position
fitting to filter body (10) and install screws securing
assembly.
d. Tighten screws evenly and safety wire.
•
5-193. DISASSEMBLY. (33701331 and F33700024
thru 33701462 and F33700055).
a. Using face spanner wrench (refer to note on
figure 5-24), remove end fitting (3).
b. Remove O-ring (9), filter disc (4), retainer,
ball (7) and spring (8) from filter body (10).
c. Discard O-ring (9).
5-194. INSPECTION OF PARTS. (Refer to paragraph
5-191.)
5-195. ASSEMBLY. (33701331 and F33700024 thru
33701462 and F33700055).
a. Install O-ring (9) on end fitting (3).
b. Insert spring (8), ball (7), retainer and filter
disc (4) in filter body (10).
c. Install end fitting on filter body and tighten with
face-spanner wrench. (Refer to note on figure 5-29.)
5-196. INSTALLATION.
a. Position fUter assembly to hydraulic lines, uncap
or unplug openings and connect lines to filter.
b. Install engine cowling.
5-197. HYDRAULIC POWER PACK.
5-30.)
(Refer to figure
•
5-198. DESCRIPTION. The hydrauliC power pack,
located in the cabin on the aft left side of the front
firewall. behind the instrument panel, is a multipurpose control unit in the hydraulic system. The
unit contains a hydraulic fluid reserVOir, valves which
control the flow of pressurized hydraulic fluid to
actuators in the landing gear and door system, and
an electrical switch, connected to the gear warning
horn and indicator lights.
5-199. OPERATION. (Refer to paragraph 5-3.)
5-200. REMOVAL. (Refer to figure 5-29. )
a. Remove front seats in accordance with procedures outlined in Section 3.
NOTE
As hydraulic lines are disconnected or removed,
cap or plug all openings to keep dirt and
foreign material out of system and components.
b. Use a protective cover over floor covering and
posi tion a gallon container under fill - and - drain
tee; loosen pressure cap and drain reservoir. Use
of a funnel and hose will simplify draining.
c. Remove drain hose, cover floor under power
•
•
OVERBOARD VENT
FILL
ELECTRICAL
CONNECTOR
DOOR SOLENOID
VALVE
DRAIN
ENGINE PUMP
SUCTION
~---
-----~=+!-_
ENGINE PUMP - - - - - - - i F R - - 4
PRESSURE
DOOR CLOSE
PRESSURE
1Ul!?J----- DOOR OPEN
PRESSURE
GEAR UP
PRESSURE
HAND PUMP----PRESSURE
GEAR DOWN
PRESSURE
HANDLE-RELEASE
PRESSURE ADJUSTMENT
OVERBOA RD VENT
•
ELECTRICAL
CONNECTOR
DOOR CLOSE-f----.-tll
PRESSURE
ENGINE PUMP
PRESSURE
--_r:1I
~:--ii~~~~~~
Hj~~
GEAR UP
PRESSURE
DOOROPEN~'
PRESSURE
GEAR DOWN
PRESSURE
HANDLE- RELEASE
PRESSURE ADJUSTMENT
•
HAND PUMP
PRESSURE
Figure 5-30.
HANDLE-DOWN
RETURN SPRING
ADJUETMENT
GEARDO~
PRESSURE
HANDLE- UP RETURN
SPRING ADJUSTMENT_
HANDLE-RELEASE ----'
PRESSURE ADJUSTMENT
Location of Power Pack Fittings (Sheet 1 of 2)
5-69
FILLER AND DRAIN
TEE-FITTING
•
PRIMARY
RELIEF
VALVE
DOOR
VENT
VALVE
PRIORITY
VALVE
•
TIME-DELAY
VALVE---'
*SECONDARY
RELIEF
VALVE---'
TOP VIEW
...
AFT
*ThiE valve is deleted for the 1968 models .
On remanufactured Power Packs, this
cavity is filled with an O-ring and plug.
Figure 5-30. Location of Power Pack Fittings (Sheet 2 of 2)
5-70
•
•
pack and disconnect all lines at power pack.
d. Remove roll pin securing gear control tube to
power pack shaft, and slide linkage clear of shaft.
e. Disconnect brake hose under power pack and
swing to one side.
f. Remove forward sections of hydraulic lines
routed to emergency hand pump. (Loosen or remove left forward upholstery panel as required
for access.)
g. Disconnect electrical plug at back of power pack.
h. Remove mounting bolts and carefully work power
pack down and aft to remove.
5-201. TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
REMEDY
GEAR CONTROL HANDLE WILL
NOT LOCK IN UP OR DOWN
DETENT.
Handle release valve plunger
setting too low or incorrect return spring adjustment.
Adjust handle release valve and
return springs.
GEAR CONTROL HANDLE
RETURNS TO NEUTRAL
BEFORE DOORS CLOSE.
Fluid low in reservoir, causing
·air in time-delay valve.
Fill reservoir and purge time-delay.
Time-delay valve stuck or will
not hold fluid charge due to faulty
time-delay valve ball seat.
Remove Power Pack and replace
time-delay valve seat.
Landing gear handle release
pressure too high.
Adjust handle release pressure.
Landing gear handle return
springs setting too low .
Adjust return springs.
Landing gear handle linkage
binding.
Remove Power Pack, repair or replace handle shaft.
Landing gear selector spool
binding.
Remove Power Pack and replace
manifold, selector spool and timedelay valve plunger as an assembly
only.
GEAR CONTROL HANDLE
FAILS TO RETURN TO
NEUTRAL AFTER DOORS
CLOSE (3 to 9 SECONDS).
•
NOTE
Extremely cold temperatures will cause a longer time delay before handle
trips after the doors close. This is normal. If landing gear handle does
not return to neutral properly, Power Pack overheating will result.
POWER PACK EXTERNAL
LEAKAGE.
SLIDING SEALS:
(Seals having a moving part. )
POWER PACK EXTERNAL
LEAKAGE.
STATIC SEALS:
(Seals with no moving parts. )
Handle release plunger.
Remove release plunger and replace
O-rings.
Landing gear selector spool.
Remove Power Pack and replace
O-ring on spool and in manifold.
Priority valve.
Remove Power Pack and replace
priority valve seals.
Hand pump plunger gland.
Remove hand pump plunger and
replace O-rings.
All fittings.
Remove and replace O-rings and
back-up rings as required.
Door solenoid.
Replace O-ring.
5-71
5-201. TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
POWER PACK EXTERNAL
LEAKAGE.
STATIC SEALS:
(Seals with no moving parts)
(Cont).
POWER PACK LOSES FLUID
WITH NO EVIDENCE OF
LEAKAGE.
REMEDY
Transfer tubes between manifold and body.
Remove Power Pack, disassemble
and replace O-rings.
Time-delay valve.
Remove Power Pack, disassemble
and replace O-rings.
Reservoir cover.
Replace seals.
Air leak at engine pump shaft
seal.
Repair or replace engine pump.
Air leak in suction line to engine
pump.
Repair or replace suction line or
fittings.
•
NOTE
Hydraulic fluid foams due to air being pumped into system and the fluid
is blown overboard through the Power Pack vent line.
5-202. DISASSEMBLY. (Refer to figure 5-31.)
NOTE
After the power pack has been removed from
the aircraft, . and all ports capped or plugged,
spray with cleaning solvent (Federal Specification P-S-66I, or equivalent) to remove all
accumulated dust or dirt. Dry with filtered
compressed air.
a. Remove reservoir cover retaining nut and
O-ring. Cover is a snug fit on reservoir. Use a
soft mallet and tap cover lightly to remove. Remove
large O-ring.
b. Remove spacer from center stud, cut safety
wire, and remove baffle from reservoir. Drain
remaining hydraulic fluid from reservoir.
c. Remove reservoir cover attaching stud (center).
This stud may be removed by using a double lock nut
at top of stud. Use care to prevent damage to stud
threads.
d. Turn Power Pack upside down so that top of
reservoir serves as a support base.
NOTE
A holding fixture (Part No. HF-I025) may
be used instead of removing the center stud
if desired. This is a plate type fixture for
use in a vise. The fixture is available from
the Cessna Service Parts Center.
e. Remove screws attaching electrical wires to
terminal strip and Power Pack. Remove small
capaCitor from beneath electrical wires and remove
terminal strip.
5-72
NOTE
All electrical wires are coded with color
stripes. Disregard color of wire terminals
or plastic sleeving. If color codes are
matched when wires are reinstalled, the
wires will be connected correctly.
f. Cut safety wire and remove screws attach~
landing gear up-down switch and bracket. Retam
washers between bracket and Power Pack.
g. Turn Power Pack over and cut safety wire at
time-delay valve.
h. Remove time-delay valve ball, spring, spacer,
and spring by removing time-delay valve retainer.
•
NOTE
Do not remove time-delay valve plunger until after manifold assembly has been removed.
i. Cut safety wire and remove screws attaching
gear and rack protective cover. Remove cover.
j. Remove clamp attaching electrical wires to
door solenoid valve and remove safety wire from
door solenoid valve.
k. Cut safety wire and remove four screws attaching manifold assembly. Work manifold assembly
from Power Pack, taking care to prevent loss of
transfer tubes between manifold and Power Pack.
1. Remove the seven transfer tubes from manifold
or Power Pack.
•
•
* Items 27 through 33 are deleted on the 1968 model
Power Packs. On remanufactured P?wer Packs,
these parts are replaced with an O-nng and plug.
*SECONDARY
RELIEF
VALVE
HAND PUMP
INLET FILTER
PRIMARY
RELIEF
VALVE
@{
35
36
~
~
'1]
•
37
I
I
l_ _ _ _ _ _ _ _ __
- - _ _ _ _ _~I
TIME-DELAY
VALVE
'~
INLET
CHECK ____________
~
VALVE
~52
8-®--54
53
DOOR
"'--VENT
VALVE
55
PRIORITY
VALVE
ADJUSTMENT
•
Figure 5-31. Reservoir and Main Body Components
5-73
References for figure 5-31.
1. Standpipe and Filter
PRIMARY RELIEF VALVE
2.
3.
4.
5.
6.
7.
8.
9.
10.
Poppet Seat
O-Ring
Back- Up Ring
Poppet
Ball
Button
Spring
Button
Retainer
11. A~justing Screw
12. Locknut
22. Vent Filter
23. Reservoir Cover O-Ring
24. Reservoir Cover
25. O-Ring
26. Cap Nut
SECONDARY RELIEF VALVE
27.
28.
29.
30.
31.
32.
33.
Adjusting Plug
Retainer
Spring
Button
Ball
Seat
Seat O-Ring
44.
45.
Center Stud
Reservoir and Body Assembly
DOOR VENT VALVE
46. Retainer
47. O-Ring
48. Spring
49. Poppet
50. Body
51. Pin
PRIORITY VALVE ADJUSTMENT
PRIORITY VALVE
34. Sight Gage
52. Button
53. Spring
54. Retainer (Adjusting Plug)
13. Poppet 0- Ring
14. Poppet
15. Poppet Seat
16. Poppet Seat O-Ring
17. Retainer 0- Ring
18. Retainer
HAND PUMP INLET FILTER
INLET CHECK VALVE
35. Snap Ring
36. Spacer
37. Filter
55. Pressure Inlet Fitting
56. Fitting O-Ring
57. O-Ring
58. Plunger
59. Spring
TIME-DELAY VALVE
19. Baffle
20. Spacer
21. Snap Ring
38.
39.
40.
41.
42.
43.
Retainer
Retainer Hex 0- Ring
Ball
Spring
Spacer
Retainer Body O-Ring
•
60. Snap Ring
61. Filler Line Filter
•
SHOP NOTES:
5-74
•
•
TRANSFER
SLEEVE
TIME-DELAY.
PLUNGER
•
HANDLE
RELEASE
VALVE
LANDING GEAR SELECTOR SPOOL
1. Screw
2. Rack
3. Laminated Shim
4. Spool
5. Spool 0- Ring
6. Manifold
7. Washer
B. Allen Screw
DOOR
SOLENOID
VALVE
13. Spool
14. O-Ring
15. Transfer Tube O-Rings
16. Transfer Tubes
HANDLE RELEASE VALVE
17. O-Ring
lB. Poppet
19. Poppet O-Ring
20. Spring
21. Retainer (Adjusting Plug)
TIME-DELAY PLUNGER
DOOR SOLENOID VALVE
9. Plunger
10. Spring
TRANSFER SLEEVE
•
11. Sleeve
12. Sleeve O-Ring
22. Plunger
23. Pin
24. Spring
25. Solenoid O-Ring
26. Solenoid
Figure 5-32. Manifold Assembly
5-75
[CAUTI~N\
As the manifold is separated from the Power
Pack body, the rack on the landing gear selector spool becomes disengaged from the
gear on the handle. This will permit the selector spool to move. Do NOT move the
selector spool from its position. Never move
it to a position that Is more than flush with the
manifold body at the end opposite the selector
spool rack. If moved beyond this position,
an O-ring w1ll become caught and the selector
spool will then be extremely difficult to remove.
5-203. MANIFOLD DISASSEMBLY.
a. Remove door solenoid by unscrewing from manifold. This solenoid is hand tightened. Use strap
wrench or strip of sandpaper to grip door solenoid
for removal. Remove plunger return spring.
b. Remove plunger and spool by carefully pulling
from manifold.
c. Using a hook formed from brass welding rod
and inserted into oil hole in transfer sleeve, withdraw sleeve from manifold.
NOTE
Be sure that end of hook is not over 1/16-inch
long, and use with care to prevent scratching
the bore in manifold. The sleeve will be hard
to withdraw due to O-ring friction.
d. Remove time-delay valve plunger, using a small
wooden dowel inserted in center of plunger. The
plunger should slide out of manifold easily.
e. Remove landing gear selector spool by grasping
rack end of spool and carefully pulling from manifold.
NOTE
Do not bend selector spool. Pull straight out.
Do not remove gear rack from selector spool
unless it is necessary to replace selector
spool and manifold. The landing gear selector
spool, time-delay plunger, and manifold are
matched, lapped parts. If it is necessary to
replace anyone of these three parts, replace
them as an assembly only.
f. Remove landing gear handle-release retainer
(adjusting plug), spring, and poppet from manifold.
The end of the poppet has a ball which should remain in the poppet. If it doesn't, remove ball from
manifold.
g. Remove caps from fittings and wash manifold
in cleaning solvent (Federal Specification P-S-661,
or equivalent) and dry with filtered compressed air.
Be sure internal passages are clean, then reinstall
caps on fittings.
5-204. DISASSEMBLY OF COMPONENTS.
5-205. SECONDARY RELIEF VALVE. (PRIOR TO
1968 MODELS. )
a. Remove adjusting plug at top of secondary relief
\'3.lvc.
5-76
b. Remove secondary relief valve retainer by unscrewing from body.
c. Remove spring, button, and ball from body.
d. Use a brass hook to remove seat from body.
Use with care to prevent scratching bore.
e. Remove O-ring from bottom of caVity.
•
5-206. PRIMARY RELIEF VALVE.
a. Loosen lock nut at top of primary relief valve.
b. Remove adjusting screw and lock nut from top
of relief valve.
c. Unscrew retainer.
d. Remove two buttons, spring, and ball.
e. Remove poppet from poppet seat by lifting out
of poppet assembly. The poppet and poppet seat are
matched parts.·
f. Using a brass hook not over 11l6-inch long, pull
poppet seat up out of body. Hook through holes in
side of seat and use care not to damage bore in body.
5-207. PRIORITY VALVE.
a. Remove priority retainer from reservoir.
b. Turn Power Pack upside down and remove retainer (adjusting plug), spring, and button from bottom of Power Pack.
c. While Power Pack is upside down, push poppet
and poppet seat into reservoir, using a punch of 1/8
inch maximum diameter. Make sure that face of
punch is square and flat.
5-208. SYSTEM INLET CHECK VALVE.
a. Remove system pressure port fitting.
b. Remove O-ring, plunger, and spring. Spring
and plunger should fall out of Power Pack after 0ring is removed. Use hook, if necessary, to remove O-ring.
5-209. STANDPIPE AND FILTER.
a. The standpipe and filter assembly should not be
removed unless it is damaged, since it is a press fit
in the reservoir.
b. Remove vent filter by remOVing the snap ring.
c. Remove fill line filter by removing the fitting
and snap ring.
d. Remove hand pump filter by removing snap ring
and spacer.
•
5-210. DOOR VENT VALVE.
a. Remove door vent valve from reservoir. The
door vent valve should not be disassembled except
for replacement of parts.
b. Remove pin from valve body a.'1d retainer. Use
care when removing pin, as the spring is under a
slight load.
c. Remove retainer, O-ring, and poppet from valve
body.
5-211. LANDING GEAR HANDLE-RELEASE MECHANISM.
a. Remove two hex-head retainers (adjusting plugs),
springs, and plungers from handle return housing.
b. Cut safety wire and remove two screws attaching handle release housing to Power Pack, and remove the housing.
c. USing a punch, drive roll pin from cam, and remove cam from landing gear handle shaft.
d. Pull handle assembly from Power Pack.
•
•
NOTE
Do not remove spacer, detent cam, or gear
from handle shaft except for replacement of
parts.
5-212. ASSEMBLY OF POWER PACK. After power
pack has been completely disassembled, remove and
discard all O-rings and gaskets. Wash all parts in
dry cleaning solvent (Federal SpeCification P-S-661,
or equivalent) and dry with filtered compressed air.
Inspect all threaded surfaces for serviceable condition and cleanliness. Inspect all parts for scratches,
scores, Chips, cracks, and indications of excessive
wear. Use new O-rings and gaskets during reassembly. Lubricate all O-rings with Dow-Corning DC-4
compound during reassembly. Lubricate all threaded
surfaces on the variOUS valves in the Power Pack
with MIL-G-S1322 grease (or equivalent) before in stalling.
5-213. DOOR VENT VALVE.
a. Install poppet in body and insert spring in body.
Be sure that spring enters poppet.
b. Lubricate and install O-ring on retainer and
insert retainer in valve body. Align holes in retainer
with holes in valve body.
c. Install pin through valve body and retainer.
d. Lubricate threads on valve body (MIL-G-S1322)
and install assembly in reservoir. Tighten securely.
•
5 -214. STANDPIPE AND FILTER.
a. If standpipe and filter assembly was removed,
press into body until standpipe bottoms.
b. Replace vent filter and snap ring.
c. Install filler line tilter and secure with snap
ring.
d. Install back-up ring and O-ring on fill and drain
tee, and install tee.
e. Install hand pump filter, spacer and snap ring.
5-215. SYSTEM INLET CHECK VALVE.
a. With pressure port up, drop spring into port.
b. Drop in plunger, making sure that small end of
plunger goes into spring. Check freeness of plunger
in body by depressing plunger against spring. Use
small wood dowel or plastic rod to depress plunger
when checking freedom of movement. Plunger must
move freely in body bore.
c. Lubricate and install O-rings on flange of fitting
and at end of fitting. Lubricate threads (MIL-G-S1322)
insert fitting, start the threads and tighten securely.
•
5-216. PRIORITY VALVE.
a. Lubricate and install O-ring on poppet and insert poppet in body through reservoir. Push poppet
down firmly. Either surface may be used as seating
surface.
b. Inspect poppet seat for sharp seating edge. Lap
as necessary to obtain a sharp seating edge. Lubricate and install O-ring on poppet seat.
c. Install poppet seat in body through reservoir,
with sharp seating edge toward poppet. Push poppet
seat down firmly against poppet.
d. Lubricate and install O-ring on retainer assembly' lubricate retainer threads (MIL-G-S1322) and
install retainer. Tighten securely.
e. Turn power pack upside down, lubricate spring
and button (MIL-G-S1322) and install body. Apply
lubricant to hold button in spring and install with button in hole first.
i. Lubricate (MIL-G-81322) threads on retainer
(adjusting plug) and install. This plug provides adjustment for the priority valve. Install flush at this
time.
5-217. PRIMARY RELIEF VALVE.
a. Inspect poppet and poppet seat for pitting or
scoring. Since they are matched parts, if either or
both are pitted or scored, replace as an assembly
only.
b. Lubricate and install O-ring and back-up ring
on seat, insert poppet in seat, and install assembly
in body.
c. Lubricate ball, buttons and spring (MIL-G-S1322).
Install with ball entering hole first. Be sure that ball
enters cavity at top of poppet.
d. Lubricate threads on retainer (MIL-G-81322) and
install over button and spring. Tighten securely.
e. Lubricate threads of adjusting screw MIL-G81322) and install at top of retainer. Turn adjusting
screw full down to lock primary relief valve closed,
but do not tighten lock nut. This is done so that the
secondary relief valve, which opens at a higher
pressure, can be adjusted before the primary
relief valve is adjusted.
5-21S. SECONDARY RELIEF VALVE. (PRIOR TO
1968 MODELS. )
a. Lubricate and install O-ring in body. Make
sure O-ring seats properly.
b. Inspect seating surface of seat. It should have
a very sharp edge. Seat may be lapped to obtain a
sharp edge.
c. Install seat in body, with sharp edge of seating
surface up.
d. Apply lubricant (MIL-G -81322) to hold ball, button and spring together, and insert in body with ball
toward seat.
e. Lubricate threads on retainer (MIL-G-81322).
Start retainer over spring and tighten securely.
f. Lubricate threads on adjusting plug (MIL-G81322) and install at top of retainer. Do not tighten
adjustirig plug. Screw it down only until spring is
contacted. This is done so that air may be bled
from valve during adjustment.
5-219. MANIFOLD ASSEMBLY.
a. Lubricate and install the O-ring on landing gear
selector spool, and the O-ring in manifold at the opposite end.
NOTE
If landing gear selector spool, manifold, and
time-delay plunger are being replaced, install
rack with a new laminated shim on selector
spool. The landing gear selector spool, timedelay valve plunger, and manifold are matched,
lapped parts. If necessary to replace, replac('
as an assembly only .
b. Insert selector spool in manifold from landing
gear handle end of manifold. Insert only until end of
fl-77
V
__ -I.....
*
OILITE BUSHING
__I..,
--',j
,\';:~V
,
,
<
"
~
....
POWER PACK
BODY (REF)
I
I
I
I
i
~
~
1
I
/ ..... )
' , ..... 1....... .....-/
~
~
,
1
TOHANDLE
LINKAGE (SEE
FIGURE 5-35
.~
I '
:
B
'>
,~t
•
/
~~'"
3
~
*This bushing installed in
1968 model Power Packs.
2
~.....-~.
-~-~
/
11
l. Shaft
2. Gear
HANDLE RETURN
SPRING ADJUSTMENTS
3. Detent Cam
4. Spacer
5. Return Cam
6. Pin
9. Plunger
10. Spring
11. Retainer (Adjusting Plug)
HANDLE
RETURN SPRING
ADJUSTMENTS
•
7. Housing
8. Allen Screw
5-33. Handle-Release Mechanism
SHOP NOTES:
5-78
•
•
selector spool is flush with solenoid end of manifold.
[~~UTIONI
If the selector spool is moved much more than
flush with the manifold at the end opposite the
rack (before the manifold is installed and the
rack engaged properly with the gear on the
landing gear handle), an O-ring will become
caught. The selector spool will then have to
be removed, the manifold cleaned to remove
all O-ring particles, and a new O-ring installed. The selector spool then must be
reinstalled correctly.
c. Check that spool slides freely.
d. Inspect door solenoid spool for freedom of movement within the transfer sleeve assembly.
NOTE
Spool and sleeve are matched parts. If necessary to replace, replace as an assembly only.
•
e. Lubricate and install O-rings on transfer sleeve
and install sleeve in manifold.
f. Attach plunger to door selector spool with pin.
g. Lubricate and install O-ring on solenoid.
h. Lubricate solenoid threads and spring (MIL-G81322) and insert into plunger, then install solenoid
over spring and plunger. Screw solenoid into manifold. Do not overtighten solenoid, but tighten securely by hand. Safety the solenoid to adjacent Power
Pack mounting lug.
5-220. LANDING GEAR HANDLE-RELEASE
MECHANISM.
a. If the landing gear handle shaft or gear was removed, the parts must be indexed and assembled as
shown in figure 5 -34.
b. Lub:icate shaft (MIL-G-81322), install spacer on
shaft with roll pin, and insert shaft into Power Pack.
c. Install cam with roll pin. Both sides of cam
surfaces are identical. Check landing gear shaft
for freedom of movement in Power Pack. Check
for slight end play in shaft. If shaft binds, remove
cam and lap inside boss of cam to obtain slight end
play in shaft with cam installed.
d. Install handle-release housing and safety attaching screws. Check landing gear handle shaft for
freedom of movement.
NOTE
Do not install plungers, springs, and hexhead retainers (adjusting plugs) at this time.
•
5-221. INSTALLATION OF MANIFOLD.
__
a. Lubricate and install 0-rings-oIrtlie-seven transfer tubes.
. __
-h. -~Inserttransfer tubes into Power Pack body.
c. Install time-delay valve plunger in manifold.
Plunger must move freely in manifold without binding.
d. Mate manifold to Power Pack body, using care
to prevent damage to O-rings on transfer tubes.
Align dowel pin in Power Pack with dowel hole in
manifold.
NOTE
When installing manifold, time the landing
gear handle shaft assembly to rack on selector spool as shown in figure 5-34. Refer to
the following steps if binding o~curs.
e. Install four manifold attaching screws and washers. Torque screws to 20-30 pound-inches. Do not
over-torque screws. as this will cause binding in
the movement of landing gear handle shaft.
NOTE
If a new landing gear selector spool, time-
delay plunger, and manifold (a matched assembly) are being installed, the rack on the
selector spool must be shimmed properly to
provide a slight backlash (free movement)
between the teeth of the rack and the teeth
of the gear on the handle. This adjustment
is provided by a laminated shim. If excessive backlash eXists, a new shim must be
used. If no backlash exists, or if a new
shim is being installed, the "trial-anderror" method should be used, since the
backlash is determined after manifold attaching screws are installed and torqued. Remove
one lamination at a time until backlash exists
when screws are torqued properly, then do
not remove any more laminations. Apply
LoctUe, Grade C, to rack retainer screws
only after final adjustment of shim has
been determined and screws are being
installed for the last time.
f. Lubricate and install two O-rings on time-delay
valve retainer.
g. Lubricate (MIL-G-81322) and insert larger
spring and spacer in body through reservoir.
h. Lubricate (MIL-G-81322) and insert ball and
smaller spring in time -delay valve retainer (ball
next to top of retainer).
i. Lubricate threads on time-delay valve retainer
(MIL-G-81322) and install reta iner in body through
reservoir. Do not overtighten time-delay valve retainer as this will cause the landing gear selector
to bind in the manifold. After tightening time-delay
valve retainer, check for freedom of movement of
landing gear handle shaft and selector spool.
j. Thoroughly lubricate handle return springs and
plungers (MIL-G-81322) and install in housing with
hex-head retainers. Do not tighten retainers at this
time.
k. Lubricate and install two O-rings on landing
gear handle release plunger and insert plunger in
body.
1. Lubricate landing gear handle -release spring
and retainer (MIL-G-81322) and install in body.
Tighten retainer (adjusting plug) until almost flush
with body.
m. Install gear and rack protective cover. Safety
attaching screws.
n. Install landing gear up-down switch and the
switch attaching bracket. Note that washers are
5-79
*This
bushing installed in
1968 model Power Packs.
•
UPPER AND LOWER SURFACES
OF CAM ARE SYMMETRICAL
*OIUTE BUSHING
THIS END OF LANDING GEAR SELECTOR
SPOOL FLUSH WITH MANIFOLD WHILE
ENGAGING GEAR WITH RACK
MARKED TOOTH DOWN
GEAR-TO-CAM
RELATIONSHIP
GEAR
THIS HOLE HORIZONTAL
•
DETENT CAM
Figure 5-34. Timing of Handle Shaft and Selector Spool
used between the bracket and Power Pack. Switch
bracket has slotted holes for switch adjustment
o. Install terminal strip and place capacitor alongside the strip. Connect electrical wires to terminal
strip and ground, clamping wires to door solenoid
valve.
NOTE
This procedure requires a minimum of test
equipment and is intended for bench-testing
the Power Pack after field repair.
5-223. PRESSURE ADJUSTMENTS.
NOTE
NOTE
Electrical wires are coded with color stripes.
Disregard color of wire terminals or plastic
sleeving. If color codes are matched when
wires are installed, the wires will be connected correctly.
p. Continue reassembly of Power Pack after pressure adjustments have been completed.
5-222. BENCH-TESTING THE HYDRAULIC POWER
PACK.
5-80
A chart of hydraulic system pressures is
provided immediately follOWing benchtesting procedures. The values contained
in the chart may be used to check opening
and reseating pressures of the power pack
valves.
5-224. TEST EQUIPMENT.
a. One hydrauliC hand pump of 2000 psi capacity.
b. One hydraulic pressure gage of 2000 psi capacity.
c. One hydraulic pressure gage of 150 psi capacity.
d. High pressure hose to attach hand pump to Power
Pack inlet fitting.
•
•
7
MODEL 337
6
15
•
MODEL 337A & ON
AND T337 SERIES
7
)
21
1.
2.
3.
4.
5.
6.
7.
18
Power Pack Shaft
Roll Pin
Control Tube
Bearing
Control Arm
Bearing
Gear Indicator Lights
14
S.
9.
10.
11.
12.
13.
14.
Neutral Barrier Bracket
Support
Bracket
Lockout Solenoid
Knob
Housing
Stop Pin
15. Retaining Pin
16. Spring Stop
17. Spring
IS. Push-Pull Rod
19. Jam Nut
20. Rod End
21. Pin
5-35. Gear Control Handle Linkage Installation
e. Drain hose to connect to Power Pack reservoir
drain fitting.
5-225. CONNECTING TEST UNIT. Use only clean
hydraulic fluid (MIL-H-5606). Install a tee at the
hand pump pressure outlet, and attach the 2000 psi
pressure gage and the pressure hose to the tee. Connect the hose from the hand pump to the Power Pack
pressure inlet fitting, labeled "PUMP." Connect
drain hose to Power Pack reservoir fill and drain
tee. Cap all other fittings with high-pressure caps.
NOTE
•
Some Hydro Test units are equipped with a
hand pump, and others are provided with a
pressure jack and provisions to install a
hand pump.
5-226. HANDLE-RELEASE MECHANISM. (Refer to
figure 5-33.) The following procedure outlines preliminary adjustments to set the handle-release detent
spring load and the handle-return spring load adjusting plugs in approximately their correct positions before installing the Power Pack in the airplane. After
it has been installed, the system must be checked and
final adjustments, if needed, made at that time.
NOTE
To complete this preliminary adjustment, use
a liS-inch punch or equivalent steel rod as a
handle in hole near the end of the shaft, to rotate shaft as required for adjustment. Use
care to prevent damage to hole in shaft.
a. With handle-return spring adjusting plugs (2
and 3) not tightened, screw in detent spring adjusting plug (1) until it is approximately flush. The
5-S1
fs~utloNl
Plug (1) should be adjusted in 1/3 turn
increments. Screwing it in too far will
result in the system relief valve opening
before suffiCient pressure is built up to
operate the release plunger.
NOTE
Remove cover under Power
Pack for access.
o
•
LOCATEDON
LEFT SIDE OF
POWER PACK
HANDLE-DOWN RETURN
SPRING ADJUSTING PLUG
/
HANDLE-RELEASE DETENT
SPRING ADJUSTING PLUG
(RELEASE PRESSURE
ADJUSTMENT)
HANDLE- UP RETURN
SPRING ADJUSTING PLUG
5-36. Handle-Release Adjustment
spring, however, must not bottom out.
b. Rotate shaft to up-detent position, then hold it
beyond this position (in overtravel).
c. Tighten forward handle-return spring adjusting
plug (2) until handle just starts to move out of overtravel, then loosen the adjusting plug one turn.
d. Rotate shaft to down-detent pOSition, then hold
it beyond this position (in overtravel).
e. Tighten aft handle-return spring adjusting plug
(3) until handle just starts to move out of overtravel,
then loosen the adjusting plug one turn.
f. Rotate shaft to up-detent position and tighten
handle-release detent spring adjusting plug (1) until
the spring bottoms out, then back the adjusting plug
out two turns.
g. Handle must hold in both detent positions, but
must return with a positive snap when manually released from either detent position. Handle-release
detent spring adjusting plug (1) may be readjusted
slighUy more or slightly less than the two turns
specified in the preceding step if necessary.
h. Refer to paragraph 5-237 for final rigging of the
handle-release mechanism after it has been installed
in the aircraft.
5-227. SECONDARY RELIEF VALVE. (PRIOR TO
1968 MODELS.)
a. With landing gear handle in either up or down
position, apply test pump pressure until fluid flows
from secondary relief valve.
b. Bleed air by cracking cap on door-open fitting.
c. Adjust retainer plug at top of valve until valve
cracks at 1900 psi. Adjusting this valve to 1900 psi
cracking pressure will give apprOXimately 1950 psi
when valve is in a flow condition. Bleed pressure
by cracking cap on door-open fitting after each adjustment.
d. Safety wire the secondary relief valve to the
5-82
time-delay valve.
5-228. PRIMARY RELIEF VALVE.
a. Loosen lock nut and back adjusting screw at top
of valve out until very little load is left on spring.
b.' With landing gear handle in neutral, apply pressure until fluid flows from primary relief valve.
c. Adjust primary relief valve until valve cracks
at 1400 psi. Adjusting this valve to 1400 psi cracking pressure will give apprOXimately 1500 psi when
valve is in a flow condition. Bleed pressure after
each adjustment by cracking cap on door-open fitting.
Tighten lock nut on adjusting screw after obtaining
correct adjustment.
5-229. PRIORITY VALVE.
a. Place landing gear handle in up position and remove cap from gear-up fitting.
b. Apply pressure and note priority valve cracking
pressure by observing pressure gage when fluid first
starts to flow from gear -up port.
c. Adjust priority valve to crack at 750 pSi. Bleed
pressure after each adjustment by cracking cap on
door -open fitting.
d. Disconnect test pump and cap all open fittings.
To complete the reassembly of the Power Pack proceed as follows:
a. Install reservoir cover attaching stud. Install
with longer threaded end down, and screw in until
stud bottoms in reservoir.
b. Install baffle and center stud spacer. Safety
wire primary relief valve lock nut to screened standpipe.
c. Lubricate and install O-ring in groove of reservoir cover.
d. Position reservoir cover on reservoir, aligning index marks on reservoir and cover. Vent
•
•
•
~ 1----~
PRIORITY VALVE (LOCATED
INSIDE POWER PACK)
~
Co
POWER PACK - -......
~}I
PRIORITY VALVE ADJUSTMENT
ADJUSTING SCREW _ _ _ _
.@(ACCESSlBLEAFTERREMOVlNG
COVER UNDER POWER PACK)
•
5-37. Priority Valve Adjustment
fitting in cover pOints to the left with Power Pack in
airplane.
ICAUTION\
Be sure that the large O-ring is positioned
properly in the groove of the reservoir cover
and that the O-ring is not pinched as the cover
is installed.
e. Lubricate and install O-ring at top of cover
around center stud.
f. Install cover retaining nut (cap nut), tighten,
and safety.
5-230. DOOR VENT VALVE.
a. Remove cap from door-open fitting on power
pack and attach pressure hose from hand pump with
150 psi pressure gage to door-open fitting.
b. Slowly apply 50 psi pressure. At 50 psi pressure continuous pumping shall be required to maintain pressure.
c. Increase pressure to 100 pSi.
d. There shall be no difficulty in maintaining pressure at 100 psi. Pressure can and should slowly
decrease as a result of leakage through the door
vent valve.
e. Relieve pressure by cracking hose fitting from
hand pump. Repeat step "b".
I. Disconnect test pump and cap all open fittings.
5 -231. RESERVOIR LEAKAGE TEST.
a. Remove filler and drain tee, and attach handtest pump and 150 psi gage to filler port.
b. Remove cap from reservoir vent fitting at top
of reservoir and operate test hand pump until reservoir is completely full, indicated by fluid coming out
of the fitting.
c. Cap reservoir vent fitting.
d. Operate test hand pump very slowly until pressure gage indicates 15 psi maximum.
e. Check for leaks. There should be no external
leakage.
f. Crack vent fitting to release pressure, remove
test equipment, drain reservoir, and cap fittings.
g. Hydraulic Power Pack is now ready to be installed in the airplane.
NOTE
After Power Pack is installed in airplane, refill reservoir.
SHOP NOTES:
•
5-83
5-232. HYDRAULIC SYSTEM PRESSURES (FOR BENCH OR FIELD TESTING).
COMPONENT
OPENING PRESSURE
RESEATING PRESSURE
Handle Release Valve.
750 to 1050 psi.
-------
Priority Valve.
750 to 800 psi.
-------
Primary Relief Valve.
1500 psi.(Max. )
1150 psi. (Min. )
Secondary Relief Valve.
1950 psi. (Max. >.
1550 psi. (Min. )
Inlet Check Valve.
10 psi. (Max. )
2 psi. (Min. )
Hand Pump Check Valves.
10 psi. (Max. )
2 pSi. (Min. )
5-233. INSTALLATION OF POWER PACK.
NOTE
When installing a new Power Pack, leave
the bulkhead nuts loose on the tubing fittings.
This will allow proper positioning of these
fittings, making it easier to align and connect hydraulic lines.
a. Work Power Pack into position and install the
four mounting bolts and washers.
b. Connect electrical plug at back of Power Pack
. and safety.
c. Install the forward sections of hydraulic lines
routing to emergency band pump.
d. Connect brake hose disconnected for access.
e. Connect all hydraulic lines to Power Pack. Make
sure fittings are properly installed and connections
are tight. Install cover and drain hose.
f. Connect landing gear handle linkage to shaft at
Power Pack, and install and safety attaching roll
pin.
g. Fill reservoir. Fill brake master cylinder and
bleed corresponding brake.
h. With airplane on jacks, use Hydro Test to operate the landing gear through several cycles to bleed
system. Check for proper operation and any signs
of hydraulic leakage.
i. Reinstall upholstery and seats.
5-234. FIELD-TESTING THE HYDRAULIC POWER
PACK (INSTALLED IN AmCRAFT).
5-235. PRIMARY AND SECONDARY RELIEF VALVE
ADJUSTMENT. If the primary or secondary relief
valve should get out of adjustment, fluid contamination, wear of parts or defective parts should be suspected. Remove the power pack, disassemble, repair and adjust as outlined in the applicable paragraphs in this Section.
5-236. ADJUSTMENT OF PRIORITY VALVE.
a. Jack aircraft and connect test unit in accordance
with applicable paragraphs.
b. Check priority valve setting in accordance with
applicable paragraph.
c. If adjustment is required, turn priority valve
adjusting screw IN to increase pressure at which valve
5-84
•
opens, and turn the adjusting bcrew OUT to decrease
pressure at which the valve opens. Adjust so that the
valve opens at 750 to 800 psi as noted on test unit gage.
d. Cycle the landing gear to check for proper operation, then lower the gear.
e. Fill reservoir and disconnect test unit in accordance with applicable paragraph.
f. Remove aircraft from jacks.
5-237. HANDLE-RELEASE ADJUSTMENT. (Refer
to figures 5-33 thru 5-36.) Adjustment of the gear
handle-release mechanism is necessary because incorrect adjustments can cause excessive pressure in
the Power Pack and can prevent free circulation of
fluid, resulting in damage to the Power Pack. If the
mechanism releases too soon, the landing gear handle
may return to neutral before the landing gear doors
are closed, if the time-delay valve should function
improperly. Pressure build-Up after the doors are
closed operates the time-delay valve. Mter the
valve opens, pressure then disengages a springloaded plunger from a detent, and a handle return
spring then pushes the handle back to neutral through
mechanical linkage. The spring load on the detent
plunger and the spring load on each handle return
spring are adjustable. To adjust the handle-release
mechanism proceed as follows:
•
NOTE
The mechanical linkage between the landing
gear control handle and the Power Pack must
be rigged properly before handle-release adjustments can be made. Refer to steps "a"
thru "c".
a. Referring to figure 5-35, adjust push-pull rod end
so that the handle will permit the detent plunger in the
power pack to engage the cam detents in both the up
and down positions of the handle, and the handle does
not contact structure in either the up or down position.
b. Roll pin (2) will be approximately horizontal when
handle is at mid-point of barrier (8).
c. After adjustments have been completed, ensure
the rod end has suffiCient thread engagement and jam
nut is tight.
d. Jack aircraft, then connect test unit.
e. If power pack is being installed, or if reservoir
fluid level has been low, fill reservoir and bleed time-
•
•
delay valve in accordance with applicable paragraph.
f. Using test unit, cycle landing gear through at
least two full cycles, unless handle Will not hold or
fails to release.
NOTE
H the handle will not hold, either the detent
spring load adjustment is set too low, the
handle-return spring load adjustments are
set too high, or the handle-return springs
are bottoming out and not permitting the
handle-release plunger to reach the detent
positions. H the plunger cannot reach the
detent positions, loosen handle-return spring
adjusting plugs (2 and 3) until the plunger
will engage the detent.
cedure checks the release pressure from
the gear up position. This is performed only
to assure satisfactory operation of other eqUipment relative to handle release operations.
e. Set test unit bypass valve full open.
f. Place landing gear handle full up.
g. Very slowly close bypass valve until handle
trips back to neutral. Read gage at point of handle
trip. This pressure should be 750 to 1050 psi. Be
sure to allow time for time-delay valve to open.
h. Refer to paragraph 5-237 for handle-release
adjustment.
1. Make sure landing gear is down and locked and
disconnect Hydro Test unit.
j. Remove airplane from jacks.
5-239. CHECKING TIME-DELAY VALVE.
H the handle will not release, either the detent
spring load adjustment is set too high (forcing
the detent plunger partially into the detent and
making it mechanically impossible for the
plunger to move completely out of the detent)
or the handle-return spring load adjustments
are set too low. Tighten detent spring load
adjusting plug (1) until detent plunger bottoms
out in detent, then loosen plug (1) approximately two full turns, until handle will release.
•
g. Using test unit, check the pressure at which the
handle-release plunger disengages the detents and
readjust handle-release detent spring adjusting plug
(item 1, figure 5-36), as necessary to obtain a release pressure of approximate~y 900 psi. Tolerance
is 750 to 1050 pSi. Use a minimum flow and ensure
time is allowed for the time-d,~lay valve to open. Cycle the landing gear between each adjustment.
h. Recheck the handle-release pressure specified
in step "g".
i. Operate lnading gear through several cycles to
check for proper operation, then lower the gear.
j. Fill reservoir and disconnect test unit in accordance With applicable paragraph,
k. Remove aircraft from jacks.
5-238. CHECKING HANDLE-RELEASE TO
NEUTRAL.
a. Cycle landing gear through two co::nplete cycles,
ending with the gear down and locked, and the doors
closed.
b. Set test unit bypass valve full open.
c. Place landing gear handle to full down.
d. Very slowly close bypass valve until handle
trips back to neutral. Read gage at point of handle
trip. This pressure should be 750 to 1050 psi. Be
sure to allow time for time-delay valve to open.
NOTE
•
One release valve serves to release the
handle from both the gear down and the gear
up positions. If the handle-return springs
are adjusted correctly, the release val-.. . e
should release the handle from both poSitions
at the same pressure. The foregoing procedure checks the release pressure from
the gear down position, and the following pro-
NOTE
The time delay between clOSing of the landing
gear doors and releasing of the landing gear
handle to neutral should be between 3 to 9
seconds at room temperature. Colder temperatures will cause a longer delay.
a. Connect test unit.
b. Set Hydro Test at approximately. 1500 pSi, with
a one gallon-per-minute flow rate.
c. With airplane master switch OFF to open the
doors, move landing gear handle to down position
and turn master switch to ON position. Note the
time delay between closing of the doors and releasing of the handle to neutral. See the preceding
"NOTE. "
d. There is no adjustment of the time-delay valve.
If it is defective, refer to applicable figure and paragraphs for disassembly and repair of the power pack.
e. Disconnect test unit.
5-240. CHECKING PRIORITY VALVE.
a. Cycle landin.!!; gear through two complete cycles.
b. With landing gear down, turn master switch
OFF to open gear doors. Leave the switch OFF to
permit doors to remain open, thereby making it
easier and faster to complete this check.
c. Open Hydro Test bypass valve.
d. Place landing gear handle full up. Very slowly
close bypass valve, observing Hydro Test pressure
gage and Hydro Test flow gage, until priority valve
opens. Priority valve should open at a pressure of
750 to 800 pSi.
NOTE
As the priority valve opens, the nose gear
downlock starts to release. Read Hydro
Test pressure gage at this point. The Hydro
Test flow gage will also aid in positively establishing opening of the priority valve. As
pressure slowly builds up in the door system,
there is practically no flow of fluid and the
flow indicator will be resting on the bottom
of the sight glass. As the priority valve opens,
the sudden increase in flow will cause the indicator to rise in the sight glass.
5-85
e. Refer to appUcable paragraph for priority valve
adjustment.
f. Ensure landing gear is down and locked, and
disconnect test unit.
g. Remove aircraft from jacks.
5-241. CHECKING PRIMARY (SYSTEM) RELIEF
VALVE.
a. Connect test unit.
b. Open test unit bypass valve.
c. Place landing gear handle full-down.
d. Slowly close bypass valve, observing pressure
b~ild-up and note point at which pressure stabilizes
on test unit gage. Stabilization indicates relief valve
setting. ReUef valve pressure Should be 1450-1500
pSi, at a flow rate of approximately o~e gallon-perminute on the test unit.
e. The power pack must be removed and partially
disassembled to adjust the primary relief setting
(refer to paragraph 5-228.)
f. Disconnect test unit.
5-242. CHECKING SECONDARY (HAND PUMP) RELIEF VALVE (PRIOR TO 1968 MODEL).
a. Place landing gear handle full down. With
master switch OFF, operate emergency hand pump
to open landing gear doors.
b. Disconnect and plug door open-line at firewall
fitting, and connect Hydro Test pressure hose to
this fi rewall fitting.
c. Close lockout valve on Hydro Test.
d. Operate emergency hand pump in airplane
observing Hydro Test pressure gage for pressure
at which secondary relief valve opens. This pressure should be 1800 to 1900 pSi.
e. The Power Pack must be removed and partially
disassembled to adjust secondary relief valve setting.
r. Open lockout valve on Hydro Test to release
the pressure, disconnect Hydro Test pressure hose,
and reconnect door open line.
g. Replenish hydraulic reservoir fluid as required.
5-243. CHECKING FOR SUCTION Am LEAKAGE.
a. Remove engine cowling as necessary for access
b. Disconnect hydraulic pump suction (larger) hose
from pump and connect Hydro Test suction (larger)
hose to the airplane suction hose, using a suitable
fitting.
c. Disconnect hydraulic pump pressure (smaller)
hose from pump and connect Hydro Test pressure
(smaller) hose to airplane pressure hose, using
a suitable fitting.
d. Connect Hydro Test vent hose to airplane reservoir vent line, protruding below lower edge of
firewall.
NOTE
BeforE" making this connection, be certain
the line is wiped clean and is free of any
dirt or foreign material which might have
worked into the line. If the line is dirty
internally, remove and flush with solvent,
then dry with compressed air and reinstall.
e. Connect Hydro Test electrical cable to appropriate electrical power source.
5-86
f. Jack the airplane and cycle the landing gear
through five complete cycles. No air should be
visible in Hydro Test sight gage.
g. Air visible in sight glass indicates leakage in
suction line, hose, or fittings. Replace defective
parts.
•
NOTE
If replacement of parts stops any visible air
in Hydro Test sight glass but air still enters
hydraulic system, engine-driven pump may
have a suction leak.
h. Make sure landing gear is down and locked and
remove airplane from jacks.
'
1. Disconnect test unit.
5-244. BLEEDING TIME-DELAY VALVE.
NOTE
The time-delay valve in the power pack
may be purged of air by operating the
engine-driven pump or the test unit may
be used.
a. Ensure reservoir is full.
b. Start engine and let run at 1000rpm, or connect
test unit.
c. Place landing gear handle in down position and
hold for apprOximately one minute, while turning the
master switch OFF until doors open, then ON until
doors close.
d. Repeat step "e" four times, waiting one minute
between each repeat. This allows time-delay valve
to refill.
e. Check that time-delay valve operates properly
by moving landing gear handle sharply to the down
position and recording time as handle returns to
neutral.
•
NOTE
The time delay between closing of the landing
gear doors and releasing the landing gear
handle to neutral should be between 3 and 9
seconds at room temperature. Colder
temperatures will cause a longer delay.
f. Shut down engine or disconnect test unit.
5-245. EMERGENCY HAND PUMP.
5-246. DESCRIPTION. The emergency hand pump is
mounted on a support beneath the floorboard just in
front of the front seats, near the center of the floorboard. The handle extends into the cabin and is enclosed by a hinged cover. The pump supplies a flow
of pressurized hydraulic fluid to open the doors and
extend the landing gear if hydrauliC pressure should
fail. The hand pump reseives a reserve supply of
fluid from the power pack reservoir and pumps the
fluid directly to the door control valve and gear pri0rity valve, then into the passages and lines used by
the regular system.
•
Ie
23
HAND PUMP OUTLET VALVE
•
24
1.
2.
3.
4.
5.
6.
Handle Stop
Lever
Handle Latch Sprlng
Handle
Knob
Pump Body
HAND PUMP INLET VALVE
7.
8.
9.
10.
Spring
Ball
Seat O-Ring
Seat
11. Snap Ring
12.
13.
14.
15.
16.
Snap Ring
Seat
Seat O-Ring
Ball
Spring
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
Plunger O-Ring
Plunger
Gland External O-Ring
Gland
Gland Internal O-Ring
Scraper
Hand Pump Bracket
Llnk
Linkage Pins
Spacer
5-38. Emergency Hand Pump
5-247. REMOVAL.
a. Loosen carpeting around hand pump, and remove
cover and pan.
b. Wedge cloth under hydraulic fittings to absorb
flufd, then disconnect the two hydraulic lines and plug
openings.
c. Remove the two mounting bolts and work hand
pump assembly out of floorboard.
•
5-248. DISASSEMBLY.
After the emergency hand pump has been removed
from the airplane and the ports are capped or plugged,
spray with cleaning solvent (Federal Specification
P-S-661, or equivalent) to remove all accumulated
dust or dirt. Dry with filtered compressed air.
To disassemble the unit, proceed as follows:
a. Remove hand pump handle by removing pivot
and linkage pins after removing cotter pins.
b. Cut safety wire and remove four Allen head
screws attaching hand pump bracket, and remove
bracket. Do not remove bushing in hand pump
braccket unless replacement of bushing is necessary.
c. Using a punch or rod in holes at end of hand
pump plunger, pull plunger from pump body.
d. Using snap ring pliers, remove snap ring at
inboard end of hand pump plunger. Remove seat,
ball, and spring from plunger by applying a sharp
5-87
blast of compressed air in the Side hole in the
plunger.
e. Remove gland and scraper from plunger.
f. Inside suction port of hand pump, remove
snap ring, seat, ball, and spring. Use a brass
hook to remove seat, ball, and spring.
g. Remove and discard all O-rings.
5-249. TROUBLE SHOOTING.
TROUBLE
REMEDY
PROBABLE CAUSE
HAND PUMP EXTERNAL
LEAKAGE.
SLIDING SEALS:
(Seals having a moving part. )
Hand pump plunger.
Remove hand pump plunger and replace O-ring.
HAND PUMP EXTERNAL
Hand pump gland.
LEAKAGE.
STATIC SEALS.
(Seals with no moving parts. )
Remove hand pump and replace
O-rings and scraper ring.
Hand pump fittings.
Remove and replace O-rings and
back-up rings as required.
5-250. INSPECTION OF PARTS.
a. Inspect seating surfaces of seats. They should
have very sharp edges. Seats may be lapped if
necessary to obtain sharp edges.
b. Inspect plunger for scores, burrs, or scratches
which could cut O-ring.. This is a major cause of
. external leakage. The plunger may be polished with
extremely fine emery paper. Never use paper
coarser than No. 600 to remove scratches or
burrs. If defects do not polish out, replace
plunger.
5-251. ASSEMBLY.
a. Insert spring and ball in pump body through
suction port.
b. Lubricate and install O-ring on seat and install seat through suction port with sharp edge of
seat next to ball. Secure seat with snap ring.
c. Install spring and ball in hand pump plunger.
d. Lubricate and install O-ring on seat and install
seat in hand pump plunger with sharp edge of seat
next to ba11. Secure seat with snap ring.
e. Lubricate and install O-ring on plunger, and
internal and external O-rings on bronze gland.
f. Install gland on plunger, and insert plunger
and gland into pump body.
g. Install scraper ring in counterbore of gland.
Install so that flat surface of scraper is in counterbore of gland, with inner protruding r;>art of scraper
faCing outward.
h. Attach hand pump bracket to pump body with
four Allen head screws. Tighten screws evenly
and install lock wire.
i. Install hand pump handle with pivot and linkage
pins. Secure pins with washers and cotter pins.
j. With a hydraulic source attached to the suction
port of pump, actuate pump to see that it operates
properly.
5-252. INSTALLATION.
a. Carefully work pump into openins in floorboard
and position pump body to mounting bracket; install
bolts and tighten.
5-88
•
b. With cloth under bydraulic fittings to absorb fluid,
uncap and connect hydraulic lines to pump. Bleed lines
and pump as connected in accordance with paragraph
5-252.
c. Install cover and pan and secure carpeting.
5-253. BLEEDING OF THE EMERGENCY HAND
PUMP. The band pump and hydraulic lines may be
bled by disconnecting the hand pump pressure (small)
line at the bottolIl of the Power Pack, operating the
hand pump until all air has been expelled from the
pump and lines, and reconnecting the line. Provide
can and drip cloth to protect carpeting. After reconnecting line, operate hand pump while master
switch is OFF until landing gear doors are fully
open. Continue to operate hand pump very slowly,
increasing pressure until the secondary relief valve
opens and all air is bled from hand pump and valve.
•
/CAUTION!
It is very important that the hand pump be-
operated very slowly as pressure is being
increased to bleed the secondary relief valve.
If the hand pump is operated rapidly, damage
to the valve can occur as air permits parts
to "slam" against each other.
5-254. DOOR CLOSE LOCK VALVE.
5-255. DESCRIPTION. Wheel door actuators are
held in the closed position by pressure trapped in the
door close line by the door close lock valve. This
enables the doors to remain in the closed position.
The valve is located in the front engine compartment,
attached to the outside of the left-hand nose gear lunnel wall.
5-256. REMOVAL.
NOTE
The doors might come open as a result of
•
•
TO FORWARD
WHEEL DOOR
ACTUATOR
FITTING-----,
. - . 0 . , ....-
TO AFT
WHEEL DOOR
,
ACTUATOR
';;:",<..
FITTING --..../'
.
LOCKOUT
VALVE - - - - '
~-+-~~--
NOSE GEAR
DOOR OPEN
FIREWALL
LEFT-HAND SIDE OF
ENGINE COMPARTMENT
A
2
•
Detail
A
4
8
1
9
10
7
1
1. Fitting
2. Packing
3. Piston
•
NOTE
Before assembly, lubricate
all packing with Petrolatum
or MIL-H-5606 hydraulic
fluid.
4.
5.
6.
7.
Housing
Check Valve
Seat
Back- Up Ring
8. Ball
9. Guide
10. Spring
5-39. Door Close Lock Valve
5-89
NOTE
changes in ambient temperature. The valve
has no provision for changes in pressure,
due to changes in ambient temperature. If
doors open, check for leakage or damage
prior to removing valve.
a. Drain power pack in accordance with applicable
paragraph.
b. Remove left hand cowling from front engine.
c. Disconnect and cap or plug hydraulic lines to
valve; remove valve.
5-257. DISASSEMBLY. (Refer to figure 5-39.)
a. Remove end fitting (1), packing (2), piston (3)
and back-up rings (7) from hO:.lsing (4).
b. Remove fitting (1), packing (2) and check valve
(5) from hO'.1sing (4) at opposite end of valve.
c. Remove seat (6) along with packing (2) and backup rings (7).
d. Remove ball (8), guide (9) and spring (10).
5-258. INSPECTION OF PARTS.
a. Insp!!ct threaded surfaces for cleanliness and
freedom of cracks and excessive wear or damage.
b. Inspect seat (6) for sharp seating edge with ball
(8). Lap as necessary to obtain a sharp seating edge.
c. Inspect piston (3) and guide (9) for cracks,
scoring, wear or surface irregularities which might
affect their function or the overall function of the
door close lock valve.
NOTE
Repair of most parts of the door close
lock valve assembly is impractical. Replace defective parts with serviceable
parts. Minor scratches may be removed
by polishing with fine abrasive crocus
cloth (Federal Sp.~cification P-C -458),
providing their removal does not affect
operation of the unit.
Install all new packing and back-up rings
during assembly. Lubricate all packing
and back-up rings with Petrolatum or
MIL-H-5606 hydraulic fluid .during assembly.
•
a. Install new packing (2) and back-up rings (7) on
piston (3); install in housing (4) with end fitting (1).
Use care to prevent damage to packing and back-up
rings.
b. Install packing (2) and check valve (5) into housing
(4) with end fitting(I).
NOTE
Install check valve with flow arrow pointing toward end fitting.
c. Install ball (8), guide (9), spring (10); install
packing (2) and back-up rings (7) on seat (6) and
install into housing (4).
5-260. INSTALLATION.
a. Connect hydraulic lines to valve.
b. Refill reservatr.
c. Jack aircraft in accordance with procedures outlined in Section 2.
d. Cycle gear several times to bleed system.
e. Check for proper operation and leakage of fluid.
f. Remove aircraft from jacks.
g. Install cowling.
5-261. LANDING GEAR ELECTRICAL CmCUITS.
5-262. DESCRIPTION. Landing gear electrical circuits are shown in figure 5-40, which shows the switches in the gear-down and locked, weight-on-gear condition. The following chart describes the function of
each electrical compoaent and what causes it to operate.
•
5-259. ASSEMBLY (Refer to figure 5-39.)
SHOP NOTES:
5-90
•
ITEM
•
OPERATED BY
FUNCTION
UP INDICATIOR SWITCH
Gear in up and locked position.
Closes circuit to gear up indicator
light, handle up-down switch, and
door solenoid valve.
DOWN INDICATIOR SWITCH
Gear in down and locked position.
Closes circuit to gear down indicator
light, handle up-down switch, and
door solenoid valve.
HANDLE UP-DOWN SWITCH
Power Pack selector spool.
"Preselects" up or down circuit.
(Completes up circuit to door solenoid
valve when gear reaches up position,
completes down circuit to door
solenoid valve when gear reaches
down position. )
DOOR SOLENOID VALVE
Completion of up circuit or down
circuit. (Handle up-down switch
and all gear indicator switches
closed. )
Shifts valve to door-close position
when energized. Spring-loaded to
door-open position. Thus, with an
electrical failure, the solenoid valve
will remain in the door-open poSition
and doors cannot be closed.
NOTE
•
Remember this rule: CLOSED circuit ;:: CLOSED doors; OPEN circuit ;:: OPEN doors. Applythis rule, the doors can be opened or closed at will by placing handle in down or down
neutral, turning master switch either on or off, and supplying pressure with the hand pump.
NOSE GEAR SAFETY SWITCH
Actuating arm on lower torque
link.
When airplane weight causes shock
strut to compress, switch opens circuit to handle lock-out SOlenoid,
which is spring-loaded to lock position. When airborne, strut extends
and closes switch, to unlock handle
from gear-down range.
HANDLE LOCK-OUT SOLENOID
Nose gear safety switch.
Prevents handle from being moved
out of gear-down range while airplane is on ground.
ICAUTIONJ
Since a fully extended strut (too much air pressure, extremely aft weight distribUtion, etc.) simulates an airborne condition, be especially careful not to move gear handle from gear-down range
under these conditions, or nose gear will retract .
•
5-91
STALL" GEAR
WARNING UNIT
,
I
:
.
•
:~~.....
ST~: t.~'.~.~~~. !.~.~ ~.~~
~
·Ul +. . .-r. )
....
:
I
T-SW
GEAR
+..·_......
................ _._.._ ... _.__...................:t.~~~:TTER
:
.:
._..--' ~J.:.
L. G. DOOR SOLENOID
•
i
. --'
T
-:;:THROTTLE
ACTUATED
SWITCHES
24V
5-------~
B
U
S
...... '
~'
STALL WARNING
HANDLE
LOCKOUT
SOLENOID
GEAR UP SWITCHES
NOSE GEAR MAIN GEAR
RIGHT LEFT
HANDLE
LIGHT TEST
CIRCllT
UP-DOWN
SWITCH
RIGHT
LEFT
PUSH- 1'0- TEST
GEAR POSITION
INn LTS
NOSE GEAR MAIN GEAR
GEAR DOWN SWITCHES
•
NOSE GEAR
SAFETY
SWITCH
5-40. Simplified Schematic of Landing Gear Circuits
5-263. SWITCH ADJUSTMENT. Lan1ing gear up indicator switches, down indicator switches, nose gear
safety switcn and handle up-down switch may be adjusted as outlined in the rigging procedures beginning
with paragraph 5-264. Adjustment of throttle actuated switches is outlined in Section 10.
,
5-264. RIGGING OF MAIN LANDING GEAR.
@~UTI~NI
When raising or lowering main landing gear
manually, avoid forcing or jerking an individual gear to prevent streSSing the universal
jOints. Apply equal pressure to each gear.
5-265. RIGGING OF ADJUSTING SUPPORT. (Refer
to figurE 5-41.) The adjusting support is bolted to the
outboard forging and forms the down stop for the
main gear. Jack the airplane and rig the adjusting
support as follows:
5-92
NOTE
The spring strut must be installed and secured before rigging the adjusting support.
a. Check for contact between flat surface of strut
and lower surface of adjusting support. Minor gaps
may exist as long as contact is made near each end
of support. Shim as required between outboard
forging and adjusting support. The following shims
are available from the Cessna Service Parts Center.
AFT
1541041-1 ................................ •
-2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. .012"
-3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . • 020"
-4 ....................... , .. .. .. .. .032"
-5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. .006"
•
SHIM AS REQUIRED FOR CONTACT
BE'IWEEN ADJU& rING SUPPORT
AND LANDING GEAR STRUT
ADJUSTING
SUPPORT
FULL CONTACT (OR
AT LEAST CONTACT
NEAREACHEND)----~--~--+__+------~--~
LOCATE AND DRILL WEDGE
FOR SLIGHT DRAG OF STRUT
TO • 010" MAX. CLEARANCE
SLIGHT DRAG OF STRUT
TO • 005" MAX. CLEARANCE
Figure 5-41. Rigging of Adjusting Support
FWD
•
1541041-6 .....•..............••.......... •
-7. . . . . . .. . • . . . . . . . . . . . . . . . . . . . . .. .012"
-8................................ .020"
-9. • . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. .032"
-10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 006"
·Sheet of • 025" laminated with ten . 002" additional
removable laminations.
b. Check that the aft edge of strut contacts adusting support (.005" maximum clearance) as Shown,
when gear is down. To shift adjusting support fore
and aft, first loosen the three bolts securing the
support (elongated holes are provided in the support),
then adjust the two jam nuts as required and retighten the three mounting bolts.
c. Check that the forward edge of strut contacts
wedge (.010" maximum clearance) as shown, when
gear is down. If adjustment is necessary, locate,
drill, and countersink a new wedge, and install with
screw, washer and nut.
NOTE
A slight drag is permissible as gear reaches
the full down position.
. The following wedges (measured at thickest part)
are available from the Cessna Service Parts Center.
•
1541029-1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. .250"
-2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. .300"
-3 ............................... , .330"
-4 ......... (Beginning with 337 -0838)
5-266. RIGGING OF DOWNLOCK MECHANISM.
360"
The downlock is a hydraulically operated pawl containing an adjustable downlock pin which wedges
under the forward edge of the strut to lock the gear
in the down position. Jack the airplane and rig as
follows:
The main gear downlock cylinders shall be
aligned at all times with the main gear downlock. The cylinders shall also be canted outboard and free from interference with structure, upholstery panels, etc., throughout their
normal operating range.
a. Check that downlock pin reaches the overcenter
pOSition shown in figure 5-42 (.03" to . 10"). Adjust
upper stop bolt as required to obtain this position.
b. Check that downlock pin reaches the retracted
position shown in figure 5-42 (.18" to .22"). Adjust
lower stop bolt as required to obtain this position.
NOTE
A downlo:k pin rigging tool, shown in figure
5-43, is available from the Cessna Service
Parts Center.
c. Check over-all length of downlock pin as shown
in figure 5-42 (snugly against strut to .005" maximum clearance), with hydraulic pressure on gear.
Downlock pin assembly must be removed to change
over-all length.
d. Disconnect actuator clevis from fuselage
bracket and use handpump to pressurize the actuator in its fully retracted position. With the actuator piston bottomed out, move the downlock and
actuator up and down as shown in figure 5-44.
5-93
•
SNUG CONTACT
TO • 005" MAX.
CLEARANCE:------~t;~~~~~~~~~~~~
DOWNLOCK PIN
SPRlNG-----
- - - . 18" TO .22"
SPRlNG-----~=_~~~
LOWER STOP
•
BOLT-~
Figure 5-42. Rigging of Downlock
Measure the minimum clearance between the actuator clevis and fuselage bracket, and install shims
as required to eliminate this clearance. Reconnect
clevis to bracket and secure. The following shims
are available from the Cessna Service Parts Center.
1512359 -1 .....................•......... • 125"
-2 ............................... .032"
e. Check that button in overcenter arm is screwed
completely in (shortened) as shown in figure 5-45, and
jam nut is tight. Check that over center arm retracts
smoothly when engaging strut and that arm is clear of
roll pin installed in down lock when gear is down and
locked.
f. Check that overcenter release bolt in upper end
of downlock extends below adjusting sl.lpport as shown,
with actuator piston bottomed out retracted and hydraulic pressure applied. See figure 5-45.
@AUTION\
Overcenter release bolt must not extend too
far, as damage to parts can be caused during
retraction, especially if gear is aided manually.
g. Release hydraulic pressure and check that overcenter stop bolt in bulkhead is adjusted so that overcenter release bolt in upper end of downlock extends
below adjusting support as shown in figure 5-45
5-94
(.06" more than straight line dimension), when
actuator is held on overcenter position against the
bulkhead stop bolt.
h. Check action of downlock switch bracket cam as
follows:
1. Place main gear in "trail position.
2. Manually push downlocks into the normally
locked position (aft).
3. Holding apprOXimately 20 pounds of force
against each wheel, extend gear to the down and
locked position. Cams on the switch brackets should
push downlocks out of the way, allowing gear to move
smoothly into the down and locked position.
4. Repeat test at least five times.
5-267. RIGGING OF UPLOCK MECHANISM. (Refer
to figure 5 -6.) The uplock is a hook which is springloaded to the locked pOSition and hydraulically-operated to the uplocked position. Jack aircraft (figure
2-1) and rig the uplock mechanism as follows:
a. (Prior to 1971 Model.) Adjust push-pull rod ends
as required to cause the hooks to release the gear
spring struts Simultaneously when operated hydraulically.
NOTE
In addition to releasing gear struts simul-
taneously, linkage must be adjusted so no
•
•
CD
MANUALLY HOLD DOWNLOCK PIN
AGAINST UPPER STOP BOLT
.
DOWNLOCK - -.....
DOWNLOCK PIN - - OVERCENTER HOLE
IN RIGGING TOOL
UPPER STOP
BOLT
7
BOLT~
DOWNLOCK PIN
RIGGING TOOL
LOWER STOP
•
® AFT
FORWARD EDGE OF TONGUE CONTACTS
EDGE OF DOWNLOCK PIN _ _ _
----I
POSITION LOWER FLANGE OF TOOL IN
FULL CONTACT WITH FLAT SURFACE
OF DOWNLOCK
OVERCENTER POSITION
OF DOWNLOCK PIN
With downlock pin depressed (1), lower bolt in
lower hole (3), lower flange flat against downlock (4), and forward edge of tongue contacting
aft edge of pin (5), upper bolt should fall within overcenter hole (2). Elongation of overcenter
hole represents tolerance permissible; adjust
upper stop bolt as required.
NOTE
Jack the airplane, retract the landing gear, and release hydraulic pressure, leaving the landing gear
doors open. Pull downlock assemblies aft for access.
•
The downlock pin rigging tool, Part No. SE772-1, is available from the Cessna Service Parts Center.
The tool is made in two halves - the left half is shown in use for the left downlock pin; the right half is
used in the same manner for the right downlock pin .
Figure 5-43. Using DoWnlock Pin Rigging Tool (Sheet 1 of 2)
5-95
CD
•
DOWNLOCK PIN RELEASED AGAINST
LOWER STOP BOLT - -....
POSITION THIS HOLE
ON THIS BOLT
RETRACTED HOLE
IN RIGGING TOOL
UPPER STOP
BOLT
~
CD FORWARD
EDGE OF TONGUE CONTACTS
AFT EDGE OF OOWNLOCK PIN
---~
OOWNLOCK PIN
RIGGING TOOL
POSITION LOWER FLANGE OF TOOL IN
FULL CONTACT WITH FLAT SURFACE
OF DOWNLOCK
•
RETRACTED POSITION
OF DOWNLOCK PIN
With downlock pin not depressed (1), lower
bolt in lower hole (3), lower flange flat against
downlock (4), and forward edge oftongue contacting aft edge of pin (5), upper bolt should
fall within retracted hole (2). Elongation of
retracted hole represents tolerance permissible; adjust lower stop bolt as required.
Figure 5-43. Using Downlock Pin Rigging Tool (Sheet 2 of 2)
5-96
•
•
MOVE UP AND DOWN TO ESTABLISH
MINIMUM CLEARANCE BETWEEN
CLEVIS AND MOUNTING BRACKET
ACTUATOR FULLY
RETRACTED BY
HYDRAULIC PRESSURE
1
ACTUATOR
•
MOUNTING
BRACKET
SHIM AS REQUIRED TO ELIMINATE
CLEARANCE BETWEEN CLEVIS
AND MOUNTING BRACKET
Figure 5-44. Shimming of Downlock Actuator
part of linkage (including up indicator switch)
contacts any part of the aircraft structure.
Actuator piston must bottom out retracted
before hydraulic fluid can be routed through
the actuator to lower the main gear.
•
b. (Prior to 1971 Model.) Vertical adjustment is
provided by Shifting washers from one side of the
hook to the other side, since the retracted pOSition
of the spring struts is not horizontal. Adjust as required to permit the spring-loaded hooks to engage
the spring struts.
c. (Prior to 1971 Models.) Inboard-outboard adjustment is provided by a slotted hole in the uplock
supporting structure. Adjust uplock in this slot so
that the gear spring always cams the hook toward the
locked pOSition.
d. (Prior to 1971 Model.) With gear up and pressurized, clearances shown in figure 5 -6 must be
attained.
e. (1971 Model.) Loosen the bolts attaching the
hangers to the supports to allow inboard and outboard adjustment.
f. (1971 Model.) With Hydro Test connected, open
the Hydro Test bypass valve to reduce hydraulic pressure to approximately 1000 psi. With gear up and
pressurized, check position of the gear stops.
g. (1971 Model.) The outboard edge of the gear
strut spring should contact the stop and the slanted
portion of the stop should be parallel to the strut
spring maintaining 20 percent contact with strut
spring.
h. (1971 Model.) The stop is adjusted to match the
angle of the gear strut spring by the addition of
shims (PiN 1541051-2) as required between the
hangers and supports.
i. (1971 Model.) Adjust push-pull rod ends as required to cause the hooks to release the gear strut
springs Simultaneously when operated hydraulically.
NOTE
In addition to releasing gear struts Simultaneously, linkage must be adjusted so no
part of linkage (including up indicator SWitch)
contacts any part of the aircraft structure.
Actuator piston must bottom out retracted
before hydraulic fluid can be routed through
the actuator to lower the main gear.
5-97
STRAIGHT LINE THRU CENTER LINES OF
PIVOT POINTS (AUTOMATICALLY FORMED
WlULE ACTUATOR IS FULLY RETRACTED
BY HYDRAULIC PRESSURE)
Z
ACTUATOR FULLY
RETRACTED BY
HYDRAULIC PRESSURE
•
ADJUSTING
SUPPORT
'----ACTUATOR
• At serial 337 -0030 and on and all spares, dimension "A" is .070" to .100". Prior to
serial 337-0030 (except spares), dimension
"A" is .20" to .22".
NOTE
At serial 337-0030 and on and all spares, the
overcenter release bolt and locknut are replaced with a bolt, washers (use as required
for adjustment), and a self locking Heli- coil
insert.
OVERCENTER STOP BOLT (Adjust
so dimension "B" is . 06 inch more
than dimension "A".)
HYDRAULIC
PRESSURE
RELEASED
•
ACTUATOR
-€::I
OVERCENTER
SPRING
BUTTON
Figure 5-45. Overcenter Adjustments of Retracted Downlock
5-98
•
•
5-268. RIGGING OF UP INDICATOR SWITCHES.
Main gear up indicator switches are mounted on
brackets attached to the uplock hooks. After jacking the airplane and retracting the landing gear until
uplock hooks are fully engaged, adjust the switches
so they are actuated with a minimum of 1/8 inch
travel of the switch plunger remaining. Switch case
must not contact any part of structure.
5-269. RIGGING OF DOWN INDICATOR SWITCHES.
Main gear down indicator switches are mounted on
brackets attached to the downlock. With landing
gear down and locked, adjust the switches so they
are positively actuated, but the leaf type switch
actuator does not contact the switch case.
5-270. RIGGING OF DOORS. After jacking the
airplane, main landing gear door adjustment is
accomplished by adjusting push-pull rod ends and
actuator rod ends as required to cause the doors to
close snugly. Doors must not close so tightly that
internal locks in actuating cylinders are not reached.
When installing new doors, some trimming and handforming at edges may be necessary to achieve a
good fit and permit actuators to lock. The doors
must clear the gear during retraction at least
1/2 inch.
•
5-271. ADJUSTMENT OF SNUBBER VALVE.
A main gear snubber valve, which restricts fluid
near the end of the gear-up cycle, is provided at the
aft end of the main gear actuator. This valve is a
hollow, contoured metering pin which forms the
hydraulic fitting at the aft end of the actuator. The
purpose of the snubber valve is to slow down action
near the end of the gear-up cycle- to cause smoother
locking action. Jack the airplane and adjust snubber
valve as follows:
a. Connect test unit.
b. Cycle the landing gear, noting the pressures
on the Hydro Test gage.
NOTE
The Hydro Test gage will indicate various
pressures during gear retraction. The first
level is the pressure needed to operate dooropen system (approximately 300 psi). The
second level is the pressure needed to retract
landing gear (apprOximately 900 psi). The
third level is the momentary pressure increase as snubbing action occurs. Pressure
should increase to system operating pressure
(1500 psi) for no longer than two seconds.
After snubbing occurs and gear up-locks,
pressure will decrease to pressure needed
to operate door-close system (approximately 300 psi), then will again build up, through
time-delay valve, until handle returns to
neutral.
•
c. If snubbing action does not occur or if pressure
does not increase momentarily to 1500 psi as previously noted, loosen jam nut and screw snubber out of
actuator as required. If pressure increases to
1500 psi and remains there for more" than two seconds, loosen jam nut and screw snubber into the
actuator as required until pressure and time limit
specified are attained. After adjustment, tighten
and safety jam nut.
[~AUTIONI
A snap ring on the snubber bottoms out against
the end of the actuator as the snubber is backed
out. Do NOT use force or damage will result.
NOTE
Another possible cause of excessive downlocklog time is a plugged or otherwise faulty restrictor valve between the main gear downlock
cylinders, as shown in the hydraulic system
schematics.
d. Fill reservoir and disconnect test unit.
e. Remove aircraft from jacks.
5-272. RIGGING OF NOSE GEAR.
NOTE
The nose gear shock strut must be properly
inflated prior to rigging of the nose gear.
5-273. RIGGING OF DOWNLOCK MECHANISM. (Re
fer to figure 5-46.) The nose gear downlock mechanism is baSically a claw hook at the piston rod end of
the nose gear actuator. The actuator contains an
internal lock to hold the claw hook mechanism overcenter. Jack the airplane and rig downlock mechanism as follows:
a. Check that the hooks and crossbar are free from
drag as illustrated. Adjustment is provided by rod
end of actuator piston rod.
[~~UTION\
The piston rod is flattened near the threads
to provide a wrench pad. Do not grip the rod
with pliers, as tool marks will cut seals in
the actuator.
b. With the gear down and locked, adjust shims
(19, figure 5-23) behind the bumper to have light
contact to .001 inch clearance. Shims should not be
allowed to increase nose gear actuator locking or unlocking pressures.
5-274. RIGGING OF UPLOCK MECHANISM. (Refer
to figure 5-19.) The uplock mechanism is a hydraulically unlocked hook that is spring-loaded to the
locked poSition. It engages a roller on the upper
left side of the nose gear. Fore-and-aft adjustment
is provided by slotted holes in the actuator mounting bracket. Adjust so the hook will positively release the nose gear from its retracted poSition hydraulically, but will securely lock the gear up. With
the gear up and locked, and hydraulic pressure released, adjust nose gear rubber bumper to contact
the gear lightly.
5-275. RIGGING OF DOWN INDICATOR SWITCH. (Refer to figure 5 -47.) The nose gear down indicator
5-99
•
.,--- CROSSBAR (Must
rotate freely)
NOSE GEAR
IN
DOWNLOCK POSITION
!:!Q.!!: Locking of internal lock is indicated by inability to lift and disengage external claw lockS manually.
Locks shall release only when hydraulic pressure is applied at anchor end port of actuator.
Figure 5-46. Rigging of Nose Gear Downlock
•
• 04" TO .06" (Remaining travel of
hooks when switch contacts close)
SWITCH - - "
Figure 5-47. Adjustment of Nose Gear Down Indicator Switch
5-100
•
•
switch is operated by an arm on the downlock mechanism. After jacking the airplane, adjust the switch
to actuate with. 06" travel of the downlock hooks
remaining, as illustrated.
5-276. RIGGING OF UP INDICATOR SWITCH. The
nose gear up indicator switch is attached to the uplock hook. After jacking the airplane, adjust the
switch so it is positively actuated as the gear retracts, but the switch plunger has at least 3/32-inch
travel remaining.
5-277. RIGGING OF SAFETY SWITCH. The safety
switch, which is electrically connected to the landing
gear handle lockout solenoid, is operated by an actuator attached to the lower torque link. Adjust the
switch to actuate when the strut is between 1/8 and
1/4 inch from the fully extended position.
•
•
5-278. RIGGING OF DOORS. (Refer to figure 5-22.)
a. Jack the airplane.
b. Adjust rod ends (20) so that bellcranks (21)
clear torque tube (16) .05":.04" when doors are
closed and internal lock in actuator is engaged.
c. Adjust actuator rod end (11) so that forward
doors are open 14.50":.50" while the actuator is
pressurized. Measure this dimension between the
lower edges of the doors at their forward hinges.
d. Recheck bellcrank clearance per step "b, " and
readjust if necessary. Doors must clear nose gear,
at closest point during extension and retraction, by
at least 1/2 inch•
e. Adjust rod ends (4 and 5) so that aft door closes
snugly .
NOTE
Doors must not close so tightly that internal
lock in actuator is not reached. Some trimming and hand-forming at edges of new doors
may be necessary to achieve a good fit and
permit actuator to lock.
f.
Remove airplane from jacks.
5-279. RIGGING OF POWER PACK SWITCH AND
LOCKOUT SOLENOID.
5-280. RIGGING OF UP-DOWN SWITCH. The handle
up-down switch is located on the power pack and is
normally adjusted during assembly of the power pack.
With landing gear at centerline of barrier, adjust so
the switch actuates at an equal distance up and down
from centerline of barrier as landing gear handle is
moved up and down.
5-281. RIGGING OF GEAR HANDLE LOCKOUT.
The handle lockout solenoid contains a plunger which
prevents the handle from being moved upward from
the gear-down range. Adjust the small nut on the
solenoid plunger so the plunger fully locks the
handle, but clears the handle when actuated, even
with slight side-pressure exerted manually on the
handle.
SHOP NOTES:
5-101
•
HYDRAUUC SYSTEM SCHEMATICS
Figure 5 -48 (sheets 1 thru 10) is effective for Serials 337 -0001 thru 337 -0755. The
secondary relief valve is installed in the power pack.
Figure 5-49 (sheets 1 thru 10) is effective for Serials 337 -0756 thru 33701462 and
F33700001 thru F33700055. The secondary relief valve is deleted. This figure also
includes the door close lock valve which is installed in aircraft Serials 33701427 thru
33701462 and F33700052 thru 33700055.
Sheet 1 shows the system "at rest" with the landing gear up. Sheets 2 through 5
show various stages of the gear-down cycle, after which, the system is again "at rest"
With the landing gear down. Sheets 6 through 9 show various stages of the gear-up
cycle, after which, the system returns to the condition shown on sheet 1. Sheet 10
shows the landing gear being extended With the emergency hand pump Without electrical
power.
•
NOTE
The door vent valve shown in these schematics is not used
in early 1966 model Power Packs. However, replacement
Power Packs (new or remanufactured) have this valve installed. The valve relieves any pressure from thermal
expansion in the door system, to keep the doors closed
while the airplane is parked.
5-102
•
•
REAR ENGINE OPTIONAL HYDRAULIC PUMP INSTALLATION
The additional hydraulic pump and filter installation at the rear
engine requires that check valves be added in both pump pressure
lines, and a check valve with oruice be installed in the nose gear up
line, as shown in the partial schematic below. The check valves are
open as long as the pumps are supplying pressure, but the applicable
check valve will close if its pump should become inoperative. The
valve in the nose gear up line slows down nose gear retraction time.
These changes do not affect color coding in the following fold-out
pages.
REAR ENGINE
FRONT ENGINE
CHECK
VALVE
t
•
=
f
CHECK VALVE
WITH ORIFICE
______________________________
~
•
___________ ____________________________________
~~~~~-----k-------------.
OO'rIH
i.:~::
~
sw .
5-103/(5-104 blank)
•
PlLLII
MAIN GUI U'LOCIt
IIIIAS! CYLINDU
CODE
'InSUIE PlOW
InUIN Plow
STATIC 'IUSUIE
STATIC IETUIN
SUPPLY
VENT
flLTU
UP L'MIT sw.
•
NOSE GEAR
DOOR ACTUATOR
\
\
\
\
\
GUI
\ CONTIOL
\ LEVU
..
·· ..
····· ....
· .
i :
\
\
\
\
\
\
\
\
NOSE GEAR
POWER PACK
\
\
\
HANDLE
UP.DOWN 5W.
DOWN L'MIT 5w.
DOWN LIMIT SW.
MAIN GUI DOWN LOCI
CYlINDEIS
GEAR UP, DOORS CLOSED, PUMP UNLOADED
••
Figure 5-48.
Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 1 of 10)
5-105/(5-108 blaDk)
•
MAIN GIAI UPLOCI
IIUA$! CYUNDII
CODE
HYD
PlnSUIE fLOW
RnURN flOW
STATIC PlnSURE
STA TIC InUIN
SUPPLY
VENT
PUMP
fiLTER
UP LIMIT 5W.
NOSE GEAR
0001 ACTUATOR
•
\
\
LANDING
\
COG~~~L
\
\
\
\
\
\
\
\
···· ....
: .
\
\
NOSE GEAR
POWER PACK
HANDLE
UP.DOWN 5W
DOWN LIMIT 5W.
DOWN LIMIT SW.
MAIN GEAI DOWNLOCI
CYUNDUS
LANDING GEAR CONTROL JUST PLACED DOWN, DOORS OPENING
Figure 5-48.
•
Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 2 of 10)
5-107/(5-108 blank)
•
MAIN GIAI UPLOCIt
IUIAS! CTlINDfi
CODE
'IUSUIE flOW
RETUIN flOW
STATIC 'IESSURE
STA TIC IETUIN
SUPPLY
VENT
'ILTll
UP LIMIT
SW
NOSE GEAI
0001 ACTUA TOI
•
.
\
\
\
\
\
\
\
\
\
~
\
\
\
\
LEfT
MAIN
liGHT
GEAR
NOSE GEAR
POWER PACK
"ANDLE
UP·DOWN 5W
DOWN LIMIT SW.
··
··
1
DOWN LIMIT SW.
MAIN GUI DOWNLOCIt
CTliNDfiS
DOORS OPEN, GEAR UNLOCKED AND EXTENDING
Figure 5-48.
•
Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 3 of 10)
5-109/(5-110 blallk)'
•
MAIN GlAR UPlOCK
IlLlASl CTLINDU
CODE
RETURN flO ....
PRESSURE fLO ....
STA TIC PRESSURE
mDlDIIIIDID"
S TA TIC RnURN
VENT
SUPPLY
HYD. PUMP
f ILTfi
•
UP liMIT S .....
NOSE GEAR
DOOR ACTUATOR
\
\\
\
LANDING
GEAR
CONTIOl
\
LEVEl
NOSE GEAI
UPLOC~
IELEASE CYLINDfi
\
\
\
\
\
\
\
\
POWER PACK
I
HANDLE
UP.DO .... N 5 ....
I
DOWN liMIT 5 ....
MAIN GEAI DOWNlOCK
CYLINDERS
•
GEAR DOWN AND LOCKED, DOORS CLOSING
Figure 5-48.
Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 4 of 10)
5-111/(5-112 blank)
M... IN 01.... U'LOCI
IIU ... U (YLiNDI.
CODE
'IUSUIi PlOW
UrulN FLOW
STATIC 'InSUIi
STATIC InUIN
SUPPLY
VENT
flLTU
UP LIMIT SW
NOU OEAR
DOOI ACTUATOI
..
•
.
\
\
LANDINO
COO~~~L
\
\
LEVU
~
\
\
\
\
\
\
\
\
\
POWER PACK
\
\
\
HANDLE
UP.DOWN SW
DOWN LIMIT SW.
MAIN OEA. DOWNLOCI
CYLINDUS
GEAR DOWN AND LOCKED, DOORS CLOSED, HANDLE RELEASE PRESSURE BUILDiNG UP
Figure 5-48.
•
Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 5 of 10)
5-113/(5-114 blank)
I.
~fILLE'
..... ,N GE .... U'LOCK
ULEA$! CnlNDEI
CODE
PRESSUIE flOW
RnURN flOW
sa TIC Plnsuu
sa TIC RETURN
SUPPLY
VENT
f IL TER
UP lI.MIT
SW
NOSE GEAR
000 .... C TU ... TOI
•
POWER PACK
\
\
HANDLE
UP-DOWN SW
DOWN LIMIT SW
DOWN LIMIT SW
DOWN LIMIT SW.
MAIN GE ... R DOWNLOCK
CYLINDERS
LANDING GEAR CONTROL JUST PLACED UP, DOORS OPENING
Figure 5-48.
•
Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 6 of 10)
5-115/(5-116 blank)
•
MAIN GEAR UP LOCI
IE LEASE CYUNDtI
CODE
RnURN flO ....
PRESSURE flO ....
STATIC PRESSURE
_.......
STATIC RnURN
SUPPLY
---
VENT
flLTU
UP LIMIT 5 .....
NOSE GEAR
\
0001 ACTUATOR
\
\
•
\
\
\
..
LANDING
C;~:R~L
\
\
..
.
LEVU
\
\
~
~
\
\
\
\
\
\
NOSE GEAR
POWER PACK
HANDLE
UP.DO .... N 5 ....
DO .... N LIMIT 5 ....
DO .... N LIMIT S .....
MAIN GEAR DO .... NLOCI
CYlINDUS
DOORS OPEN, GEAR UNLOCKED AND RETRACTING
Figure 5-48.
•
Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 7 of 10)
5-117/(5-118 blank)
MAIN GIA' UPLOCK
IIUASI CYLINDU
CODE
PRESSURE flOW
RETURN
t8
flOW
STATIC IETURN
'B,
VENT
SUPPL Y
filTER
UP
lI~IT
SW
NOSE GEAR
0001 ACTUATOR
•
NOSE GEAI
UPLOCK
..
.: !:
MAIN
GEAR
NOSE GEAR
POWER PACK
HANDLE
UP.DOWN SW
DOWN liMIT SW
DOWN 1I .. ,T SW
"A IN GfAI DOWNL'OCK
CYlINDUS
GEAR UP AND LOCKED, DOORS CLOSING
•
Figure 5-48.
Hydraulic System Schematic 337-0001 thru 337-07·55 (Sheet 8 of 10)
5-119/(5-120 blank)
•
..AIN GEA. UPLOCIt
ULEASE CYLINDU
CODE
PRESSURE flOW
RETURN
flOw
:.r:.r:.I:I-.
ST A TIC PRESSURE
STA TIC RETURN
lII:I_E:IIIC;:;AGIII:
SUPPL Y
VENT
, IllER
uP
LIMIT' SW
NOSE GEA.
DOOR ACTUATOR
: : MAIN
•
!/ """,0'
..
,
LANDING
GEA.
CONTROL
LEVU
,
'
,,
,,
GEA.
\
\
\
\
\
NOSE GEAR
POWER PACK
HANDLE
UP.DOWN 5W
DOWN LIM'T SW.
DOWN LIMIT SW
MAIN GEA. DOWNLOCK
CYLINDUS
GEAR UP AND LOCKED, DOORS CLOSED, HANDLE RELEASE PRESSURE BUILDING UP
Figure 5-48. Hydraufic S'ystem Schematic 337-0001 thru 337 -0755 (Sheet 9 of 10)
•
5-121/(5-122 blank)
~ flLLU
c::xOOI
MAIN GUI UPLOCl
RELIASE CYLiNDII
CODE
RETURN flOW
PRESSURE flOW
STA TIC PRESSURE
III"ma_IIl'III"
STATIC
IE TURN
vENT
MAIN GEAR
DOOR ACTUATOR
f ILTlI
UP LIMIT
•
SW
NOSE GEAR
DOOR ACTuA TOR
~. '4~!~lJJMIjl--"
\\
\
\
LANDING
GEAR
CONTROL
LEVEl
....
··· ...
\
\
\
\
\
\
\
\
.
NOSE GEAR
..
!\
· ..
y
! \
NOSE GEAR
POWER PACK
HANDLE
UP·DOWN SW
DOWN LIMIT SW
DOWN LIMIT Sw
DOWN LIMIT SW
MAIN GEAR DOWNLOCl
CYliNDERS
'EMERGENCY CONDITION. ENGINE PUMP FAILURE, NO ELECTRICAL POWER.
DOORS OPENED, GEAR UNLOCKED AND BEING EXTENDED BY HAND PUMP PRESSURE
•
Figure 5-48. Hydraulic System Schematic 337-0001 thru 337-0755 (Sheet 10 of 10)
5-123/(5-124 blank)
•
PlUU
MAIN GEAI UPlOCI
IIUASE CYUNDII
CODE
i:E:ElC--=-K3II-
'IUSUIt flOW
IETUIN flOW
STATIC' If SSUII
STA TIC IETUIN
SUPPLY
-==
VfNT
Figure 5-49 (sheets 1 thru 10) :s effective for Serials
337-0756 thru 33701462 and F3 ;700001 thru F33700055.
The secondary relief valve is deleted. This figure also includes the door close lock valve which is installed
in aircraft Serials 33701427 anli F33700052 thru 33701462 and F33700055.
'Ilrn
I' WOIT SW.
•
HOse GfAI
DOOI ACTUAT,.,.
\
\
\
\
\
lANDING
GUI
\ CONTIOI
\ UVU
\
\
\
\
\
\
\
\
POWER PACK
\
\
! iii i
ii!
\
HANDU UP.DOWN SW.
DOWN liMIT 5W
DOWN liMn Sw.
iii
MAIN OIAI DOWN lOCI
CYUNDUS
GEAR UP, DOORS CLOSED, PUMP UNLOADED
•
FlcUr! 6-49. Hydnullc Sy.em SchemIWc (Sbeft 1 01 10)
6-12$/(5-121 biaaIr.)
I.
MA.N GeAI U'LOCI
IILiASI C'L1NOU
CODE
ENGINE.DIIYeN
PlUSUIl flOW
UTUIN flOW
STAnc 'IESSUIE
STAroC IETUIN
SUPPLY
VENT
Figure 5-49 (sheets 1 thru ]0) is effective for Serials
337 -0756 thru 33701462 and F33700001 thru F33700055.
The secondary rellef valve is deleted. This figure also includes the door close lock valve which is installed
in aircraft Serials 33701427 and F33700052 thru 33701462 and F33700055.
•
"UU
U' LIMIT SW
NOSE GEAR
DOOI AC TUA TO.
\
\
\
\
\
.."
..
·· ..:
\
\
\
\
\
\
\
\
NOSE GEAR
POWER PACK
".
i
i
i!
i
!
.. ANOLI U'.DOWN sw.
r---------~~OO-------------------------------------------------------o~
DOWN LIMIT SW
DOWN LIMIT Sw.
.AIN GIA' DOWNLOCIt
C'L1NOUS
•
LANDING GEAR CONTROL JU!;T PLACED DOWN, DOORS OPENING
FlIW'e
~49.
Hydra.oI1c System Schematic (Sheet 2 of 10)
~
12'1/(5-121 bla.lll<)
•
.. ... IN Ge .... U'IOCIt
Ule... ,( CYLINDU
CODE
now
_____
'.EISUI!
c::.3CElES:I-
ST... TIC PrESsu.e
RETU.N PlOW
ST ... TlC UTU.N
VENT
Figure 5-49 (sheets 1 thru 10) Is effective for Serials
337 -0756 thru 33701462 and F33700001 thru F33700055.
The secondary relief valve is deleted. This figure also includes the door close lock valve which is installed
in aircraft Serials 33701427 lnd F33700052 thru 33701462 and F33700055 .
•
NOS! Of ....
DOO .... CTU'" TO.
.
MAIN
lou.
"-CTU ... TO.
\
\
\
.I
\
\
\
\
\
.
\
\
\
\
lffT
MAIN
'IOHT
GEAR
NOSE GEAR
POWER PACK
DOWN II .. IT SW.
1
DOWN II .. IT SW
..... IN Of .... DOWN10CIt
CYIINDUS
•
DOORS OPEN, GEAR UNLOCKED AND EXTENDING
FltlUre S-49. HydnUllc System SchemaUc (Sheet 3 o( (0)
S-128/CS-13D bIaDt)
•
MAIN C!AI UPIOCK
IELIASI cnlNDU
CODE
=_==i~\
'IU5UIE HOW
RETUIN HOW
STATIC PlUSUI!
5TA TIC IETUIN
SUPPlY
VINT
Figure 5-49 (sheets 1 thru 10) is effective for Serials
337 -0756 thru 33701462 and F33700001 thru F33700055.
The secondary relief valve is deleted. This figure also includes the door close lock valve which is installed
in aircraft Serials 33701427 and F33700052 thru 33701462 and F33700055.
•
Fllua
UP liMIT SW
NOS! CI,u
0001 ACTUATOI
\
\ LANDING
\ C,,u
\ CONTIOI
(lYU
\
\
\
\
\
\
\
\
\
lifT
MAIN
GEAR
POWER PACK
HANDlI UP.DOWN 5W.
DOWN LIMU
$W
MAIN CUI DOWNIOCK
CnlNDUS
•
GEAR DOWN Af'iD LOCKED, DOORS CLOSING
Flg\II'e
~49.
Hydraulic Syslem Schemauc (Sheet. of 10)
~ 131/(~·132
blank)
•
MAIN OIAI UPLOCK
IILIAII CTUNDn
CODI
'Insun flOW
UTUIN flOW
STATIC 'InSUII
5T A TIC InUIN
VENT
Figure 5-49 (sheets 1 thru 10) is effective for Serials
337-0756 thru 33701462 anI! F33700001 thru F33700055.
The secondary relief valve is deleted. This figure also includes the door close lock valve which is installed
in aircraft Serials 33701427 and F33700052 thru 33701462 and F33700055.
•
"Llil
U, LIMIT SW
NOSE GEAI
DOOI ACTUATOI
.
.
~
POWER PACK
HANDl! U'.DOWN SW.
DOWN lI .. 1T SW.
MAIN GUI DOWNLOCK
CTUNDns
•
GEAR DOWN AND LOCKED, DOORS CLOSED, HANDLE RELEASE PRESSURE BUILDING UP
FIIw'e $-49. Hydra..uc System SchemaUc (Sheet 501 101
~
1S3/($- 134 blukl
•
MAIN GEAI U'LOCI
ULIAU cnlNDU
CODE
'USSUU 'LOW
IHUIN fLOW
STATIC 'Insuu
STATIC UTUIN
SU'PL'
VENT
Figure 5-49 (sheets 1 thru 10) is effective for Serials
337 -0756 thru 33701462 and F33700001 thru F33700055.
The secondary relief valve is deleted. This figure also includes the door close lock valve which is installed
in aircraft Serials 3370142'1 and F33700052 thru 33701462 and F33700055 •
•
UP LIMIf
sw.
NOSE GEAI
0001 ACTUATOI
LEfT
MAIN
liGHT
GEAR
POWER PACK
\
\
HANDLE UP·DOWN SW.
DOWN LIMIT SW
DOWN
DOWN LIMIT SW
LIMIT SW.
MAIN GEAI DOWNLOCI
cnlNDUS
•
LANDING GEAR CONTROL JUST PLACED UP, DOORS OPENING
Figure So-49.
HYdraulic System SchemaUc (Sheet 6 01 10)
$0 135/($0138 blaJlk)
•
MAIN OIAI UPIOCI
IIIIASI CYIINDII
CODE
PUSSUII
now
UTUIN flOW
STATIC PlUSUU
STATIC UTUIN
SUPPLY
VfNT
Figure 5-49 (sheets 1 thru 10) is effective for Serials
337 -0756 thru 33701462 and F33700001 thru F33700055.
The secondary relief valvl' is deleted. This figure also includes the door close lock valve which is installed
In aircraft Serials 337014!!7 and F33700052 thru 33701462 and F33700055.
•
UP liMIT
sw .
"OSI GfAI
0001 ... CTUATOI
\
\
\
\
\
.
\
L..... O' .. G
COG~~~L
\
\
LlVU
\
\
\
\
\
\
\
\
~
NOSE GEAR
POWER PACK
H..... DU UP.DOW .. SW.
DOW .. LIMIT SW.
DOW .. LIMIT SW
M""" GfAI DOWNIOCI
CYIINDUS
DOORS OPEN, GEAR UNLOCKED AND RETRACTING
•
Figure ~48.
Hydralll1c System SchemaUc (Sheet 7 of 10)
~ 1:n /(5·138 blank)
•
MAIN GlAI U'lOCII
IlllASl C YlINDIi
CODE
'InSUll flOW
IfTUIN flOW
Sf ATIC PlUSUIE
STATIC .nUIN
SUPPLY
VINT
Figure 5-49 (sheets 1 thru 10) is effective for Serials
337-0756 thru 33701462 and F33700001 thru F33700055.
The secondary relief valve is deleted. This figure also includes the door close luck valve which is installed
in aircraft Serials 33701427 and F33700052 thru 33701462 and F33700055.
•
MAIN GIAI
0001 ACTUATOI
U' LIMIT SW
NOSI GIAi
0001 ACTUATOI
hOSE GEAR
POWER PACK
"ANDlt U'·DOWN SW.
DOWN LIMIT SW
DOWN LIMIT sw.
MAIN GIAI DOWNlOCI
CYlINDIiS
•
GEAR UP ANI) LOCKED, DOORS CLOSING
FIpI'e ~·19. Hydrallllc System Scbemauc (S....t 8 of 101
~ 1311/1&- 140
blutkl
•
MAIN GEAI UP lOCI
IIUASE cnlNDlI
CODE
•
Plnsuu flOW
IETUIN flOW
STATIC PlnSUIl
STATIC UTUIN
SUPPLY
YENT
Figure 5-49 (sheets 1 thru 10) is effective for Serials
337-0756 thru 33701462 and F33700001 thru F33700055.
The secondary relief valve is deleted. This figure also includes the door close lock valve which is installed
in aircraft Serials 33701427 and F33700052 thru 33701462 and F33700055 .
•
UP LIMIT sw.
NOS! GEAI
0001 ACTUATOI
\ lANDING
\
GUI
, CONUOl
LEVU
\
\
\
\
\
\
\
\
\
POWER P'ACK
HANDLE UP.DOWN SW.
DOWN LIMIT SW.
DOWN LIMIT SW
MAIN GEAI DOWN lOCI
CYLINDIiS
•
GEAR UP AND LOCKED, DOORS CLOSEtl , HANDLE RELEASE PRESSURE BUILDINC:; UP
FIpre >411. Hy<irallllc Systelll Scllematic (Sh. .t II 01 10)
~1.l/(5-142111Uk)
•
.. ... 'N GI .... UP lOCI[
IIlIASI CTIINon
CODE
,
!MIIGINCT
H ... NO PUM'
•
P.f55UU flOW
'ETU'N flOW
ST ... TlC PUSSUIE
ST ... TIC IETU.N
SUPPLY
YENT
Figure 5-49 (sheets 1 thru 10) is effective for Serials
337 -0756 thru 33701462 and F33700001 thru F33700055.
The secondary relief valvl! is deleted. This figure alSO includes the door close lock valve which is installed
in aircraft Serials 337014:!7 and F33700052 thru 33701462 and F33700055.
uP LIMIT sw
..---
NOSE GI ... .
000 .... CTU ... TO'
~.~
\\
L"'ND'NG
GU.
CONT'OL
\
\
·
·
·
uvu
..
!...\.
\
\
\
\
\
:
y
~
\
\
\
POWER PACK
H... NOLI UP·OOWN SW.
DOWN LIMit $W.
DOWN lI .. 1T SW.
DOWN II .. IT SW
.. ... 'N GU' DOWN LOCI[
CTIINOIIS
•
EMERGEN CY CONDITION. ENGINE PUMP FAILURE, NO ELECTRICAL POWER.
DOORS OPENED, GEAR UNLOCKED AND BEING EXTENDED BY HAND PUMP PRESSURE
Flpre 1;049. Hydl'8'll1c System SchemaUc (Sheet 10 of 10)
~ IU/(~ 144 bIaIIk)
•
PART 2
(BEGINNING WITH SERIALS 33701399 AND F33700046)
·'WARNINGa
Before working in landing gear wheel wells, PULL-OFF
hydraulic pump circuit breaker. Circuit breaker know
is located in circuit breaker panel. The hydro-electric
power pack system is designed to pressurize the landing
gear "DOOR CLOSE" system to 1500 psi at any time the
master switch is turned on. Injury might occur to someone working in wheel well area if mas ter switch is turned
on for any reason.
5-282. LANDING GEAR SySTEM.
•
•
5-283. DESCRIPTION. A hydraulically-operated retractable landing gear is employed on the aircraft.
The source of hydraulic power is obtained from a
hydra-electric power pack, installed in the lower
part of the control quadrant, immediately below the
instrument panel. The power pack consists of an
electric motor, driving a hydraulic pump With adequate valving to properly control the now to actuators
at the landing gear. The operation of the system is
controlled by an electrical landing gear switch located
to the left of the pedestal quadrant on the instrument
panel.
5-284. OPERATION. When the aircraft master
switch is closed, the hydraulic power pack is ready
to operate. When the gear-up position is selected
with the selector handle, the gear valve solenoid
connects the gear-up line to system pressure, and
the gear down line to return. At the same time, the
electric motor that powers the hydraulic pump is
turned on. The hydraulic pressure is passed through
a filter, and then is divided to the gear valve and door
valve. Before hydraulic pressure can reach the gear·
valve, a priority valve must open. The priority valve
can open only under two conditions
1. There can be no pressure in the door close
line, because door close pressure is applied to a
piston to hold the priority valve closed.
2. System pressure must build up to 750 psig
.before the valve can open. Pressure, therefore,
must go to the door open line. Pressure in the door
close line is prevented from returning by the door
close lock check valve, and the valve is opened bV a
piston that senses door open pressure. When the
pressure reaches 400 psig, the door close lock check
valve opens and the doors on the aircraft open. At
750 psig, the priority valve opens and the landing
gear begin to retract. As soon as the landing gear
is locked into the UP position, the landing gear up
limit switches sequence the door solenoid valve to the
door close position. When pressure in the door close
line reaches 1500 psig, the pressure switch shuts off
the motor, and the GEAR-UP cycle is complete. The
GEAR-DOWN cycle is similar to the GEAR-UP cycle,
except the gear solenoid is not energized during the
gear-down cycle. The system has been designed so
that anytime during system operation, the direction
of system operation may be reversed. Under these
conditions, the first operation of the system after the
selector handle is moved, is to completely open the
doors, and then move the gear into the newly- selected
position, after which, the doors will close again.
There is no danger of interference between the gear
and doors of the aircraft, since the gear does not
receive hydraulic pressure unless the doors are in
the fully-opened position.
5-285. MAIN GEAR SYSTEM.
5-286. DESCRIPTION. The main landing gears consist of two leaf type spring steel legs attached to rotating pivot castings mounted in the structure at an
angle of apprOximately 45°. The wheels, hydraulic
brakes and tires are attached to the lower end of the
legs by bolt attaChing an axle assembly. The main
landing gears are retracted hydraulically up and aft
into the belly of the fuselage with the Wheels extending past the rear firewall into the engine compartment area. Each gear has a separate linear-rotary
hydraulic actuator. The actuators consist of a linear
acting piston assembly, the shaft of which is also a
rack, a matching pinion, bearings, a rotary output
shaft to the actuator to the pivot casting. Downlock
linkages are used to secure the gears in the down
position. These pawls are secured in place by small
hydraulic linear actuators which also move the locking pawls out of the way before the gear retracts.
The gears are also locked in the up position by an
uplockpawl, each with a common hydraulic actuator.
5-145
5-287. TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
REMEDY
Fluid level low in reservoir.
Refill reservoir.
Motor pump failure.
Repair or replace pump.
Faulty check valve.
Repair or replace check valve.
No fluid in reservoir.
Refill reservoir.
Broken gear or door line.
Repair or replace hydraulic line.
Door or gear solenoid valve jammed
or sticking at mid-travel.
Repair or replace valve.
Master switch not on.
Turn master switch on.
Defective limit switch circuit.
Repair defective component in
limit switch circuit.
Circuit breaker tripped.
Reset circuit breaker.
Defective gear selection switch
or wiring circuit.
Repair or replace defective switch
or wiring.
Defective door solenoid.
Replace solenoid
Door solenoid valve stuck.
Remove power pack; repair or
replace solenoid valve.
GEAR OPERATES PROPERLY
BUT INDICATOR LIGHT DOES
NOT ILLUMINATE.
Lamp burned out.
Replace lamp.
PUMP OPERATES BUT
DOORS WILL NOT OPEN.
Door solenoid valve jammed or
stuck in door-closed position.
Repair or replace solenoid valve.
Repair any damage to doors and
linkage.
GEAR UNLOCKS BEFORE
. DOORS ARE FULL OPEN.
Priority valve setting too low.
Replace valve spring.
Priority valve leaking or stuck
open.
Repair or replace valve.
Gear solenoid valve jammed or
stuck in pOSition.
Repair or replace solenoid valve.
Priority valve setting too high
or stuck closed.
Repair or replace valve.
MOTOR PUMP WILL NOT
OPERATE GEAR BUT
EMERGENCY HAND PUMP
WILL OPERATE GEAR.
PUMP OR EMERGENCY
PUMP WILL NOT BUILD
PRESSURE IN SYSTEM.
DOORS WILL NOT CLOSE,
GEAR INDICATOR LIGHT
NOT ILLUMINATED.
DOORS WILL NOT CLOSE,
GEAR INDICATOR LIGHT
IS ILLUMINATED.
I
DOORS OPEN BUT GEAR
DOES NOT OPERATE.
5-146
•
•
•
•
5-287. TROUBLE SHOOTING (Cont).
PROBABLE CAUSE
TROUBLE
.
HAND PUMP OOES NOT BUILD
UP PRESSURE BUT ELECTRIC
PUMP OPERATES GEAR
PROPERLY.
LANDING GEAR OPERATION
EXTREMELY SLOW.
•
REMEDY
P.JWER PACK EXTERNAL
LEAKAGE.
POWER PACK LOSES
FLUID WITH NO
EVIDENCE OF LE.\KAGE.
Faulty hand pump plunger check
valve or O-ring.
Remove and inspect hand pump
plunger; replace parts as needed.
Faulty system inlet check valve
or hand pump inlet check valve.
Repair or replace check valves.
Reservoir fluid level low.
Refill reservoir.
Downlock rod adjustment
incorrect (Mainly LH rod).
Adjust rod end to lengthen
actuator one turn.
Pump failure or internal leakage.
Repair or replace pump.
Air leakage around pump suction
tube.
Seal suction tube.
Fluid leak in door or gear line.
Tighten or replace lines.
Defective piston seal in gear
or door r:ylinder.
Repair or replace defective parts.
Excessive internal power pack
leakage.
Remove and repair or replace
power pack.
(Static seals) All fi ttings.
Remove and replace O-rings and·
back-up rings as needed.
Gear solenoid.
Replace O-rings.
Door solenoid.
Replace 0- rings.
Transfer tubes between manifold
and power pack body.
Disassemble; replace O-rings.
Reservoir cover.
Remove power pack and remove
cover. Replace seals.
Air leak at pump shaft seal.
Repair or replace pump.
NOTE
Refer to the trouble shooting chart in paragraph 5-6
for additional procedures not covered in paragraph 5-287.
•
5-147
•
•
Actuator installation is described in
paragraph 5-15. Align index marks
shown in view "A-A" in accordance
with step "a" of that paragraph.
1.
2.
3.
4.
Bulkhead Forging
Actuator
Tunnel
Angle
View
A-A
Figure 5-50. Main Gear Actuator Removal
5-288. MAIN LANDING GEAR STRUT REMOVAL
AND INSTALLATION. Refer to figure 5-1 and paragraphs 5-7 thru 5-8 for removal and installation of
the main landing gear struts.
5-289. MAIN LANDING GEAR ACTUATOR.
5-290. DESCRIPTION. The main landing gear actuator consists of a linear acting piston assembly, the
shaft of which is also a rack, a matching pinion, bearings, a rotary output shaft to the actuator to the pivot
casting.
5-2!n. REMOVAL. (Refer to figure 5-50.)
a. Remove center seat.
b. Jack aircraft in accordance with instructions
outlined in Section 2.
c. Remove floor panel above tWIne I (3) area and
5-148
above actuator (2) to be removed.
d. Place landing gear control handle UP, with master switch OFF, and operate emergency hand pump
Wltil main gear downlock releases.
e. Disconnect and cap or plug hydraulic lines at
actuator (2).
f. Remove angle (4) on side of tunnel adjacent to
actuator.
g. Remove three bolts attaChing actuator mOWlting
flange to bulkhead forging (1).
h. Work l..ctuator free of forging and pivot assembly
and remove actuator.
NOTE
It may be necessary to disconnect lines in
the tWlnel area to facilitate removal of
actuator.
•
•
NOTE
9
Lubricate sector, piston rack gears and
all bearings with MIL-G-21164 lubricant
during assembly of actuator. .
11
•
~
22
17
1.
2.
3.
4.
5.
6.
7.
8.
Retainer
Washer
Cap
Bearing
Roller
CyHnder Body
O-Ring
Piston
9. O-Ring
Screw
Metering Pin
Nut
End Gland
O-Ring
O-Ring
Retainer
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
Shaft
Sector
Setscrew
Washer
Bearing
End Cap
Washer
Bolt
Figure 5-51. Main Landing Gear Actuator
•
5-292. DISASSEMBLY. (Refer to figure 5-51.)
a. Remove screw (10) and remove end gland (13)
and metering pin (11) by Unscrewing end gland from
cylinder body (6).
b. Remove cap end (22) and remove cap (3) by pulling from cylinder body (6). Using a small rod, push
piston (8) from cylinder body (6).
c. Remove cap (3) from shaft (17) by removing
reta.tn~ (1-) and washer (2).
d. Remove shaft (17), sector (18) and washer (20)
from cylinder body (6).
e. Remove setscrew (19) from sector (18) and remove sector from shaft (17).
NOTE
Unless defective, do not remove name plate,
bearings (4 and 21) or roller (5).
f. Remove and discard O-ring (7) from cylinder
body (6).
g. Remove retainer ring (16) and loosen locknut
(12) and remove metering pin (11) from end gland.
Remove and discard O-rings (14 and 15) from end
gland.
h. Remove and discard O-ring (9) from piston (8)
1. Thoroughly clean all parts in cleaning solvent
(Federal Specification P-S-661, or equivalent).
5-293. INSPECTION OF PARTS.
a. Inspect all threaded surfaces for cleanliness,
cracks and evidence of wear.
b. Inspect cap (3), washers (2 and 20), sector (18),
shaft (17), piston (8), roller (5) and cylinder body
(6) for cracks, Chips, scratches, scoring, wear or
surface irregularities which might affect their func5-149
tion or the overall operation of the actuator.
c. Inspect bearings (4 and 21) for freedom of motion, scores, scratches and Brinnel marks.
5-294. REPLACEMENT/REPAm OF PARTS.
a. Repair of small parts of the actuator is usually
impractical. Replace defective parts with serviceable parts. Minor scratches or score marks may be
removed by polishing with abrasive crocus cloth
(Federal Specification P-C-458), providing their removal does not affect the operation of the unit.
b. Install all new O-rings during assembly.
Install new O-rings (14) and (15) on end gland
i.
(13).
j. Install metering pin (11) in end gland. Install
retainer (16) on metering pin.
k. Install end gland and metering pin assembly in
cylinder and tighten until end of gland is flush with
end of cylinder body. Install and tighten screw (10).
1. Install end cap (22) at end of actuator assembly.
5-296. INSTALLATION.
a. With main gear pivot assembly rotating freely,
match pivot and actuator markings and slide actuator
into place.
5-295. ASSEMBLY. (Refer to figure 5-51.)
NOTE
Lubricate roller (5), bearings (4 and 21) and
sector (1S) with MIL-G-21164 high and low
temperature grease when installing parts in
cylinder body (6).
a. Press one bearing (4) into cylinder body (6) unInstall roller (5) and press other bearing
(4) in place to hold roller. Use care to prevent damage to bearings and roller.
b. Press bearing (21) into cap (3) until flush.
c. Assemble sector (IS) on shaft (17) with index
marks on shaft and sector aligned. Install setscrew
(19), aSSuring that setscrew enters shaft.
d. ~Osition washer (20) and cap (3) on shaft (17),
then lDstall washer (2) and retainer (1) on shaft
noting that end of shaft with fitting is positioned in
cap (3).
til flush.
NOTE
Use AN316-4R nut on bolt (24) to hold assembled cap and shaft to cylinder body.
e. Install cap and shaft assembly on cylinder body
using bolts and nuts.
'
NOTE
Lubricate all O-rings with Petrolatum or
MIL-H-5606 hydraulic fluid during assembly.
f. Install new O-ring (7) in cylinder body bore and
install new O-ring (9) on piston CS).
g. Rotate shaft (17) so that teeth on sector (IS) are
toward cylinder body.
h. Slide piston (S) into cylinder body, rotating shaft
(17) as necessary to engage first tooth on sector with
first tooth on piston rack. Use care to prevent damage to O-rings in cylinder body bore and on piston.
NOTE
Lubricate sector and piston rack gears with
MIL-G-21164 high and low temperature grease
sparingly during assembly. Over-greasing
might cause contamination of hydraulic cylinder area of cylinder body (6), past O-ring (7).
5-150
•
NOTE
Make sure index marks are aligned.
b. Install three bolts attaChing mounting flange to
bulkhead forging. Torque bolts to 50-70 pound-inches.
c. COMect hydraulic lines at actuator.
d. Install angle on side of tunnel, adjacent to actuator.
e. Check rigging of main landing gear as described
in applicable paragraph of this section.
f. Remove aircraft from jacks; install floor panels,
carpeting and center folding seat.
5-297. LINKAGE.
5-29S. DESCRIPTION. Each main land~ng gear actuator attaches directly to a shaft, which in turn
rotates its own main landing gear. The landing gear
strut and pivot shaft are fastened together by a saddle and rotate in bearings contained in inboard and
outboard main landing gear support forgings.
5-299. SADDLE AND PIVOT SHAFT REMOVAL.
(Refer to figure 5-52.)
a. Remove main landing gear strut as outlined in
paragraph 5-7.
b. Remove main landing gear actuator in accordance with instructions outlined in paragraph 5-291.
c. Remove three bolts attaChing saddle to pivot
shaft.
d. Pull pivot shaft inboard until clear of support
bearings. Allow saddle, thrust bearing, bearing
race and spacers to slide outboard as shaft is pulled
inboard. When shaft is clear of bearings, lift outboard end and slide saddle off shaft. Remove remaining bearing parts from shaft.
e. Pull pivot shaft inboard to remove.
•
5-300. SADDLE AND PIVOT SHAFT INSTALLATION.
(Refer to figure 5-52.)
a. Position pivot shaft through inboard forging and
slide spacers, thrust bearing race, thrust bearing
and saddle onto shaft.
NOTE
The ,pacers are used as required to remove
end play from the pivot shaft, without causing
it to bind.
•
•
•
Washers (3) and spacers (13) are used
as required to eliminate end play from
pivot shaft.
1.
2.
3.
4.
5.
6.
Outboard Support
Bearing
Washer
Saddle
Bolt
Landing Gear Strut
7.
8.
9.
10.
11.
Strut Clamp
Bolt
Bolt
Spacer
Thrust Bearing
12.
13.
14.
15.
16.
17.
Thrust Race
Spacer
Pivot Shaft
Inboard Support
Bearing
Barrel Nut
Figure 5-52. Main Landing Gear Linkage
b. Position outboard end of pivot shaft in bearing
in outboard support forging, check for end play of
shaft and adjust spacers as noted.
c. Install bolts securing saddle to pivot shaft.
d. Reinstall main landing gear actuator as outlined
in paragraph 5-296.
e. Reinstall main landing gear strut in accordance
with instructions outlined in paragraph 5-8.
•
To remove or install main gear downlock system
components, refer to paragraphs 5-35 thru 5-37 and
figure 5-8. To disassemble, inspect or assemble
the downlock actuator, refer to paragraphs 5 -28 thru
5-30 and figure 5-7.
5-303. MAIN LANDING GEAR DOOR SYSTEM. To
remove or install main wheel doors, refer to paragraphs 5-41 thru 5-51 and figure 5-9.
5-301. MAIN LANDING GEAR UPLOCK SYSTEM.
To remove or install main gear uplock system components, refer to paragraphs 5-27 thru 5-31 and
figure 5-6. To disassemble, inspect or assemble
the uplock actuator, refer to paragraphs 5-28 thru
5 -30 and figure 5 -7 .
5-304. MAIN WHEEL DOOR ACTUATOR. To remove, disassemble, inspect, assemble or install
main wheel door actuators, refer to paragraphs 5 -41
thru 5-51 and figure 5-10 .
5 -302. MAIN LANDING GEAR DOWNLOCK SYSTEM.
5-305. MAIN GEAR WHEELS AND TIRES. To remove, disassemble, inspect, repair, assemble or
5-151
install main landing gear wheels and tires, refer to
paragraphs 5-55 thru 5-59 and figure 5-11.
5-306. MAIN WHEEL AND AXLE REMOVAL AND
INSTALLATION. To remove or install main wheels
and axles, refer to paragraphs 5-60 and 5-61.
5-307. MAIN WHEEL ALIGNMENT. For information regarding main wheel alignment, refer to paragraph 5-62 and figure 5-12.
5-308. WHEEL BALANCING, For wheel balancing
information, refer to paragraph 5-63.
5-309. BRAKE SYSTEM. For information regarding
system trouble shooting, master cylinder removal,
disassembly, repair and installation; brake system
bleeding; wheel brake removal, inspection, repair,
assembly and installation and brake lining checking
and replacement, refer to paragraphs 5-65 thru 578 and figures 5-13 and 5-14.
5-310. PARKING BRAKE SYSTEM. For information
regarding the parking brake system, refer to paragraphs 5-79 thru 5-93 and figure 5-14.
5-311. NOSE GEAR SYSTEM. For a description,
operational description and nose gear trouble shooting, refer to paragraphs 5-94 thru 5-97 and figure
5-15.
5-312. NOSE GEAR ASSEMBLY. To remove the
nose gear assembly, refer to paragraph 5 -98 and
figure 5-15.
to paragraphs 5-118 and 5-120.
5-320. DISASSEMBLY, INSPECTION OF PARTS
AND ASSEMBLY. Refer to paragraphs 5-27 thru
5-30 and figure 5-7.
5-321. NOSE GEAR DOOR SYSTEM. For a description and operational information, refer to paragraphs
5-133 thru 5-134.
5-322. REMOVAL AND INSTALLATION OF NOSE
WHEEL DOORS. Refer to paragraphs 5-135 thru 5139 and figure 5-22 for procedures for removing
and installing nose gear doors.
5-323. NOSE WHEEL STEERING SYSTEM. Refer
to paragraphs 5-141 thru 5-145 and figure 5-23 for
description, removal, installation and rigging of
components of the nose Wheel steering system.
5-324. NOSE GEAR WHEEL. To remove, disassemble, inspect, repair, assemble and install nose wheels,
refer to paragraphs 5-149 thru 5-153 and figure 5-24.
5-325. LANDING GEAR HYDRAULIC POWER.
Refer to paragraphs 5-283 and 5-284 for a description and operational information.
5-326. HYDRAULIC TOOLS AND EQUIPMENT.
Refer to paragraphs 5-158 thru 5-168 for description and operational procedures while using hydraulic
system test eqUipment. Refer to figures 5-25 and
5-26 for Hydro Test Unit information.
5-327. HYDRAULIC POWER SYSTEM COMPONENTS.
5-313. NOSE GEAR STRUT. To disassemble and
assemble the nose gear strut assembly, refer to
paragraphs 5-100 thru 5-102 and figure 5-16.
5-314. SHIMMY DAMPENER. To remove, disassemble, assemble and install nose gear shimmy
dampeners, refer to paragraphs 5-106 thru 5-110
and figure 5-17.
5-315. TORQUE LINKS. For information regarding
removal, disassembly, assembly and installation of
nose gear torque links, refer to paragraphs 5-113
thru 5-114 and figure 5-18.
5-316. NOSE GEAR UPLOCK MECHANISM. To
remove or install nose gear up lock components, refer to paragraphs 5-118 thru 5-120 and figure 5-19.
5-317. NOSE GEAR DOWNLOCK MECHANISM. To
remove and install components of the nose gear downlock system, refer to paragraphs 5-124 thru 5-131
and figure 5-20.
5-318. NOSE GEAR ACTUATOR. To remove, disassemble, inspect, assemble and install nose gear
actuators, refer to paragraphs 5-125 thru 5-131
and figure 5-21.
5-319. REMOVAL AND INSTALLATION OF NOSE
GEAR UPLOCK AND RELEASE ACTUATOR. Refer
5-328. GENERAL DESCRIPTION. The hydraulic
power system includes equipment required to provide a flow of pressurized hydrauliC fluid to the retractable landing gear system. Main components of
the hydraulic power system include the power pack
and the emergency hand pump.
•
5-329. HYDRAULIC COMPONENT REPAIR. Since
emphasis here is on repair and not overhaul of the
basic components of the hydraulic system, it is
unlikely that the mechanic will go through all of the
procedures outlined. Instead, he will repair the
particular item which is causing the difficulty.
5-330. REPAIR VERSUS REPLACEMENT. Often,
the moderate trade-in price for a factory-rebuilt
component is less than the accumulated cost of labor,
parts and (often time-consuming) trial and error adjustment. Repair or replacement of a component
will depend on the time, equipment and skilled labor
that is locally available.
5-331. REPAIR PARTS AND EQUIPMENT. Repair
parts may be ordered from the applicable Parts
Catalog. Test equipment may be ordered from The
Special Tools and Support Equipment Catalog. Both
publications are available from the Cessna Service
Parts Center.
5-332. EQUIPMENT AND TOOLS.
5-152
•
•
•
TEST FITTING
•
Figure 5-53. Hydraulic Lines Installation
5-333. HAND TOOLS. The following hand tools are
necessary for repair work on the power pack and
other hydraulic components.
Snap Ring Pliers
Strap Wrench (for removing door solenoids and
various cylinder barrels of the hydraulic actuators)
Needle-nose Pliers
Duck-bill Pliers
Pin Punches
Box and Open -end Wrences
•
Locally-fabricated items, handy for power pack repair, are various 1/4-inch aluminum rods, ground
to a gradual taper, and hooks formed from brass
welding rod, to extricate small plungers from hydraulic ports. Hooks formed from brass welding
rod must not be over IllS-inch in length, so as not
to scratch or score the bore. Various sizes of Allen
wrences h
wrenches may be welded to "T" handles for use when
removing, installing or adjusting the various internal
wrenching plugs or valves.
5-334. COMPRESSED Am. The easiest method of
removing SODie hydraulic parts in inaccessible galleries of the power pack is a quick blast of compressed air from behind. Parts can be blown out in
seconds, which would otherwise take endless ''fishing" operations to extricate. An air hose and nozzle
are common-sense -tools.
5-335. POWER PACK.
5-33S. DESCRIPTION. The hydraulic power pack,
located in the pedestal, is a multi-purpose control
unit in the hydraulic system. It contains a hydraulic
reservoir and valves which control flow of pressurized fluid to the various actuators in the door and
landing gear system.
5-337. REMOVAL.
5-153
NOTE
As hydraulic lines are-disconnected or removed, plug or cap all openings to prevent
entry of foreign material in the lines or
fittings.
a. Remove front seats in accordance with instructions outlined in Section 3 and roll back carpet from
control pedestal.
b. Remove lower decorative cover by removing
screw s around cover.
c. Remove floorboard panel at aft side of pedestal.
d. Position gallon container under test fitting at
bracket on aft side of power pack.
e. Remove cap from test fitting and attach drain
hose.
f. Using hand pump, drain reservoir fluid into container.
g. Disconnect and cap or plug all hydraulic lines at
power pack.
h. Disconnect wiring at pressure switch.
i. Remove six screws attaching power pack support
to floorboard.
j. Work power pack aft out of pedestal.
5-338. DISASSEMBLY. (Refer to figure 5-54.)
a. Remove fittings from body assembly (41) and
place body assembly in vise.
b. Remove nut (30), washer (29) and packing (2) at
attaChing stud (38) at bottom of reservoir; remove
reservoir.
NOTE
If reservoir will not disengage from body
assembly, replace fittings removed from
body assembly and cap or plug all fittings
except vent fitting. Attach air hose at
vent fitting and apply pressure (not to exceed 15 psi - reservoir proof pressure);
remove reservoir. A strap clamp is not
recommended as clamp may damage reservoir.
c. Remove door manifold assembly and gear manifold assembly from body assembly of power pack.
d. Remove pressure switch and dipstick from body
assembly.
e. Remove large packing from bottom of body assembly.
f. Remove baffle (36), spacers (34) and washer (33).
g. Remove union (19), paCking, retainer ring (7)
and screen (31).
h. Remove motor and pump assembly (10) from
body assembly.
i. Remove packings and back-up ring from pump
assembly (10); remove coupling (11).
j. Remove return tubes (37) and packings from body
assembly.
k. Remove relief valve assembly from body assembly.
NOTE
Suction screen assembly (39) need not be
removed from body assembly to be cleaned.
However, if screen assembly is damaged,
it should be removed in accordance with
step "1" of this paragraph, observing the
following caution.
Use extreme caution in removing suction
screen assembly. Damage to screen assembly or clearance between screen assembly and body will cause slow landing gear
retraction.
1. Working through center hole in top of body assembly, and using a drift or punch made of soft material, tap out suction screen assembly (39).
m. Remove fittings from body assembly, if still installed, union (19), packing, retainer ring (7) and
screen (8) from body assembly.
n. Remove thermal relief valve and inlet check
valve from body assembly.
5-339. INSPECTION.
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661, or equivalent) and dry with
filtered air.
b. Inspect seating surfaces. They should have very
sharp edges. Seats may be lapped, if necessary, to
obtain sharp edges.
c. Inspect all threaded surfaces for serviceable
condition and cleanliness.
d. Inspect all parts foI' scratches, scores, chips,
cracks and indications of excessive wear.
•
5-340. ASSEMBLY. (Refer to figure 5-54.)
NOTE
Use all new packing and back-up rings for
reassembly. Before assembly, lubricate
all packings and back-up rings with MILH-5606 hydraulic fluid or Petrolatum.
Lubricate all threads with Petrolatum.
a. Assemble and install thermal relief valve and
inlet check valve in body assembly.
b. Install screen (8), retainer ring (7), packing and
union (19) in body assembly.
c. Install suction screen (39), if removed.
ICAUTION\
Use extreme caution when installing suction
screen assembly. Damage to screen assembly or clearance between screen assembly
and body will cause slow landing gear retraction.
d. Install relief valve assembly in body assembly.
e. Install packings and return tubes (37) in body
assembly.
f. Install packings and back-up ring on pump assembly (10); install coupling (11).
5-154
•
•
•
NOTE
Before assembly, lubricate all
packing with Petrolatum or
MIL-H-5606 Hydraulic fluid.
16
t:Y
U PRESSURE
,
•
•
1. Check Valve
1A. Thermal Relief Valve
2. Packing
3. Spacer
4. Self-Relieving Filter
5. Back-Up Ring
6. Retainer
7. Retainer Ring
8. Screen Assembly
9. Dipstick
10. Pump Assembly
11. Coupling
12. Spring
13. Piston
14. Nut
15. Fitting
16. Cap
17. Switch
18. Housing
19. Union
20. Adapter
21. Orifice
22. Seat
23. Poppet
24. Ball
25. Spring Guide
26, Housing
27. Setscrew
28. Nut
29. Reservoir Washer
30. Nut
31. Screen
32. Reservoir
33. Washer
34. Spacer
35. NamepWe
36. Baffle
37. Return Tube
38. Stud
39. Suction Screen Assembly
40. Plug
41. Body Assembly
SWITCH
17
41
21
22
2
3
RELIEF
VALVE
35
34
3332
31
,:~[/
~2
7"-30 29
Figure 5-54. Hydraulic Power Pack
5-155
1'f~UTIONI
To avoid damage to parts prior to assembly,
turn pump assembly (10) upside down and
lubricate shaft. Turn pump shaft by hand,
circulating oil.
.g. Install pump assembly (10) and motor on body
assembly.
h. Install screen (31), retainer ring (7), packing
and union (19).
i. Install washer (33), spacers (34) and baffle (36).
j. Install large packing on bottom of body assembly.
k. Install dipstick, pressure switch, door manifold assembly and gear manifold assembly on body
assembly.
1. Attach reservoir (32) to body assembly with
packing, washer (29) and nut (30).
reset, if required.
g. Lower landing gear, remove external power
source and remove aircraft from jacks.
5-344. GEAR MANIFOLD ASSEMBLY. (Refer to
figure 5-55.)
•
5-345. DISASSEMBLY.
NOTE
Mter the manifold has been removed from
the body assembly of the power pack, seat
(2) will remain in body assembly. Ball (4)
will fall free.
a. Remove seat (2) from body assembly of power
pack; remove two packings from seat.
NOTE
5-341. INSTALLATION.
a. Work power pack into pOSition and install six
screws attaching power pack support to floorboard.
b. Connect all hydraulic lines to power pack fittings.
Make sure fittings are properly installed, with jam
nuts tight, after lines are tightened.
c. Attach pressure switch wiring.
d. Fill reservoir through dipstick hole with clean
hydraulic fluid.
e. Jack aircraft in accordance with instructions
outlined in Section 2. Using Hydro Test unit, operate
landing gear through several cycles to bleed system.
Check for proper operation and any signs of hydraulic
fluid leakage.
f. Install floorboard panel at aft side of pedestal,
lower decorative pedestal cover, replace carpet and
install front seats.
5 -342. PRESSURE SWITCH. When installed in the
aircraft, the pressure switch is mounted on the lefthand aft side of the power pack installed on the floorboard inside the control pedestal. This switch opens
the electrical circuit to the pump solenoid when the
main gear fully retracts and pressure in the system
increases to apprOximately 1500 psi. The pressure
switch will continue to hold the electrical circuit open
until pressure in the system drops to approximately
1100 psi at Which time the pump will again operate to
build up pressure to apprOximately 1500 psi as long
as the gear control is in the UP position. With the
gear control handle in the DOWN position, the pressure switch has no effect on the system.
5 -343. PRESSURE SWITCH ADJUSTMENT. (Refer
to figure 5-54.)
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. Attach external power source and install pressure gage in landing gear UP line. (Refer to figure
5-2. )
c. Loosen jam nut on switch and back off switch
housing (18).
d. Retract landing gear and apply pressure to
1500±50 PSI.
e. Tichten switch housing until snap action switch
actuates, then tighten jam nut against housing.
f. Recheck operating pOint of 1500±50 PSI, and
5-156
Difficulty may be encountered in remOving
poppet (5) and spring (6). It may be necessary to apply air pressure at port "A"
(View A-A) to force spring and poppet from
port "B".
b. Remove back-up rings and packing from grooves
in poppet.
c. Remove packing from bottom of manifold assembly; remove spring (6).
d. Cut safety wire and remove solenoid (9).
e. Using a hook, formed from brass welding rod,
and inserted into oil hole in selector valve (8), withdraw selector valve from manifold.
[f~UTIONI
•
Be sure that end of hook is not over 1/16-inch
long. Use with care to prevent scratching
bore in manifold. Removal of selector valve
will be difficult due to friction caused by
packings.
f.
Remove packings from selector valve.
5-346. INSPECTION.
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661, or equivalent) and dry with
filtered air.
b. Inspect seating surfaces. They should have very
sharp edges. Seats may be lapped, if ne~essary,
with No. 1200 lapping compound.
c. Inspect all threaded surfaces for serviceable
condition and cleanliness.
d. Inspect all parts for scratches, scores, chips,
cracks and indications of excessive wear.
5-347. ASSEMBLY. (Refer to figure 5-55. )
NOTE
Use all new packing and back-up rings for
reassembly. Before assembly, lubricate
all packings and back-up rings with MILH-5606 hydraulic fluid or Petrolatum.
Lubricate all threads with Petrolatum.
•
•
SAFETY
WlRE
_ ........ SHOULDER (REF)
•
View
A-A
14
NOTE
Before assem~ly. lubricate all
packing with Petrolatum or
MIL-H-5606 hydraulic fluid.
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
•
Packing
Seat
Back-Up Ring
Ball
Poppet
Spring
Gear Solenoid Assembly
Selector Valve
Solenoid
Retainer Ring
End Gland
Piston
Door Manifold Assembly
Plug
Check Valve
Transfer Tube
DOOR M.'NIFOLD
ASSEMBLY
GEAR MANIFOLD
ASSEMBLY
Figure 5-55. Hydraullc Power Pack Manifold Assemblies
5-157
a. Install packings on selector valve (8).
b. Install packings in bottom of manifold.
c. Install spring and selector valve (8) in manifold.
d. Install packing on solenoid (9), install solenoid
on manifold and safety wire as shown in view A-A.
e. Install spring in bottom of manifold.
f. Install packing and back-up rings on poppet (5).
g. Install poppet in manifold.
IC~UTIO~I
Use extreme caution when installing poppet
(5). Shoulder, referenced in view A-A will
cut packings on poppet.
h. Install packings on seat (2); install ball (4) and
seat (2) in manifold.
5-348. DOOR MANIFOLD ASSEMBLY. (Refer to
figure 5-55.)
5-349. DISASSEMBLY.
a. As door manifold assembly is removed from
body of power pack, transfer valve (16) will fall free.
b. Remove packings from transfer tube.
c. Remove packings from bottom of manifold, and
remove check valve (15).
d. Remove spring (6).
e. Cut safety wire and remove solenoid (9); remove
packing from solenoid.
f. Using a hook, formed from brass welding rod,
and inserted into oil hole in selector valve (8), withdraw selector valve from manifold.
Be sure that end of hook is not over 1/16-inch
long. Use with care to prevent scratching
bore in manifold. Removal of selector valve
will.be difficult due to friction caused by packings.
g. Remove packings from selector valve.
h. Remove retainer ring (10).
i. Remove end gland (11).
j. Remove piston (12).
k. Remove packings and back-up rings from end
gland and piston.
NOTE
Use all new packing and back-up rings for
reassembly. Before assembly, lubricate
all packings and back-up rings with MILH- 5606 hydraulic fluid or Petrolatum.
Lubricate all threads with Petrolatum.
•
a. Install new packings and back-up rings on gland
(11), piston (12), selector valve (8) and transfer tube
(16).
b. Install packings and check valve (15) in bottom
of manifold.
c. Install spring (6) and selector valve (8) in manifold.
d. Install packing on solenoid (9).
e. Install solenoid on manifold and safety wire as
shown in view A-A.
f. Install piston (12) and end gland (11) in manifold.
g. Install retainer ring (10).
h. Prior to installing manifold on body of power
pack, install transfer tube (16) in body of power pack.
5-352. EMERGENCY HAND PUMP. (Refer to figure
5-56. )
5-353. DESCRIPTION. The emergency hand pump is
mounted on a support beneath the floorboard just in
front of the front seats, near the center of the floorboard. The handle extends into the cabin and is enclosed by a hinged cover. The pump supplies a flow
of pressurized hydrauliC fluid to open the doors and
extend the landing gear if hydraulic pressure should
fail. The hand pump receives a reserve supply of
fluid from the power pack reservoir and pumps the
fluid through passages and lines to the door control
valve and gear priority valve in the manifold and
through the remainder of the system.
•
5-354. REMOVAL.
a. Loosen carpeting around hand pump and remove
cover and pan.
b. Wedge cloth under hydrauliC fittings to absorb
fluid, then disconnect hydraulic lines at hand pump
and plug openings.
c. Remove two mounting bolts and work hand pump
out of floorboard opening.
5-355. DISASSEMBLY. (Refer to figure 5-56.)
5-350. INSPECTION.
a. Wash all parts in cleaning solvent (Federal
Specification P-S-661, or equivalent) and dry with
filtered air.
b. Inspect seating surfaces. They Should have very
sharp edges. Seats may be lapped, if necessary, to
obtain sharp edges.
c. Inspect all threaded surfaces for serviceable
condition and cleanliness.
d. Inspect all parts for scratches, scores, chips,
cracks and indications of excessive wear.
5-351. ASSEMBLY. (Refer to figure 5-55.)
5-158
NOTE
After hand pump has been removed from
aircraft, and ports are capped or plugged,
spray with cleaning solvent (Federal Specification P-S-661, or equivalent) to remove
all accumulated dust or dirt. Dry with filtered compressed air. To disassemble the
unit, proceed as follows:
a. Remove handle (3) by removing pins (19) and
washers after removing cotter pins (4).
b. Place pump in vise with fitting (8) at top.
c. Unscrew fitting (8) and remove, along with
washer (9).
•
•
NOTE
7
Before assembly, lubricate
all O-Rings and Back-Up
Rings with Petrolatum or
MIL-H-5606 hydraulic
fluid.
11
1&
2
12
NOTE
•
1.
2.
3.
4.
5.
6.
Roll Pin
7.
8.
9.
10.
Knob
Stop
Fitting
Handle
Washer
Cotter Pin
O-Ring
11. Check Valve
Fork
Spring Lock 12. Back-Up Ring
13. Setscrew
14.
15.
16.
17.
18.
19.
Spacer
KEP-O-SEAL Valve
Piston
Pump Body
Union and Gasket
Pin
During assembly, prime parts with
Primer T. Fill first three threads
of fitting (8) with Loctite Hydraulic
Sealant. Install fitting in pump body
(17), and allow parts to set for one
hour at 72°F. Pump should be held
vertically, with fitting (8) at top,-during setting-up of sealant.
Figure 5-56. Emergency Hand Pump
NOTE
Use caution when removing fitting (8) as check
valve (11) will fall free.
d. Remove pump from vise and push piston (16) out
of pump body (17). Push from handle end of piston.
A slight drag will be experienced until piston clears
back-up ring and packing inside pump body.
e. Remove setscrew (13) from piston (16) and remove spacer (14), O-ring (10) and KEP-O-SEAL
valve (15).
f. Remove union and gasket (18).
g. Remove and discard back-up ring and O-ring
from inside pump body (17) and fitting (8).
•
5-356. INSPECTION.
a. Inspect seating surfaces. They should have very
sharp edges. Seats may be lapped, if necessary, to
obtain sharp edges.
b. Inspect piston (16) for scores, burrs or scratches
Which might cut O-rings. This is a major cause of
external leakage. The piston may be polished with
extremely fine emery paper. Never use paper coarser than No. 600 to remove scratches or burrs. If
defects do not polish out, replace piston.
c. The threads on fitting (8) and in pump body (17)
are coated with Loctite Sealant. This sealant shoult:
be cleaned from the threads with a wire brush. After
threads are cleaned out, inspect for damage.
5-357. ASSEMBLY.
(Refer to figure 5-56. )
NOTE
Lubricate O-rings and back-up rings with
Petrolatum or MIL-H-5606 hydraulic
fluid before assembly.
a. USing all new O-rings and back-up rings, install
back-up rings and O-rings inside pump body (17).
NOTE
Assure that check valve (11) is inserted correctly in order to seat inside fitting (8).
b. Insert KEP-O-SEAL valve (15), O-ring and spacer (14) into piston (16). Install setscrew (13). Install back-up rings and O-ring in grooves on piston
(16).
5-159
•
SHIM AS REQUIRED FOR CONTACT
BETWEEN ADJUSTING SUPPORT
AND LANDING GEAR STRUT
FULL CONTACT (OR
AT LEAST CONTACT
FORE-AND-AFT
ADJUSTMENT
NEAREACHEND)----~--~~~_+------~--~
LOCATE AND DRILL WEDGE
FOR SLIGHT DRAG OF STRUT
TO • 010" MAX. CLEARANCE
SLIGHT DRAG OF STRUT
TO • 005" MAX. CLEARANCE
Figure 5-57. Rigging Adjustable Support
c. Line up piston in pump body (17). Carefully insert piston into pump body. Use extreme caution to
avoid cutting packing inside pump body.
NOTE
A "pumping" back and forth motion must be
employed to get piston pOsitioned inside
pump body.
d. Install washer (9).
e. Fill first three threads of fitting (8) with Loctite
Hydraulic Sealant. Install fitting in pump body (17),
and allow parts to set for one hour at 72°F. Pump
should be held vertically, with fitting (8) at top during setting-up of sealant.
f. Install union and gasket (~8).
g. Line up holes in piston (16) and pump body (17)
with holes in fork (5). Iristall pins (19), washers and
cotter pins (4).
5-358. INSTALLATION. (Refer to figure 5-56.)
a. Position pump between brackets in floorboard
opening.
b. Install two mounting bolts.
c. Attach hydraulic lines at hand pump.
d. Bleed all air from hand pump and hand pump
lines by loosening pressure cap, located at aft of
power pack, and pumping the hand pump until all
air is expelled: retorque test fitting's pressure cap.
e. Install cover and pan; reinstall carpeting.
5-359. RIGGING MAIN LANDING GEAR.
5-360. RIGGING ADJUSTING SUPPORT. (Refer to
figure 5-57.) The adjusting support is bolted to the
outboard forging and forms the down stop for the main
gear.
a. Jack aircraft as outlined in Section 2.
NOTE
Spring strut must be installed and secured
before rigging the adjusting support.
b. Check for contact between flat surface of strut
and lower surface of adjusting support. Minor gapping may exist as long as contact is made near each
end of support. Shim as required between outboard
forging and adjusting support to obtain required contact. Shims are available from the Cessna Service
Parts Center. The following shims are available for
installation at the forward end of support.
1541041-6 .
-7
-8
-9
-10
•
.012"
.020"
· 032"
.006"
The following shims are available for installation at
the aft end of support.
1541041-1
-2 .
-3
-4
-5 .
5-160
•
•
.012"
· 020"
· 032"
· 006"
•
•
MOVE UP AND DOWN TO ESTABLISH
MINIMUM CLEARANCE BETWEEN
CLEVIS AND MOUNTING BRACKET
ACTUATOR FULLY
RETRACTED BY
HYDRAULIC PRESSURE
CLEVIS
ACTUATOR
•
MOUNTING
BRACKET
~SHIM
AS REQmRED TO ELIMINATE
CLEARANCE BETWEEN CLEVIS
AND MOUNTING BRACKET
Figure 5-58. Main Gear Downlock and Actuator Alignment and Actuator Shimming
*Sheet of .025" laminated with ten. 002" additional
removable laminations.
c. Check that aft edge of strut contacts adjusting
support (.005" maximum clearance) as shown in
figure 5- 30, when gear is down. To shift adjusting
support fore and aft, first loosen three bolts securing
support (elongated holes are provided in the support),
then adjust the two jam nuts as required and retighten
the three mounting bolts.
d. Check that forward edge of strut contacts wedge
(.010" maximum clearance) as shown in figure 5-30,
when gear is down. A slotted hole in the adjusting
support will allow moving the wedge to obtain the required clearance. IT necessary, remove attaching
hardware and install a new wedge.
NOTE
A slight drag is permissible as gear reaches
the full down pOSition.
•
The wedges listed in the following chart are available
from the Cessna Service Parts Center. The dimensions listed are measured at the thickest part of the
wedge.
1541029-1
-2
-3
-4
· 250"
.300"
· 330"
.360"
5-:~61. RIGGING DOWNLOCK MECHANISM. (Refer
to figures 5-58 thru 5-62.)
a. Disconnect actuator clevis from fuselage bracket
and use hand pump to pressurize the actuator in its
fully-retracted position. With the actuator piston
bottomed out, position the downlock so a straight line
is formed through actuator pivot point, piston rod
pivot point and downlock pivot point as shown in figure
5-31. Measure the clearance between actuator clevis
and fuselage bracket and install shims as required to
eliminate this clearance. Connect clevis to bracket
and secure. The shims listed in the following chart
are available from the Cessna Service Parts Center.
1512359-1
-2
.125"
· 032"
b. Check that downlock pin reaches the over center
position shown (.03" to .10"). Adjust upper stop
bolt as required to obtain this position. (Refer to
figure 5-32.)
c. Check that downlorJc pin reaches retracted position shown (.18" to . 22"). Adjust lower stop bolt as
required to obtain this position. (Refer to figure 5-59. )
5-161
r.
03" TO .1~' (OVERCENTER)
•
DOWNLOCK PIN
SPRlNG----
- - - • 18" TO .22"
•
SPRlNG------~~~~,d1
LOWER STOP BOLT --~
Figure 5-59. Rigging of Main Gear Downlo:::k
NOTE
A downlock rigging tool, PiN SE772-1, shown
in figure 5-33, is available from the Cessna
Service Parts Center.
d. Check over-all length of downlock pin as shown
(snugly against strut to .005" maximum clearance),
with hydraulic pressure on gear. Downlock pin assembly must be removed to change over-all length.
(Refer to paragraph 5-59.)
e. Check that overcenter release bolt in upper end
of downlock extends below support as shown (.070" to
. 100") when the actuator piston is bottomed out
retracted, with hydraulic pressure applied. (Refer
to figure 5-61.)
f. Release hydraulic pressure and check that overcenter stop bolt in bulkhead is adjusted so that overcenter release bolt in upper end of downlock extends
below adjusting support as shown (. 06" more than
dimension "A") when actuator is held in overcenter
position against bulkhead stop bolt. (Refer to figure
5-61. )
g. Check that button in overcenter arm is screwed
completely in (shortened) as shown, and jam nut is
tight. Check that overcenter arm retracts smoothly
5-162
when engaging strut and that arm is clear of roll pin
installed in downlock when gear is down and locked.
(Refer to figure 5-62.)
h. Check action of cam on main gear downlock
switch bracket as follows:
1. Place main gear in "trail" position.
2. Manually push downlocks into normally locked
pOSition (aft).
3. Holding approximately 20 pounds of force
against each wheel, extend gear to the down and
locked position. Cams on the switch brackets should
push downlocks out of the way, allowing gear to move
smoothly into the down and locked pOSition.
4. Repeat test at least five times .
5-362. RIGGING UPLOCK MECHANISM. (Refer to
figure 5-6.)
a. Jack aircraft in accordance with procedures outlined in Section 2.
b. Loosen bolts attaChing hangers (6) to supports (9)
to allow inboard and outboard adjustment.
c. With Hydro Test connected, open test stand bypass valve to reduce hydraulic pressure to approximately 1000 pSi. With gear up a'ld pressurized, check
position of gear stops (8).
d. Outboard edge of gear spring strut (17) should
•
•
CD
MANUALLY HOLD DOWN LOCK PIN
AGAINST UPPER STOP BOLT
POSITION THIS HOLE
ON THIS BOLT
UPPER STOP
BOLT
~ DOWNLOCK
PIN
RIGGING TOOL
.7
BOLT~
LOWER STOP
® AFT
FORWARD EDGE OF TONGUE CONTACTS
EDGE OF DOWNLOCK PIN _ _ _
--1
•
POSITION LOWER FLANGE OF TOOL IN
FULL CONTACT WITH FLAT SURFACE
OF DOWNLOCK
OVERCENTER POSITION
OF DOWNLOCK PIN
With downlock pin depressed (1), lower bolt in
lower hole (3), lower flange flat against downlock (4), and forward edge of tongue contactin~
aft edge of pin (5), upper bolt should fall within overcenter hole (2). Elongation of overcenter
hole represents tolerance permissible; adjust
upper stop bolt as required.
NOTE
Jack the aircraft, retract the landing gear, and release hydraulic pressure, leaving the landing gear
doors open. Pull downlock assemblies aft for access.
•
The downlock pin rigging tool, Part No. SE772-1, is available from the Cessna Service Parts Center.
The tool is made in two halves - the left half is shown in use for the left downlock pin; the right half is
used in the sar.le manner for the right downlock pin .
Figure 5-60.
Using Main Gear Downlock Pin Rigging Tool (Sheet 1 of
2)
5-163
•
G)OOWNLOCK PIN RELEASED AGAINST
LOWER STOP BOLT - - - , .
POSITION THIS HOLE
ON THIS BOLT
OOWNLOCK
DOWNLOCK PIN
RETRACTED HOLE
IN RIGGING TOOL
~
7
LOWERSTOV
BOLT
CD AFT
FORWARD EDGE OF TONGUE CONTACTS
EDGE OF DOWNLOCK PIN - - - - - '
DOWNLOCK PIN
RIGGING TOOL
POSITION LOWER FLANGE OF TOOL IN
FULL CONTACT WITH FLAT SURFACE
OF OOWNLOCK
•
RETRACTED POSITION
OF DOWNLOCK PIN
With downlock pin not depressed (I), lower
bolt in lower hole (3), lower flange flat against
downlock (4), and forward edge of tongue contacting aft edge of pin (5), upper bolt should
fall within retracted hole (2). Elongation of
retracted hole represents tolerance permissible; adjust lower stop bolt as required.
Figure 5-60. Using Main Gear Downlock Pin Rigging Tool (Sheet 2 of 2)
15-164
•
•
/
ACTUATOR FULLY
RETRACTED BY
HYDRAULIC PRESSURE
STRAIGHT LINE THRU CENTER LINES OF
PIVOT POINTS (AUTOMATICALLY FORMED
WHILE ACTUATOR IS FULLY RETRACTED
BY HYDRAUUC PRESSURE)
Z
'~
, x>
/«>,
,
. ~~.
::
I
. .
.I\.COP'"'
... ~. :t:.. :
.... --.- ..,: ____ ji
ADJUSTING
SUPPORT
«----ACTUATOR
OVERCENTER STOP BOLT (Adjust
so dimension "B" is . 06 inch more
than dimension "A".)
•
ACTUATOR
•
Figure 5-61. Overcenter Adjustments of Main Gear Retracted Downlock
5-165
•
~-------LOCKNUT
BUTTON
Figure 5-62. Checking Main Gear Overcenter Arm Button
contact stop (8) and slanted portion of stop should be
;»arallel to spring strut, maintaining 20 percent contact with spring strut.
e. stop (8) is adjusted to match angle of gear spring
strut by the addition of shims (7) (P/N 1541051-2) as
required between hangers (6) and supports (9).
f. Adjust push-pull rod ends (12) as required to
cause hooks (11) to release gear spring struts simultaneously when operated hydraulically.
NOTE
In addition to releasing gear struts simultaneously, linkage must be adjusted so no part
of linkage (including up indicator switch) contacts any part of the aircraft structure. Actuator piston must bottom out retracted before
hydraulic fluid can be routed through the actuator to lower the main gear.
5-363. RIGGING UP INDICATOR SWITCHES. (Refer
to figure 5-6.) Main gear up indicator switches (16)
are mounted on brackets (15) attached to the uplock
hooks (11). After jacking aircraft in accordance
with procedures outlines in Section 2, retract landing
gear until uplock hooks are fully engaged. Adjust
switches so they are actuated with a minimum of 1/8
inch travel of the switch plunger remaining. Switch
case must not contact any part of structure.
(WARNING'
Before working in landing gear wheel wellS,
PULL-OFF hydraulic pump circuit breaker.
Circuit breaker knob is on circuit breaker
panel. The hydro-electric power pack system is designed to pressurize the landing
gear "DOOR CLOSE" system to 1500 psi at
any tim4=! the master switch is turned on.
Injury might occur to someone working in
wheel well area.
5-364. RIGGING DOWN INDICATOR SWITCHES.
(Refer to figure 5-8.) Main gear down indicator
switches (6) are mounted on brackets (5) attached to
downlocks (9). With landing gear down and locked,
adjust switches so they are positively actuated, but
5-166
leaf-type switch actuator does not contact switch case.
5-365. RIGGING DOORS. (Refer to figures 5-9 and
5-22.) Jack aircraft in accordance with procedures
outlined in Section 2. Adjust push-pull rod ends and
actuator rod ends as required to cause doors to close
snugly. Doors must not close so tightly that internal
locks in actuating cylinders are not reached. When
installing new doors, some trimming and hand-forming at edges may be necessary to achieve a good fit
and permit actuators to lock. Doors must clear the
gear at least 1/2 inch during retraction.
5-3£i6. ADJUSTMENT OF SNUBBER VALVES.
(Refer to figure 5-51.) A main gear snubber valve,
which restricts fluid near the end of the gear-up
cycle, is provided at the aft end of each main gear
actuator. These valves are hollow, contoured
metering pins which form the hydraulic fittings at
the aft end of the actuators. The purpose of the
snubber valves is to slow down action near the end
of the gear-up cycle to cause smoother locking
action. Position of the snubber (screw in or out)
shall be fixed such that:
a. Snubbing occurs during the final 1/2 to one
second of gear-up travel.
b. Both main gears lock in up pOsition simultaneously.
c. The gear struts do not strike uplock stops with
sufficient force to jar the structure or jar the aircraft
or cause objectional noise.
•
5-367. RIGGING OF NOSE GEAR. Refer to paragraph 5-272 for nose gear rigging procedures.
5-368. HYDRAULIC AND ELECTRICAL SYSTEM
SCHEMATICS. (Refer to figure 5-63.} The following seven pages contain coded schematic diagrams
of the aircraft hydraulic system. A conplete geardown cycle is illustrated, from selecting the gear
down position to the condition where the gear is down
and locked and the master switch is OFF. Incorporated into the hydraulic system scbematic is the electrical wiring diagram which shows switch positions,
lights, solenoids and other components of the system,
and their condition during the gear-down cycle.
•
•
SWITCHES
NOSE
I
COM
LEFT
RIGHT
'N~'I
)"6
"""
L-o
S-GDI-CO-'NO'
j' ~M
8
e
Cd'M
I S-GD9J
DOWNlOCK SWITCHES
L-------S-GDI§---------r.~--------------JT~------~
HITE
X
•
CODE
.... ml
ICIIAial
Ill. PlI'
•
•
GEAROOWN
GEAR UP
*
*
DOORS OPEN
DOORS CLOSED
GEAR DOWN SELECTED - DOORS UNLOCKING
•
_
III
III SUCTION
•
STAnc
•
CD
UGHT ON
STAl1C
1'{(41 PRESSURE
RETURN
PRESSURE
FLOW
Figure 5-63. Hydraulic and Electrical System Schematic (Sheet 1 of 7)
5-167
•
SWITCHES
RIGHT
GDI-lD-'NO'
j~M
R
LEFT
Cd'M
9
Is-oD,J
DOWNLOCK SWITCHES
Hill
X
U__.r----rtllill llCI
lAm
•
CODE
l1li IlIl
aC"aTlI
lall
•
•
GEARDOWN
GEAR UP
1111 IEII ..llCI ICIIUI.
"I'
**
DOORS OPEN
•
STAnc
•
PRISSURE
PRESSURE
FLOW
•
"SUC'l10N
•
STAnc
•
RETURN
@ LIGHT ON
DOORS CLOSED
GEAR DOWN SELECTED - DOORS OPENING
Figure 5-63. Hydraulic and Electrical System Schematic (Sheet 2 of 7)
5-168
•
•
SWITCHES
RIGHT
LEft
DOWNlOCK SWITCHES
.--------;:::;..11GI1GCI 'IlIE
•
CODE
.111
•
GEAR DOWN
•
GEAR UP
IAII CEIl lPlacl IClI11II
•
PI.,
**
OOORSOPEN
OOORS CLOSED
_
_
STAl1C
PRES';IURE
III
PRESSURE
111
SUcnON
•
STAl1C
•
RETURN
®
LIGHT ON
FLOW
DOORS OPEN - GEAR UNLOCKING
Figure 5-63. Hydraulic and Electrical System Schematic (Sheet 3 of 7)
5-169
•
S-GE3--::::;:;::o8\.I~)
S-
un
RIGHT
I-ID-'NO'
,2M
j~M
R e
Is_oD,J
DOWNLOCk SWITCHES
HITE
X
•
CODE
I ... lUI
aC1I1I1.
IIII "I'
1111 CEIl "lOCI ImalOI
•
GEAR DOWN
•
GEAR UP
**
DOORS OPEN
DOORS CLOSED
STATiC
PR~URE
•
PRESSURE
•
•
SUCTION
•
STAnc
•
RETURN
®
UGHT ON
GE.AR EXTENDING
Figure 5-63. Hydraulic and Electrical System Schematic (Sheet 4 of 7)
5-170
FLOW
•
•
UP
DOWNlOCK SWITCHES
~------S-GD15--------~------------
5-0024
__-4I~______~
•
•• '1 m.
SPI'I' DOli
IIU.U I(u£f IIlIE
IYIIIIUC
rllal:lall::=--~
'lin 'ICI
••'Inn
III
PRESSURE
FLOW
II SUCTION
•
STAnc
•
0
UGHT ON
mOUDI
II.....,
••,1 CUI
•
moci manoa
•
GEAR DOWN
•
GEAR UP
*
OOORSOPEN
*OOORS CLOOED
•
STAnc
PRESSURE
RETURN
GEAR DOWN - DOOR CLOSING
Figure 5-63. Hydraulic and Electrical System Schematic (Sheet 5 of 7)
5-1n
•
OS-GD6~U----'
UPLOCK SWITCHES
LEFT
RIGHT
DOWN LOCK SWITCHES
•
1111 Cli.
om ....
ICTlIII.S
CODE
**
1111 'II'
•
GEAR DOWN
•
GEAR UP
DOORS OPEN
DOORS CLQlED
•
STAl1C
•
PRESSURE
PRESSURE
FLOW
•
SUC110N
•
STAl1C
•
RETURN
~ LIGHT ON
GEAR DOWN - DOORS CLOSED & MOTOR TURNING OFF
Figure 5-63. Hydraulic and Electrical System Schematic (Sheet 6 of 7)
5-172
•
•
NOSE
RIGHT
LEFT
5- GDI-a::r-'N
DOWN
•
IIII CUI 11m .111 mUlTIlS
CODE
.1.1 "I'
IIII CEil IPLICK leTllUI
•
•
GEAR DOWN
•
GEAR UP
**
DOORS OPEN
DOORS CLOSED
1m
PRESSURE
II SUCTION
•
STAl1C
•
0
LIGHT ON
11m
_
STAl1C
PRESSURE
RETURN
FLOW
SYSTEM COMPLETE (AIRCRAFT MASTER SWITCH OFF)
Figure 5-63. Hydraulic and Electrical System Schematic (Sheet 7 of 7)
5-173/(5-174 blank)
•
SECTION 6
AILERON CONTROL SYSTEM
TABLE OF CONTENTS
Page
AILERON CONTROL SYSTEM
. 6-1
Description . .
. 6-1
Trouble Shooting . . . .
. 6-1
Control Column . . • .
. 6-2
Description . . . .
. 6-2
Removal and Installation
. 6-2
Control Wheel Tube - Rear Section. 6-2
Control Wheel Tube - Forward
Section . . . . . .
.6-3
Repair . . . . . . . . .
.6-3
Bearing Roller Adjustment
.6-3
.6-3
Bellcranks . . . . . . . . .
Removal and Installation .
.6-3
Repair . . . . . . . . .
Ailerons . . . . . . . . .
Removal and Installation
Repair . . . . . . . .
Aileron Trim Tabs . . . . .
Removal and Installation
Adjustment . . . . . .
Cables and Pulleys . . . . .
Removal and Installation
Direct Cable - Inboard
Direct Cable - Outboard.
Carry-Thru Cable
Rigging
.6-3
. 6-3
. 6-3
.6-8
.6-8
.6-8
.6-8
.6-8
.6-8
.6-8
.6-9
.6.-9
.6-9
comprised of push-pull tubes, bellcranks, cables,
pulleys, sprockets and components forward of the
instrument panel, all of which, link the control
wheels to the ailerons.
6-1. AILERON CONTROL SYSTEM. (Refer to figure 6-1.)
6-2. DESCRIPTION. The aileron control system is
6-3. TROUBLE SHOOTING.
NOTE
•
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 6-20.
TROUBLE
LOST MOTION IN CONTROL
WHEEL.
RESISTANCE TO CONTROL
WHEEL MOVEMENT.
•
PROBABLE CAUSE
REMEDY
Loose control cables.
Check cable tension. Adjust
cables to proper tension.
Broken pulley or bracket,
cable off pulley or worn
rod end bearings.
Check visually. Replace worn or
broken parts, install cables
correctly.
Deformed bellcrank or
pulley bracket.
Check visually. Replace
deformed parts.
Loose chains.
Adjust chains in accordance
with paragraph 6- 20.
Cables too tight.
Check cable tension. Adjust
cables to proper tension.
Pulleys binding or cable off.
Observe motion of the pulleys.
Check cables visually. Replace
defective pulleys. Install cables
correctly.
Bellcrank distorted or
damaged.
Check visually. Replace defective
bellcrank.
Clevis bolts in system too tight.
Check connections where used.
Loosen, then tighten properly.
6-1
6-3. TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
REMEDY
Defective bearing in bearing
blocks at sprockets.
Disconnect chains and check for
binding. Replace defective parts.
Rusty chain.
Check visually. Replace chain.
Chain binding with
sprockets.
Check freedom of movement.
Replace defective parts.
Defective bearings in
sleeve weld assembly
on control wheel tube.
Disconnect chains and check for
binding. Replace defective parts.
Nuts securing shaft
in bearing blocks on
firewall too tight.
Loosen nuts the least
amount required to
eliminate binding and
align cotter pin hole,
but not over • 030"
maximum clearance.
Improper adjustment of
chains or cables.
Adjust in accordance with
paragraph 6-20.
Improper adjustment of
aileron push-pull tubes.
Adjust push-pull tubes to
obtain proper alignment.
DUAL CONTROL WHEELS
NOT COORDINATED.
Chains not properly
adjusted on sprockets.
Adjust in accordance with
paragraph 6-20.
INCORRECT AILERON
TRAVEL.
Push-pull tubes not adjusted
properly.
Adjust in accordance with
paragraph 6-20.
Incorrect adjustment of travel
stop bolts.
Adjust in accordance with
paragraph 6-20.
RESISTANCE TO CONTROL
WHEEL MOVEMENT (Cont).
PILOT CONTROL WHEEL NOT
LEVEL WITH AILERONS
NEUTRAL.
6-4. CONTROL COLUMN. (Refer to figure 6-2.)
6-5. DESCRIPTION. Rotation of the control wheel
rotates four bearing roller assemblies (8) on the end
of the control wheel tube (2), which in turn, rotates
a square control tube assembly (17) inside and extending from the control wheel tube. Attached to this
square tube (17) is a sprocket (23) which operates
the aileron system. This same arrangement is provided for both control wheels and synchronization of
the control wheels is obtained by the crossover
chains (26) and turnbuckles (27). The forward end of
the square control tube (17) is mounted in a bearing
block (20) on the firewall and does not mo.,,·e fore-andaft, but rotates with the control Wheel. The four
bearing roller assemblies (8) on the end of the control wheel tube reduce friction as the control wheel
is moved fore-and-aft for elevator system operation.
A sleeve weld assembly (6), containing bearings
which permit the control wheel tube to rotate within
it, is secl.&red to the control wheel tube by a sleeve
6-2
•
•
and retaining rings in such a manner it moves foreand-aft with the control wheel tube. This movement
allows the clamp blocks (7) attached to the sleeve
weld assembly (6) to move the elevator cable. When
dual controls are installed, the copilot's control
wheel is linked to the aileron and elevator control
systems in the same manner as the pilot's control
Wheel.
6-6. REMOVAL AND INSTALLATION.
a. CONTROL WHEEL TUBE - REAR SECTION.
1. THRU AIRCRAFT SERIALS 33701398 AND
F33700045. Remove lower screw securing decorative collar (33), slide collar toward instrument panel
and remove remainder of screw s securing control
Wheel (32) to control Wheel tube (2).
2. BEGINNING WITH AIRCRAFT SERIALS
33701399 AND F33700046. Slide cover (58) toward
instrument panel to expose adapter (57) and remove
screws securing adapter (57) to rear section of
tube (2).
•
•
•
•
3. ALL AIRCRAFT. Disconnect electrical wiring at connector (34).
NOTE
On aircraft equipped with the ribbon wire (39),
mark the ribbon wire and the connector at the
control wheel for reference on reinstallation.
IT IS POSSmLE TO PLUG THIS CONNECTION
BACKWARDS.
4. Carefully remove control wheel.
5. Remove screw securing adjustable glide plug
(15) to control tube assembly (17) and remove plug
and glide.
6. THRU AIRCRAFT SERIAL 337-0239. Cut
safety wire, remove bolts (4) and remove clamp
halves (5) to detach the rear section of control wheel
tube (2) from the forward section. Pull rear section
of tube (2) aft, out through the instrument panel to
remove.
7. BEGINNING WITH AIRCRAFT SERIAL 3370240. Cut safety wire and remove studs (9) from
collar (10) to detach the rear section of control wheel
tube (2) from the forward section. Pull rear section
of tube (2) aft, out through the instrument panel to
remove. Do not drop collar (10) into tunnel area.
8. Reverse the preceding steps for reinstallation. Safety wire all items previously safetied,
check rigging of aileron system and rig, if necessary,
in accordance with paragraph 6-20.
b. CONTROL WHEEL TUBE - FORWARD SECTION.
l. Complete steps 1 thru 7 of subparagraph "a. "
2. Remove bolt securing shaft (19) and forward
control column stop (18) to control tube assembly
(17). Pull tube assembly aft, out through the instrument panel.
3. Remove bolts securing clamp blocks (7) and
slide blocks out of sleeve weld assembly (6).
4. THRU AIRCRAFT SERIAL 33701193. Disconnect microphone cable (37) terminals at terminal
block (38) and carefully work forward section of control wheel tube (2) out from under instrument panel.
5. BEGINNING WITH AIRCRAFT SERIALS
33701194 AND F3370000l. Cut sta-strap securing
ribbon wire to forward section of control wheel tube
(2) and carefully pull ribbon wire out of tube. Carefully work forward section of control wheel tube (2)
out from under instrument panel.
6. Reverse the preceding steps for reinstallation. Safety wire all items previously safetied,
check rigging of aileron system and rig, if necessary, in accordance with paragraph 6-20.
7. If control column works hard, or drags foreand-aft, loosen screw securing adjustable glide plug
(15).
8. The nuts (25) securing shafts (19) to the firewall should be tightened snugly, then loosened the
least amount required to eliminate binding and to
align a cotter pin hole, but not more than. 030"
maximum clearance.
6-7. REPAIR. Worn, damaged or defective shafts,
bearings, sprockets, roller chains or other compo-
nents should be replaced. Refer to Section 2 for
lubrication requirements .
6-8. BEARING ROLLER ADJUSTMENT. Each bearing roller (29) has an 0.062" eccentric adjustment
when installed, for adjustment of the control wheel
tube (2), control tube assembly (17) and bracket (28).
For adjustment, proceed as follows:
a. Adjust bearing rollers (29) until control wheel
tube (2) is centered in bracket (28).
b. Operate ailerons and elevators through several
cycles and check for binding. If binding is evident,
readjust bearing rollers individually until binding is
eliminated.
6-9. BELLCRANKS. (Refer to figure 6-l. )
6-10. REMOVAL AND INSTALLATION.
a. Remove access plate adjacent to bell crank (25)
on underside of wing and remove plug button for
access to pivot bolt (23).
b. Remove wing strut fairings or headliner as necessary to gain access to turnbuckle (7, 9 or 13).
c. Remove safety wire and relieve tension at turnbuckle.
d. Disconnect cables (14 and 15) at bellcrank.
e. Disconnect push-pull tube (24) at bellcrank.
f. Remove safety wire from pivot bolt (23) and
remove bolt.
g. Remove bellcrank through access opening, using
care that bushing (26) is not dropped from bellcrank.
NOTE
Brass washers (21) may be used as shims
between upper and lower ends of bellcrank
and brackets (18 and 22). Retain these
shims. Tape open ends of bellcrank to
prevent dust and dirt from entering bellcrank needle bearings (20).
h. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 6-20,
safety wire turnbuckle and pivot bolt and reinstall
all items removed for access.
6-11. REPAIR. Repair of bellcranks is limited to
replacement of defective bearings and bushings. If
needle bearings are dirty or in need of lubrication,
clean thoroughly and lubricate as outlined in Section
2.
6-12. AILERONS. (Refer to figure 6-3.)
6-13. REMOVAL AND INSTALLATION.
a. Run flaps to full DOWN position for access to
inboard hinge bolt.
b. Remove wing tip for access to outboard binge
bolt.
c. Remove access plate (8) and plug buttons from
underside of aileron.
d. Remove bolt (7) securing push-pull tube (6) to
aileron .
e. Remove pivot bolts (3) and pull aileron aft to
remove.
6-3
•
DetauB
Detail
A
NOTE
DetailC
NOTE
•
14. Cable (Carry-Thru)
NOTE
15. Cable (LH Outboard Direct)
16. Pulley (Carry-Thru Cable)
Refer to figure 4-2 for
17. Bushing
cable routing through
18. Upper Bracket
wing strut fairleads.
19. Travel Stop Bolt
20. Bearing
21. Brass Washer
I~AUT~ONI
22. Lower Bracket
MAINTAIN PROPER CONTROL
23. Pivot Bolt
CABLE TENSION.
24. Push-Pull Tube
25. Bellcrank
26. Bushing
CABLE TENSION:
30 ± 10 LBS (AT AVERAGE TEMPERATURE FOR THE AREA. )
Safety wire these items.
REFER TO FIGURE 1-1 FOR TRAVEL.
1. Cable Guard
2. Pulley (RH Direct Cable)
3. Pulley (Elevator Down)
4. Pulley (Elevator Up)
5. Bracket
6. Pulley (LH Direct Cable)
7. Turnbuckle (Carry-Thru Cable)
8. Cable (RH Outboard Direct)
9. Turnbuckle (RH Direct Cable)
10. Cable (RH Inboard Direct)
11. CleviS
12. Cable (LH Inboard Direct)
13. Turnbuckle (LH DirecfCable)
*
Figure 6-1. Aileron Control System (Sheet 1 of 2)
6-4
•
•
5
Detail
0
Detail
E
5
•
Detail
F
Detail
G
NOTE
Direction of stop bolts (19) may
be reversed if rigging interference occurs.
'.
Detail
H
*
Safety wire these items.
DETAn..S D THRU H ARE TYPICAL
FOR LEFT AND RIGHT HAND SIDES
Figure 6-1. Aileron Control System (Sheet 2 of 2)
6-5
1. Single Controls Cover
Control Wheel Tube
Collar
Bolt
Clamp Halves
6. Sleeve Weld Assembly
7. Terminal Blocks
8. Roller Assembly
9. Stud
10. Collar
11. Retainer Ring
12. Thrust Bearing Race
13. Thrust Bearing
14. Needle Bearing
15. Adjustable Glide Plug
16. GIlde Assembly
17. Control Tube Assembly
18. Forward Control Column
Stop
19. Shaft
20. Bearing Block
21. Retainer Ring
22. Dowel Pin
23. Sprocket
24. Teflon Thrust Washer
25. Nut
26. Crossover Chain
27. Crossover Chain
Turnbuckle
28. Bracket
29. Roller
30. Direct Chain
31. Aft Control Column Stop
32. Control Wheel
33. Decorative Collar
34. Connector
35. Tube
36. Tie Strap
2.
3.
4.
5.
61. Autopilot Disengage
Microphone Cable
Switch
Terminal Block
62. Support
Cable Assembly
Relay
Socket
.Safety wire these items.
Fuse
3
Microphone Switch
Plug
Setscrew
Electric Trim Switch
Circuit Board Shield
Circuit Board Assembly
Insulator
Bracket
Pad
Map Light Assembly
Map Light Rheostat
Plate
Spacer
Rubber Cover
Adapter
Cover
Electric Trim
Disengage Switch
11
60. Housing
Beginning with aircraft serials 33701463
and F33700056
37.
38.
39.
40.
41.
42.
43.
44.
45.
46.
47.
48.
49.
50.
51.
52.
53.
54.
55.
56.
57.
58.
59.
15
16
28
B
.---1
OBeginning with aircraft serial 337 -0240
.
I
NOTE
25
Use teflon thrust washers (24) as
required to remove free play. Use
a minimum of 1 forward and aft of
bearing blocks (20) and sprockets (23).
Figure 6-2. Control Column Installation (Sheet 1 of 2)
6-6
•
-(:( Thru aircraft serial
337-0239
/'"
Detail
•
•
•
•
This section (\f microphone cable
must be tight to clear structure.
NOTE
From centerline of adapter (57) to
centerline of plate (54) lower sc:"ew
should be 2.38 " as shown.
Collar (33) must be attached
with bottom screw.
THRU AIRCRAFT
SERIAL 33701193
36
•
\\111
41
42
AIRCRAFT SERIALS 33701317 THRU
33701398 AND F33700025 THRU F33700045
45
57
34
39
41
42
'.
BEGINNING WITH AIRCRAFT
SERIALS 33701399 AND F33700046
2.38 "
• Microphone switch position when electric trim and
oxygen systems are both installed •
•• Electric trim switch position when oxygen system
and electric trim system are installed. Microphone
switch position when oxygen system is installed and
electric trim system is NOT installed.
Figure 6-2. Control Column Installation (Sheet 2 of 2)
6-7
•
5
1. Bearing
2. Inboard Hinge
3. Pivot Bolt
4. Trim Tab
5. Aileron
6. Push- Pull Tube
7. Mounting Bolt
8. Access Plate
9. Center Hinge
10. Wing
11. Outboard Hinge
•
\
18
11
Figure 6-3. Aileron Installation
f.
Reverse the preceding steps for reinstallation.
6-19. REMOVAL AND INSTALLATION.
If rigging was correct and push-pull tube rod end
adjustment was not disturbed, it should not be necessary to re-rig system. Check aileron travel and
alignment, re-rig if necessary, in accordance with
paragraph 6-20. Install all items removed for access.
6-14. REPAIR. Aileron repair may be accomplished
NOTE
•
The following procedures are written for
cables on the left side of the aircraft.
Cables on the right side are removed in
a similar manner.
in accordance with instructions outlined in Section 16.
6-15. AILERON TRIM TABS. (Refer to figure 6-3.)
6-16. REMOVAL AND INSTALLATION.
a. Remove screws from lower side of tab.
b. Drill out rivets on upper side of tab.
c. Reverse the preceding steps for reinstallation.
6-17. ADJUSTMENT. Adjustment is accomplished
by loosening the screws, shifting the tab trailing
edge UP to correct for a wing-heavy condition or
DOWN for a wing-light condition, then tightening the
screws. Beginning with aircraft serial 337-0240
divide the correction equally on both tabs. When installing a new Wing or aileron, set the tabs in neutral
and adjust as necessary after flight test.
6-18. CABLES AND PULLEYS. (Refer to figure
6-l. )
6-8
a. DIRECT CABLE-INBOARD.
l. Remove seats and access plates as necessary
to expose Details B and D.
2. Remove wing strut fairings as necessary to
expose turnbuckle (13).
3. Remove safety wire and relieve cable tension
at turnbuckle (13). Disconnect cable (12) end from
turnbuckle barrel.
4. Disconnect cable (12) at clevis (11).
5. Remove cable guards and pulleys as necessary to work cable free of aircraft.
NOTE
To ease routing of cable, a length of wire
may be attached to the end of cable before
being withdrawn from the aircraft. Leave
wire in place, routed through structure;
then attach the cable being installed and
use wire to pull 'cable into position.
•
•
AVAILABLE FROM CESSNA SERVICE PARTS CENTER (TOOL NO. SE 716)
Figure 6-4. Inclinometer for Measuring Control Surface Travel
•
•
6. After cable is routed in position, install pulleys and cable guards. Ensure cable is poSitioned
properly through strut fairleads and in pulley grooves
before installing guards.
7. Re-rig aileron system in accordance with
paragraph 6-20, safety turnbuckle (13) and reinstall
all items removed for access.
b. DIRECT CABLE-OUTBOARD.
1. Remove access plates as necessary to expose
Details E, G and H.
2. Remove wing strut fairings as necessary to
expose turnbuckle (13).
3. Remove safety wire and relieve cable tension
at turnbuckle barrel.
4. Disconnect cable (15) at bellcrank (25).
5. Complete step 5 of subparagraph "a. "
6. After cable is routed in position, install pulleys and cable guards. Ensure cable is poSitioned
properly through strut fairleads and in pulley grooves
before installing guards.
7. Re-rig aileron system in accordance with
paragraph 6-20, safety turnbuckle (13) and reinstall
all items removed for access.
c. CARRY-THRU CABLE.
1. Remove wing root fairings and access plates
as necessary to expose Details E, F and H.
2. Remove headliner as necessary to expose
turnbuckle (7).
3. Remove safety wire and relieve cable tension
at turnbuckle (7). Disconnect cable (14) end from
turnbuckle barrel.
4. Disconnect cable (14) at bellcrank (25).
5. Complete step 5 of subparagraph "a."
6. After cable is routed in poSition, install pulleys and cable guards. Ensure cable is positioned in
pulley grooves before installing guards.
7. Re-rig aileron system in accordance with·
paragraph 6-20, safety turnbuckle (7) and reinstall
all items removed for access.
6-20. RIGGING. (Refer to figure 6-1.)
a. Remove access plates and the outer plug button
adjacent to bellcranks (25) on underside of wings ..
b. Remove wing strut fairings and headliner as
necessary to gain access to turnbuckles (7, 9 and 13).
c. Run flaps to full UP position.
d. With aileron faired (aileron trailing edge aligned
with flap trailing edge), loosen jam nuts and adjust
push-pull tube (24) so the nut securing the push-pull
tube to the bellcrank is centered above the plug button
hole. A 3/8" deep-socket, long enough to extend
through the plug button hole when placed on the attaChing nut, may be used.as a rigging tool. Tighten
jam nuts.
e. Complete step "d" for opposite push-pull tube.
r. Install the control lock to place pilot's control
wheel in neutral position.
g. (Refer to figure 6-2.) Check that the direct
chain (30) is engaged on the forward sprocket (23)
and that the chain has apprOximately an equal number of links extending from the sprocket on both
sides. If necessary, loosen the direct cable turnbuckles and reposition chain on sprocket.
h. (Refer to figure 6-1.) With the control lock still
in place, adjust the direct and carry-thru cable turnbuckles to align both ailerons in neutral position and
to obtain proper cable tension. Results to turnbuckle
adjustments are as follows:
1. Loosening the carry-thru cable turnbuckle
(7) and tightening .either direct cable turnbuckle
(9 or 13) will move the aileron for that particular
side down.
6-9
2. Loosening the carry-thru cable turnbuckle
(7) and tightening both direct cable turnbuckles (9
and 13) will move both ailerons down.
3. Loosening either direct cable turnbuckle
(9 ar 13) and tightening carry-thru cable turnbuckle
(7) will move the aileron for that particular side up.
4. Loosening both direct cable turnbuckles (9
and 13) and tightening the carry-thru cable turnbuckle
(7) will move both ailerons up.
i. (Refer to figure 6-2.) To synchronize the copilot's control wheel with the pilot's control wheel,
adjust crossover chain turnbuckles (27) so that both
control wheels are in neutral pOsition with the control lock installed. Chain tension should be the minimum required to remove slack from chain.
j. (Refer to figure 6-1.) Remove control lock and
adjust stop-bolts (19) at each bellcrank (25) to degree
of travel specified in figure 1-1. If the thickness of
the stop-bolt heads should interfere with rigging, the
stop-bolts may be reversed in their nutplates.
NOTE
•
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
k. Safety wire all turnbuckles, tighten all jam nuts
and reinstall all items removed for access.
IWARNING'
Be sure ailerons move in the correct direction when operated by the control wheels.
SHOP NOTES:
•
•
6-10
SECTION 7
•
WING FLAP CONTROL SYSTEM
TABLE OF CONTENTS
•
WING FLAP CONTROL SYSTEM .
Description . . . • . . . .
Operational Check . . . . .
Trouble Shooting . . . . . . . . . .
Flap Motor, Transmission and Actuator
Assembly . . . . . . . .
Removal and Installation
Repair . . . . . . . .
Flaps . . . . . . . . . . .
Removal and Installation
Repair.
BeUcranks . . . . . . . .
Page
7-1
7-1
7-2
7-3
7-5
7-5
7-10
7-10
7-10
7-10
7-10
7-1. WING FLAP CONTROL SYSTEM. (Refer to
figure 7-1. )
'.
7-2. DESCRIPTION.
a. THRU AIRCRAFT SERIAL 337-0239 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK337-19. (Refer to figure 7-2, sheet 1.) The wing flap control
system consists of an electric motor, transmission
and actuator assembly, three interconnected bellcranks in each wing, synchronizing push-pull tubes,
push-pull rods, control cables, pulleys, a down-limit
switch located in the wing and a control switch mounted in the instrument panel. The transmission con-
Removal and Installation
Repair . . . . . . . .
Flap Position Transmitter
Removal and Installation
Adjustment . . . . . .
Flap Control Lever. . . • .
Removal and Installation
Cables and Pulleys . . . • .
Removal and Installation
Rigging . . • • • . . . • .
Flap/Elevator Trim Interconnect
7-10
7-10
7-10
7-10
7-10
7-12
7-12
7-12
7-12
7-12
7-15
verts the rotary motion of the motor to the push-pull
motion needed to operate the flaps and will freewheel
at each end of its stroke, but a down-limit switch
opens the motor circuit just before free-wheeling
occurs in the down poSition. Overrunning at intermediate flap settings is minimized by a solenoidreleased brake at the flap motor. A three-position,
momentary-on switch, spring-loaded to the center
OFF poSition, operates the flap control system. The
position indicator is operated electrically by a transmitter which is linked mechanically to the left inboard
flap bellcrank.
7-1
b. BEGINNING WITH AIRCRAFT SERIALS 337-0240
AND F33700001. (Refer to figure 7-2, sheet 2.) The
wing flap control system consists of an electric
motor, transmission and actuator assembly, three interconnected bellcranks in each wing, synchronizing
push-pull tubes, push-pull rods, control cables, pulleys and a follow-up control. The transmission converts the rotary motion of the motor to the push-pull
motion needed to operate the flaps and will NOT freeWheel at each end of its stroke, the limit switches
(2 and 26) MUST be adjusted properly to stop the
motor at the flap travel extremes or structural damage will result. Electrical power to the motor is controlled by two micro switches mounted on a "floating"
arm. The position indicator is mechanically linked
to the actuator and "floating" arm by the follow-up
control. (Refer to figure 7-3.) Switches (9 and 10)
at the instrument panel actuate the system and control all mid-range flap settings While the limit switches on the actuator de-actuate the system at either
travel extreme. As the control lever (5) is lowered
to the desired flap setting, cam (6) contacts microswitch (9) actuating the motor. As the flaps move
down, the follow-up control (4) pivots arm (2) until
micro switch (9) clears cam (6) breaking the circuit.
As the control lever (5) is raised, cam (6) contacts
micro switch (10) actuating the motor in the reverse
direction, raising the flaps in a similar manner.
Refer to Section 8 for the flap/elevator trim interconnect system.
c. THRU AIRCRAFT SERIAL 337-0239 WHEN
MODIFIED IN ACCORDANCE WITH SK337-19.
(Refer to figure 7-1, sheet 2.) This wing flap control
system consists of the motor, transmission, actuator
and limit switches described in subparagraph "b"
utilizing the three-position switch and transmitting
system described in subparagraph "a. "
7-3. OPERATIONAL CHECK.
a. Operate flaps through their full range of travel
observing for uneven or jumpy motion, binding and
lost motion in system. Ensure all flaps move simultaneously through their full range of travel.
b. THRU AIRCRAFT SERIAL 337-0239 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK337-19. Run
flaps to full down position unW down-limit switch
breaks circuit, then run flaps full up and overrun
motor to check that transmission freewheeling occurs
in UP position only.
c. BEGINNING WITH AIRCRAFT SERIALS 337-0240,
F33700001 AND ALL AIRCRAFT MODIFIED IN ACCORDANCE WITH SK337-19. Check for positive
shut-off of motor at flap travel extremes. FLAP
MOTOR MUST SHUT-OFF.
d. Check ::hat flaps are not sluggish in operation.
It should take apprOximately 6 to 8 seconds for the
flaps to extend or retract fully.
e. Stop flaps at various settings during extension
and retraction to check that flaps do not coast.
f. Raise flaps and check each flap manually for full
up position.
•
NOTE
At least one roller on each flap should contact the end of flap track slot with flaps in
the full up position.
g. With flaps full UP, mount an inclinometer on
one flap and set to 0 0 • Lower flaps to DOWN poSition and check flap angle as specified in figure 1-1.
Raise flaps to 1/3 pOsition, check that inclinometer
reads apprOximately 8 0 and that position indicator
reads approximately 1/3 (thru aircraft serial 3370239) or that pointer indicates 1/3:t1/16 inch (beginning with aircraft serial 337-0240 and F33700001).
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
h. Remove nap well gap seal panels and access
plates and attempt to rock bellcranks to check for
bearing wear.
i. Inspect flap rollers and tracks for evidence of
binding and defective parts.
j. Install elevator control lock or rigging tool to
keep elevator in neutral, lower flaps to DOWN position and place elevator trim control in full NOSE UP
position (trim tab full DOWN).
k. Mount an inclinometer (refer to note in step "g")
on trim tab, raise flaps and check that trim tab
moves from FULL DOWN position to degree of travel
specified in figure 1-1 for that specific aircraft model.
Refer to Section 8 for details of the flap/elevator
trim interconnect system.
•
SHOP NOTES:
•
7-2
•
7-4. TROUBLE SHOOTING •
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 7-21.
TROUBLE
FLAPS FAIL TO MOVE.
•
BINDING IN SYSTEM AS FLAPS
ARE RAISED AND LOWERED.
PROBABLE CAUSE
REMEDY
Popped circuit breaker.
Check visually and reset
breaker. If breaker pops
again, determine cause
and correct.
Defective circuit breaker.
Check continuity. Replace
breaker.
Defective limit-switch.
Check continuity. Replace
switch.
Defective motor.
Remove and bench test. Replace
motor.
Broken or disconnected wires.
Check continuity. Connect or
replace wiring.
Defective or disconnected
transmission or actuator
assembly .
Connect or replace transmisSion
or actuator assembly. Remove
and bench test if necessary.
Disconnected cables.
Check visually. Connect cables.
Follow-up control disconnected
or slipping. (Beginning with
aircraft serials 337-0240 and
F33700001. )
Check visually. Secure control
or replace if defective.
Three-position switch on
instrument panel defective.
(Thru aircraft serial 337-0239.)
Check continuity. Replace
defective switch.
Cables not riding on pulleys.
Check visually. Route cables
correctly over pulleys. Check
cable guards.
Bind in bellcranks.
Check visually. Repair or
replace bellcranks.
Broken or binding pulleys.
Check visually. Replace
defective pulleys.
Frayed cable.
Check visually. Replace
defective cable.
Flaps binding on tracks.
Check visually. Replace
defective parts.
Solenoid brake not releaslng
completely. (Thru aircraft
serial 337-0239.)
Check brake operation. Adjust brake properly or replace
if defective.
7-3
7-4. TROUBLE SHOOTING (Cont).
TROUBLE
FLAPS ON ONE WING FAIL
TO MOVE.
INCORRECT FLAP TRAVEL.
FLAPS COASTING.
(Tbru aircraft serial
PROBABLE CAUSE
Disconnected or broken cable.
Check visually. Connect or
replace cable.
Broken attachment to
actuator.
Check visually. Replace
defective parts.
Defective bellcranks or
linkage to flaps.
Check visually. Replace
defective parts.
Incorrect rigging.
Refer to paragraph 7-21.
Defective limit-switch.
Check continuity. Replace
switch.
Follow-up control disconnected
or slipping. (Beginning with
aircraft serials 337-0240 and
F33700001. )
Secure control or replace if
defective.
Solenoid brake defective
or improperly adjusted.
Check brake operation.
Adjust or replace brake
as required.
Popped circuit breaker.
Check visually. Reset
breaker. If it pops out
again, determine cause
and correct.
Defective Circuit breaker.
Check continuity. Replace
defective breaker.
Defective Wiring.
Check continuity. Repair
Wiring.
Defective position transmitter.
Disconnect "hot" wire
to transmitter. Check
transmitter for varying
resistance as transmitter
arm is moved. Replace
defective transmitter.
Defective position indicator.
If there is voltage to
Follow-up control slipping
Check visually. Connect
or secure control. Replace
if defective.
337-0239. )
FLAP POSITION INDICATOR FAILS TO
RESPOND. (Thru aircraft serial 337-0239.)
FLAP POSITION INDICATOR FAILS TO
RESPOND OR READlNGS ERRONEOUS.
(Beginning with aircraft serials 337-0240
and F33700001. )
7-4
REMEDY
in clamps or broken or
disconnected control.
Pointer bent or broken.
•
•
the indicator, continuity
through wires and transmitter is good. Replace
defective indicator.
Check visually. Repair
or replace pointer.
•
•
7-4. TROUBLE SHOOTING (Cont) .
TROUBLE
PROBABLE CAUSE
FLAP POSITION INDICATOR READINGS
ERRONEOUS. (Thru
aircraft serial 337-0239.)
FLAPS FAIL TO EXTEND.
(Beginning with aircraft
serials 337-0240 and
F33700001. )
•
FLAPS FAIL TO RETRACT.
(Beginning with aircraft
serials 337-0240 and
F33700001. )
Position transmitter not
adjusted properly.
Refer to paragraph 7-21.
Defective position transmitter.
Substitute a known-good
transmitter and check
operation. Replace defective transmitter.
Defective position indicator.
Substitute a known-good
indicator and check
operation. Replace defective indicator.
Loose electrical connection.
Check connections and
tighten as required.
Defective, "loose, or improperly adjusted forward
operating switch.
Check security, adjustment
and operation of switch.
Adjust, secure or replace
switch as required.
Follow-up control slipping,
broken or disconnected.
Check visually. Connect and
secure control. Replace if
defective .
Defective down-limit switch.
Check continuity. Replace
defective switch.
Defective loose or improperly adjusted aft operating
switch.
Check security, adjustment
and operation of switch.
Adjust, secure or replace
switch as required.
7-5. FLAP MOTOR, TRANSMISSION AND ACTUATOR ASSEMBL Y.
7-6 . REMOVAL AND INSTALLATION.
a. THRU AIRCRAFT SERIAL 337-0239 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK337-19.
NOTE
The flap motor, brake, transmission and
actuator assembly are normally removed
as a unit. However, the motor and/or
solenoid brake may be removed separately
if desired.
•
REMEDY
1. Run flaps to DOWN position.
2. Disconnect battery terminals as a safety
precaution.
3. Remove headliner and soundproofing as
necessary to gain access.
4. (Refer to figure 7-1.) Remove flap well gap
seal panel and flap well access plate aft of bellcrank
assembly (16) in each wing, remove safety wire (17)
and relieve tension on cable (10) by loosening adjustment nut (18).
5. (Refer to figure 7-2.) Remove bolts securing
cables to actuator (10).
6. Disconnect the electrical wiring to motor
assembly.
7. Remove bolts (12, 13 and 14) attaching actuator to support structure (15) and carefully remove
assembly from aircraft.
8. Reverse the preceding steps for reinstallation. Rig system in accordance with paragraph 7-21.
NOTE
If the motor and transmission were separated
for any reason, refer to figure 7-5 during
reassembly.
7-5
•
5
'---- REFER TO FIGURE 7-3
•
11
& 7
Detail
A
12
Detail
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
Interconnect Control
Aileron Assembly
Follow-Up Control
Control Lever
Upper Cabin Skin
Cable Guard
Bracket
Bushing
Pulley
Extend Cable (Inboard)
Retract Cable
Extend Cable (Outboard)
Lower Bellcrank Bracket
Brass Washer
Bearing
Bellcrank Assembly
(Inboard Flap)
Safety Wire
Adjustment Nut
19. Bushing
20. Upper Bellcrank
Bracket
21. Pivot Bolt
22. Push-Pull Rod
23. Bracket
24. Cable Guard
25. Busbing
26. Pulley
27. Upper Doubler
28. Inboard Bellcrank Assembly (Outboard Flap)
29. Synchronizing PushPull Tube
30. Lower Bracket
31. Attach Bracket
32. Outboard Bellcrank Assembly (Outboard Flap)
33. Lower Doubler
B
NOTE
All details shown are for LEFT wing.
RIGHT wing opposite.
{CAUTION\
M.!\INTAIN PROPER CONTROL
CABLE TENSION.
CABLE TENSION:
30 ± 10 LBS (AT AVERAGE TEMPERATURE FOR THE AREA. )
REFER TO FIGURE 1-1 FOR TRAVEL.
Figure 7-1. Wing Flaps Control System (Sheet 1 of 2)
7-6
•
•
" ~,
23
19
r-'I--___
""
' ""'.
'5
"
Detail
~21
17
t!r - 27
r-
C
29
• BEGINNING
SERIAL 337 -~r: AIRCRAFT
11
Detail
E
23
\./
Detail
F
RIGGING
PIN HOLE
(TYPICAL)
14-~~
Detail
'.
Detail
G
Figure 7-1.
W~~
H
* Install w"th
.
'
head down
•
aps Control System (Sheet 2 of 2)
7-7
3
Slotted holes are provided In - - - - - -.......
bracket for brake adjustment.
'\.---2
•
• NUT TO BE FINGER TIGHT
WHEN PIN IS INSTALLED
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
Support
DOWN-LIMIT Switch
Brake Solenoid
Brake Lining
Grommet
Motor Assembly
Coupling
Bushing
Transmission Assembly
.......
10.
11.
12.
13.
......
Actuator Assembly
Setscrew
Bolt
Bolt
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
Bolt
Support Structure
Electrical Connector
Spacer
Bolt
Follow-Up Control
Bolt
Bellcrank
Plate Assembly
Bracket
Insulator
Actuator
UP-LIMIT Switch
Switch Actuating Cam
AmCRAFT SERIALS 337 -0001 THRU 337 -0239 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK337-19. REFER TO
SHEET 2 FOR AmCRAFT WHICH HAVE BEEN MODiFIED.
Figure 7 -2. Flap Motor, Transmission and Actuator Installation (Sheet 1 of 2)
7-8
•
B
•
23--"-'
23
o BEGINNING WITH AIRCRAFT SERIALS
~24.
~
~25
•
~~
A
, I
Detail
337-1067 AND F33700001
• BEGINNING WITH AIRCRAFT SERIALS
337 -1022 AND F33700001
BEGINNING WITH AIRCRAFT SERIALS
337 -0240 AND F33700001
.2.-0
2~!'-"/:~
I
;
,
\
0~
25~
AIRCRAFT SERIALS 337 -0001 THRU
337 -0239 WHEN MODIFIED IN ACcoRDANcE WITH SK337-19
Detail
B
Figure 7 -2. Flap Motor, Transmission and Actuator Installation (Sheet 2 of 2)
7-9
9. The solenoid brake must be adjusted with
the solenoid actuated. The minimum clearance
between the brake lining and any part of the coupling
is .001 inch and the maximum clearance is .010 inch.
b. BEGINNING WITH AIRCRAFT SERIAL 337-0~40
AND F33700001.
NOTE
Remove motor, transmission, actuator
and support as a unit.
1. Complete steps 1 thru 5 of subparagraph "a. "
2. Remove bolt (18) attaching bellcrank (21) to
actuator (10).
3. Disconnect electrical connector (16) and remove switch (26) from mounting bracket (23). DO
NOT DISCONNECT WIRING FROM SWITCH.
4. Remove bolts (12, 13 and 14) attaching actuator to support structure (15) and carefully remove
assembly from aircraft.
5. Reverse the preceding steps for reinstallation. Rig system in accordance with paragraph 7-21.
c. THRU AIRCRAFT SERIAL 337-0239 WHEN
MODIFIED IN ACCORDANCE WITH SK337-19.
1. Use same procedure outlined in subparagraph
"b" omitting step 2.
7-7. REPAIR. Repair consists of replacement of
motor, transmission, coupling, brake, actuator
parts and associated hardware. Lubricate as outlined in Section 2.
7-8.
FLAPS. (Refer to figure 7-4.)
7-9. REMOVAL AND INSTALLATION.
a. Run flaps to DOWN pOSition.
b. Disconnect push-pull rods at attach brackets (8)
on flap to be removed.
c. Remove access plates (9) at top leading edge of
flap.
d. Remove bolts (6) at each flap track. As flap is
removed from wing, all spacers, rollers and bushings will fall free. Retain these for reinstallation.
e. Reverse the preceding steps for reinstallation.
If the push-pull rod adjustment is not disturbed, rerigging of the system should not be necessary.
Check flap travel and rig in accordance with paragraph 7-21, if necessary.
7-10. REPAIR. Repair may be accomplished in
accordance with instructions outlined in Section 16.
7-11. BELLCRANKS. (Refer to figure 7-1.)
7-12. REMOVAL AND INSTALLATION.
a. BELLCRANK ASSEMBLY. (INBOARD FLAPDETAIL B.)
1. Run flaps to DOWN position.
2. Remove flap well gap seal panel and access
plate.
3. Disconnect push-pull rod (22) at bellcrank.
4. Remove safety wire (17), remove adjustment
nuts (18) and remove cables from bellcrank.
7-10
5. Disconnect position transmitter link (index
11, figure 7-4) at bellcrank (thru aircraft serial 3370239).
6. Remove pivot bolt (21) attaching bellcrank to
wing structure.
7. Using care, remove bellcrank through access
opening, being careful not to drop bushing (19). Retain brass washer (14) between bellcrank and lower
wing structure for use on reinstallation. Tape open
ends of bellcrank after removal to protect bearings
(15).
8. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 7-2l.
b. INBOARD BELLCRANK ASSEMBLY. (OUTBOARD FLAP-DETAIL D.)
1. Complete steps 1 thru 4 in subparagraph
•
"a."
2. Disconnect flap/elevator trim interconnect
control from bellcrank (right wing only).
3. Disconnect synchronizing push-pull tube (29)
at bellcrank (28).
4. Complete steps 6 thru 8 in subparagraph
"a."
c. OUTBOARD BELLCRANK ASSEMBLY. (OUTBOARD FLAP-DETAIL G.)
1. Run flaps to DOWN position.
2. Remove flap well gap seal panel and access
plate.
3. Disconnect synchronizing push-pull tube (29)
at bellcrank (32).
4. Disconnect push-pull rod (22) at bellcrank.
5. Remove pivot bolt (21) attaching bellcrank to
doublers (27 and 33).
6. Complete steps 7 and 8 in subparagraph
"a."
•
7-13. REPAIR. Repair is limited to replacement
of bearings. Cracked, bent or excessively worn
bellcranks must be replaced. Lubricate as outlined
in Section 2.
7-14. FLAP POSITION TRANSMITTER. (THRU
AIRCRAFT SERIAL 337-0239.) (Refer to figure 7-4.)
7-15. REMOVAL AND INSTALLATION.
a. Remove access plates from bottom of left wing
below inboard flap bellcrank.
b. Remove screw s and nuts securing transmitter
(13).
c. Remove the cotter pin, washer and spacer
securing the flap position transmitter wire rod (12)
to the link rod (11).
d. Disconnect the transmitter electrical wires at
the quick-disconnects and remove the transmitter.
e. Reverse the preceding steps for reinstallation
and adjust in accordance with paragraph 7-16.
7-16. ADJUSTMENT.
a. Remove access plates from bottom of left wing
below inboard flap bell crank.
b. Mount an inclinometer on trailing edge of flap
and adjust to 0
Lower flaps to 8 and adjust transmitter as necessary so that indicator reads 1/3.
Slotted holes are provided at transmitter mounting
0
0
•
•
•
.'
........... .
.... .....:;~.........................
-
.......................
.'
. :::=.............
,---REFER TO FIGURE 7-2
...............
REFER TO FIGURE 8-10
•
3
1-
Bushings
2. Switch Mounting Arm
3. Spring
Follow- Up Control
Lever Assembly
Cam
Washers
Knob
Flaps DOWN Operating Switch
10. Flaps UP Operating Switch
II. POSition Indicator
12. Bolt
13. Bracket
4.
5.
6.
7.
8.
9.
13
NOTE
Beginning with aircraft serial 337 -1090
insulators are installed between switches
(9 and 10) and switch mounting arm (2).
•
Beginning with aircraft serials 337 -1121
and F33700001, the washers and nuts
which secured switches (9 and 10) to
switch mounting arm (2) were replaced
by a nut plate assembly to ease removal
and installation of switch~s .
~--APPLY
Detail
A
LOCTITE GRADE C
OR CE UPON INSTALLATION
OF KNOB (8)
BEGINNING WITH AmCRAFT SERIAL 337 -0240
Figure 7-3. Control Lever Installation
7-11
screws. H additional adjustment is necessary, the
wire rod on transmitter may be bent slighUy.
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Senice Parts Center. Refer to figure 6-4.
7-17. FLAP CONTROL LEVER. (BEGINNING
WITH AIRCRAFT SERIAL 337-0240 AND F33700001.)
(Refer to figure 7-3.)
7-18. REMOVAL AND INSTALLATION.
a. Disconnect battery terminals as a safety precaution.
b. Disconnect follow-up control (4) at switch mounting
arm (2).
c. Remove flap operating switches (9 and 10) from
switch mounting arm (2). DO NOT disconnect electrical wiring at switches.
d. Remove knob (8) from control lever (5).
e. Remove remaining items by removing bolt (12).
f. Reverse the preceding steps for reinstallation.
Do not overtighten bolt (12) causing lever (5) to bind.
Rig system in accordance with paragraph 7-21.
7-19. CABLES AND PULLEYS. (Refer to figure
7-1. )
7-20. REMOVAL AND INSTALLATION.
a. EXTEND CABLE (INBOARD).
1. Run flaps to DOWN position.
2. Remove flap well gap seal panels and access
plates as necessary to expose components in Details
A and B.
3. Remove headliner as necessary to expose
actuator assembly (figure 7-2).
4. Remove safety wire (17) and remove adjustment nut (18) from control cable (10) 10 Detail B.
5. Disconnect cables at actuator (index 10, figure
7-2).
6. Remove cable guards and pulleys as necessary to work cable free of aircraft.
NOTE
To ease routing of cable, a length of wire
may be attached to the end of cable being
withdrawn from the aircraft. Leave wire
in place, routed through structure; then
attach the cable being installed and use
wire to pull cable into position.
7. Reverse the preceding steps for reinstallation.
8. After cable is routed in position, install pulleys and cable guards. Ensure cable is installed in
pulley grooves before installing guards. Re-rig system in accordance with paragraph 7-21, safety cable
ends and reinstall all items removed for access.
b. EXTEND CABLE (OUTBOARD).
1. Run flaps to DOWN position.
2. Remove flap well gap seal panels and access
plates as necessary to expose components in Details
B, C and D.
7-12
3. Remove safety wire (17) and remove adjustment nuts (18) from control cable (12) in Details B
andD.
4. Remove pulley (index 9, Detail C).
5. Refer to "note" in step 6 of sub-paragraph "a."
6. Reverse the preceding steps for reinstallation.
7. After cable is routed in position, install pulley. Ensure cable is installed in pulley groove and
that cable guard (24) is installed. Re-rig system in
accordance with paragraph 7-21, safety cable ends
and reinstall all items removed for access.
c. RETRACT CABLE.
1. Run flaps to DOWN position.
2. Remove flap well gap seal panels and access
plates as neces~ry to expose components in Details
A, D, E and F.
3. Remove headliner as necessary to expose
actuator assembly (figure 7-2).
4. Remove safety wire (17) and remove adjustment nut (18) from control cable (11) in Detail D.
5. Disconnect cables at actuator (index 10, figure
7-2).
6. Remove cable guards and pulleys as necessary to work cable free of aircraft.
7. Refer to "note" in step 6 of subparagraph
•
"a."
8. Reverse the preceding steps for reinstallation.
9. After cable is routed in position, install
pulleys and cable guards. Ensure cable is installed
in pulley grooves before installing guards. Re-rig
system in accordance with paragraph 7-21, safety
cable ends and reinstall all items removed for access.
7-21. RIGGING.
•
NOTE
The following procedures ouUine COMPLETE
nap system rigging. All steps of these procedures should be noted, although individual
circumstances may not requh e that all steps
be completed.
a. THRU AIRCRAFT SERIAL 337-0239 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK337-19.
1. (Refer to figure 7-1.) Run flaps to DOWN
position.
2. Remove flap well gap seal panels and access
plates as necessary to expose Details A thru H on
BOTH wings.
3. Remove headliner as necessary to expose
actuator assembly (figure 7-2).
4. Disconnect all flap push-pull rods (22) at the
bell c ranks.
5. Loosen all cables (10, 11 and 12) at bellcranks (16 and 28) in both wings by loosening adjustment nuts (18).
6. Run flap motor to the full UP position.
7. Tighten adjustment nuts (18) on cables (10,
11 and 12) evenly to position bellcranks (16 and 28)
in each wing so that rigging pins will engage at each
bellcrank while maiiltaining 30± 10 pounds cable tension.
•
•
"
1~'"
2
:--'-~-,
".
3~.@1
~
;-;:?~ / ~
d~/
'.
IJJffi
A'1
. 11/
5
". . . . .
A
......
. . ··f!
....
1
1;2
•
f'-,
~
I
r;~'>
'''-'",';/
'7tt,~
'"
'--
5
Detail
C
, 13
6
2
NOTE
Detail
'.
1.
2.
3.
4.
5.
6.
7.
Nut
Washer
Spacer
Roller
Bushing
Bolt
Rub Button
0
Attach Bracket
Access Plate
Left Inboard Bellcrank
Link
Transmitter Wire Rod
13. Position Transmitter
14. Flap Roller Slot
8.
9.
10.
11.
12.
Beginning with aircraft serial 337 -0239 the
metal spacers were replaced by nylon spacers
(3). The number of these spacers (3) may be
altered on either side of the flap tracks to aid
alignment of flaps, providing sufficient clearance is maintained for free movement of the
roller assemblies.
All details shown are for the LEFT wing.
RIGHT wing opposite.
Figure 7-4. Flap Installation
7-13
* PRIOR TO AIRCRAFT SERIAL 337 -0037
FLAP MOTOR
.0 AIRCRAFT SERIALS 337 -0037
THRU
337-0239 AND ALL PRIOR SPARES
o BEGINNING WITH AIRCRAFT SERIALS
SETSCREW
•
337 -0240 AND F33700001
,>~BRAKE DRllM.
COUPLING~*
. ""
~ .~
BONDED
AND BRAKE DRUM
(Jry .
.................
~
@--'D
.
/~
COUPLING 0
~,
./
.X
.>./'<:-.
". . .
-
...............,.
...........
NOTE
...............
............
Alignment of the flap motor shaft and the transmission
shaft is important. After reassembly, the coupling
assembly must turn freely. It is permissable to enlarge the holes illustrated to a maximum of .250 " to
obtain proper alignment.
Apply LOCTITE sealant, grade C or CE to threads of
setscrew upon final installation.
HOLES (ENLARGE AS REQUIRED
TO .250 " MAXIMUM)
Figure 7 -5. Flap Motor and Transmission Alignment
NOTE
Ensure that the cables are in their pulley
grooves and correct bellcrank tracks before and after completion of step 7.
• The rigging pins may be fabricated from any
suitable 3/16 inch diameter material such as
steel rod or bolts. The length of the outboard
bell crank pins should be approximately 6 inches and 2 inches for the inboard bellcrank
11. (Refer to figure 7-2, sheet 1.) Loosen
screWs attaching DOWN-LIMIT switch (2) to bracket
and slide the switch aft in the slotted holes as far as
possible.
12. Run the flap motor to the full DOWN pOSition.
13. Manually hold one inboard flap in the full UP
poSition (snug, but not tight). Mount an inclinometer
on the trailing edge of flap and set to 0°. Lower the
flap manually to full DOWN poSition and adjust pushpull rod (22) to align with bellcrank attaching hole.
Connect push-pull rod and tighten jam nuts.
•
pins.
NOTE
8. Disconnect push-pull tubes (29) at outboard
bellcranks (32).
9. Install rigging pins in bell cranks (32) and adjust push-pull tubes (29) to align with bellcranks.
Connect push-pull tubes and tighten jam nuts.
I~AUTION1
DO NOT run flap motor While rigging pins
are installed.
10.
Remove all rigging pins.
NOTE
IT rigging pins cannot be removed from bellcranks with only slight effort, repeat steps
7 thru 9.
7-14
An inclinometer for measuring control surface travel is available from the Cessna Service Parts Center. Refer to figure 6-4.
14. Repeat step 13 for the remainder of flaps.
15. Run the flap actuator back from the full
DOWN position. 050 inch.
16. Slide the DOWN-LIMIT switch forward in the
slotted holes unW the switch just actuates. Secure
SWitch in this position.
17. Operate the flaps several times, checking
that the switch opens the circuit .050 inch BEFORE
freewheeling occurs.
18. Adjust the pOSition transmitter in accordance
with paragraph 7-16.
19. Perform an operational checkout of the flap
system in accordance with paragraph 7-3, install all
•
•
safeties and reinstall all items removed for access.
b. BEGINNING WITH AIRCRAFT SERIAL 337-0240
AND F33700001.
Do not use aircraft power to operate the flap
motor until the limit- switches on the actuator
assembly have been adjusted or damage may
occur due to overtravel. Separate the electrical connector at the flap motor and connect
jumper Wires from a 24-volt power source to
operate the flap motor. The leads may be reversed to change motor direction or a 3-position switch (spring-loaded to center OFF position) may be used. Use caution when approaching travel extremes as there is no provision
for freewheeling in the transmission.
1. Complete steps 1 thru 5 of subparagraph "a. "
2. Disconnect the follow-up control clevis
(index 19, figure 7-2) from bellcrank (index 21,
figure 7-2).
3. Disconnect battery terminals as a safety precaution. Using jumpers and an external power source,
carefully run flap motor to full UP position.
4. Complete steps 7 thru 10 of subparagraph
"a."
5. Using jumpers and external power source,
carefully run flap motor to full DOWN position.
6. Complete steps 13 and 14 of subparagraph
•
"a."
7. (Refer to figure 7-2.) With flap motor in the
full DOWN position, adjust DOWN-LIMIT switch
(index 2, sheet 2) to the ACTUATED position and secure switch.
8. Using jumpers and external power source,
carefully run flap motor to the full UP position. Adjust up-limit switch (26) to DEACTUATE flap motor
When inclinometer reads 0 and secure switch.
9. Cycle flaps several times and check degree
of travel as specified in figure 1-1. Check cable tension at various mid-range settings and at travel extremes. Readjust down-limit switch as necessary to
obtain proper travel.
10. Connect follow-up control (19) to bellcrank
(21). Run flaps through full range of travel and observe pointer movement. Adjust follow-up control
clevis in slot of bellcrank (21) and rod end at instrument panel as necessary to obtain full pointer travel
in indicator slot.
11. Carefully run flaps to full UP position, then
disconnect and remove the jumpers and external
power source from flap motor.
12. Connect electrical connector at flap motor
and connect battery terminals.
13. (Refer to figure 7-3.) Move control handle
(5) to the full UP position, move switch mounting
arm (2) until cam (6) is centered between switches
(9 and 10).
14. Adjust switches (9 and 10) in slotted holes
until switch rollers just clear cam (6) and secure.
15. Turn master switch ON and run flaps through
various mid-range settings to the full DOWN position.
Check that the limit-switches on the actuator de-actuate system at the travel extremes.
16. Run flaps to full UP poSition. Mount an inclinometer on one flap and set to 0 0 • Move control
lever (5) to 1/3 position, check that flaps stop at 8
and that the pointer indicates 1/3 pOSition (± 1/16
inch).
17. Check all rod ends and clevis ends for sufficient thread engagement, all jam nuts are tight,
safety wire all cable ends and reinstall all items removed for access.
18. Flight test aircraft and check that follow-up
control does not cause automatic cycling of flaps. If
cycling occurs, readjust switches (9 and 10) as necessary per steps 13 and 14.
c. AIRCRAFT SERIALS 337-0001 THRU 337-0239
WHEN MODIFIED IN ACCORDANCE WITH SK337-19.
0
[C~uT~~~1
Do not use aircraft power to operate the flap
motor until the limit- switches on the actuator
assembly have been adjusted or damage may
occur due to overtravel. Separate the electrical connector at the flap motor and connect
jumper wires from a 24-volt power source to
operate the flap motor. The leads may be
reversed to change motor direction or a 3position switch (spring-loaded to center OFF
position) may be used. Use caution when
approaching travel extremes as there is no
provision for freewheeling in the transmission.
0
•
1. Complete steps 1 thru 5 of subparagraph "a. "
2. Complete step 3 of subparagraph "b. "
3. Complete steps 7 thru 10 of subparagraph
"a."
4. Complete step 5 of subparagraph "b. "
5. Complete steps 13 and 14 of subparagraph
"a. "
6. Complete steps 7 thru 9 of subparagraph
''b. "
7. Complete steps 11 and 12 of subparagraph
''b. "
8. Complete steps 18 and 19 of subparagraph
"a."
7-22. FLAP/ELEVATOR TRIM INTERCONNECT.
Refer to Section 8 for removal, installation and
rigging of flap/elevator trim interconnect.
7-15/(7-16 blank)
•
SECTION 8
ELEVATOR. ELEVATOR TRIM AND FLAP
ELEVATOR TRIM INTERCONNECT SYSTEMS
TABLE OF CONTENTS
•
I
Page
ELEVATOR CONTROL SYSTEM
Description . .
Trouble Shooting . . . . .
Control Column . . . . .
Elevator. . . . . . . . .
Removal and Installation
Repair . . . . . . . . .
Bellcrank . . . . . . . '.
Removal and Installation
Cables and Pulleys . . . . .
Removal and Installation
Forward Cables
Aft Cables . . . . .
Rigging . . . . . . . . . .
ELEV ATOR TRIM CONTROL SYSTEM
Description . . . . . . . .
Trouble Shooting . . . . . .
Trim Tab . . . . . . . . .
Removal and Installation
Trim Tab Actuator . . . . .
Removal and Installation
Disassembly • . • • • •
Cleaning, Inspection and Repair
Reassemblv . • • • • • . .
Trim Tab Free-Play Inspection •.
Trim Tab Bellcrank . . . .
Removal and Installation
Trim Tab Control Wheel
Removal and Installation
Cables and Pulleys . . . . .
Removal and Installation
Forward Cable. . .
Aft Cable . . . . .
Center Cable - Tab Up
Forward Section .
Aft Section
Center Cable - Tab Down
Rigging . . . . . . . . . . . .
Electric Trim Assist Installation.
Description . . . . . .
Trouble Shooting . . . . . .
Removal and Installation
Clutch Adjustment . . . . .
FLAP/ELEVATOR TRIM INTERCONNECT
SYSTEM . . . . . . . . .
Description . . . . . .
Trouble Shooting . . . .
Removal and Installation
Rigging . . . . . . . .
8-1
8-1
8-1
8-3
8-3
8-3
8-3
8-3
8-3
8-3
8-3
8-3
8-6
8-8
8-9
8-9
8-10
8-11
8-11
8-11
8-11
8-11
8-11
8-11
· 8-14
· 8-14
· 8-14A
· 8-14A
·8-14A
· 8-14A
·8-14A
8-15
8-16
8-16
8-16
8-16
8-19.
8-20
8-20
8-20
8-20
8-21
8-21
8-21
8-21
8-23
8-23
8-14
8-1. ELEVATOR CONTROL SYSTEM. (Refer to
figure 8-1. )
through pulleys and fairleads to a beUcrank in the
left vertical fin. This bellcrank operates a push-pull
tube connected to the left balance weight arm of the
elevator.
8-2. DESCRIPTION. The elevator is controlled by
a system of cables routed from the control column.
8-3. TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to're-rig system. refer to paragraph 8-12.
TROUBLE
N(' RESPONSE TO CONTROL
WHEEL FORE-AND-AFT
MOVEMENT.
•
PROBABLE CAUSE
REMEDY
Push- pull tube disconnected.
Check visually. Attach push-pull
tube correctly.
Cables disconnected.
Check visually. Attach cables and
rig system in accordance with
paragraph 8-12.
Cables not clamped to control
column.
Check Visually. Secure cables to
control column.
Change 1
8-1
8-3. TROUBLE SHOOTING (Cont).
TROUBLE
BINDING OR JUMPY MOTION
FELT IN MOVEMENT OF ELEVATOR SYSTEM.
PROBABLE CAUSE
REMEDY
Defective bearing in elevator
bell crank or balance weight
arm.
Check visually. Replace defective bearings.
Cables slack.
Rig system in accordance with
paragraph 8-12.
Cables not riding correctly on
pulleys.
Check visually. Route cables
correctly over pulleys.
Defective elevator hinge
bearings.
Check Visually. Replace defective bearings.
Defective control column roller
bearings.
Check that roller bearings will
rotate freely. Replace defective
bearings.
Push-pull tube clevis bolts too
tight.
Check visually. Readjust to
eliminate binding.
Adjustable glide plug on aft end
of control square tube adjusted
too tightly.
Remove control wheel and check
glide for binding. Loosen screw
in end of glide enough to eliminate binding.
Control column needs lubrication.
Check visually. Lubricate in
accordance with Section 2.
Defective pulleys or cable guards.
Check visually. Replace
defective parts and install
guards properly.
Incorrect rigging.
Rig system in accordance with
paragraph 8-12.
Teflon tape too thick in collar on
control wheel tube. (Thru
aicraft serials 33701462 and
F33700055. )
Check thickness of tape.
Replace . 07" thick tape
with . 06" thick tape.
Eccentric bearings at control
column adjusted too tightly or
defective. (Beginning with
aircraft serials 33701463 and
F33700056. )
Readjust or replace defective
bearings.
Defective bearings in elevator
bob weight mechanism. (Thru
aircraft serial 337-0755.)
Check bearings. Replace
defective bearings.
•
•
•
8-2
•
8-3. TROUBLE SHOOTING (Cont) .
PROBABLE CAUSE
TROUBLE
ELEVATORS FAIL TO ATTAIN
PRESCRIBED TRAVEL.
SLIGHT UNDULATION OF
TAIL DURING FLIGHT.
Stops incorrectly set.
Rig system in accordance with
paragraph 8-12.
Cables tightened unevenly.
Rig system in accordance with
paragraph 8-12.
Interference at instrument
panel.
Rig system in accordance with
paragraph 8-12.
Excessive lateral movement
of elevator bellcrank.
Check clearance with feeler gage
(.005" max). Add brass shims
as required. (Refer to figure 8-1,
sheet 2.)
Cable tension low.
Rig system in accordance with
paragraph 8-12.
8-6. REMOVAL AND INSTALLATION.
a. Remove rudders as outlined in Section 9.
b. Remove access plates as necessary from left
vertical fin.
c. Disconnect elevator push-pull tube (15) from
left balance weight arm (12).
d. Remove safety wire and relieve cable tension at
either turnbuckle (8).
e. Disconnect cables (18 and 22) at bellcrank.
1. Remove bellcrank pivot bolt and shims (19),
noting number and position of shims on each side, of
bellcrank.
g. Remove bellcrank through leading edge access
hole.
h. Reverse the preceding steps for reinstallation.
Rig elevator system in accordance with paragraph
8-12, safety turnbuckle and reinstall all items removed for access.
NOTE
NOTE
Do not disturb push-pull tube length to maintain elevator system rigging.
The elevator down spring, linkage and pushpull tube can be removed from aircraft
without removing bellcrank.
8-4. CONTROL COLUMN. (Refer to figure 6-2.)
Section 6 outlines removal, installation and repair
of the control column.
8- 5. ELEVATOR. (Refer to figure 8-1. )
d. (Refer to figure 8-7.) Disconnect trim tab links
(2) from actuator screw end (1). Wire screw end and
clamp trim control wheel so they cannot be turned to
maintain trim control system rigging.
e. (Refer to figure 8-4.) Remove hinge bolts (6)
and pull elevator aft. Guide balance weight arms (1)
out of fins as elevator is removed.
f. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 8-12 and
reinstall all items removed for access.
8-7. REPAIR. Repair may be accomplished as outlined in Section 16. If repair has affected static balance, check and rebalance as required.
8-8. BELLCRANK. (Refer to figure 8-1.)
•
REMEDY
8-9. REMOVAL AND INSTALLATION.
a. Remove access plates and leading edge section
from left vertical fin.
b. Disconnect elevator downspring (23) fro 1: iJellcrank linkage. (Thru aircraft serial 337-0755.)
c. Disconnect elevator push-pull tube (15) at bellcrank (17) and lower elevator gently.
8-10. CABLES AND PULLEYS.
8-11. REMOVAL AND INSTALLATION.
a. FORWARD CABLES.
1. (Refer to figure 8-1.) Remove pilot's seat,
carpeting and access plates in floorboard area as
necessary to expose Details A, B, and C.
2. Remove left wing strut fairings as necessary
to expose turnbuckles (8).
3. Remove safety Wire, relieve cable tension
and disconnect turnbuckles (7 and 8).
4. (Refer to figure 8-2.) Remove bolts (9) seCuring clamp blocks (7) to sleeve weld assembly (8)
and remove cable swaged balls from blocks.
5. (Refer to figure 9-1.) Remove safety wire
and relieve rudder control system cable tension at
turnbuckle (9).
6. (Refer to figure 8-1.) Mark or tag cables
and pulleys in Details Band C and remove bolts securing pulleys (4, 5 and 6) to brackets (2).
7. Remove cable guards from Detail A and control column as necessary to work cables free of aircraft.
8-3
•
2
..
a ••••····,'
,.I~\.t, ,"
Detail
B
./
,"
"
"
,.I,,.
,
I
:':' REFER TO FIGURE 8-4
"
Detail
A
...
........
,,
,, ,
"- ,,
8-2) ', ',. . . . .
..\ .....,~.
"....
.o'
''',
REFER TO FIGURE
......?
,. , ,
" ,,,,"...
"
,'.
......" , j , ' ...........:..........
"l":::~
E
<."." . . . .. ':::::;~::::::/
././'S..... "
.......
,......... ";.':,..-.
....•
.
......" -
REFER TO
FIGURE 8-3
1.
2.
3,
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
\ .... ,.
1 h,
" fit ....'
!.....
....
",
•
"
..OlIo ....
. .... .
"
B C
Cable Guard
Bracket
Pulley (Aileron Cable)
Pulley (Elevator Down Cable)
Pulley (Elevator Up Cable)
Pulley (Rudder Cable)
Turnbuckle
Turnbuckle
Cover
Elevator
Balance Weight
Balance Weight Arm
Torque Tube
NOTE
14. Up-Stop (Contacted by balance
weight arm)
15. Push-Pull Tube
16. Bearing
17 . Bellcrank
18. Cable (Elevator Down)
19, Brass Shim
20. Down-Stop (Contacted by stop
on bellcrank)
21. Link
22. Cable (Elevator Up)
23. Elevator Downspring
24. Bushing
Locate turnbuckles of
adjacent cables so they
do not meet, cross or
rub.
Shaded pulleys are used
in this system only,
Refer to figure 4- 2 for
cable routing thr(\ugh
wing strut fairleads.
@AUTION}
M...~INT AlN
PROPER CONTROL
CABLE TENSION.
CABLE TENSION:
20 + 10 - 0 LBS (AT AVERAGE TEMPERA TURE FOR THE AREA. )
REFER.TO FIGURE 1-1 FOR TRAVEL.
Figure 8-1. Elevator Control System (Sheet 1 of 2)
8-4
•
•
6 5
~~
4
THRU AIRCRAFT
SERIALS 33701462
AND F33700055
6
"
Detail
Detail
2
4
D
BEGINNING WITH AIRCRAFT SERIALS 33701463 AND F33700056
C
NOTE
Elevator torque tube (13) and balance
weight arm (12) are matched parts,
drilled on assembly. When replacing
components of the balance weight assembly, rebalance in accordance with Section 16.
•
12
13
Add brass shims (19) as required to
reduce lateral movement of belle rank
(17) to . 005 .. maximum.
..-
All components of DETAIL E are located inside left fin. Similar balance
weight is located inside right fin.
/
.I
14--i1
16
~
* 14. 30 .. APPROX.
014.12" APPROX.
I
15
....~,.
.21
castellated nuts and pins
thru aircraft serial 3370958
.~//'
. s
.---.J"""''¥I-..
1&
Detail
E
*Torque to 15 to 40 lb-in when torque
wrench is attached to nut. Torque to
10 to 40 lb-in when torque wrench is
attached to bolt head.
• THRU AIRCRAFT SERIAL 337 -0755
•
*
THRU AIRCRAFT SERIAL 337 -0978
o BEGINNING WITH AIRCRAFT SERIALS
337 -0979 AND F33700001
Figure 8-1. Elevator Control System (Sheet 2 of 2)
8-5
9
. Detail
A
3
1. Control Column
2. Bracket
3. TurDbuckle
Elevator DOWN Cable
Elevator UP Cable
Swaged Ball
Clamp Block
8. Sleeve Weld Assembly
9. Bolt
4.
5.
6.
7.
•
Figure 8-2. Elevator Cable Routing of Control Column
NOTE
To ease routing of cables, a length of wire
may be attached to the end of cable being
withdrawn from aircraft. Leave wire in
place, routed through structure; then attach the cable being installed and use wire
to pull the cable into position.
8. Reverse the preceding steps for reinstallation and install pulleys and cable guards. Ensure
cables are positioned in pulley grooves before installing guards.
9. Re-rig elevator and rudder control systems
in accordance with paragraphs 8-12 and 9-16 respectively, safety turnbuckles and reinstall allltems removed for access.
b. AFT CABLES.
1. (Refer to figure 8-1.) Remove access plates
from lower left vertical fin as necessary to expose
Detail E.
8-6
2. Remove left wing strut fairings as necessary
to expose Detail D and turnbuckles (8).
3. Remove safety Wire, relieve cable tension
and disconnect turnbuckles (8).
4. (Refer to figure 9-1.) Remove safety wire
and relieve rudder control system cable tension at
turnbuckle (9).
5. (Refer to figure 8-1.) Mark or tag cables
and pulleys in Detail D and remove bolt securing
pulleys (4, 5 and 6) to bracket (2).
6. Disconnect cables (18 and 22) at bellcrank
(17).
NOTE
To ease routing of cables, a length of wire
may be attached to the end of cable being
withdrawn from aircraft. Leave wire in
place, routed through structure; then attach
the cable being installed and use wire to pull
the cable into position.
•
•
THRU AIRCRAFT SERIAL 337 -0239
3
5
3
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
Spacer
Support
Bearing
Link
Control Column
Sleeve Weld Assembly
Arm
Bob-Weight
Shield
Bellcrank Assembly
Bearing Block
Link
Guard
7
8
5
NOTE
Rig elevator control system prior
to rigging bob-weight.
12
8
AIRCRAFT SERIALS 337 -0240 THRU 337 -0755
•
Adjust links (4) equally so forward
top corner of bob-weight (8) is approximately even with top edge of
support (2) when elevator is in the
full up position .
Figure 8-3. Elevator Bob-Weight Installation
8-7
THRU AIRCRAFT
SERIAL 337 -0978
*Torque to 20 - 28 lb-in.
•
REFER TO FIGURE 8-7
0979
A~
F33700001
Detail
A
*
1.
2.
3.
4.
5.
6.
Balance Weight Arm
Balance Weight
Trim Tab
Elevator
Hinge Bracket (TYP)
Mounting Bolt
•
Figure 8-4. Elevator Installation
7. Reverse the preceding steps for reinstallation
and install pulleys.
8. Re- rig elevator and rudder control systems
in accordance with paragraphs 8-12 and 9-16 respectively, safety turnbuckles and reinstall all items removed for access.
1.
DOWN
2.
DOWN
ward.
Loosen turnbuckle (7) and tighten
cable turnbuckle to move copilot's
Tighten turnbuckle (7) and loosen
cable turnbuckle to move copilot's
elevator
wheel aft.
elevator
wheel for-
NOTE
8-12. RIGGING. (Refer to figure 8-1.)
a. Remove access plates from left vertical fin as
necessary to expose Detail E.
b. Remove left wing strut fairings as necessary to
expose turnbuckles (8).
.
c. Lock pilot's control colwnn in neutral poSition
in accordance with instructions in figure 8- 5.
d. Adjust push-pull tube (15) to dimension specified in figure 8-1, tighten jam nuts and connect tube
to balance weight arm (12).
e. Streamline elevator and stabilizer to set elevator in neutral position •.
f. Adjust turnbuckle forward of the instrwnent
panel and turnbuckles at the left wing strut to position the copilot's control Wheel the same distance
from the instrument panel as the pilot's control wheel
and also to obtain the proper cable tension as follows:
8-8
When dual controls are not installed, the ball
ends swaged on the elevator cables should be
used as reference points during rigging sequence.
g. Readjust push-pull tube (15), if necessary, to
streamline elevator in neutral position and tighten
jam nuts.
h. Mount an inclinometer on trailing edge of elevator and set at 0 0 with elevator in neutral position.
i. Remove control lock or neutral rigging tool from
pilot's control colwnn and adjust travel stops (14 and
20) to obtain degree of travel specified in figure 1-1.
With the left balance weight arm (12) resting on the
up-stop (14), adjust-the overtravel stop in the right
vertical fin 1/16" from right balance weight arm.
•
•
AmCRAFT SERIALS 337-0756
AND F33700001 THRU 33701398
AND F33700045
,
BEGINNING WITH AIRCRAFT
SERIALS 33701399 AND F33700046
3
2 4
Collar
Neutral Rigging Tool
Instrument Panel
Control Wheel Tube
Decorative Collar
Control Wheel
Cover
1.
2.
3.
4.
5.
6.
7.
7
Fabricate from 1/4 " steel rod.
NOTE
Thru aircraft serial 337 -0755, installation of the control
lock positions the control column in the neutral pOSition.
Beginning with aircraft serials 337 -0756 and F33700001
the control lock hole is moved farther aft in the control
wheel tube and installation of the control lock results in
a nose down attitude .
•
Figure 8-5. Control Column Neutral Rigging Tool
NOTE
An inclinometer for measuring control surface travel is available from the Cessna
Service Parts Center. Refer to figure 6-4.
j. Safety turnbuckles and reinstall all items removed for access.
IWARNING'
Be sure elevator moves in the correct direction when operated with the control wheels.
8-13. ELEVATOR TRIM CONTROL SYSTEM.
fer to figure 8-6.)
(Re-
8-14. DESCRIPTION. The elevator trim tab, located on the right hand trailing edge of the elevator,
is controlled by a trim wheel mounted in the lower
center section of the instrument panel. Power to
operate the tab is transmitted from the trim control
wheel by means of roller chains, cables, an actuator
assembly, bellcrank and push-pull channel. A mechanical pointer adjacent to the control Wheel indicates tab pOSition. A "nose-up" setting results in a
tab-down pOSition. The small bellcrank, mounted
inside the elevator, links the actuator to the pushpull channel which operates the tab. The bellcrank
provides a differential rate of movement of the tab
to furnish more rapid movement to-and-from the tab
down (nose-up) position and slower movement to-andfrom the tab up (nose-down) poSition. Beginning with
aircraft serials 337-0526 and F33700001, an extra
length of control cable is installed in the trim tab up
cable, the cable stops are located farther aft in the
right tail boom and the flap/elevator trim interconnect control is attached farther aft. The extra cable
is installed in the tab up cable to facilitate installation
of the optional electric elevator trim control system.
'.
8-9
•
8-15. TROUBLE SHOOTING.
Due to remedy procedures in the following trouble shooting chart, it
may be necessary to re-rig the system. Refer to paragraph 8-26.
TROUBLE
.TRIM CONTROL WHEEL MOVES
WITH EXCESSIVE RESISTANCE.
LOST MOTION BETWEEN
CONTROL WHEEL AND
TRIM TAB.
TRIM INDICATOR FAILS TO
. INDICATE CORRECT TRIM
POSITION.
18-10
Change 1
PROBABLE CAUSE
REMEDY
Cable tension too high.
Check and adjust cable tension .
Pulleys binding or rubbing.
Check visually. Install cable s
correctly.
Cables not in place on pulleys.
Check routing. Install cables
correctly.
Trim tab hinge or linkage
binding.
Disconnect actuator and check
resistance to tab movement.
Check bearings in bellcrank
and tab arm. Lubricate or
replace hinge or linkage as
necessary.
Defective trim tab actuator.
Disconnect chain and linkage
from actuator and operate
actuator with fingers. Replace
defective actuator.
Rusty or excessively worn chain.
Check visually.
chain.
Replace rusty
Damaged or worn sprocket.
·Check visually.
Replace sprocket.
Bent sprocket shaft.
Check visually. Replace bent
sprocket shafts.
Chain guard rubbing chain.
Check visually.
Defective bearings at control
wheel shaft.
Lubricate bearings; replace if
defective.
Cable tension too low.
Check and adjust cable tension.
Broken pulley.
Replace defective pulley.
Cables not in place on pulleys.
Check visually. Install cables
correctly.
Worn trim tab actuator or linkage.
Check actuator for excessive
play. Move trailing edge of
trim tab and observe linkage.
Replace actuator or worn linkage.
Actuator attachment loose.
Check attachment. Secure
actuator properly.
•
Free chain guard.
Indicator incorrectly engaged
on wheel track.
Check visually. Reset indicator .
Indicator bent.
Check visually. Straighten or
replace indicator.
•
•
8-15. TROUBLE SHOOTING (Cont) .
TROUBLE
INCORRECT TRIM TAB TRAVEL.
Stop blocks loose or
incorrectly adjusted.
Rig system in accordance with
paragraph 8-26.
Flap/elevator trim
interconnect improperly
rigged.
Rig system in accordance
with paragraph 8-36.
Incorrect rigging.
Rig system in accordance with
paragraph 8-26.
8-16. TRIM TAB. (Refer to figure 8-7.)
8-17. REMOVAL AND INSTALLATION.
a. Disconnect push-pull channel (8) from arm (7)
on trim tab (6).
b. Remove safety wire securing hinge pin (5) at the
outboard end, defiect rudder to the right, pull pin out
and remove tab.
c. Reverse the preceding steps for reinstalia.t1on.
nicks and dents.
d. Check bearings (3'. screw (6) and screw assemblv (10) for excessive wear and scoriD!l.
Examine screw assembly (10) and screw (6) for
damaged threads or dirt particles that might Impair
smooth operauon.
r. Check sprockets (2) for broken. chipped and/or
worn teeth.
g. Check bearing (111 for snloothness of operation.
e.
8-18. TRIM TAB ACTUATOR. (Refer to figure
8-7A. )
•
8-18A. REMOVAL AND INSTALLATION.
a. Relieve tension on elevator trim control system by loosening an elevator trim cable turnbuckle
in right taU boom.
b. Loosen chain guard and disengage chain from
actuator sprocket.
c. Disconnect links from aft end of actuator screw.
d. Remove bolts, clamps and spacers securing actuator and remove actuator through access hole.
e. Reverse this procedure to install the actuator.
Rill the elevator trim tab system in accordance with
paragraph 8-12.
8-18B. DISASSEMBLY. CRefer to figure 8-7A.)
;\. Remuve chain guard (1) if nut pre\'iousl~' rCIllO\'cd
in par3~r3ph 8-l8A.
b. Usi~ suitable punch and hammer. l'emO\'e roll
pin securin~ sprocket (2) to screw (6): remuve sprllcket from screw.
c. Remove screw assembly (10) frolll actuatur.
d. RenlUve Ilroove pins (12) securinJ,! bearinl!s (31 at
ends of hOUSing (71.
e. LighU~' tap screw (6) in oppusite direction from
sprockEt end: remove bearin!: (31. packin!! (9) and
collar /41.
NOTE
Do not attempt to repair damaJ,!ed or worn
parts of the actuator assembl\·. Discard
all defective items and install new parts
during reassembly.
8-180. REASSEMBLY. (Refer to figure 8-7A.)
a. Always discard the followi"1% compunents :Inc!
install new parts durmg reassemblv: bea rinl!s / J l.
alll/:roove pins. packin~ (9) and nuts 1131.
b. During reassemblv. lubricate collars (4 I. Sf.: re\\
(6' and screw ~lssembly (10) :IS shown In Section ;2 nf
this manual.
c. Press sprocket (21 into end of screw 161. alll:.n
ptn holes In sprocket and screw. and Install pins.
d. Slip bearinc (3) and collar 141 on scrc\\' (61 and
slide them down against sprocket (21.
e. Ingert screw '(61. With assembled parts. into
housing (7). until beartn~ (3) is flush with end of
housing.
NOTE
When inserting screw (61 into housin~ (7),
locate sprocket (2) at end of housing which
is farthest away from groove for retaining
ring (8).
NOTE
NOTE
It is not necessary to remove ril1!!s (81.
•
REMEDY
PROBABLE CAUSE
8-18C. CLEANING, INSPECTION AND REPAm.
a. Do not remove bearing (11) from screw assembly
(10) unless replacement of bearinll is necessary.
b. Clean all components. except bearing (11' in
Stoddard solvent or equivalent.
c. Inspect all componentS for obvious indications of
dam;u::e. such as stripped threads. cracks. deep
Bearings e3l are not pre-drilled, and must
be drilled on assembly. Pins are 1, 16- inch
in diameter. therefore. requiring a 1/16
(0.0625) inch drill.
f.
With bearing (3' flush with end of housing (7).
drill bearmg so the drill will emerge from
hole on opposite side of housing (7),
c~lrefully
Change 1
B-ll
•
NOTE
I
Shaded pulleys are used
in this system only.
5--~
2
Detail
A
Detail
B
Detail
REFER TO FIGURE 8.,12 _ _--.
C
REFER TO FIGURE 8.,10
REFER TO FIGURE 8.,12
Detail
•
D
A
'.
REFER TO FIGURE 8.,8
NOTE
Locate turnbuckles of adjacent
cables so they do not meet,
cross or rub.
Refer to figure 4-2 for cable
routing through wing strut
fairleads.
1. Cable Guard
2. Bracket
3. Pulley (Tab Up Cable)
4. Pulley (Tab Down Cable)
5. Tab Up (Nose Down) Cable
6. Tab Down (Nose Up) Cable
7. Pulley (Rudder)
8. Turnbuckle (Tab Down Cable)
9. Turnbuckle (Tab Up Cable)
10. Clevis (Tab Up Cable)
11. Spacer
12.
13.
14.
15.
16.
17.
18.
19.
20.
Clevis (Tab Up Cable)
Clevis (Tab Down Cable)
Cover
Rub Block
Auxiliary Spar
Actuator
Clamp
Spacer
Roller Chain Sprocket
21. Roller Chain Guard
22. Stabilizer Rear Spar
f~~UTIONI
MAINTAIN PROPER CONTROL
CABLE TENSION.
CABLE TENSION:
20 ± 5 LBS (AT AVERAGE TEMPERATURE FOR THE AREA. )
REFER TO FIGURE 1-1 FOR TRAVEL.
Figure 8-6. Elevator Trim Tab Control System (Sheet 1 of 2)
8-12
•
•
14
~,
2~"~"
,
~~""
.
\
,'"
THRU
ts~~701462
RAFT
~:'~33700055
,~
,
1
3
4
4
WITH Am- ,
BEGINNING IALS 337 SERF33700056
CRAFT
01463 AND
,
Detail
E
1
,
2
•
- (JI'
4
~
Detail
F
.... WITH 2 CLAMPS (18)
•
23) 20 to 25 lb-in.,
Torque
bolt~'~e
and
apply
w 1 lacquer putty.
Detail
H
Figure 8-6. Elevator Tri m Tab Control System (Sheet 2 of 2)
Chauge 1
8-13
3
•
1C-U
4
I
"tlj~;:~:::i?=C:i~.h~-:r;>
----.------._.---JU{.~~l
•
8
2
•
THRU AmCRAFT SERIAL 337 -0978
3
4
\
] Jd
..
-- ------. ----
....
_----- ..~
BEGINNING WITH AmCRAFT SERIALS
337-0979 AND F33700001
•
7
FORCE
\lAXI~n:~l
DOWN ~ DEFLECTION
(FREE-PLA YI
-- -----___ --L
Multiply dimension A by 0 025 to delE'rllllne
maximum allowable freE'-play. Free-plav IS
measured at the left end of tri III tab.
Safety wire hinge pin (5) at
outer end of elevator.
:;
---.~
•
.Torque nut to 15 - 401b-in
or bolt head to 10 - 40 lb-in.
-~
-~;~A
jT
......... ' . :. _
... ,L.....
"
.........
-------1. Actuator Screw End 5. Hinge Pin
2. Link
6. Trim Tab
FORCE
3. Bellcrank
7. Arm
UP
4. Elevator
8. Push-Pull Channel
•
Figure 8-7. Elevator Trim Tab Linkage and Free-Play Inspection
NOTE
Do not
~
enlar~e
holes in housin!!:.
Press new groove pins (12 I into pin holes.
Insert collar (4 I. new packing (91 and bearing
(31 into opposite end of housing (71.
i. Conlplete steps 'T' and "g" for bearing (31 just
installed.
).
U new bearing 111 I is required. a new bearing
mav be pressed into boss. Be sure force bears
agaanst outer race of bearing.
k. Screw screw assembly (101 into screw (61.
I
Install retaininlr rinlts (81. if removed.
m. Test actuator assemblv bv rotatill\t sprocket (2 I
with finl!ers while holding screw assemblv (101.
Sc rew assemblv should travel in and out smoothlv,
WIth no indication of binding
.
h.
8-19. TRIM TAB FREE-PLAY INSPECTION.
• a. Placp plp.vator!: :lne! trim tab in neutral position.
b. Restram elevator. and manually de(1ect tab at
the trallin!!: edge at the point where the actuator pushpull rod lS located_
c. Deflect t:lb in one direction to the point of positive stop. :md measure the deflection from neutral.
8-14
Change 1
using the elevator surface as a. reierence.
d. Measure the deflectlon from neutral In the IJPPOsite direction.
e. The sum of the two deflections must nut eleceed
the result of the formula: Multiplv dimenSIOn '·A"
(refer to figure 8-4) by 0.025.
f. If the sum of the two deflections exceed the
figure attained from the formula. replace AN bolts
with NAS464 bolts of E'quivalent diameter :lnd j!rip
length in the push rod and recheck.
g. If this does not obtain desired results. reDlace
bearings in rod end and recheck.
h. If this does not obtain desired results. replace
trim tab horn bearing and recheck.
L U this does not obtain desired results. overhaul
or replace trim tab actuator and insure that all areas
are properly saflied.
8-20. TRIM TAB BELLCRANK. (Refer to figure
8-7. )
8-21. REMOVAL AND INSTALLATION.
a. Remove access plate below bellcrank (3).
b. Disconnect push-pull channel (8) from aft end
of bellcrank.
c. Disconnect links (2) from forward end of bell-
•
11
&
2
•
1. Guard
2. Sprocket
3. Bearing
4. Collar
5. Ring
6. Screw
7. Housi~
8. Ring
9. Packing
10. Screw Assemblv
11. Gearin!:
12. Groo\'e Pin
* Lubricate collars (4) and screw
housing (7) as shown in Section
2 of this Manual.
12
Figure 8-7A. Elevator Trim Tab Actuator Assembly
crank. Secure links (2) and trim control wheel so
they cannot be turned to maintain control system
rigging.
d. Remove bellcrank pivot bolt and remove bellcrank thro\lllth access opening.
e. Reverse the preceding steps for reinstallation.
8-22. TRIM TAB CONTROL WHEEL. (Refer to
figure 8-8.)
•
8-23. REMOVAL AND INSTALLATION.
a. Disconnect battery cables and insulate terminals
as a safety precaution.
b. Remove access plates from right tall boom as
necessary to expose turnbuckles (index 8 or 9,
figure 8-6), remove safety wire and relieve cable
tension.
c. Remove switch mounting nuts, switches, etc.
as necessary to remove covers from left side of
instrument panel.
d. Remove pin (14) and washer (12) securing trim
wheel shaft to support bracket (13).
e, Remove screws securing support bracket (5) to
instrument panel structure and move control wheel
(1) outboard. Remove spacer (9) from shaft and disengage chain (7) from sprocket (8).
f. Remove control wheel (1), bracket (5) and indicator (3) as an assembly. Position indicator (3) may
be removed from assembly after removal from the
aircraft.
g. Reverse the preceding steps for reinstallation.
Rig system in accordance with paragraph 8-26 and
reinstall all items removed for access.
8-24. CABLES AND PULLEYS.
8-25. REMOVAL AND INSTALLATION.
a. FORWARD CABLE.
1. (Refer to figure 8-6.) Remove copilot's seat.
carpeting and access plates in floorboard area as
necessary to expose Details C, D and E.
2. Remove right wing strut fairings as necessary
to expose Detail F and clevises (12 and 13).
3. Remove access plates from inboard side of
right tall boom as necessary to expose turnbuckles
(8 and 9).
4. Remove safety wire and relieve cable tenSIon
from either turnbuckle (8 or 9).
Change 1
8-14A/C8-14B Blank)
•
NOTE
*t,~,_
.
Thru aircraft serial 337 -0239. the elevator
trim wheel cover is an integral part of the
console cover. Refer to Section 10 for removal procedures. Beginning with aircraft
serials 337 -0240 and F33700001. the trim
wheel cover is independent of the console
cover and extends left to cover the switch
panel. Applicable knobs and controls must
be removed before the cover can be removed.
THRU AmCRAFT SERIALS _____ /
33701462 AND F33700055 l~
~
/.
*RIVET END OF PIN (2) THROUGH
SUPPORT BRACKET (5). DRILL
OUT PIN TO REMOVE INDICATOR
1
,.~;~
*2
L':flJ....
4'
:/~·i 11
;':: :~&~
~~1. .:';:, , .
~.l·
~ ....
•
?
'V/?<i..·
/
5
11
~&
/..
BEGINNING WITH AmCRAFT (
SERIALS 33701463 AND F337 - . -..;:
00056
~\
~,~
~/
•
)
7
,
/" / '
13
'-':.:"--...,
' / , . . . / \3
"J..,-G...f.;..o..
...
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
Trim Control Wheel
Pin
Position Indicator
Screw
Support Bracket
Roll Pin
Roller Chain
Sprocket
Spacer
Screw
Bearing
Washer
Support Bracket
Pin
Figure 8-8. Elevator Trim Control Wheel Installation
5. (Refer to figure 9-1.) Remove safety wire
and relieve rudder control system cable tension at
turnbuckle (8).
6. (Refer to figure 8-6.) Disconnect cables at
clevises (12 and 13).
7. Disengage roller chain from trim control
Wheel sprocket at instrument panel.
8. Mark or tab cables and pulleys in Details C
and E and remove bolts securing pulleys (3, 4 and
7) ~o brackets (2).
9. Remove cable guards from Details A, B and
D as necessary to work cables free of aircraft.
NOTE
•
To ease routing of cable, a length of Wire may
be attached to the end of the cable being withdrawn from the aircraft. Leave wire in place,
routed through structure; then attach the cable
being installed and pull the cable into position.
10. Reverse the preceding steps for reinstallation and install pulleys and cable guards. Ensure
cables are positioned in pulley grooves before installing guards.
11. Re-rig elevator trim and rudder control systems in accordance with paragraphs 8- 26 and 9-16
respectively, safety turnbuckles and reinstall all
items removed for access.
b. AFT CABLE ..
1. (Refer to figure 8-6.) Remove access plates
from inboard side of right tail boom as necessary to
expose turnbuckles (8 and 9).
2. Remove access plates from lower right vertical fin and stabilizer as necessary to expose Details
G and H.
3. Remove safety wire, relieve cable tension and
disconnect turnbuckles (8 and 9), leaving the turnbuckle barrelS on the forward cables.
4. (Refer to figure 8-12.) Remove safety wire.
remove screws and remove travel stop block (8) from
cable (7).
Change 1
5. Disconnect nap/elevator trim control assembly (5) at clamp (9) and remove clamp.
6. (Refer to figure 8-6.) Remove chain guard
(21) and disengage roller chain from sprocket (20).
7. Remove cable guards from Detail G as necessary to work cable free of aircraft.
h. Re-rig elevator trim and rudder control
system in accordance with paragraphs 8- 26 and 9-16
respectively, safety turnbuckles and reinstall all
items removed for access.
2. AFT SECTION.
•
NOTE
NOTE
H electric trim assist is installed, refer to
To ease routing of cable, a length of wire may
be attached to the end of the cable being withdrawn from the aircraft. Leave wire in place,
routed through structure; then attach the cable
being installed and pull the cable into position.
8. Reverse the preceding steps for reinstallation and install cable guards. Ensure cables are
positioned in pulley grooves before instamng guards.
9. Re-rig trim and interconnect systems in accordance with paragraphs 8-26 and 8-36 respectively,
safety turnbuckles and travel stop screws and reinstall all items removed for access.
c. CENTER CABLE-TAB UP.
paragraph 8-30 for removal of this cable.
a. (Refer to figure 8-10.) Remove access
plates from inborad side of right tail boom as necessary to expose Detail A.
b. Remove safety wire, relieve cable tension and disconnect turnbuckle (7), leaving barrel
attached to cable (8).
c. Disconnect cables (28 and 29) at clevis.
d. Remove safety Wire, remove screws
securing travel stop blocks (5 and 30) to cable (29)
and remove stop blocks.
e. Remove cable (29) from aircraft.
NOTE
NOTE
Beginning with aircraft serials 337-0526
and F33700001, the center TAB- UP
cable consists of two sections. The forward section begins at clevis (12) and
ends at clevis (10). The aft section which
is replaced with the electric trim servo
cable when the electric trim installation
is installed, begins at clevis (10) and ends
at turnbuckle (9).
1. FORWARD SECTION.
a. (Refer to figure 8-6.) Remove access
plates from inboard side of right tail boom as necessary to expose clevis (10) and turnbuckle (9).
b. Remove right wing strut fairings as
necessary to expose Detail F and clevis (12).
c. Remove safety wire and relieve cable
tension at turnbuckle (9).
d. (Refer to figure 9-1.) Remove safety
wire and relieve rudder control system cable tension at turnbuckle (8).
e. (Refer to figure 8-6.) Disconnect
clevises (10 and 12).
f. Mark or tag cables and pulleys in Detail
F and remove bolt securing pulleys (3, 4 and 7) to
bracket (2).
NOTE
To ease routing of cable, a length of wire
may be attached to the end of the cable
being withdrawn from the aircraft. Leave
wire in place, routed through structure;
then attach the cable being installed and
pull the cable into position.
g. Reverse the preceding steps for reinstallation and install pulleys and cable guards. Ensure cables are installed in pulley grooves before installing guards.
8-16
To ease routing of cable, a length of wire
may be attached to the end of the cable
being withdrawn from the aircraft. Leave
wire in place, routed through structure;
then attach the cable being installed and
pull the cable into position.
f. After cable is routed in pOSition, re-rig
trim system in accordance with paragraph 8-26,
safety turnbuckle (7) and reinstall all items removed
for access.
d. CENTER CABLE - TAB DOWN.
1. (Refer to figure 8-6.) Remove access plates
from inboard side of right tail boom as necessary to
expose turnbuckle (8).
2. Remove right wing strut fairings as necessary
to expose Detail F and clevis (13).
3. Remove safety wire, relieve cable tension
and disconnect turnbuckle (8) leaving the turnbuckle
barrel attached to the aft cable.
4. (Refer to figure 9-1.) Remove safety wire
and relieve rudder control system cable tension at
turnbuckle (8).
5. (Refer to figure 8-6.) Disconnect clevis (13).
6. Mark or tag cables and pulleys in Detail F
and remove bolt securing pulleys (3, 4 and 7) to bracket (2).
•
NOTE
To ease routing of cable, a length of wire may
be attached to the end of the cable being withdrawn from the aircraft. Leave wire in place,
routed through structure; then attach the cable
being installed and pull the cable into position.
7. Reverse the preceding steps for reinstallation and install pulleys and cable guards. Ensure
cables are installed 'in pulley grooves before installing guards.
•
•
*Used when electric trim is NOT installed•
TAB-DOWN
RESTRICTED
POSITION
FLAP INTERCONNECT)
INTERCONNECT
CONTROL
• Used when electric trim IS installed.
STA. 110.50
TAB-UP
CABLE
TAB-DOWN
CABLE
TAB-UP
STOP
STANDARD SYSTEM
THRU AIRCRAFT
SERIAL 337 -0525
,
TAB DOWN
RESTRICTED
POSITION STOP
(FLAP INTERCONNECT)
INTERCONNECT CONTROL
STA. 110.50
o
,.FWD
o
•
TAB-OOWN / '
STOP:::::..:/"
* BRACKET
¢
...
STANDARD SYSTEM
BEGINNING WITH AmCRAFT SERIALS 3370526 AND F33700001
Trim cable is free to pass
back and forth through bushing inside clamp.
NOTE
Safety wire screws on travel stop blocks.
Install stop blocks with rounded corners
towards the cable and block assemblies
perpendicular to the cable.
TAB-OOWN
RESTRICTED
POSITION STOP
(FLAP INTERCONNECT)
TAB-UP
STOP
TAB-OOWN
STOP
INTERCONNECTC~~ROLJ7
STA. 117.50
""""----1
• GUIDE ASSEMBLY
•
,.FWD
D [Q]
I=----f
o
..
ELECTRIC TRIM
BEGINNING WITH AmCRAFT SERIALS 3370526 AND F33700001
Figure 8-9. Elevator Trim Travel Stops
8-17
•
NOTE
.Use as required to remove
end play and ensure positive
connection between pin on
clutch (23) and stop assembly
(25). Stop assembly (25)
should be deflected approximately • 021 " when clutch is
installed.
The drum gJ,"oove and cable
must be free of grease and oil.
Detail
A
BEGINNING WITH AIRCRAFT SERIAL 337 -0526
• Safety wire screws on
travel stop blocks.
,,
'.
NOTE
Install stop blocks with rounded corners
towards the cable and block assemblies
perpendicular to the cable.
The stop blocks (5 and 30) contact the
housing covers (4) at travel extremes.
Roll Pin
Support Assembly
Sprocket
Housing Cover
Tab Up stop Block
Grommet
Turnbuckle (Tab Up Cable)
Tab Up Cable (Aft)
Guide Assembly
Motor Support
Roller Chain
Motor Assembly
Motor Cover
Clutch Cover
Bearing
Washer
Bushing
Spanner Nut
Washer
Friction Washer
Drive Drum
Shaft
Clutch Assembly
Mounting Bolt
Clutch Stop
Tab Down Cable
Rudder Cable
Tab Up Cable (Fwd)
Tab Up Cable (Center)
Tab Down Stop Block
Housing
Figure 8-10. Electric Elevator Trim Control System
8-18
•
•
•
SOCKET WITH 3/8" OR 1/4 " DRIVE _ _ _ _ _.-----.
TO ACCEPl' TORQUE WRENCH
I
BRAZE SOCKET TO PLATE - - - - - - - -
...;;+I11uJJ.~-4=I-t_t
-t
3/16 " STEEL
(4130 NORMALIZED OR
EQUIVALENT)
AN122693 PIN (2) OR EQUIVALENT _ - - . J
(SPACE TO MATCH CLUTCH SPROCKET)
1-3/8 " DIAMETER
TOP VIEW
TORQUE
WRENCH
ADAPTER
SPRING SCALE
(FISH SCALE)
.251 " HOLE (CLEARANCE FOR SHAFT)
•
Figure 8-11. Electric Trim Servo Adjustment Tools
8. Re-rig elevator trim and rudder control systems in accordance with paragraphs 8- 26 and 9-16
respectively, safety turnbuckles and reinstall all
items removed for access.
8-26. RIGGING.
NOTE
The elevator trim and flap system are interconnected, therefore, the flaps must be in the DOWN position while rigging the trim
control system.
g. Rotate trim wheel to full nose down (tab up) position, then back 1 1/4 turns (approximate neutral
position).
h. (Refer to figure 8-7.) Place elevator and trim
tab both in neutral (streamlined) position. (Refer to
figure 8-5.) Adjust actuator screw end (1) OUT or
IN as necessary to align with links (2) and install
bolt.
1. Mount an inclinometer on trailing edge of trim
tab and check tab for sufficient travel as specified
in figure 1-1. If travel is insufficient in either direction, readjust actuator screw end (1).
NOTE
•
a. Remove access plates from inboard side of right
tail boom.
b. (Refer to figure 8-6.) Remove safety wire and
relieve cable tension at turnbuckles (8 and 9).
c. Remove safety wire and loosen screws securing
travel stop blocks (index 5 and 30, figure 8-10).
d. (Refer to figure 8-7.) Disconnect actuator screw
end (1) at links (2).
e. (Refer to figure 8-8.) Rotate trim wheel (1) to
the mid-range poSition. Check that roller chain (7)
ends extend the same distance from sprocket (8). If
necessary, disengage roller chain and re-engage
chain on sprocket.
f. (Refer to figure 8-6.) Adjust turnbuckles (8 and
9) evenly to proper tension and safety.
An inclinometer for measuring control sur-
face travel is aVailable from the Cessna
Service Parts Center. Refer to figure 6-4.
j. (Refer to figure 8-9.) Rotate trim wheel to position tab at specified UP travel, slide tab up-stop block
on cable against stop bracket, secure stop and safety
wire screws.
NOTE
When electric trim assist is installed the
stop blocks will strike the housing covers
as a stop at travel extremes.
8-19
k. Rotate trim wheel to po siUon tab at specified
DOWN travel, slide tab down stop block on cable
against bracket, secure stop and safety wire screws.
1. Check and rig interconnect system in accordance with paragraph 8-36, if necessary.
m. Reinstall all items removed for access.
8-28. DESCRIPTION. Beginning with aircraft serial
337-0526, an electric elevator trim assist may be
installed. This system is operated by a switch on the
left side of the pilot's control wheel. The servo unit,
installed in the right tail boom, includes a motor and
an adjustable chain-driven, solenoid-operated clutch.
A section of the trim tab UP cable is removed and replaced with the servo cable which enters the housing
and double wraps around a drive drum. This drum
is secured to and driven by the clutch. When the
clutch is not energized, the drive drum "free wheels"
so that manual operation of the trim system is not
affected. In case of malfunction, the manual system
and interconnect system will override the servo
clutch.
IWARNING'
Be sure trim tab moves in the correct direction when operated with control wheel.
8-27. ELECTRIC TRIM ASSIST INSTALLATION.
(Refer to figure 8-10. )
•
8-29. TROUBLE SHOOTING.
NOTE
When de-actuated, the electric trim system should not affect
the manual system; therefore, the standard trouble shooting
chart also applies to the electric trim system. The remedy
procedures in the following trouble shooting chart may require
re-rigging of trim system, refer to paragraph 8-26.
TROUBLE
SYSTEM INOPERATIVE.
TRIM MOTOR OPERATING TRIM TAB FAILS TO MOVE.
PROBABLE CAUSE
Circuit breaker out.
Check visually.
Defective Circuit breaker.
Check continuity. Replace breaker.
Defective Wiring.
Check continuity. Repair wiring.
Defective trim switch.
Check continuity. Replace switch.
Defective trim motor.
Remove and bench test.
motor.
Defective clutch solenoid.
Check continuity. Replace solenoid
Improperly adjusted clutch
tension.
Adjust tension in accordance
with paragraph 8-31.
Disconnected or broken
cable.
Check continuity. Connect or
replace cable.
Defective actuator.
Check actuator operation.
Replace actuator.
8-30. REMOVAL AND INSTALLATION. (Refer to
figure 8-10.)
a. Remove access plates from inboard side of
right tail boom as required.
b. Remove safety wire, relieve cable tension and
disconnect cable from turnbuckle (7), leaving barrel
attached to cable (8). Slide cable (29) out through
grommet (6) in aft guide (9).
8-20
REMEDY
Reset breaker.
•
Replace
c. Disconnect cables (28 and 29) at cleviS.
d. Remove screws securing forward guide (9) to
forward housing (4).
e. Disconnect ALL electrical wiring from trim
unit.
f. Remove safety Wire, remove screws securing
stop block (30) to cable (29) and remove stop block.
Slide cable (29) out through grommet (6) in forward
guide (9).
•
•
g. Remove mounting bolts (24) and remove unit
from aircraft •
h. Reverse the preceding steps for reinstallation.
Rig trim system in accordance with paragraph 8-26,
safety wire all items previously safetied and reinstall all items removed for access.
8-31. CLUTCH ADJUSTMENT. (Refer to figure
8-10.)
a. SERVO UNIT REMOVED FROM THE AmCRAFT
BUT STILL INSTALLED IN THE HOUSING.
1. Remove servo unit from aircraft in accordance with paragraph 8-30.
2. Remove forward housing cover (4) to gain
access to clutch assembly.
3. Loosen outside locking sparmer nut (18) so
that tension can be adjusted with inside spanner nut.
4. Connect spring scale (fish scale) to forward
end of cable (29). (Refer to figure 8-11 for spring
scale. )
5. Energize clutch assembly using a 24-volt
power source.
6. Hold opposite end of cable (29) to prevent
slippage of cable on drum (21).
7. Pull cable (29) with spring scale until clutch
slips, noting pounds required to slip clutch.
8. Adjust inside spanner nut (18) until clutch
slips at 28 to 32 lbs tension. Tighten outside locking
•
8-32. FLAP/ELEVATOR TmM INTERCONNECT
SYSTEM. (Refer to figure 8-12. )
8-33. DEScmPTION. The flap/elevator trim interconnect system restricts the amount of nose up trim
available with the flaps up. As the flaps are raised
from the full down position, the interconnect system
automatically removes full nose up trim to a restricted poSition.
8- 34. TROUBLE SHOOTING .
NOTE
The flap control system and elevator trim control system must
be correctly rigged to ensure proper operation of the interconnect system.
TROUBLE
INTERCONNECT DOES
NOT MOVE TRIM TAB
FROM FULL DOWN POSITION AS FLAPS ARE
RAISED.
•
spanner nut against inside nut.
b. CLUTCH ASSEMBLY REMOVED FROM HOUSING.
1. Loosen outside locking spanner nut (18) so
that tension can be adjusted with inside spanner nut.
2. Clamp clutch assembly in a vise at the drum
(21) with sprocket (3) in the UP position.
3. Energize clutch assembly using a 24-volt
power source.
4. Connect torque wrench (lb-in) and adapter
over shaft on sprocket (3) so the pins of the adapter
engage between teeth of sprocket. (Refer to figure
8-11 for adapter. )
5. Apply torque to clutch assembly noting tension required to slip clutch.
6. Adjust inside spanner nut (18) until clutch
slips at 25±3 lb-in. Tighten outside locking spanner
nut against inside nut.
INTERCONNECT DOES
NOT MOVE TRIM TAB
FAR ENOUGH .
PROBABLE CAUSE
REMEDY
Disconnected or broken interconnect control.
Check visually. Connect control;
replace if broken.
Control casing not secured
to structure.
Check security of attaching clamps.
Position control casing and tighten
clamps.
Trim tab up stop loose or
improperly located.
Check stop for security and proper
location. Locate stop for proper
tab travel and tighten.
Interconnect control attached
around wrong trim cable.
Check visually. Attach around
tab up cable in proper position.
Control not rigged correctly.
Rig in accordance with paragraph
8-36.
Control casing slipping in
clamps.
Check visually. Position control
casing and tighten clamps.
Control not rigged correctly.
Rig in accordance with paragraph
8-36.
8-21
•
o Tighten nut until it bottoms
out, bolt must turn freely.
eTHRU AIRCRAFT
SERIAL 337 -0020
OBEGINNING WITH
AIRCRAFT SERIAL /
337-0021
I
I
I
i~O
i f..
/
2
I
I
'-
J
3~' ~_5
.
~
;
I
I
"
A< ~'5
/
/ '"
'/
A
DeWl
~~----~/
A
,
10
THRU AIRCRAFT' ; '
'SERIAL 337 -0239 /
----~
,
3-~·
!
..
1
/
/
*
Detail
B
: Bend wire around clamp-bolt
after routing through bolt.
//'
.05" MAX.
BEGINNING WITH AIRCRAFT"
5
SERIALS 337 -0240 AND F33700001
4
ONE COMPLETE
TURN (MIN)
'.
VIEW
•
A-A
Flap Push-Pull Rod
Bellcrank
Synchronizing Push- Pull Tube
Bracket
Interconnect Control
Clamp
Trim Tab Up Cable
Travel Stop Block
Clamp
Bushing
Spacer
.. Safety wire screws to each other.
* Install bolt with head down.
'* Clamps are identical at both ends of control
Figure 8-12. Flap/Elevator Trim Interconnect System
8-22
(5).
•
•
8-35. REMOVAL AND INSTALLATION. (Refer to
figure 8-12.)
a. Remove access plates from inboard side of right
tail boom.
b. Run flaps to DOWN position.
c. Remove flap well gap seal panel and access plate
at right outboard flap, inboard bellcrank (Detail A).
d. Disconnect control wire at clamp (9).
e. Remove bolts securing bracket (4) to bellcrank
(2).
f. Remove clamps (6) securing control assembly
(5) to aircraft structure.
g. Tie a guide wire to the aft end of control assembly (5) and pull control out through bellcrank access
opening. Leave guide wire in place to aid in reinstallation of control assembly.
NOTE
H a new control wire is to be installed in
casing, bend the forward end of wire as
illustrated in figure 8-12, lubricate wire
with MIL-G-23827, slide wire through
washer and bracket (4) and insert wire
into casing.
•
h. Using guide wire pull control assembly through
structure, iTl place and disconnect guide wire.
i. Secure control casing in clamps (6) with approximately I-inch extending beyond clamp at each end.
j. Secure bracket (4) to bellcrank (2).
k. Pull aft on control wire to remove slack, rig
system in accordance with paragraph 8-36, bend wire
180 around clamp bolt before tightening bolt and
reinstall all items removed for access.
0
IWARNING'
Do nor reuse the wire inside control casing
if it has been removed by straightening the
ends or bent severely and then straightened.
The wire becomes brittle and will break
from work hardening.
8-36. RIGGING.
(Refer to figure 8-12.)
NOTE
The following rigging procedure should be
completed ONLY if the flap and elevator
trim control systems are properly rigged
•
and if an interconnect control assembly
has been installed in accordance with
paragraph 8-35.
a. Loosen bolt securing clamp (6) at aft end of
control (5).
b. Raise flaps to full UP position.
c. Place elevator in neutral (streamlined) position.
(Refer to figure 8-5.)
d. Rotate trim control wheel to place trim tab in
neutral position (streamlined with elevator), mount
an inclinometer on tab and adjust to 0 0 •
NOTE
An inclinometer for measuring control surface
travel is available from the Cessna Service
Parts Center. Refer to figure 6-4.
IWARNING'
Do not use the wire inside control casing if
the ends have been straightened and then
rebent, or if the wire has been bent severely
and restraightened. The wire becomes
brittle and will break.
e. Pull aft on control wire to remove slack, then
slide control assembly (5) through clamp (6) either
forward or aft to position clamp (9) firmly against
the restricted position stop (8). Refer to figure 8-9
for minimum position of stop (8). Tighten bolt
securing clamp (6) .
f. If binding occurs after initial installation or
during service use, an effort should be made to
relieve this condition by realignment or by repositioning the assembly through the aircraft structure
rather than by removing the control wire from casing.
g. Check that full elevator trim tab travel can still
be obtained with flaps in the DOWN position. Check
that the tab moves from the full DOWN pOSition to the
restricted position when the flaps are raised. Refer
to figure 1-1 for specified travel.
NOTE
Trim tab travel is not restricted until the
flaps are raised from full down to approximately the 2/3 down position. From 2/3
down to full up position, the trim tab is
gradually restricted to degree specified in
figure 1-1.
8-23/(8-24 blank)
SECTION 9
••
RUDDER AND RUDDER TRIM CONTROL SYSTEMS
TABLE OF CONTENTS
•
RUDDER CONTROL SYSTEM
Description . . . . . .
Trouble Shooting . . . .
Rudder Pedal Assembly.
Removal and Installation
Repair . . . . . . . .
Rudders . . . . . . . . . .
Removal and Installation
Repair . . . . . . . .
Bellcranks . . . . . . . .
Removal and Installation
Rudder Bungee. . . . . . .
Removal and Installation
Page
9-1
9-1
9-2
9-3
9-3
9-3
9-3
9-3
9-3
9-3
9-3
9-3
9-5
9~5
9-7
9-10
9-10
9-10
9-13
9-13
9-13
9-13
9-13
9-3
9-1. RUDDER CONTROL SYSTEM.
9-2. DESCRIPTION. The rudder control system
consists of the rudder pedal installation, cables,
pulleys, push-pull rods and rudder beUcranks. The
•
Cables and Pulleys . . • . . . . .
Removal and Installation . . .
Rigging - Rudder, Rudder Trim and
Nose Wheel Steering Systems
RUDDER TRIM CONTROL SYSTEM
Description . . . . . • . .
Trouble Shooting . . . . . .
Trim Control Wheel . . • .
Removal and Installation
Rigging . . . . . . . . . .
Console and Quadrant Covers
Removal and Installation
rudder bars are connected to the forward rudder
bellcrank by a push-pull rod and rudder trim actuator. Nose gear steering is controlled by the rudder
pedals through a bungee, beUcrank and push-pull
rod.
9-1
•
9-3. TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 9-16.
TROUBLE
PROBABLE CAUSE
REMEDY
RUDDERS DO NOT RESPOND TO PEDAL MOVEMENT.
Broken or disconnected cables
or push-pull rods.
Check visually. Connect cables
and push-pull rods. Replace if
broken.
BINDING OR JUMPY
MOVEMENT OF RUDDER
PEDALS.
Incorrect cable tension.
Check and adjust cable tension.
Cables not routed properly on
pulleys.
Check cable routing. Route
cables properly.
Defective pulleys or cable
guards.
Check visually. Replace
defective parts and install
guards properly.
Rudder bars binding.
Visually inspect rudder bars.
Install bearing blocks properly
and lubricate bearing surfaces.
Replace defective parts.
Defective rudder hinge bearing or bell crank bearings.
Replace defective bearings.
Clevis bolts too tight.
Readjust to eliminate binding.
Incorrect rigging.
Rig in accordance with paragraph
9-16.
Defective rudder trim bungee.
Disconnect bungee and check
operation of rudder system.
Replace defective bungee.
Defective nose gear.
Disconnect bungee and check
nose gear manually. Repair
or replace nose gear.
Incorrect rigging.
Rig in accordance with paragraph
9-16.
Bent push-pull rods.
Check visually. Replace
push-pull rods.
Weak or binding bungee.
Improperly rigged bungee.
Friction in rudder system.
Repair or replace bungee. Re-rig
bungee in accordance with paragraph 9-16. Check cable tension.
Check for correct installation and
routing of cables.
Rudder trim system.
Check riggin~ of trim system in
accordance with paragraph 9-16.
RUDDER TRAVEL INCORRECT.
RUDDER PEDALS DO NOT
RETURN TO NEUTRAL.
9-2
•
•
•
9-4. RUDDER PEDAL ASSEMBLY .
9-5. REMOVAL AND INSTALLATION.
a. Remove lower section of control quadrant cover.
b. (Refer to figure 9-5.) Remove safety Wire,
relieve chain tension at either turnbuckle (6) and
disconnect rod end (23) at rudder bar arm (21). DO
NOT TURN ROD END.
c. (Refer to figure 9-2.) Disconnect push-pull rod
(18) at rudder bar arm.
d. Discc;»nnect steering bungee (6) at rudder bar
arm.
e. Disconnect master cylinders (7) at rudder bar
arms (1).
f. Remove bolts securing bearing blocks (4).
g. Carefully work rudder bars down and aft to
remove.
NOTE
If additional clearance is desired, depending
on the equipment installed, complete step
"h."
•
h. Disconnect pedal supports (17) and brake links
(16) at rudder bars.
i. Reverse the preceding steps for reinstallation.
If the trim actuator rod end was not turned, rerigging should not be necessary, although it is advisable to check for proper rudder travel and tension.
j. Rig rudder trim system, if necessary, in accordance with paragraph 9-16, safety turnbuckle and reinstall all items removed for access.
9-6. REPAffi. Repair of rudder bar assemblies consists of attaching parts replacement as necessary.
Lubricate as outlined in Section 2.
9-7. RUDDERS.
9-8. REMOVAL AND INSTALLATION. (Refer to
figure 9-3.)
a. Remove access plate from top of stabilizer adjacent to vertical fin to expose rudder bellcrank
(index 12, figure 9-1).
b. Disconnect push-pull rod at bellcrank (index 12,
figure 9-1).
c. Remove hinge bolts (5) and carefully work the
lower end of rudder inboard as the upper end of rudder is worked outboard until rudder clears the vertical fin structure, then work rudder inboard and aft
until push-pull rod and arm assembly (7) clears
vertical fin.
NOTE
If additional clearance is required, rudder
tip (2) and weight assembly (1) and its bracket may be removed.
d. Reverse the preceding steps for reinstallation.
•
If adjustment of push-pull rod was not disturbed, re-
rigging of system should not be necessary. Rig system, if necessary, in accordance with paragraph
9-16 and reinstall all items removed for access.
9-9. REPAIR. Repair may be accomplished as outlined in Section 16.
9-10. BELLCRANKS.
9-11. REMOVAL AND INSTALLATION.
a. FORWARD. (Refer to figure 9- 5.)
1. Remove wing strut fairings as necessary to
expose turnbuckle (index 8 or 9, figure 9-1).
2. Remove safety wire and relieve cable tension
at turnbuckle.
3. Disconnect cable (3) at each end of bellcrank
(4).
4. Remove safety wire and relieve chain tension at either turnbuckle (6). DO NOT ALLOW ROD
END (23) TO TURN.
5. Remove bolt (5) securing rod end (23) to
bellcrank (4).
6. Remove bolt securing push-pull rod (index
18, figure 9-2) to bellcrank (4).
7. Remove bellcrank pivot bolt and remove
bellcrank from under instrument panel. Use care
not to drop parts.
8. Reverse the preceding steps for reinstallation. Rig rudder and trim systems in accordance
with paragraph 9-16, safety turnbuckles and reinstall
all items removed for access.
b. AFT. (Refer to figure 9-1.)
1. Remove access plate from top of stabilizer
adjacent to vertical fin to expose rudder bellcrank
(12).
2. Remove access plate from top of stabilizer"
to expose turnbuckle (7).
3. Remove safety wire and relieve cable tension
at turnbuckle (7).
4. Disconnect cables (10 and 14) at bellcrank
(12).
5. Disconnect push-pull rod at bellcrank.
6. Remove pivot bolt and remove bellcrank
through access opening.
7. Reverse the preceding steps for reinstallation.
Rig rudder system in accordance with paragraph 9-16,
safety turnbuckle and reinstall all items removed for
access.
9-12. RUDDER BUNGEE. (Refer to figure 9-6.)
9-13. REMOVAL AND INSTALLATION.
a. Remove lower console cover.
b. Remove bolt (10) securing rod end (8) to bellcrank (9).
c. Remove bolt (4) securing bungee (5) to rudder
bar arm (2) and remove bungee.
d. Reverse the preceding steps for reinstallation.
Adjust bungee to dimension shown on installation,
rig system in accordance with paragraph 9-16 and
reinstall all items removed for access.
NOTE
Before installation of a new bungee, a complete rudder and rudder trim system operational check should be accomplished. Refer
to paragraph 9-16.
9-3
•
THRU AIRCRAFT SERIALS
33701462 AND F33700055
BEGINNING WITH AIRCRAFT SERIALS 33701463 AND F33700056
Detail
REFER TO
FIGURE 9-3
B
7
..........,'
,
. ...."
......
,"', ..
............
"
DatallC
•
B
c-
.......... ' ......... (~~
..
,
I
•. .-.. "
.'....'
'. \
I
,.,
~
.........
"',
,, .' ........--_....
\
~ ...
..•.
...l",l,'\'·
........": ~".,\ \\..\...... ."........, . ..............
,
........., ,~.,.: .
~
NOTE
"
9
"" ........ '-\l.~:··
2
Detail
D
1. Bracket
2. Cable Guard
3. Pulley (Tab Down Cable)
4. Pulley (Tab Up Cable)
5. Pulley (Rudder)
6. Cover
7. Turnbuckle (Interconnect)
8. Turnbuckle
Locate turnbuckles of
adjacent cables so they
do not meet, cross or
rub •
Shaded pulleys are used
in this system only.
D
REFER TO FIGURE 9-2
Refer to figure 4-2 for
cable routing through
wing strut faideacls.
9.
10.
11.
12.
13.
14.
15.
Turnbuckle
Cable (Left Interconnect)
Stop Bolt
Bellcrank
Rudder Arm
Cable (Left Aft Rudder)
Bearing
MAINTAIN PROPER CONTROL
CABLE TENSION.
CABLE TENSION:
30 :t 10 LBS (AT AVERAGE TEMPERATURE FOR THE AREA.)
REFER TO FIGURE 1-1 FOR TRAVEL.
Figure 9-1. Rudder Control System (Sheet 1 of 2)
9-4
•
•
RIGGING PIN HOLE
12
Detail
•
E
80 lb-in.
Figure 9-1. Rudder Control System (Sheet 2 of 2)
9-14. CABLES AND PULLEYS.
9-15. REMOVAL AND INSTALLATION.
a. FORWARD CABLES. (Refer to figure 9-1.)
NOTE
The folloWing procedure is written for removal of BOTH forward cables. If ONE
is to be removed, use only the steps necessary for that particular cable.
•
*Torque to 60 -
1. Remove seats, carpeting and access plates
as necessary to expose Details B, C, and D.
2. Remove wing strut fairings as necessary to
expose turnbuckles (8 and 9).
3. Remove safety Wire, relieve cable tension
and disconnect turnbuckles (8 and 9).
4. Remove safety wire and relieve elevator control system cable tension at turnbuckles (index 8,
figure 8-1).
5. Remove access plates from inboard aft side
of right tail boom as necessary to expose turnbuckles
(index 8 and 9, figure 8-6). Remove safety wire and
relieve cable tension.
6. Disconnect cables (index 3, figure 9-5) from
bellcrank (index 4, figure 9-5).
7. Mark or tag cables and pulleys in Details B
and C and remove bolts securing pulleys to brackets
(1).
9. Work cables free of aircraft by routing cables
from under floorboards and out of Wing struts.
NOTE
To ease routing of cables, a length of wire
may be attached to the end of the cable
being withdrawn from the aircraft. Leave
wire in place, routed through structure;
then attach the cable being installed and
pull the cable into position.
10. Reverse the preceding steps for reinstallation and install pulleys and cable guards. Ensure
cables are positioned in pulley grooves before installing guards.
11. Re-rig system in accordance with paragraph
9-16 and safety turnbuckles.
12. Re-rig elevator and elevator trim systems in
accordance with paragraphs 8-12 ~nd 8- 26 respectively, safety turnbuckles and reinstall all items removed for access.
b. CENTER CABLES. (Refer to figure 9-1. )
NOTE
The following procedure is written for removal of BOTH center cables. If ONE is
to be removed, use only the stepG necessary for that particular cable.
8. Remove cable guards (2) from bracket (1) in
Detail D.
9-5
•
&
REFER TO FIGURE 9-5
14
•
1&
15
1. Brake Actuating Arm
2. Brake Actuating Torque Tube
3. Right Rudder Bar
4. Bearing Block
5. Bellcrank
6. Steering Bungee
7. Master Cylinder
8. Spacer
9. Bracket
10. Left Rudder Bar
11. Bearing
12. Anti-Rattle Spring
13. Pedal
14. Pivot Shaft
15. Support
16. Brake Link
17. Support
18. Push-Pull Rod
NEUTRAL PEDALS
AT STATION 73.42
Figure 9-2. Rudder Pedals Installation
9-6
7
•
•
•
1. Remove wing strut fairings as necessary to
expose Detail A and turnbuckles (8 and 9).
2. Remove access plates as necessary to expose
Detail E.
3. Remove safety wire, relieve cable tension
and disconnect turnbuckles (8 and 9).
4. Complete steps 4 and 5 of subparagraph "a. "
5. Disconnect cable (14) from forward side of
bellcrank (12).
6. Mark or tag cables and pulleys in Detail A
and remove bolt securing pulleys to bracket (1).
7. Remove cable guards (2) from bracket (1) in
Detail E.
8. Complete "NOTE" in step 9 of subparagraph
"a."
9. Work cables free of aircraft by routing cables
through tail booms and out of wing struts.
10. Reverse the preceding steps for reinstallation and install pulleys and cable guards. Ensure
cables are positioned in pulley grooves before installing guards.
11. Re-rig system in accordance with paragraph
9-16 and safety turnbuckles.
12. Complete step 12 of subparagraph "a. "
c. INTERCONNECT CABLE. (Refer to figure 9-1.)
1. Remove access plates from stabilizer and
vertical fins as necessary to expose Detail E and
turnbuckle (7).
2. Remove safety wire and relieve cable tension
at turnbuckle (7).
3. Disconnect cable (10) at forward end of bellcranks (12).
4. Complete "NOTE" in step 9 of subparagraph
"a."
5. Work cable free of aircraft by routing cable
out through bellcrank access opening.
6. Reverse the preceding steps for reinstallation. Rig system in accordance with paragraph 9-16,
safety turnbuckle (7) and reinstall all. items removed
for access.
9-16. RIGGING-RUDDER, RUDDER TRIM AND
NOSE WHEEL STEERING SYSTEMS.
NOTE
Since rudder, rudder trim and nose wheel
steering systems are interconnected, adjustments to one system may affect the
others. The following procedure outlines
rigging, in proper sequence, for all three
systems.
•
a. (Refer to figure 9-1.) Remove quadrant covers,
wing strut fairings and stabilizer access plates as
necessary to expose turnbuckles (7, 8 and 9), rudder trim system and steering bungee.
b. (Refer to figure 9-5.) Disconnect steering bungee (18) from right rudder bar (20).
c. Clamp rudder pedals in neutral position.
d. Remove safety Wire, relieVe chain tension at
turnbuckles (6) and disengage chain (8) from actuator
sprocket (24). Adjust trim control wheel (16) so
position indicator (14) is neutral and an equal number
of chain links are between turnbuckles (6) and trim
wheel sprocket (9). Re-engage chain on sprocket if
necessary .
NOTE
The actuator MUST be installed with the left
hand threaded rod end at top and approximately .IS" exposed threads at each end. If
necessary, disconnect a.::tuator at bellcrank
(4) and rotate sprocket (24) to extend actuator
to 4.23" between rod ends and reconnect actuator to belle rank (4). (Refer to VIEW A-A.)
e. While maintaining the actuator dimensions required in step "d" and bell crank (4) in the horizontal
position, re-engage chain on sprocket (24). Make
sure the chain (8) has an equal nwnber of links as
outlined in step "d. "
f. Connect turnbuckles (6) and adjust chain tension.
NOTE
Remove clamps from rudder pedals. Holding
full right rudder and maintaining neutral position of trim wheel and actuator, tighten chain
turnbuckles (6) evenly to remove slack from
chains without binding. Safety turnbuckles,
then reclamp the rudder pedals in neutral
position.
g. (Refer to figure 9-1.) Remove safety wire and
loosen turnbuckles (7, Sand 9).
h. Install 3/16 inch diameter rigging pins at least
five inches long in rudder bellcranks (12), adjust
rudder push-pull rods to place rudders in neutral
(streamlined) position and remove rigging pins.
i. Adjust turnbuckles (7, 8 and 9) to obtain proper
cable tension while keeping the rudders in the neutral
position. Results of adjusting the turnbuckles are as
follows:
1. Loosening turnbuckles (8 and 9) and tightening turnbuckle (7) will move both rudder trailing
edges inboard.
2. Loosening turnbuckle (7) and tightening turnbuckles (S and 9) will move both rudder trailing edges
outboard.
3. Loosening turnbuckle (7) and tightening turnbuckle (8 or 9) will move the rudder trailing edge for
that particular side outboard.
4. Loosening turnbuckle (8 and 9) and tightening
turnbuckle (7) will move the rudder trailing edge for
that particular side inboard.
j. Safety turnbuckles (7, 8 and 9).
k. Remove clamps from rudder pedals. Adjust
stop bolts (11) at both bellcranks (12) to degree of
travel speCified in figure 1-1. Adjust inboard travel
first, then outboard travel to ensure no interference
between rudders and elevator. Refer to figure 9-4
when adjusting travel.
1. Jack the nose gear free of ground and make sure
the centering lug on the upper torque link seats
firmly against flap spot on strut, locking nose gear
in neutral .
m. (Refer to figure 9-6.) Adjust push-pull rod (12)
to 11.09 ± .03 inches between centers of rod end
holes, tighten jam nuts and reinstall.
9-7
•
5
2
4
&
Detail
A
•
Detail
B
B
1.
2.
3.
4.
5.
6.
7.
8.
Weight
Rudder Tip
Hinge Assembly
Hinge Bracket
Pivot Bolt
Bearing
Arm Assembly
Rudder Assembly
Figure 9-3. Rudder Installation
9-8
•
•
4.35 "
10.75 "
26.25 "
\ . - 11" MAXIMUM
TRAVEL
INBOARD
27° MAXIMUM
TRAVEL
OUTBOARD
1
17.35 "
•
....-----1f-.75 " (TYPICAL)
.25 " (TYPICAL)
'_--AFT FIN SPAR
RIVET PATTERN
CENTERLINE
-.. r'
2° 30' (TYPICAL)
50 " (TYPICAL)
l.50"
RADIUS
(TYPICAL)
. 25 ,0 (TYPICAL)
•
~
FRONT FIN SPAR
RIVET PATTERN
-CENTERLINE
(
C AL..
. 75 "TYPI
3° (TYPICAL)
Position template parallel
with rivet line of rudder rib
at middle hinge.
template edge, measure
distance from trailing edge
of rudder to template edge.
Using • 23 inch equals 1 degree, convert the inch measurement to degrees and rerig rudder system as required to meet these tolerances •
9.60 "
T
Rudder travel is measured
perpendicular to hinge line.
If rudder does not contact
1-----6.60" - - - -
t
NOTE
MATERIAL: 2024- T3 CLAD SHEET
(MAXIMUM THICKNESS .10 INCH)
ALTERNATE: 1/4 OR 3/8 INCH PLY-
I~'
-.j i
33 "
7.25 " - - - -...
• (TYPICAL)
WOOD
Figure 9-4. Measuring Rudder Travel
9-9
operation of trim wheel applies correct
rudder trim.
n. Position rudder pedals in neutral position
(streamlined), adjust bungee (5) to dimension shown
by adjusting rod end (8) to align with rudder bar arm
(2) and reinstall hardware. DO NOT uncover safety
inspection hole on shaft (7).
o. Inspect rudder, rudder trim and nose wheel
steering systems for neutral positions. If all systems are not neutral, repeat rigging procedures.
p. Lower the nose gear to ground, check all components are secure and safetied as required and reinstall all items removed for access.
9-17. RUDDER TRIM CONTROL SYSTEM.
9-18. DESCRIPTION. The rudder trim control system is operated by a control Wheel, mounted in the
control console and is connected by chains to a trim
actuator located between the right rudder bar and the
forward rudder control bellcrank. As the trim wheel
is rotated, the actuator is lengthened or shortened,
causing the bellcrank to pivot against the force of the
bungee, effecting rudder offset. The bungee serves
as a rudder trim bungee when airborne and a steering
bungee when on the ground.
IWARNING,
Be sure rudders move in the correct direction when operated by the pedals and that
9-19.
•
TROUBLE SHOOTING.
NOTE
Due to remedy procedures in the following trouble shooting
chart it may be necessary to re-rig system, refer to paragraph 9-16.
TROUBLE
NO RESPONSE TO TRIM
WHEEL MOVEMENT.
BINDING OR JUMPY
MOVEMENT OF
TRIM WHEEL.
REVERSE TRIM APPLIED
WHEN SYSTEM IS OPERATED.
INSUFFICIENT RUDDER
TRIM IMMEDIATELY
AFTER TAKE-OFF.
(Refer to figure 9-6.)
9-10
PROBABLE CAUSE
REMEDY
Broken or disconnected chain.
Check visually. Connect if
disconnected. Replace
defective parts.
Defective actuator.
Disconnect actuator and check
manually. Replace defective
actuator.
Incorrect chain adjustment.
Check and adjust tension.
Defective actuator.
Disconnect actuator and check
manually. Replace actuator.
Defective trim wheel
bearings.
Check and replace defective
bearings.
Inverted actuator.
Disconnect actuator and check
manually. Replace actuator
with left hand threads UP.
Idler bellcrank (9) attach
Jack aircraft and partially
retract gear. Check visually.
Tighten bolt. Add washers
as necessary.
bolt loose.
Steering cam lock (15) bolt
loose.
Jack aircraft and partially retract
gear. Check visually. Torque
bolt until cam lock is free of
lateral movement, but sUll free
to move up or down.
Improper cable tension.
Check and adjust cable tension.
Rudders improperly aligned.
Rig rudders in accordance
with paragraph 9-16.
•
•
•
o
Screws and clinch nuts beginning
1. Spacer
with
aircraft serials 33701195
2. Chain Guard
and F33700001
3. Cable (Right Fwd)
4. Bellcrank
5. Bolt
6. Turnbuckle
4.23 "
7. Chain Stop
NOTE
8. Chain
9. Sprocket
Dimensions shown are for
10. Upper Support Assembly
no rudder, no trim and no
.18 "
11. Bearing
nose steedng condition.
12. Chain Guard
13. Roll Pin
14. ~:ition Indicator (Typical)
t
15.
16. Trim Control Wheel
17. Lower Support Assembly
18. Steering Bungee
VIEW
19. Bolt
20. Right Rudder Bar
21. Arm
22. Bolt
LEFT HAND THREADS
23. Rod End
• RIGHT HAND THREADS
24. Sprocket
1
TYpl
ArA
*
8
•
16
11
11
A
•
Figure 9- 5. Rudder Trim Control System
9-11
5
3
NOTE
Dimensions shown are for
no n1dder, no trim and no
nose steering condition.
&
•
7
2
----
........
11.82 "
II
!
~"
--~
j\~
rx:::-1
..-=!.:. . -~~
r~
•
15
THRU AmCRAFT SERIAL 337-0500 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK337-4
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
Right Rudder Bar
Arm
Bushing
Bolt
Steering Bungee
Safety Wire
Shaft
Rod End
Bellcrank
Bolt
Spacer
Rod Assembly
Boot
Bolt
Steering Cam Lock
Steering Cam
Spring
BEGINNING WITH AIRCRAFT SERIALS 337-0501,
F33700001 AND ALL AmCRAFT MODIFIED IN
ACCORDANCE WITH SK337-4
Figure 9-6. Nose Wheel Steering and Rudder Trim System Check Points
9-12
•
•
9-20. TRIM CONTROL WHEEL. (Refer to figure
9-5. )
9-21.. REMOVAL AND INSTALLATION.
a. Remove console covers as necessary in accordance with paragraph 9- 24.
b. Remove chain guard (12) from upper support
assembly (10).
c. Remove safety wire, relieve chain tension and
disengage chain from sprocket (9).
d. Remove screws securing upper support assembly (10) to console structure. Lift upper support,
trim indicator and control wheel assembly out of
structure. Use care not to lose spacer (1).
e. Sprocket (9) and control wheel (16) may be removed from upper support (10) by driving out roll
pin (13).
f. Trim indicator may be removed by drilling out
pin (15) •
g. Reverse the preceding steps for reinstallation.
Rig trim control system in accordance with paragraph 9-16, safety turnbuckles and reinstall all items
removed for access.
9-22. RIGGING. The rudder, rudder trim and nose
wheel steering systems are interconnected and adjustments to one system may affect the others. A
complete rigging procedure, in proper sequence, for
all three systems Is outlined in paragraph 9-16.
9-23. CONSOLE AND QUADRANT COVERS.
9-24. REMOVAL AND INSTALLATION. (Refer to
figure 10-4, sheet 2.)
SHOP NOTES:
•
•
9-13/(9-14 blank)
SECTION 10
•
ENGINES
(NON- TURBOCHARGED)
TABLE OF CONTENTS
•
•
ENGINE COWLING
Description
Front.
Rear
.
Removal and Installation
Front.
Rear
.
..
Cleaning and Inspection
Repair
ENGINES
•
Description
Engine Data
Trouble Shooting
Removal
Front.
.
Rear
..
Cleaning
Accessories Removal .
Inspection .
Build-Up
Installation
Front.
.
Rear
Flexible Fluid Hoses
Pressure Test
Replacement
Engine Baffles .
Description
Cleaning and Inspection
Removal and Installation
Repair
•.•.•
Engine Mount
.
.
•.
Description
•.
Removal and Installation
Repair
., .
Engine Shock- Mount Pads
COWL FLAPS
.,.
..
DeSCription
...
.
Trouble Shooting •
•
Removal and Installation
Front
Rear
..
.
Page
10-2
10-2
10-2
10-2
10-3
10-3
10-3
10-3
10-3
10-3
10-3
10-4
10-5
10-8
10-8
10-9
10-10
10-11
10-11
10-11
10-11
10-11
10-12
10-13
10-13
10-13
10-13
10-13
10-13
10-13
10-14
10-14
10-14
10-14
10-14
10-14
10-14
10-14
10-16
10-21
10-21
10-21
Rigging
Front.
Rear
•
CONTROL QUADRANT
Description
Removal and Installation
Disassembly and Reassembly
ENGINE CONTROLS
.
Description
Removal and Installation
Rigging
.
Throttle-Operated Gear Warning
Switches
Description
Rigging .
INDUCTION.AIR SYSTEM .
Description
Removal and Installation
Cleaning Induction Air Filter
FUEL INJECTION SYSTEM
Description
Trouble Shooting.
.
Fuel-Air Control Unit
Description
Removal and Installation
Adjustments .
.
Fuel Manifold Valve (Fuel
Distributor)
•
Description
•
Removal and Installation
Cleaning
•...
Fuel Discharge Nozzles
Description
•.
Removal.
.
Cleaning and Inspection
Installation
.
Fuel Injection Pump
.
•
Description
. .
Removal and Installation
Adjustments
EXHAUST SYSTEMS
Description
10-21
10-21
10-23
10-26
10-26
10-26
10-26
10-26
10-26
10-30
10-30
10-30
10-30
10-30
10-31
10-31
10-32
10-32
10-32
10-32
10-34
10-35
10-35
10-33
10-36
10-36
10-36
10-36
10-36
10-36
10-36
10-37
10-37
10-37
10-37
10-37
10-38
10-38
10-40
10-40
10-1
Front Engine .................. .
Rear Engine ................... .
Removal .......................... .
Inspection ........................ .
Installation ....................... .
OILSYSTEM .......................... .
Description ....................... .
Trouble Shooting .................. .
Full-Flow Oil Filter ................ .
Description .................... .
Element Removal and Installation
Adapter Removal .............. .
Adapter Disassembly, Inspection
and Reassembly ............ .
Adapter Installation ............ .
Oil Dilution System ................ .
Removal of Oil Dilution System ..... .
Forward Engine ............... .
Rear Engine ................... .
Installation of Oil Dilution System .. .
Forward Engine ............... .
Rear Engine ................... .
IGNITION SYSTEM ................... .
Description ....................... .
10-40
10-40
10-40
10-40
10-40
10-40
10-40
10-42
10-45
10-45
10-46
10-45
10-47
10-47
10-47
10-48
10-48
10-48
10-48
10-48
10-48
10-48
10-48
Trouble Shoot.ing .................. .
Magnetos ......................... .
Description .................... .
Removal and Installation ....... .
Internal Timing ................ .
Magneto-to-Engine Timing ...... .
Magneto Check ................ .
Maintenance ................... .
Tachometer Breaker Point
Adjustment ................. .
Spark Plugs ....................... .
STARTING SYSTEM ................... .
Description ........................ .
Trouble Shooting .................. .
Starter Motor
Removal and Installation ....... .
Primary Maintenance .............. .
EXTREME WEATHER MAINTENANCE .
Cold Weather ...................... .
Hot Weather ...................... .
Seacoast and Humid Areas .......... .
Dusty Areas ....................... .
Ground Service Receptacle .......... .
Hand Cranking .................... .
10-49
10-51
10-51
10-51
10-51
10-51
10-52
10-53
I
•
10-53
10-53
10-53
10-53
10-54
10-55
10-55
10-55
10-55
10-55
10-56
10-56
10-56
10-56
•
10-1. ENGINE COWLING.
10-2. DESCRIPTION.
a. FRONT. The front engine cowling is divided into
four removable sections. The right and left nose caps
are fastened to the lower section and to each other
with screws. The right and left upper cowl sections
are secured with quick-release fasteners and either
section may be removed individually. The left cowl
section has two access doors located at the rear.
The upper door provides access to the engine oil filler
neck and the lower door provides access to the oil dipstick and fuel strainer drain control. The lower portion of the cowl is an extension of the fuselage, enclOsing the retractable nose wheel and providing the
engine mount structure.
10-2
Change 1
b. REAR. The rear engine cowling is divided into
five removable sections. The upper and lower tail
caps are fastened to the lower section and to each
other with screws. The right and left side panels are
secured witt: quick- release fasteners and the upper
cowl is attached with screws. Access to the oil filler
cap is gained through a door in the upper cowl. The
oil dipstick and fuel strainer drain control are located
behind a door in the right side panel, directly above
the cowl flap. The lower portion of the cowl is an extension of the fuselage, enclosing the retracted main
landing gear. An air scoop, secured to the upper aft
part of the fuselage, directs ram air to the rear engine. Air flow is controlled by laterally mounted cowl
flaps, one in each side panel. Thru aircraft serials
33701316 and F33700024 a course-mesh screen is installed at the tail cap to protect the propeller.
•
•
10-3. REMOVAL AND INSTALLATION.
a. FRONT •
1. Loosen the quick-release fasteners on the
right and left upper cowl sections and remove sections.
2. Remove screws securing the nose cap halves
together.
3. Disconnect the induction air and heat exchanger air ducts from nose caps. (Turbocharged aircraft
only. )
4. Remove screws securing the nose cap halves
to the lower section and remove nose cap halves.
5. Reverse the preceding steps for reinstallation.
Make sure the vertical baffle seals fold forward and
the side baffle seals fold upward to ensure proper
engine cooling.
b. REAR.
1. Turn master switch ON, run cowl flaps to the
OPEN position and disconnect cowl flap push-pull rod
ball joints at the cowl flaps. 00 NOT DISTURB ROD
END ADJUSTMENT.
NOTE
If the cowl flaps operate normally, proceed
to step 6, if the cowl flaps cannot be opened
electrically, complete steps 2 through 5.
(Refer to figure 10-2, sheets 3 and 4.)
I.
I
2. Reach through the access door on the right
side panel and disengage the vertical push-pull rod
(49) from arm (24) at end of torque tube (48), at
quick-disconnect (30).
3. Manually open the right cowl flap (56), disengage the horizontal push-pull rod (53) from right
cowl flap (56) at quick-disconnect (30) and remove the
right side panel.
4. Reach across firewall and disengage the stationary link (20) from cowl flap motor arm (31) at
quick-disconnect (30).
5. Rotate the complete torque tube assembly
(48) to open the left cowl flap, disengage the horizontal push-pull rod (53) from left cowl flap at quickdisconnect (30) and remove left side panel.
6. Loosen the quick-release fasteners on the
right and left cowl panels and remove panels.
7. (THRU AIRCRAFT SERIALS 33701316 and
F33700024.) Remove bolts securing the three coursemesh screens together and remove screens.
8. To remove the upper and lower tail caps or
the air scoop, remove the screws securing them to
each other and to the fuselage.
9. Reverse the preceding steps for reinstallation.
Make sure the vertical baffle seals fold forward and
the side baffle seals fold upward to ensure proper engine cooling. Check cowl flaps for proper operation.
10-4. CLEANING AND INSPECTION. Wipe the inner
surfaces of the cowling segments with a clean cloth
•
saturated with cleaning solvent (Stoddard or equivalent). If the inside surface of the cowling is coated
heavily with oil or dirt, allow solvent to soak until
foreign material can be removed. Wash painted surfaces of the cowling with a solution of mild soap and
water and rinse thoroughly. After waShing, a coat of
wax may be applied to the painted surfaces to prolong
paint life. After cleaning, inspect cowling for dents,
cracks, loose rivets and spot welds. Repair all defects to prevent spread of damage.
10-5. REPAIR. If cowling skins are extensively
damaged, new complete sections of the cowling
should be installed. Standard insert-type patches
may be used for repair if repair parts are formed to
fit contour of cowling. Small cracks may be stopdrilled and small dents straightened if they are reinforced on the inner surface with a doubler of the
same material as the cowling skin. Damaged reinforcement angles should be replaced with new parts.
Due to their small size, new reinforcement angles
are easier to install than to repair the damaged part.
10-6. ENGINES.
10-7. DESCRIPTION. Air cooled, wet sump, six
cylinder, horizontally-opposed, fuel-injected, Continental, 10-360 series engines are installed on the
aircraft. Both engines are located on the fuselage
centerline, one forward and one aft of the cabin.
The engines themselves are similar, although their
front (propeller) ends point in opposite directions. A
conventional tractor propeller is required for the
front engine and a pusher propeller is required for
the rear engine. Each propeller rotates in the same
direction in relation to its engine, but rotate in opposite directions in relation to each other. Cooling for
the rear engine is obtained by an overhead air scoop
and laterally mounted cowl flaps. Refer to paragraph
10-8 for engine data. For repair and overhaul of the
engines, accessories and propellers, refer to the
appropriate publications issued by their manufacturer's.
These publications are available from the Cessna Service Parts Center.
NOTE
Since the installed engines face in opposite
directions, some confusion might arise
from terms such as "right, " "left, " "front"
and "rear." Except where further clarified
in the text, these terms shall be applied to
the rear engine as though it were removed
from the aircraft and viewed from its accessory case end. Rear engine baffles, cowling
and firewall are not considered part of the
basic engine and shall be identified as "right, "
"left, " etc., in relation to the aircraft.
10-3
10-8. ENGINE DATA.
MODEL (Continental)
Thru aircraft serial 337-0755
Aircraft serials 337-0756 thru 33701462
and F33700001 thru F33700055
Beginning with aircraft serials
33701463 and F33700056
1Q-360-G (Front and Rear)
BHPatRPM
210 at 2800
Number of Cylinders
6-Horizontally-Opposed
Displacement
Bore
Stroke
360 Cubic Inches
4.438 Inches
3.875 Inches
Compression Ratio
8.5:1
Magnetos
Right Magneto
Left Magneto
Bendix-Scintilla S6LN- 25
Fires 20° BTC Upper Right and Lower Left
Fires 20° BTC Upper Left and Lower Right
Firing Order
1-6-3-2-5-4
Spark Plugs
18MM x • 750-20 Thread Connection
(Refer to current Continental active
factory approved spark plug chart. )
Torque Value
IO-360-C (Front) IO-360-D (Rear)
IO-360-C (Front and Rear)
330±30 Lb-In.
Fuel Metering System
Unmetered Fuel Pressure
Continental Fuel Injection
6 to 8 PSI at 600 RPM FRONT or 650 RPM REAR
25 to 27 PSI at 2800 RPM
Oil Sump Capacity
With Filter Element Change
10 U.S. Quarts
11 U.S. Quarts
Tachometer
Electric (Operated by Magneto Pick- Up)
Oil Pressure (PSI)
Minimum Idling
Normal
Maximum (Cold Oil Starting)
Connection Location
10
30 to 60
100
Between No.2 and No.4 Cylinders (Front and Rear)
Oil Temperature
Normal Operation
Maximum Permissible
Within Green Arc
Red Line (240°F)
Cylinder Head Temperature
Normal Operating
Maximum
Probe Location (Front Engine)
Probe Location (Rear Engine)
Approximate Dry Weight
10-4
•
•
Within Green Arc
Red Line (460°F)
Lower side No.3 Cyl. (Thru aircraft serial 337-0755.)
Lower side No. 2 Cyl. (Aircraft serials 337-0756 thru
33701316 and F33700001 thru F33700024.)
Lower side No.6 Cyl. (Beginning with aircraft
serials 33701317 and F33700025.)
Lower side No. 2 Cyl. (Thru aircraft serials
33701316 and F33700024.)
Lower side No. 6 Cyl. (Begir.ru.ng with aircraft
serials 33701317 and F33700025.)
327 lbs. (Weight is apprOximate and will vary with
optional accessories installed. )
•
•
•
•
10-9. TROUBLE SHOOTING •
TROUBLE
ENGINE FAlLS TO START.
PROBABLE CAUSE
REMEDY
Improper use of starting
procedure.
Review starting procedure.
Refer to Owner's Manual.
Defective aircraft fuel system.
Refer to Section 11.
Spark plugs fouled or defective.
Remove and clean. Check gaps and
:r.sulators. Use neW gaskets. Check
cables to persistenly fouled plugs.
Replace if defective.
Defective magneto switch or
grounded magneto leads.
Check continuity, repair or replace
switch or leads.
Defective ignition system.
Refer to paragraph 10-92.
Excessive induction air leaks.
Check visually. Correct cause of
air leaks.
Dirty screen in fuel control unit
or defective fuel control unit.
Check screen visually. Check fuel
flow through control unit. Replace
defective fuel control unit.
Defective electric fuel pump.
Refer to Section 11.
Defective fuel manifold valve
or dirty screen .
Check fuel flow through valve.
Remove and clean. Replace if
defective.
Clogged fuel injection lines or
discharge nozzles.
Check fuel through lines and nozzles.
Clean lines and nozzles. Replace if
defective.
Fuel pump not permitting fuel
from auxiliary pump to bypass.
Check fuel flow through engine-driven
fuel pump. Replace engine-driven
pump.
Vaporized fuel in system.
Refer to paragraph 10-103.
Fuel tanks empty.
Visually inspect tanks. Fill with
proper grade and quantity of gasoline.
Fuel contamination or water in
fuel system.
Open fuel strainer drain and check
for water. Drain all fuel and flush
out fuel system. Clean all screens,
fuel lines, strainer, etc.
Mixture control in the IDLE
CUT-OFF position.
Move control to the full RICH
position.
Engine flooded.
Refer to paragraph 10-103.
Fuel selector valves in OFF
position.
Place selector valves in the ON
position to tanks known to contain gaSOline.
Magneto impulse coupling
failure.
Repair or Install neW coupling.
10-5
10-9. TROUBLE SHOOTING (Cont).
TROUBLE
ENGINE STARTS BUT
DIES, OR Wll..L NOT
IDLE PROPERLY.
PROBABLE CAUSE
REMEDY
Idle stop screw or idle mixture
lncorrecUy adjusted.
Refer to paragraph 10-55.
Spark plugs fouled or improperly
gapped.
Remove, clean and regap plugs.
Replace if defective.
Water in fuel system.
Open fuel strainer drain and check
for water. If water is present,
drain fuel tank sumps, lines and
strainer.
Defective ignition system.
Refer to paragraph 10-92.
Vaporized fuel. (Most likely to
occur in hot weather with a hot
engine. )
Refer to paragraph 10-103.
Induction air leaks.
Check visually. Correct the
cause of leaks.
Manual primer leaking.
Disconnect primer ouUet line.
If fuel leaks through primer,
repair or replace primer.
Dirty screen in fuel control unit
or defective fuel control unit.
Check screen visually. Check
Fuel flow through control unit.
Clean screen. Replace fuel
control unit if defective.
Defective manifold valve or
clogged screen.
Check fuel flow through valve.
Replace if defective. Clean screen.
Defective engine-driven fuel
pump.
If engine continues to run with
electric pump turned on, but
stops when it is turned off, the
engine-driven pump is defective.
Replace pump.
Defective engine.
Check compression. Listen for
unusual engine noises. Engine
repair is required.
Propeller control set in high
pitch position (low rpm).
Use low pitch (high rpm) position
for all ground operation.
Defective aircraft fuel system.
Refer to Section 11.
Restricted fuel injection lines
or discharge nozzles.
Check fuel flow through lines and
nozzles. Clean lines and nozzles.
Replace if defective.
Obstructed air intake.
Check visually. Remove obstruction;
service air filter, if necessary .
•
•
•
10-6
•
•
10-9. TROUBLE SHOOTING (Cont).
TROUBLE
ENGINE RUNS ROUGHLY,
WILL NOT ACCELERATE
PROPERLY, OR LACKS
POWER.
POOR IDLE CUT-OFF.
PROBABLE CAUSE
Propeller control in high pitch
(low rpm) position.
REMEDY
Use low pitch (high rpm) for
all ground operations.
Restriction in aircraft fuel
system.
Refer to Section 11.
Restriction in fuel injection
system.
Clean system. Replace any
defective units.
Engine-driven fuel pump pressure improperly adjusted.
Refer to paragraph 10-68.
Worn or improperly rigged
throttle or mixture control.
Check visually. Rig properly.
Replace worn linkage.
Spark plugs fouled or improperly
gapped.
Clean and regap. Replace if
defective.
Defective ignition system.
Refer to paragraph 10-92.
Defective engine.
Check compression. Listen for
unusual engine noises. Engine
repair is required.
Propeller out of balance.
Check and balance propeller.
Worn or improperly rigged
mixture control.
Rig properly. Replace worn
linkage.
Defective or dirty manifold valve.
Operate electric fuel pump and
check that no fuel nows through
manifold valve with mixture control in IDLE CUT-OFF. Remove
and clean. Replace if defective.
Fuel leakage through primer.
Repair or replace primer.
Auxiliary fuel pump ON.
Turn to OFF position.
Defective fuel control unit.
U none of the preceding causes
"
.
corrects the problem, the control unit is probably at fault.
Replace control unit.
•
10-7
10-9. TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
HIGH CYLINDER HEAD
TEMPERATURE.
Defective cylinder head temperature indicating system.
Refer to Section 14.
Improper use of cowl naps.
Refer to Owner's Manual.
Defective cowl nap operating
system.
Refer to paragraph 10-31.
Engine baffies loose, bent or
missing.
Check visually. Install baffies
properly. Repair or replace
if defective.
Dirt accumulated on cylinder
cooling fins.
Check visually. Clean
thoroughly.
Incorrect grade of fuel.
Drain and refill with proper fuel.
Incorrect ignition timing.
Refer to paragraph 10-90.
Defective fuel injection system.
Refer to paragraph 10-51.
Improper use of mixture control.
Refer to Owner's Manual.
Defective engine.
Repair as required.
HIGH OR LOW OIL
TEMPERATURE
OR PRESSURE.
Refer to paragraph 10-76.
10-10. REMOVAL. U an engine Is to be placed in
storage or returned to the manufacturer for overhaul,
proper preparatory steps should be taken for corrosion prevention prior to beginning the removal procedure. Refer to Section 2 for storage preparation.
The routing and location of wires, cables, lines,
hoses and controls wID vary With optional equipment
installed, however, the following general procedure
may be followed.
a. FRONT. The front engine may be removed as a
complete unit with the accessories installed, however,
the exhaust system must be disconnected.
I~AUTIONl
Place suitable padded stands under the tail
boom tie-down rings before removing front
engine. The loss of front engine weight will
cause the aircraft to be tail heavy.
NOTE
Tag each item when disconnected to aid in
identifying wires, hoses, lines and control
linkages when engine is reinstalled. Likewise, shop notes made during removal will
10-8
REMEDY
•
•
often clarify reinstallation. Protect openings, exposed as a result of removing or
disconnecting units, against entry of foreign
material by installing covers or sealing with
tape.
1. Place all cabin switches in the OFF position.
2. Place fuel selector valves in the OFF position.
3. Remove engine cowling and nose caps in accordance with paragraph 10-3.
4. Disconnect battery cables, remove battery
and battery box for additional clearance, if desired.
5. Drain fuel strainer and lines with strainer
drain control.
NOTE
During the following procedures, remove
any clamps which secure controls, wires,
hoses or lines to the engine, engine mounts
or attached brackets, so they will not interfere with the engine removal. Some of
these items listed can be disconnected at
more than one place. It may be desirable
to disconnect some of these items at other
•
•
than the places indicated. The reason for
engine removal should be the governing factor in deciding at which point to disconnect
them. Omit any of the items which are not
present on a particular engine installation.
6. Place propeller control in high rpm position.
Release unfeathering accumulator pressure through
the filler valve and disconnect hose at accumulator.
7. Drain the engine oil sump and oil cooler.
8. Disconnect magneto primary lead wires at
magnetos.
!WARNING'
The magnetos are in a SWITCH ON condition
when the switch wires are disconnected.
Ground the magneto points or remove the high
tension wires from the magnetos or spark
plugs to prevent accidental firing.
•
•
9. Remove the spinner and propeller in accordance with Section 12. Cover the exposed end of the
crankshaft to prevent entry of foreign material.
10. Unclamp exhaust stacks from both sides of
engine.
11. Disconnect throttle, mixture and propeller
governor controls. Remove clamps attaching controls to engine and pull controls aft clear of engine.
Use care to avoid bending controls too sharply.
12. Disconnect oil temperature wire at sending
unit .
13. Disconnect tachometer pick-up from bottom
of right magneto.
I~AUTION1
When disconnecting starter cable do not
permit starter terminal bolt to rotate.
Rotation of the bolt could break the conductor between bolt and field coils causing the starter to be inoperative.
14. Disconnect starter electrical cable at starter.
15. Disconnect cylinder head temperature wire at
probe.
16. Disconnect electrical wires and wire shielding
ground at alternator.
17. Disconnect electrical wires at throttle-operated switch.
18. Disconnect exhaust gas temperature wires at
probe leads.
19. Disconnect ground strap and any other electrical wiring not previously noted which may be damaged during engine removal.
20. Disconnect fuel strainer drain control wire
at strainer. Remove control housing lock nuts securing housing to nose gear tunnel and pull control and
housing from tunnel area.
21. Disconnect vacuum hose at suction relief
valVe.
22. Disconnect manifold pressure line at firewall.
23. Disconnect fuel supply hose at nose gear tunnel and vapor return hose at firewall.
!WARNING'
Residual fuel and oil draining from disconnected lines and hoses constitutes a fire
hazard. Use caution to prevent accumulation of such fuel and oil when lines or hoses
are disconnected.
24. Disconnect fuel-flow gage hose at firewall.
25. Disconnect oil pressure hose at firewall.
26. Disconnect cylinder fuel drain line at hose
connection on each side of engine.
27. Disconnect engine primer line at fireWall.
28. Disconnect hydraulic hoses at firewall.
29; Carefully check the engine again to ensure
ALL hoses, lines, Wires, cables and clamps are disconnected or removed which would interfere with the
engine removal. Ensure all Wires, cables and engine controls have been pulled aft to clear the engine.
30. Attach a hoist to the lifting eye at the top center of the engine crankcase. Lift engine just enough
to relieve the weight from the engine mounts.
31. Remove bolts attaching engine to engine
mounts and slowly hoist engine and pull it forward.
Checking for any items Which would interfere with
the engine removal. Balance the engine by hand and
carefully guide the disconnected parts out as the engine is removed.
32. Remove the engine shock-mounts.
b. REAR. The rear engine may be removed as a
complete unit with the accessories installed, however,
the exhaust system must be disconnected.
NOTE
Tag each item disconnected to aid in identifying Wires, hoses, lines and control linkages
when the engine is reinstalled. Likewise,
shop notes made during removal will often
clarify reinstallation. Protect openings, exposed as a result of removing or disconnecting
units, against entry of foreign material by installing covers or sealing with tape.
1. Place all cabin SWitches in the OFF position.
2. Place fuel selector valves in the OFF poSition.
3. Remove ALL engine cowling in accordance
with paragraph 10-3.
4. Remove front engine left upper cowl section,
disconnect battery ground cable and insulate terminal
as a safety precaution.
5. Drain fuel strainer and lines with strainer
drain control.
NOTE
During the follOWing procedures, remove
any clamps or lacings which secure controls, wires, hoses or lines to the engine,
engine mount or attached brackets, so they
will not interfere with engine removal.
Some of the items listed can be disconnected
at more than one place. It may be desirable
10-9
to disconnect some of these items at other
than the places indicated. The reason for
engine removal should be the governing
factor in deciding at which point to disconnect them. Omit any of the items which
are not present on a particular engine.
6. Drain the engine oil sump and oil cooler.
7. Disconnect magneto primary lead wires at
magnetos.
IWARNING,
The magnetos are in a SWITCH ON condition
when the switch wires are disconnected.
GroWld the magneto points or remove the
high tension wires from the magnetos or
spark plugs to prevent accidental firing.
8. Remove the spinner and propeller in accordance with Section 12. Cover the exposed end of the
crankshaft to prevent entry of foreign material.
S. Disconnect throttle, mixture and propeller
governor controls. Remove any clamps attaching
controls to engine and pull controls clear of engine.
Use care to avoid bending controls too sharply.
10. Disconnect propeller synchronizer control at
actuator.
11. Disconnect oil temperature wire at sending
unit.
12. Disconnect tachometer pick-up from bottom
of right magneto.
ISAUTIONl
When disconnecting starter cable do not permit starter terminal bolt to rotate. Rotation'
of the bolt could break the conductor between
bolt and field coils causlng the starter to be
inoperative.
13. Disconnect starter electrical cable at starter.
14. Disconnect cylinder head temperature wire at
probe.
15. Disconnect electrical wires and wire shielding ground at alternator.
16. Disconnect electrical wires at throttle-operated switch.
17. Disconnect exhaust gas temperature wires at
probe leads.
18. Disconnect groWld strap and any other electrical wiring not previously noted which may be damaged during engine removal.
19. Disconnect fuel strainer drain control wire at
strainer and remove control housing lock nuts securing housing to fuselage structure. Pull control and
housing from structure area.
20. Disconnect vacuum hose at firewall.
21. Disconnect manifold pressure line at firewall.
22. Disconnect fuel supply hose at auxiliary pump
and vapor return hose at firewall.
10-10
IWARNING,
Residual fuel and oil draining from disconnected lines and hoses constitutes a fire
hazard. Use caution to prevent accumulation of such fuel and oil when lines or
hoses are disconnected.
23. Disconnect fuel-flow gage hose at firewall.
24. Disconnect oil pressure hose at engine.
25. Disconnect cylinder fuel drain line at hose
connection at each side of engine and engine-driven
fuel pump drain line.
26. Disconnect hydraulic hoses at pump.
27. Disconnect engine primer line at firewall
fitting or at rear baffle.
28. Place propeller control in high-rpm position.
Release unfeathering accumulator pressure through
the filler valve and disconnect hose at accumulator.
29. Disconnect and remove flexible ducts as required.
30. Unclamp exhaust stacks at both sides of engine.
31. Carefully check the engine again to ensure
ALL hoses, lines, Wires, cables and clamps are disconnected or removed which would interfere with the
engine removal. Ensure all wires, cables and engine controls have been pulled forward to clear the
engine.
32. Attach a hoist to the lUting eye at the top center of the engine crankcase. Lift engine just enough
to relieve the weight from the engine mOWlt assembly.
ICAUTIONI
•
•
Be sure there is clearance at the top of the
tail section, as the tail section of the air-
craft will rise with the loss of engine weight.
33. Remove bolts attaching engine to engine
mount and slowly hoist engine and pull it aft. Checking for any items which would interfere ,with the engine removal. Balance the engine by hand and carefully guide the disconnected parts out as the engine
is removed.
34. Remove the engine shock-mounts.
10-11. CLEANING. The engine may be cleaned with
Stoddard solvent or equivalent, then dried thoroughly.
I~AUTIONI
Particular care should be given to electrical
equipment before cleaning. Cleaning fluids
should not be allowed to enter magnetos,
starter, alternator, etc.' Protect these components before saturating the engine with solvent. All other openings should also be covered before cleaning the engine assembly.
Caustic cleaning solutions should be used
cautiously and should always be properly
neutralized after their use.
•
•
10-12. ACCESSORIES REMOVAL. Removal of engine accessories for overhaul or for engine replacement involves stripping the engine of parts, accessories and components to reduce it to the bare engine.
During the removal process, removed items should
be examined carefully and defective parts should be
tagged for repair or replacement with new components.
1. Hoist the engine to a point just above the
nacelle.
2. Install engine shock-mount pads as illustrated in figure 10-1.
3. Carefully lower engine slowly into place on
the engine mount pads. Route controls, lines, hoses
and wires in place as the engine is positioned on the
engine mounts.
NOTE
NOTE
Items easily confused with similar items
should be tagged to provide a means of
identification when being installed on a
new engine. All openings exposed by the
removal of an item should be closed by
installing a suitable cover or cap over the
opening. This will prevent entry of foreign
material. If suitable covers are not available, tape may be used to cover the openings.
•
10-13. INSPECTION. For specifiC items to be inspected refer to the engine manufacturer's manual.
a. Visually inspect the engine for loose nuts, bolts,
cracks and fin damage.
b. Inspect baffles, baffle seals and brackets for
cracks, deterioration and breakage.
c. Inspect all hoses for internal swelling, chafing
through protective plys, cuts, breaks, stiffness,
damaged threads and loose connections. Excessive
heat on hoses will cause them to become brittle and
eaSily broken. Hoses and lines are most likely to
crack or break near the end fittings and support
points.
d. Inspect for color bleaching of the end fittings or
severe discoloration of the hoses.
NOTE
Avoid excessive flexing and sharp bends
when examining hoses for stiffness.
e. All flexible fluid carrying hoses in the engine
compartment should be replaced at engine overhaul
or every five years, whichever occurs first.
f. For major engine repairs, refer to the manufacturer's overhaul and repair manual.
10-14. BUILD-UP. Engine build-up consists of installation of parts, accessories and components to the
basic engine to build up an engine unit ready for installation on the aircraft. All safety wire, lockwashers, nuts, gaskets and rubber connections should
be new parts.
10-15. INSTALLATION.
a. FRONT. Before installing the front engine on
the aircraft, install any items which Were removed
from the engine or aircraft after the engine was removed.
NOTE
•
Remove all protective covers,
and identification tags as each
nected or installed. Omit any
present on a particular engine
plugs, caps
item is conitems not
installation.
Be sure engine shock-mount pads, spacers
and washers are in place as the engine is
lowered into position.
4. Install engine mount bolts, washers and nuts,
then remove the hoist and tail boom support stands.
Torque bolts to 450-500 lb-in.
5. Route throttle, mixture and propeller governor controls to their respective units and cOlUlect.
Secure controls in position with clamps.
6. Connect hydraulic hoses at firewall.
7. Cmmect engine primer line at firewall.
8. Connect cylinder fuel drain lines at hose connection on each side of engine.
9. Connect oil pressure hose at firewall.
10. Connect fuel-flow gage hose at firewall.
11. Connect fuel supply hose and vapor return
line at tunnel and firewall.
NOTE
Throughout the aircraft fuel system, from
the fuel tanks to the engine-driven fuel
pump, use RAS-4 (Snap-On Tools Corp. ,
Kenosha, Wisconsin), MIL-T-5544 (Thread
Compound, Anti seize, Graphite-Petrolatum)
or equivalent, as a thread lubricant or to
seal a leaking connection. Apply sparingly
to male fittings only, omitting the first two
threads. Always ensure that a compound,
the residue from a previously used compound or any other foreign material cannot
enter the system. Throughout the fuel injection system, from the engine-driven pump
through the discharge nozzles, use only a
fuel soluble lubricant, such as engine lubricating oil, on fitting threads. Do not use
any other form of thread compound on the
injection system fittings.
12. Connect manifold pressure line at firewall.
13. Connect vacuum hose at suction relief valve.
14. Install all clamps and lacings securing hoses
and lines to the engine or structure.
15. Connect ground strap to engine mount.
16. Connect exhaust gas temperature wires to
probe leads. Be sure wires are not crossed.
17. Connect electrical wires to throttle-operated
switch.
18. Connect wires and wire shielding ground to
alternator.
19. Connect cylinder head temperature wire to
probe .
10-11
ItAUT(ON\
When connecting starter cable, do not permit
starter terminal bolt to rotate. Rotation of
the bolt could break the conductor between
bolt and field coils causing the starter to be
inoperative.
20. Connect starter electrical cable at starter.
21. Connect tachometer pick-up at bottom of
right magneto.
22. Connect oil temperature Wire at sending unit.
23. Install all clamps and lacings securing Wires
and cables to the engine or structure.
24. Connect exhaust stacks at both sides of engine.
25. Route the fuel strainer drain control through
the nose gear tunnel structure to the strainer, install
the lock nuts to secure housing and connect control
wire to strainer control bellcrank.
26. Install propeller and spinner in accordance
with instructions outlined in Section 12.
27. Complete a magneto switch ground-out and
continuity check, then connect primary lead wires to
the magnetos. Remove the temporary ground or connect spark plug leads, whichever procedure was used
during removal.
IWARNING,
Be sure magneto switch is in OFF poSition
when connecting switch wires to magnetos.
28. Connect unfeathering accumulator hose at
accumulator and service accumulator in accordance
with Section 12.
29. Clean induction air filter and install filter
and induction air inlet duct.
30. Service engine with proper grade and quantity
of engine oil. Refer to Section 2 if engine is new,
newly overhauled or has been in storage.
31. Check all switches are in the OFF position,
install battery box and battery and connect cables.
32. Rlg engine controls in accordance with paragraph 10-41.
33. Inspect engine installation for security, correct routing of controls, lines, hoses and electrical
wiring, proper safetying and tightness of all components.
34. Install engine cowling in accordance with
paragraph 10-3.
35. Check cowl flaps and rig in accordance with
paragraph 10-33, if necessary.
36. Perform an engine run-up and make final adjustments on the engine controls.
b. REAR. Before installing the rear engine on the
aircraft, reinstall any items which were removed
from the engine or aircraft after the engine was removed.
NOTE
Remove all protective covers, plugs, caps
and identification tags as each item is connected or installed. Omit any items not
present on a particular engine installation.
10-12
1. Hoist the engine assembly to a point near the
engine mount and carefully route controls, lines,
hoses and wires in place as the engine is positioned
in the mount. Be sure the shock-mount pads are in
place as the engine is lowered into position. (Refer
to figure 10-1.)
2. Install engine mount bolts, washers and nuts,
then remove the hoist. Torque bolts to 450- 500 lb. in.
3. Route throttle, mixture and propeller governor controls to their respective units and connect.
Secure controls in position with clamps.
4. Connect engine primer line at firewall.
5. Connect cylinder fuel drain line at hose connections on each side of engine and engine-driven fuel
pump drain line.
6. Connect oil pressure hose at firewall.
7. Connect fuel flow gage hose at firewall.
•
NOTE
Throughout the aircraft fuel system, from
the fuel tanks to the engine-driven fuel
pump, use RAS-4 (Snap-On Tools Corp.,
Kenosha, Wisconsin), MlL-T-5544 (Thread
Compound, Anti seize, Graphite-Petrolatum)
or equivalent, as a thread lubricant or to
seal a leaking connection. Apply sparingly
to male fittings only, omitting the first two
threads. Always be sure that a compound,
the residue from a previously used compound or any other foreign material cannot
enter the system. Throughout the fuel injection system, from the engine-driven fuel
pump through the discharge nozzles, use
only a fuel soluble lubricant, such as engine
lubricating oil, on the fitting threads. Do
not use any other form of thread compound
on the injection system fittings.
8. Connect fuel supply hose to auxiliary pump
and vapor return hose at firewall.
9. Connect manifold pressure line at firewall.
10. Connect vacuum pump hose at firewall.
11. Connect hydraulic hoses at pump.
12. Connect propeller unfeathering accumulator
hose at accumulator and service accumulator in
accordance with instructions ouUined in Section 12.
13. Install all clamps and lacings securing hoses
and lines to engine, engine mount or structure.
14. Route the strainer drain control through fuselage structure to the strainer, install control housing
lock nuts securing housing to structure and connect
control wire to strainer.
15. Connect ground strap to engine mount.
16. Connect exhaust gas temperature Wires at
probe leads. Be sure wires are not crossed.
17. Connect electrical wires at throttle-operated
switch.
18. Connect electrical wires and wire Shielding
ground at alternator.
19. Connect cylinder head temperature wire at
probe.
•
•
•
When connecting starter cable, do not permit starter terminal bolt to rotate. Rotation of the bolt could break the conductor
between bolt and field coils causing the
starter to be inoperative.
20. Connect starter electrical cable at starter.
21. Connect tachometer pick-up at bottom of
right magneto.
22. Connect oil temperature wire at sending unit.
23. Connect propeller synchronizer control at
actuator.
24. Connect exhaust stacks at both sides of engine.
25. Install all clamps and laCings securing wires
and cables to engine, engine mount or structure.
26. Install propeller and spinner in accordance
with instructions outlined in Section 12.
27. Complete a magneto switch ground-out and
continuity check, then connect primary ground or
connect spark plug leads, whichever procedure was
used during removal.
IWARNING,
Be sure magneto switch is in OFF position
when connecting primary leads to magnetos.
•
28. Install all flexible ducts.
29. Service air filter and install .
30. Service engine with proper grade and quantity
of engine oil. Refer to Section 2 if engine is new,
newly oJerhauled or has been in storage.
31. Check all switches are in the OFF position
and connect battery ground cable.
32. Rig engine controls in accordance with paragraph 10-41.
33. Check engine installation for security, correct routing of controls, lines, hoses and electrical
wiring, proper safe tying and tightness of all components.
34. Install engine cowling in accordance with
paragraph 10-3.
35. Check cowl flaps and rig in accordance with
paragraph 10-33, if necessary.
36. Perform an engine run-up and make final adjustments on the engine controls.
4. Hoses found leaking should be replaced.
5. After pressure testing fuel hoses, allow sufficient time for excess fuel to drain overboard from
the engine manifold before attempting an engine start.
6. Refer to paragraph 10-13 for detailed inspection procedures for flexible hoses.
10-18. REPLACEMENT.
a. Hoses should not be twisted on installation.
Pressure applied to a twisted hose may cause failure
or loosening of the nut.
b. Provide as large a bend radius as possible.
c. Hoses should have a minimum of one-half inch
clearance from other lines, ducts, hoses or surrounding objects or be butterfly clamped to them.
d. Rubber hoses Will take a permanent set during
extended use in service. Straightening a hose with a
bend having a permanent set will result in hose cracking. Care should be taken during removal so that
hose is not bent excessively, and during reinstalla- .
tion to assure hose is returned to its original position.
e. Refer to AC 43.13-1, Chapter 10, for additional
installation procedures for flexible fluid hose assemblies.
10-19. ENGINE BAFFLES.
10-20. DESCRIPTION. The sheet metal baffles installed on the engine direct the flow of air around
the cylinders and other engine components to provide
optimum cooling. These baffles incorporate rubberasbestos composition seals at points of contact with
the engine cowling and other engine components to
help confine and direct the airflow to the desired area.
It is very impOrtant to engine cooling that the baffles
and seals are in good condition and installed correctly. The vertical seals must fold forward and the side
seals must fold upwards. Removal and installation
of the various baffle segments is possible with the
cowling removed. Be sure that any new baffles seal
properly.
10-21. CLEANING AND INSPECTION. The engine
baffles should be cleaned with a suitable solvent to
remove oil and dirt.
NOTE
The rubber-asbestos seals are oil and grease
resistant but should not be soaked in solvent
for long periods.
10-16. FLEXmLE FLUID HOSES.
•
10-17. PRESSURE TEST.
a. After each 50 hours of engine operation, all flexible fluid hoses in the engine compartment should be
pressure tested as follows:
1. Place mixture control in the idle cut-off position.
2. Operate the auxiliary fuel pump in the high
position.
3. Examine the exterior of hoses for evidence
of leakage or wetness.
Inspect baffles for cracks in the metal and for loose
and/or torn seals. Repair or replace any defective
parts.
10-22. REMOVAL AND INSTALLATION. Removal
and installation of the various baffle segments is possible with the cowling removed. Be sure that any replaced baffles and seals are installed correctly and
that they seal to direct the airflow in the correct direction. Various lines, hoses, wires and controls
are routed through some baffles. Make sure that
these parts are reinstalled correctly after installation of baffles.
10-13
10-23. REPAIR. Repair of an individual segment of
engine baffle 1s generally impractical, since, due to
the small size and formed shape of the part, replacement 1s usually more economical. However, small
cracks may be stop-drilled and a reinforcing doubler
installed. Other repairs may be made as long as
strength and cooling requirements are met. Replace
sealing strips if they do not seal properly.
Inspect the metal parts for cracks and excessive
wear due to aging and deterioration. Inspect the
rubber pads for separation between the pad and
metal backing, swelling, cracking or a pronounced
set of the pad. Install new parts for all parts that
show evidence of wear or damage.
10-24. ENGINE MOUNT. (Refer to figure 10-1.)
10-30. DESCRIPTION.
a. (THRU AIRCRAFT SERIAL 337-0978.) The
front and rear cowl flaps are electrically operated
by small motors attached to torque tubes which actuate the cowl flaps through mecbanicallinkage. The
cowl flap control levers at the instrument panel
operate "floating" switch assemblies located just
forward of the panel. As either lever is moved
DOWN, a microswltch is contacted, which closes
the motor circult, thus closing the cowl flaps. As
the torque tube rotates, mechanical linkage (followup control) from the torque tube causes the microswitch mounting arm to pivot away from the cam
unW the circult is opened. As either lever is
moved UP, a second micro switch is contacted, which
reverses the motor direction and opens the cowl flaps
in a similar manner. BEGINNING WITH AIRCRAFT
SERIAL 337-0756, a second set of switches are installed on the motor arm of the REAR cowl flap motor
to prevent overtravel.
b. (BEGINNING WITH AIRCRAFT SERIALS 3370979 AND F33700001.) The front and rear cowl
flaps are electrically operated by small motors attached to torque tubes which actuate the cowl flaps
through mechanical linkage. Two three-poSition
switches with indicator lights are located on the
lower instrument panel left of the elevator trim control wheel. Full open and closed poSitions of the
cowl flaps are controlled by limit switches on the
cowl flap motor. To operate the cowl flap at an intermediate position, place the switch to the OFF position before the cowl flaps reach their extreme limits.
When the cowl flaps are in operation, a blue indicator
light is on; when the cowl flap reaches the full open or
closed position, the blue indicator light turns off.
There is one indicator light for each cowl flap. The
indicator light for the front cowl flap is to the left of
the switch and the indicator light for the rear cowl
flap is to the right of the switch.
10-25. DESCRIPTION. The rear engine mount is
composed of sections of steel tubing welded together
and reinforced with gussets. The mount forms a truss
structure, fastened to the fuselage at four points,
which supports the engine through a cradle arrangement. This contrasts with the front engine mown,
which 1s an integral part of the lower nose section.
Both engines are attached to the engine mounts with
shock-mount assemblies which absorb engine vibrations. The rear engine mount is so designed that a
severe forward impact, such as in a crash landing,
will cause the rear engine to fall below the cabin.
10-26. REMOVAL AND INSTALLATION. Removal
of the rear engine mount is accomplished by removing the engine, then removing the mount from the
fuselage. On reinstallation torque the engine-tomount bolts to 450-500 lb-in. Torque the mount-tofuselage bolts to 160-190 Ib-in.
10-27. REPAIR. Repair of the rear engine mount
shall be performed carefully as outlined in Section
16. The mount shall be painted with heat-resistant
black eoamel after welding or whenever the original
finish has been removed. This will prevent corroalon.
10-28. ENGINE SHOCK-MOUNT PADS. (Refer to
figure 10-1.) The bonded rubber and metal shockmounts are designed to reduce transmission of engine vibrations to the airframe. The rubber pads
should be wiped clean with a clean dry cloth.
NOTE
Do not clean the rubber pads and dampener
assembly with any type of cleaning solvent.
10-14
•
10-29. COWL FLAPS.
•
•
•
1. Engine Mounting Lug
2.
3.
4.
5.
6.
Mount
Spacer
Spacer
Engine Mounting Pad
Washer
7. Bolt
8. Rear Engine Mounting Structure
9. Firewall
NOTE
FRONT ENGINE (TYPICAL 4 PLACES)
o NON-TURBOCHARGED AmCRAFT ONLY
Attach one end of the ground strap for each engine
under an alternator mounting nut and the opposite
end to the nearest engine mount pad bolt.
Torque engine-to-mount bolts to 450 -500 Ib-in.
Torque rear engine mount-to-firewall bolts to
160 -190 Ib-in•
•
ENGINE- TO-MOUNT
(TYPICAL 4 PLACES)
•
REAR ENGINE
MOUNT-TO-FmEWALL
(TYPICAL-UPPER AND LOWER)
Figure 10-1. Engine Shock-Mounts
10-15
10-31. TROUBLE SHOOTING.
TROUBLE
COWL FLAPS DO NOT
OPERATE.
PROBABLE CAUSE
REMEDY
Battery or master switch in OFF
position.
Check visually. Turn switch ON.
Circuit breaker popped or fuse
blown.
Check visually. Reset breaker.
If it pops again, determine cause
and correct.
Defective circuit breaker.
Check continuity. Replace
circuit breaker.
Defective wiring or defective
switch at instrument panel.
Pull continuity check on wiring
and switch. Replace wiring,
•
replace switch.
INTERMITTENT OR
ERRATIC OPERATION.
CIRCUIT BREAKER
POPS REPEATEDLY
OR FUSE BLOWS.
COWL FLAPS DO NOT
CLOSE COMPLETELY.
10-16
Defective, loose or improperly
adjusted operating switches.
Replace, adjust or secure
switches as required.
Defective cowl nap motor.
Check voltage to motor.
Replace motor.
Disconnected or broken
linkage.
Check visually. Correct or
replace linkage.
Follow-up control slipping in
clamps, or broken or disconnected control.
Check visually. Connect and
secure control. Replace if
defective.
Loose electrical connection.
Tighten loose connections.
Defective, loose or improperly
adjusted operating switches.
Replace, adjust or secure
switches as required.
Control housing loose in clamp
blocks.
Check visually. Adjust and
secure controls.
Defective cowl nap motor.
Use jumper wires to test motor.
Replace motor.
Disconnected or broken
actuator linkage or follow-up
control.
Check visually. Connect or
replace actuator linkage or
follow-up control.
Cowl flaps close too tightly.
Flaps should be adjusted to
close snugly. Rig in accordance with paragraph 10-33.
Defective or improperly adjusted
operating switches.
Replace or adjust operating
switches.
Incorrect rigging.
Refer to paragraph 10- 33.
Incorrectly adjusted cowl
nap push-pull rods.
Rig in accordance with
paragraph 10-33.
•
•
•
1. Follow-Up Control
2. Switch Mounting Arm (Front)
3. Spacer
4. Switch Mounting Arm (Rear)
5. Spring
6. Follow-Up Control (Rear)
7. Switch Actuating Cam (Rear)
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.
34.
35.
36.
37. Torque Tube (Front)
38. Instrument Panel
39. Indicator Light
40. Cowl Flap Switch (Front)
41. Cowl Flap Switch (Rear)
42. Clinch Plate
43. Switch (OPEN-LIMIT)
44. Switch Mounting Bracket
45. Switch Mounting Bracket
46. Switch (CLOSED-LIMIT)
47. Switch Actuating Bracket
48. Torque Tube (Rear)
49. Vertical Push-Pull Rod
50. Engine Mount
51. Stop
52. Bellcrank
53. Horizontal Push-Pull Rod
54. Grommet
Control Lever (Rear)
Knob
Cowl Flaps CLOSED Switch
Cowl Flaps OPEN Switch
Bolt
Bracket
Detent Spring
Switch Actuating Cam (Front)
Control Lever (Front)
Clamp
Bracket
CleviS
Stationary Link
Bracket
Bearing
21
Washers
Torque Tube Arm
Rod End
Shock-Mount
Bracket
Clevis
Push-Pull Rod
Ball Joint
Motor Arm
Pin
TURBOCHARGED
Clamp
AIRCRAFT
Motor Assembly
Torque Tube Arm
Detail
Rod End
20
55.
56.
57.
58.
59.
60.
Shield
Cowl Flap
Stop
Adjllstment Screw
Switch Mounting Bracket
Switch Mounting Plate
o Use washers (23) as required
at each end to align bellcranks.
SLOTTED HOLE .
B
22 __--7/L,,;I
023_--7//.
24
25
29
33
H~.:
•
~27
34
35
Detail
Detail
A
B
NON-TURBOCHARGED
AIRCRAFT
THRU AIRCRAFT SERIAL 337-0978
FRONT COWL FLAPS
'* excess
Use as required to shim out
play.
Figure 10-2. Cowl Flap Installation (Sheet 1 of 4)
10-17
38
Detail
20
21
•
39 40 41 39
A
19
.010 " max. between actuator
leaf and actuator pin
22--~
VIEW
A-A
o 23----ut
.80" max.
(Open position)
24------.[~
e:..
•
,---O_ _0--..l.-L
25
47
29
~
.
28
°
Detail
B
• • 05 " min. at
position
011"
mak~ J
Switch body to be parallel with cam (47)
.Due to the cowl flap motor overrun, switch rollers must be rigged to allow roller to continue
travel down cam (47) and along BEGINNING WITH AIRCRAFT SERIALS
337-0979 AND F33700001
fiat inboard side of cam after
FRONT COWL FLAPS
switch makes.
Figure 10-2. Cowl Flap Installation (Sheet 2 of 4)
10-18
BEGINNING WITH
AmCRAFT SERIALS
33701463 AND F33700056
VIEW
SaB
•
•
,.
NOTE
.••....•..•.•...•..•........•......
Access to screws securing stop (57) may be
gained by unscrewing wing flap control knob,
removing cowl flap control knobs (8 and 16),
removing screws securing left end of panel
cover and Hexing cover aft to expose stop
screws (58). Thru aircraft serial 337-0239,
remove wing flap control knob, cowl flap
control knobs, screws securing right end of
console cover and flex cover aft to expose
stop screws. Beginning with aircraft serial
337 -0526 and F33700001, non-adjustable
stops are used.
~
.»........
......
........
:'..
.
;
--...
.'.: : : : : :. . . . . . . . . . . .~~:::.:~:;.::r. . .,-
".
:
/
,
...
.
'
....!
.......... ..
\...........
.......•~" ...•.•....
~
'.
.....
'
~.:.:.:.:.:.'..,:.:.:.... ....•..•...•.
..:....
...,;~::::::.:::::::
............. .
.. ----,
"'( / c
-----~.:---
-:
.......
~:! ~··~::::)t,"··"
-..
....
.......
................
.':'Qi
......
(.or..·.i
• THRU AIRCRAFT SERIAL 337-0978
'.
•
B
32
ADJUSTMENT SLOTS
8
•
",
Detail
C
THRU AmCRAFT SERIAL 337-0755 WHEN NOT
MODIFIED IN ACCORDANCE WITH SK337-8
REAR COWL FLAPS
Figure 10-2. Cowl Flap Installation (Sheet 3 of 4)
10-19
...... .. '
......
. ...... .
: ~ : : ~:-.~.: :~: :..~,..............::.....:,;;.:" ,!
),.,.
..~.::::.,.....,..
c
..... ..
A
&)-.. , .~S·""
~ .2(.. .·.......· ·,. ·::L:·{·::·,··, .................
··················fi······
"....:.. ..., ,
...... ,
.
,/::::.::. .-:::.:,(
,'......
.
I.,. "',
~".'.,'
..\,\\, \,
.
\,
'-- ........
," ~.......
\~. \
\
...::><..., B
\\,.\. :.\......
.•.
.............
,."......
"'~.,........
.........
'~\.,.~
\~
33
to''\'",;'·:,)
47
"l..~ •
IT
,.~\/.'
u-I
AIRC~F;' ~;RIALS
337 -0756 THRU 337 _
0906 AND ALL AIRCRAFT MODIFIED IN
ACCORDANCE WITH
SK337-8
32
59
...........
........... ,
~~:s
43
AIRCRAFT SERIAL 3370907 THRU 337-1130
'>
42
_Jr~~
,
.;X!J
•
21
30
~
53
rJY)~
Detail
C
18
20
49
•
56
3&
BEGINNING WITH AIRCRAFT SERIALS
337-1131 AND F33700001
,
"~~50
~~2 ~
~~
.r
.
Insulators are installed between
switches
and 46) and brackets
(59) beginning with aircraft serials
337p1463 and F33700056
Figure 10-2. Cowl Flap Installation (Sheet 4 of 4)
10-20
NOTE
.,
51
52
53
REAR COWL FLAPS
VIEWC<
1--22
(4~
•
10-32. REMOVAL AND INSTALLATION. (Refer to
figure 10-2.)
a. FRONT.
1. Run flaps to OPEN pOsition.
2. Disconnect push-pull rods (29) at shockmounts (26).
3. Remove screws securing cowl flap hinges to
lower fuselage structure.
4. Reverse the preceding steps for reinstallation. Rig flaps, if necessary, in accordance with
paragraph 10-33, check that flaps move freely and
clear all adjacent parts.
b. REAR.
1. Run flaps to OPEN position.
2. Disconnect push-pull rods (53) at cowl flaps
(56).
3. Drill out rivets securing hinges and remove
cowl flaps and hinges as a unit.
4. Reverse the preceding steps for reinstallation. Rig flaps, if necessary, in accordance with
paragraph 10-33, check that flaps move freely and
clear all adjacent parts.
10-33. RIGGING.
a. FRONT. (THRU AIRCRAFT SERIAL 337-0978.)
(Refer to figure 10-2, sheet 1.)
I"fAUTION]
•
The battery/master switch must be turned
off before rigging the cowl flaps. If electrical power is applied before rigging has
been completed, the protective fuse will
blow or damage may occur to the cowl flap
motor (34), motor arm (31) or torque tube
(37).
1.
Ensure battery/master switch is in OFF posi-
8. Using jumper wires and a 24-volt dc external
power supply, operate motor to place motor arm (31)
in the position illustrated in figure 10-3.
/CAUTION\
Do not try to use master switch before rigging has been completed. When using the
jumper wires, connect only one wire to a
motor lead and "strike" the other jumper
wire against the remaining motor lead.
The motor arm moves rapidly. If it does
not move in the desired direction, reverse
jumper leads.
9. Remove jumper wires and connect electrical
leads separated in step 7 making sure the leads are
not connected in reverse. The direction of movement
will be reversed and the protective fuse will blow if
the system is operated with crossed leads. Insulate
quick-disconnect with plastiC tubing and tie wires .
back to prevent interference with other equipment.
10. Place the front cowl flap control lever in
CLOSED position. Move the lower end of switch
mounting bracket (2) away from instrument panel
until the switch nearest the firewall is de-activated
(switch just breaks the clOSing cirCuit). Adjust the
follow-up control rod end (36) to align with attaching
hole in torque tube arm (35) and install.
NOTE
Opening of the microswitch may be determined
by listening for a faint "click, " or continuity
may be checked. Follow-up control may be
adjusted by slipping it in its clamps (17) or adjusting rod end (36). If rod end is adjusted,
maintain sufficient thread engagement.
tion.
•
2. Remove right and left cowling panels for access.
3. Center the FRONT cowl flap control lever
cam (15) between switch rollers (the switch nearest
the instrument panel is mounted in a slotted hole for
adjustment), with rollers just clearing cam (15).
4. Disconnect push-pull rods (29) between torque
tube arms (24) and cowl flaps, at the torque tube
arms. Tape the rods where they cannot cause damage as the linkage is moved.
5. Disconnect follow-up control (1) at torque
tube arm (35).
6. Disconnect stationary link (20) between cowl
flap motor arm (31) and firewall bracket (18). Adjust link to the length specified in figure 10- 3, tighten
jam nuts and reinstall stationary link.
7. On aircraft without a radio noise iilter installed at the cowl flap motor, tag and disconnect the
motor leads at the motor assembly by separating the
quick-disconnects. On aircraft equipped with a radio
noise filter installed at the cowl flap motor, separate
one motor lead at the fuse holder and cut the other
motor lead in the approximate area of the fuse holder .
Install a quick-disconnect connector on each end of
'the cut wire.
11. Place battery/master switch in the ON position and operate cowl flap motor through several
cycles, checking for interference between torque tube
and linkage.
NOn:
Check that follow-up control (1) does not
cause automatic cycling (continuous opening and closing), which indicates that switch
rollers are set to actuate too soon. Manually "rock" torque tube (37) to remove any lost
motion in torque tube and follow-up control
(1). If automatic cycling occurs, readjust
switches as specified in step 3, setting switch
rollers with slightly more clearance from cam,
as necessary to prevent automatic cycling,
then repeat the rigging procedure.
12. Operate the cowl flaps to the CLOSED pOSition and place battery/master switch in OFF position.
13. Manually holding cowl flaps closed (snugly),
adjust push-pull rods (29) to align with attaching holes
in torque tube arms (24) and install bolts.
10-21
f
•
3.25 + .25 - . 00 ..
4
5
3--Ii-UC
\"--.A J. _ _../.,
2--T.-H
•
6---1
FRONT
l.
2.
3.
4.
5.
6.
Front Firewall (Station 65.00)
Stationary Link Rod
Torque Tube
Torque Tube Arm
Cowl Flap Motor Arm
Rear Firewall (Station 186.00)
Figure 10-3.
10-22
Cowl Flap Rigging
•
•
NOTE
The front cowl fiaps are streamlined with the
fuselage in the CLOSED position, and the
cowl flaps are open 4. 50±. 25 inches in the
OPEN position, measured at the midpoint of
the flap trailing edge to a corresponding
point on lower edge of the fuselage.
14. Check that all rod ends and clevis ends have
sufficient thread engagement, all jam nuts are tight
and all safeties are installed. Reinstall cowling.
NOTE
Refer to Section 3 for rigging of cowl flap
doors on non-turbocharged aircraft equipped with a cargo pack.
b. FRONT. (AIRCRAFT SERIALS 337-0979 THRU
33701462 AND F33700001 THRU F33700055.) (Refer
to figure 10-2, sheet 2.)
1. Complete steps 2, 3, 4, 6 and 7 of subparagraph "a. "
2. Using jumper Wires and a 24-volt dc external
power supply, operate motor to place motor arm (31)
in the position illustrated in VIEW A-A.
•
Do not try to use master switch before rigging has been completed. When using the
jumper wires, connect only one wire to a
motor lead and "strike" the other jumper
wire against the remaining motor lead.
The motor arm moves rapidly. If it does
not move in the desired direction, reverse
jumper leads.
3. Loosen CLOSED-LIMIT switch bracket (45),
adjust CLOSED-LIMIT switch (46) and bracket (45)
toward switch actuating bracket (47) unW switch just
de-actuates. Secure bracket and switch in this position.
6. Loosen OPEN-LIMIT switch bracket (44),
adjust OPEN-LIMIT switch (43) and switch bracket
(44) toward switch actuating bracket (47) unW switch
justs de-actuates. Secure bracket and switch in this
position.
NOTE
Opening of the micro switch may be determined by listening for a "click" or by
checking continuity.
7. Complete step 9 of subparagraph "a. "
8. Turn master switch ON and operate cowl flaps
through several cycles. Check position indicating
lights for operation. Stop cowl flaps at intermediate
positions to check toggle switch. Check for interference between torque tube and linkage.
9. Check that all rod ends and clevis ends have
sufficient thread engagement, all jam nuts are tight,
all safeties are installed and install cowling.
c. FRONT. (BEGINNING WITH AIRCRAFT SERIALS 33701463 AND F33700056.) (Refer to figure 10-2,
sheet 2.)
1. Complete steps I, 2, 4, 6 and 7 of subparagraph "a. "
2. Complete step 2 of subparagraph ''b.''
3. Loosen CLOSED-LIMIT switch bracket (45),
adjust CLOSED-LIMIT switch (46) and bracket (45) to
dimensions shown in VIEW B-B and secure bracket
and switch.
NOTE
Opening of the micro switch may be determined by listening for a "click" or by
checking continuity.
4. Complete steps 4 and 5 of subparagraph "b. "
5. Loosen OPEN-LIMIT switch bracket (44),
adjust OPEN-LIMIT switch (43) and switch bracket
(44) to dimensions shown in VIEW B-B and secure
bracket and switch.
NOTE
NOTE
Opening of the micro switch may be determined
by listening for a "click" or by checking continuity.
4. Hold cowl flaps closed (snugly), adjust pushpull rods (29) to align with torque tube arm (24) attaching holes and install bolts.
5. Using j'.lJDper Wires am external power supply, run cowl flaps to full OPEN position, observing
"CAUTION" in step 2.
NOTE
•
The front cowl flaps are streamllned with the
fuselage when in the CLOSED poSition and
4. 50±. 25" when in the OPEN position, measured at the midpoint of the flap trailing edge
to a corresponding point on the lower edge of
the fuselage.
Opening of the micro switch may be determined by listening for a "click" or by
checking continuity.
6, Complete step 9 of subparagraph "a."
7 • Complete steps 8 and 9 of subparagraph "b. "
d. REAR. (THRU AIRCRAFT SERIAL 337-0755
WHEN NOT MODIFIED IN ACCORDANCE WITH
SK337-8.) (Refer to figure 10-2, sheet 3.)
@AUT~oNI
The battery/master switch must be turned
OFF before rigging the cowl flaps. If electrical power is applied before rigging has
been completed, the protective fuse will
blow or damage may occur to the cowl flap
motor (34), motor arm (31) or torque tube
(48).
10-23
1. Ensure battery/master switch is in OFF
position.
2. Disconnect horizontal push-pull rods (53) at
cowl naps (56).
3. Remove right and left cowling side panels.
4. Center the REAR cowl flap control lever cam
(index 7, sheet 1) between switch rollers (the switch
nearest the instrument panel is mounted in a slotted
hole for adjustment), with rollers just clearing cam
5. Disconnect vertical push-pull rods (49) at the
torque tube arms (24). Tape rods to prevent damage
When linkage is moved.
6. Disconnect follow-up control (1) at torque
tube (24).
7. Disconnect stationary link (20) between cowl
flap motor arm (31) and firewall bracket (18). Adjust
link to the length specWed in figure 10-3, tighten jam
nuts and reinstall stationary link.
8. On aircraft without a radio noise filter installed at the cowl flap motor, tag and disconnect the motor leads at the motor assembly by separating the
quick-disconnects. On aircraft equipped with a radio
noise filter installed at the cowl flap motor, separate
one motor lead at the fuseholder and cut the other
motor lead in the approximate area of the fuseholder.
Install a quick-disconnect connector on each end of
the cut wire.
9. Using jumper wires and a 24-volt dc external
power supply, operate motor to place motor arm (31)
in the pOSition illustrated in figure 10-3.
To not try to use master switch before rigging has been completed. When using the
jumper wires, connect only one wire to a
motor lead and "strike" the other jumper
wire against the remaining motor lead.
The motor arm moves rapidly. If it does
not move in the desired direction, reverse
jumper leads.
10. Remove jumper wires and connect electrical
leads separated in step 8 making sure the leads are
not connected in reverse. The direction of movement
will be reversed and the protective fuse will blow if
the system is operated with crossed leads. Insulate
quick-disconnect with plastic tubing and tie Wires
back to prevent interference With other equipment.
11. Place the REAR cowl flap control lever in
CWSED position. Move the lower end of switch
mounting bracket (index 4, sheet 1) away from instrument panel until the switch nearest the firewall is deactuated (switch just breaks the closing circuit). Adjust the follow-up control rod end to align with attaching hole in torque tube arm (24) and install.
NOTE
Opening of the micro switch may be determined
by listening for a faint "click, " or continuity
may be checked. Follow-up control may be adjusted by slipping it in its clamps or adjusted,
maintain sufficient thread engagement.
10-24
12. 1'1ace battery/master switch in the ON position and operate cowl flap motor through several
cycles, checking for interference between torque
tube and linkage.
NOTE
•
Check that follow-up control does not cause
automatic cycling (continuous opening and
closing), which indicates that switch rollers
are set to actuate too soon. Manually "rock"
torque tube (48) to remove any lost motion in
torque tube and follow-up control (1). If automatic cycling occurs, readjust switches as
specified in step 4, setting switch rollers
the minimum amount of clearance from cam
necessary to prevent automatic cycling, then
repeat the rigging procedure.
13. Operate the cowl flaps to the OPEN poSition
and place battery/master switch in the OFF position.
14. Connect vertical push-pull rods (49) to the
torque tube arms (24).
15. Install right and left cowling side panels and
connect horizontal push-pull rods (53) to cowl flaps
(56).
16. Place battery/master switch in the ON position and slowly close cowl flaps, checking that linkage is not adjusted to close the cowl flaps tight enough
to cause damage.
17. Operate cowl flaps to the single-engine position and measure travel at trailing edge. The cowl
flaps should open 5.50+.25-.12 inches, but still close
snugly. Readjust. push-pull rods (49 and 53) to bellcranks (52) and cowl flaps (56) and select a different
hole in the bellcranks as required for a snug fit and
proper travel. Check that the stops (51) on the bellcranks just clear the engine mount tubes. Lower
wing flaps cautiously with rear cowl flaps full OPEN
and check for at least 3/8 inch clearance in any wing
flap position.
18. Place rear cowl nap control lever in the NORMAL OPEN pOsition (twin-engine operation) and check
that the cowl flaps are open 3. 50 ± .25 inch, measured
at the trailing edges. Thru aircraft serial 337-0525,
it may be necessary to readjust control lever stop as
shown on sheet 3. Beginning with aircraft serial
337-0526, non-adjustable stops are used.
19. Check that all rod ends and clevis ends have
sufficient thread engagement, all jam nuts are tight
and all safeties are installed.
e. REAR. (AIRCRAFT SERIALS 337-0756 THRU
337-0978 AND ALL AIRCRAFT MODIFIED IN ACCORDANCE WITH SK337-8.) (Refer to figure 10-2.)
•
I~AUTIONI
The battery/master switch must be turned
OFF before rigging the cowl flaps. If
electrical power is applied before rigging
has been completed, the protective fuse
will blow or damage may occur to the cowl
flap motor (34), motor arm (31) or torque
tube (48).
.
•
•
1. Remove upper cowling sections.
2. Make sure the limit switches (index 43 and
46, sheet 4) do not actuate, limiting travel of the
cowl flap motor during the following rigging procedures. Readjust switches to clear motor arm (31),
if necessary.
3. Complete steps 1 thru 18 of subparagraph
"d."
4. After rigging the rear cowl flaps as outlined,
rig the limit switches as follows:
NOTE
When moving the flap motor arm (31) up or
down to adjust limit switches (43 and 46),
disconnect motor lead at quick-disconnect
or fuse holder, place flap control in cabin
in up or down position and control the motor
by momentarily making contact at quickdisconnect or fuse.
•
5. Place battery/master switch in the ON position and rear cowl flap control in the full OPEN position. After motor reaches limit of travel, adjust
switch (43) until switch roller makes contact with
switch actuating bracket (47). Secure switch in this
position.
6. Place rear cowl flap control in CLOSED position. After motor reaches limit of travel, adjust
switch (46) until switch roller makes contact with
switch actuating bracket (47). Secure switch in this
position .
NOTE
When making this adjustment, do not move
switch beyond the point at which the roller
first makes contact with the switch actuator.
These limit switches come into cperation
only in event of failure of control lever
mounted switches.
7. Complete step 19 of subparagraph "a" and
reinstall upper cowl section.
f. REAR. (AIRCRAFT SERIALS 337-0979 THRU
33701462 AND F33700001 THRU F33700055.) (Refer
. to figure 10-2, sheet 4.)
I~~UTION\
The master switch must be turned OFF before
rigging the cowl flaps. If electrical power is
applied before rigging has been completed, the
protective fuse will blow or damage may occur
to the cowl flap motor (34), motor arm (31) or
torque tube (48).
•
1. Remove upper cowling section.
2. Complete steps 1, 2, 3, 5, 7, 8 and 9 of subparagraph "d. "
3. Loosen screws on CLOSED-LIMIT switch (46)
and adjust switch toward actuating bracket (47) until
switch just de-activates. Secure switch in this position.
NOTE
Opening of the microswitch may be determined
by listening for a faint "click, " or continuity
may be checked.
4. Install RIGHT HAND cowl panel, connect
horizontal push-pull rod (53) to cowl flap (56) and
connect vertical push-pull rod (49) to torque tube
arm (24). The cowl flap (56) must be faired with
cowl panel in the closed position. If not, readjust
push-pull rods as necessary.
5. Using jumper Wires and external power supply, run the right hand cowl flap to open 5.50+.25-. 12
inches, measured from the trailing edge of cowl flap
to aft edge of cowl flap opening.
ICAUTIONI
Do not use master switch before rigging has
been completed. When using jumper wires
connect only one wire to motor lead and
"strike" the other jumper wire against
the remaining motor lead. If motor does
not move in the correct direction, reverse jumper leads.
6. Loosen screws on OPEN-LIMIT switch (43)
and adjust switch toward actuating bracket (47) until
SWitch just de-actuates.
NOTE
Opening of the microswitch may be determined
by listening for a faint "click, " or continuity;
may be checked.
7. Complete step 10 of subparagraph "d. "
8. Install LEFT HAND cowl panel, connect vertical push-pull rod (49) to torque tube arm (24) and
horizontal push-pull rod (53) to cowl flap (56). The
cowl flap should be open 5.50+.25- .12 inches in the
open position. If not, readjust push-pull rods as necessary.
9. Place master switch in the ON position and
using cowl flap toggle switch (index 41, sheet 1),
slowly run the cowl flaps to the CLOSED position
and check that the LEFT cowl flap is faired with the
cowl panel. If not, readjust the push-pull rods as
necessary.
NOTE
In all cases, the final result of rigging, is
that the cowl flaps are to be faired with
the cowl panels in the CLOSED position
and are to be open 5.50+.25-.12 inches in
the OPEN position.
10. Using toggle switch (index 41, sheet 1), run
cowl flaps through several cycles. Check pOSition
indicating lights for operation. Stop cowl flaps at
intermediate-openingsto ·check toggle switch operation.
10-25
11. Check that all rod ends and clevis ends have
sufficient thread engagement, all Jam nuts are tight,
all safeties are installed and reinstall upper cowling
section.
g. REAR. (BEGINNING WITH AIRCRAFT SERIALS
33701463 AND F33700056.) (Refer to figure 10-2,
sheet 4.)
If~UTIONl
The master switch must be turned OFF
before rigging the cowl flaps. If electrical power is applied before rigging
has been completed, the protective fuse
will blow or damage may occur to the cowl
flap motor (34), motor arm (31) or torque
tube (48).
1. Remove upper cowling section.
2. Complete steps 1, 2, 3, 5, 7, 8 and 9 of subparagraph "d. "
3. Complete steps 3 and 4 of subparagraph "f."
4. Using jumper Wires and external power supply, run the right hand cowl flap to open 6.00+.25-.00
inches, measured from the trailing edge of cowl flap
to aft edge of cowl flap opening.
[~~UTIO:NI
Do not use master switch before rigging has
been completed. When using jumper wires
connect only one wire to motor lead and
"strike" the other jumper wire against
the remaining motor lead. If motor does
not move in the correct directton, reverse jumper leads.
5. Complete step 6 of subparagraph "f."
6. Complete step 10 of subparagraph "d."
7 . Install LEFT HAND cowl panel, connect vertical push-pull rod (49) to torque tube arm (24) and
horizontal push-pull rod (53) to cowl flap (56). The
cowl flap should be open 6.00+.25-.00 inches in the
open position. If not, readjust push-pull rods as
necessary.
8. Place master switch in the ON position and
using cowl flap toggle switch (index 41, sheet 1),
slowly run the cowl flaps to the CLOSED position and
check that the LEFT cowl nap is faired with the cowl
panel. If not, readjust the push-pull rods as necessary.
10-35. DESCRIPTION. Throttle, mixture and propeller controls for each engine are contained in the
control quadrant. The throttle levers are located at
the left, the propeller levers are in the center and
the mixture levers are at the right. The left lever
of each pair is for the front engine and the right is
for the rear engine. Each pair of knobs has its own
shape and can easily be distinguished from the others.
A knurled friction knob at the right end of the quadrant may be rotated to increase or decrease the
amount of friction on the levers.
•
10-36. REMOVAL AND INSTALLATION. (Refer to
figure 10-4.)
a. Remove console cover in accordance with sheet
2.
b. Remove control lever knobs.
c. Remove slotted cover from quadrant.
d. Disconnect throttle, propeller and mixture controls from quadrant levers. Do not disturb rod end
adjustments. Note which direction each pin and bolt
point, so that they may be reinstalled in the same
position for clearance. Also note which mounting
hole in control lever is used.
e. Remove the quadrarit assembly as a unit by removing the two bolts securing the end plates (1 and
22) at each end of the quadrant.
f. Reverse the preceding steps for reinstallation.
All four quadrant attaching bolts must be installed
with their heads pointing to the right for clearance
with adjacent parts.
10-37. DISASSEMBLY, INSPECTION AND REASSEMbly. (Refer to figure 10-4.)
NOTE
Since the quadrant assembly contains numerous spacers, washers and friction discs~ in
addition to the control levers, note their relative positions before disassembling to aid during reassembly. After removal of the quadrant
assembly, use figure 10-4 as a guide for disassembly and reassembly.
•
Clean all metal parts with a solvent-dampened cloth
(Stoddard or equivalent), then wipe with a clean, dry
cloth. Lubricate only the control levers by applying
a thin film of petrolatum to each side of the levers
within a one-inch radius of their pivot holes. Replace
any defective parts and reassemble the quadrant, positioning the parts as noted during disassembly.
NOTE
10-38. ENGINE CONTROLS.
In all cases, the final result of rigging, is
that the cowl flaps are to be faired with
the cowl panels in the CLOSED position and
are to be open 6.00+.25-.00 inches in the
OPEN position.
9. Complete steps 10 and 11 of subparagraph
"f."
10-34. CONTROL QUADRANT.
10-26
10-39. DESCRIPTION. The throttle, propeller and
mixture controls are located in the control quadrant.
Each set of controls is characterized by different
shaped mobs. A spring-loaded feathering mechanism
is built into the handle of each propeller control. The
propeller control must be lifted and pulled aft to feather a propeller. Controls for the rear engine are
routed to the front firewall, then down into the tunnel
from the front firewall to the rear firewall beneath
•
•
NOTE
'~
When removing propeller knobs (7),
hold down on propeller levers (S and
9) tu prevent loss of internal parts.
I
1 2
•
Z4
22
",(:)
" ~\
14
15
21
1•
.1.1-_-19
1.
2.
3.
4.
5.
6.
7.
S.
9.
10.
11.
12.
13.
•
End Plate
Washer
Roll Pin
Front Throttle Lever
Throttle Knob
Rear Throttle Lever
Propeller Knob
Front Propeller Lever
Rear Propeller Lever
Mixture Knob
Front Mixture Lever
Rear Mixture Lever
Washer
14.
15.
16.
17.
IS.
19.
20.
21.
22.
23.
24.
25.
26.
Spring
Spacer
Friction Disc
Spacer
Spring
Shaft
Arm
Friction Knob
End Plate
Spacer
Stud
Hub
Strap
........--20
Detail
A
Figure 10-4. Control Quadrant (Sheet 1 of 2)
10-27
•
LANDING GEAR
INDICATOR LIGHTS
NOTE
To remove the console cover, the
items listed must be removed
from the locations shown.
ENGINE PRIMERS
-_~
COWL FLAP
CONTROL KNOBS
AUTOPILOT CONTROL
(OPTIONAL)
..
FRICTION KNOB
THRU AmCRAFT SERIAL 337-0239
•
ENGINE PRIMERS --'o::::::::--_~
AUTOPILOT CONTROL
(OPTIONAL)
..
FRICTION KNOB
BEGINNING WITH AmCRAFT
SERIAL 337-0240 (TYPICAL)
Figure 10-4. Control Quadrant (Sheet 2 of 2)
10-28
•
•
SCHEMATIC DIAGRAM OF
TYPICAL ENGINE CONTROL
ENGINE COMPONENT ARM
(TYPICAL)
MECHANICAL STOPS
FIXED BRACKETS
CONTROL MOUNTING
BRACKET (REF)
CONTROL ARM (REF)
THROTTLE, MIXTURE AND PROP
EFFECT OF ADJUSTMENTS:
•
CONTROL~
1.
Lengthening control at either end will furnish more cushion at full forward position of quadrant lever.
2.
Shortening control at either end will furnish more cushion at full aft position of quadrant lever .
3.
Lengthening one end and shortening the other end an equal amount will have no effect on cushion at
quadrant lever. However, this may be necessary to attain full travel before jam nut contacts swivel
end of control.
4.
Control levers in quadrant contain four holes where controls attach. Rig front engine controls with
as short a leverage as possible. Rig rear engine controls with whatever leverage w1ll cause the
quadrant levers to move the same distance in the quadrant as the front levers.
5.
Throttle and mixture control arms at their corresponding engine components may be repOSitioned
on their shafts if necessary. Make sure the countersunk side of the arm faces the serrated portion
of its shaft. If throttle arms are repOSitioned, check rigging of landing gear warning system cam
and microswitch.
I~AUTIONI
Whenever a fuel pump arm or fuel-air control unit arm is removed or installed, always use a wrench at the wrench pads on the arm when removing
or installing attaching nut. This will prevent twisting the shaft or other
damage which might be caused.
RESULTS TO BE ACCOMPLISHED:
•
1.
Arm at engine component must attain full travel, contacting mechanical stops in both directions.
2.
Cushion must be provided at both travel limits of quadrant lever to assure that mechanical stops at
engine component are actually limiting travel during flight.
3.
Quadrant lever knobs should align within one-half knob at cruising power.
4.
Adjust control cable mounting bracket, if required, to prevent the swivel on the control cable
from exceeding 7.5 0 movement from centerline .
Figure 10- 5. Rigging Engine Controls
10-29
access covers which form the center floorboards.
Thru aircraft serial 337-0755, the intake heater controls are located in the right side of the switch panel.
The heater controls are equipped with thumb-button
locks which must be depressed to operate controls.
Beginning with aircraft serials 337-0756 and F33700001, the manual intake heater controls are deleted
from the switch panel. Tbe manual controls are replaced by an automatic device. If the air filter should
become clogged, suction from tbe engine will open a
spring-loaded door in the induction airbox.
10-40. REMOVAL AND INSTALLATION.
a. Remove seats, carpeting and tunnel cover plates
as necessary for access.
b. Remove console cover in accordance with sheet
2.
c. Remove control lever knobs.
d. Remove slotted cover from quadrant.
e. Disconnect throttle, propeller and mixture controls from quadrant levers. Note which direction
each pin and bolt point, so that they may be reinstalled in the same position for clearance. Also note
which mounting hole in control lever is used.
f. Release the multiple clamp securing the controls
between the quadrant and the firewall.
g. Disconnect each control from its respective engine component and remove rod ends, jam nuts and
rubber boots from engine end of controls.
h. Loosen the shield through which the rear engine
controls pass at the horizontal firewall.
1. Remove the controls from clamps and brackets
in the engine compartments and from clamps and tiestraps securing the controls along their routing to the
control quadrant.
j. Pull the controls into the cabin area to remove.
k. Remove either intake heater control as follows:
1. Disconnect the control at the induction airbox lever and straighten the bend in the control wire.
2. Loosen or remove any clamps securing the
control along its routing to the instrument panel.
3. Remove the nut, 10ckWasher and tapered
spacers where the control passes through the instrument panel and pull the entire control assembly into
the cabin to remove.
l'fAuYIONl
Do not pull the control out of its housing While
1t is disconnected. To do so would permit intricate parts of the locking mechanism to fall
out and possibly be lost.
1. Reverse the preceding steps for reinstallation,
then rig each control. Note that large safety washers
are used where the throttle, mixture and propeller
controls are attaChed to the engine component arm.
10-41. RIGGING. (Refer to figure 10-5.) The throttle, propeller and mixture controls are equipped with
adjustable rod ends at the engines and at the quadrant.
Each control contains a small inspection hole through
which the control itself must be visible to ensure suf-
10-30
ficient thread engagement. Since it is easier to adjust rod ends at the engines, attempt rigging at this
end first. However, if correct rigging results cannot be attained by this method, the rod ends at the
quadrant may also have to be adjusted. Figure 10- 5
shows a tyPical engine control, explains what happens
as adjustments are made and gives results which must
be accomplished by rigging.
•
NOTE
Refer to the inspection chart in Section 2 for
inspection and/or replacement intervals for
the throttle, propeller and mixture controls.
Thru aircraft serial 337-0755 the intake heater controls are rigged as follows:
a. Loosen tbe clamp securing control to the airbox.
b. Push the control full forward, then pull it out
apprOximately 1/8 inch for cushion.
c. Shift the control housing in its clamp at the airbox so that the air valve lever is pushed as far as it
will move, with the valve seating inside the airbox.
d. Pull the control out and check that the air valve
inside the airbox seats in the opposite direction.
e. Check that the end of the wire is secured, that
the attaching bolt will swivel and that the wire tip is
bent 90°.
NOTE
When installing a new control or rigging a control whose wire tip has broken off, it may be
necessary to shorten the control hoUsing, although sh1ft1ng the housing in its clamp usually
will permit proper travel.
•
10-42. THROTTLE-OPERATED GEAR WARNING
SWITCHES.
10-43. DESCRIPTION. The landing gear warning
horn will blow whenever either throttle is retarded
while the landing gear is not down and locked. Cams
(one attached to each throttle shaft) actuate microswitches as the throttles are retarded to a manifold
pressure of approximately 13 inches of mercury.
10-44. RIGGING. (Refer to figure 10-6.) If the horn
will not blow after correct rigging, check continuity of
switches and electrical circuit. Adjust the cams and
microswitches as follows:
a. Perform an initial ground adjustment on front
engine as follows:
1. Close throttle and adjust cam as shown in
detail "A. "
2. Refer to detail "B" and set micro switch to be
actuated on the peak of the cam and de-actuated on the
flat portion. Be sure roller arm clears switch body
in actuated position.
3. Start and run engine to approximately 2000
rpm, then reduce power slowly until horn sounds,
noting rpm setting. (Allow tachometer and manifold
pressure needles to stabilize before taking readings. )
•
•
• BETWEEN fl. OF
IDLE STOP AND
CAM PEAK
ADJUSTING SLOTS
(Some early aircraft do
not have slots. It is permissible to slot holes as
needed for switch adjustment. )
____- - INSU LA TOR
IDLE STOP
--.jI~-1
~~~~~t-- CLINCH PLATE 0
/16 inch •
ASSY (Far side)
rIDLESCREW
CAM---...
/
(Against stop)
THROTTLE CONTROL --~
ATTACH POINT
•
o BEGINNING WITH AmCRAFT
SERIALS 337-0906 AND F33700001
DETAIL B
THROTTLE ARM
(Full Retard)
DETAIL A
POSITION COUNTERSUNK FACE OF CONTROL
ARM TOWARD TAPERED PORTION OF CONTROL
SHAFT TO PREVENT 'IWISTING OF SHAFT.
WHEN TIGHTENING NUT, HOLD CONTROL ARM
AT WRENCH PADS.
Figure 10-6. Rigging the Gear Warning System Switches
NOTE
Because the gear is down and locked, it will
be necessary to depress gear-down (green)
indicator light approximately one-hall its
travel distance before warning horn will
sound.
•
4. If horn does not blow between 1650 and 1750
rpm, run engine to 1700 rpm and tighten friction knob
to hold throttle at this setting, then stop engine using
miXture control.
5. Adjust microswitch to actuate at this setting.
6. When desired results are achieved, repeat
procedure on rear engine.
b. Perform flight test at 2500 feet pressure altitude
as follows:
1. Set both propellers at 2300 rpm.
2. Slowly reduce power on front engine until
horn blows and note manifold pressure reading .
(Again allow needles to stabilize.) Horn should blow
between 12.5 and 14 inches of mercury manifold pressure.
3. If horn actuation does not fall within this tolerance, mark throttle at 13 inches of mercury manifold pressure for ground reference.
4. Repeat procedure for rear engine.
NOTE
After flight testing, if required results Were
not obtained, set throttles as marked and readjust micro switches to actuate horn at this
setting. Repeat flight test until desired results are obtained.
10-45. INDUCTION AIR SYSTEM.
10-46. DESCmPTlON. Air to the engine induction
system enters the cylindrical air filter and flows
through the airbox, through the air throttle body, into
the intake manifolds. The complete air induction
system, including the intake manifold. are located on
10-31
the top side of the engine. The airbox contains a
valve, operated by the intake heater control (thru
aircraft serial 337-0755) in the cabin, which permits
air from an exhaust-heated source to be selected.
The valve is spring-loaded and will open if the air
filter should become obstructed. A spring-loaded
nash-back valve, located at the heated air entrance
to the airbox, will close automatically in case of
engine backfire wblle using the heated air source.
Beginning with aircraft serials 337-0756 and F33700001, the manual intake heater control is replaced
by an alternate air valve which opens by engine suction in the event the air filter should become obstructed. This permits the engine to draw heated, unfiltered air from within the engine compartment. The alternate air valve should be checked periodically for
freeness of operation and complete closing. The induction filters should be cleaned, inspected and replaced as ouUined in Section 2.
10-47. REMOVAL AND INSTALLATION.
a. Remove safety wire and loosen wing nut at outer
end of filter. Unhook the air filter hook and remove
filter from airbox assembly.
b. (THRU AIRCRAFT SERIAL 337-0755.) Disconnect intake heater control from air valve lever,
loosen clamp on control and pull control free of airbox.
c. Disconnect flexible duct at airbox.
d. Disconnect lever return spring.
e. Remove the four bolts and nuts securing airbox
to air throttle body and remove airbox. Lay parts of
gear warning system to one side.
f. Reverse the preceding steps for reinstallation.
Rig intake heater control in accordance with paragraph 10-41. Check rigging of gear warning system
and rig, if necessary, in accordance with paragraph
10-44. Do not overtighten wing nut on air filter hook
and resafety hook.
NOTE
The air throttle body is a part of the fuel-air
control unit, which is included in the fuel injection system discussed later.
g. Removal of various intake manifold sections is
accomplished by loosening hose clamps, sliding hoses
back and removing nuts attaching those segments
which are secured to engine cylinders. Disconnect
any lines or hoses interfering with removal. Reverse
this procedure to install the intake manifold.
10-32
10-48. CLEANING INDUCTION AIR FILTER. Refer
to Section 2.
10-49. FUEL INJECTION SYSTEM. (Refer to figure
10-7. )
10-50. DESCRIPTION. The fuel injection system is
a low-pressure system of injecting metered fuel into
the intake valve ports in the cylinders. It is a multinozzle, continuous-now system which controls fuel
flow to match engine airflow. Any change in throttle
position, engine speed or a combination of both,
causes changes in fuel flow in the correct relation to
engine airflow. A manual miXture control and a fuelflow indicator are provided for leaning at any combination of altitude and power setting. The four major
components of the system are: the fuel injection
pump, fuel-air control unit, fuel manifold (distributor)
valve and the fuel discharge nozzles. Since the intake
manifolds are installed on the top side of the cylinders,
drain lines are installed in the bottom side of the intake ports to drain fuel which may have accumulated
in the intake port during engine shut-down.
•
NOTE
Throughout the aircraft fuel system, from
the tanks to the engine-driven fuel pump,
use Never-Seez RAS-4 (Snap-On Tools
Corporation, Kenosha, Wisconsin) or MILT-5544 (Thread Compound, Antiseize,
Graphite-Petrolatum) or equivalent, as a
thread lubricant or to seal a leaking connection. Apply sparingly to male fittings only,
omitting the first two threads on the fitting.
Always be sure that a compound, the residue
from a previously used compound or any other
foreign material cannot enter the system.
Throughout the fuel injection system, from
the engine-driven fuel pump through the discharge nozzles, use only a fuel soluble lubricant, such as engine lubricating oil, on the
fitting threads. Do not use any other form of
thread compound on the injection system
fittings.
•
IWARNING'
Residual fuel draining from lines and hoses
constitutes a fire hazard. Use caution to
prevent accumulation of fuel when lines or
hoses are disconnected throughout the fuel
injection system.
•
•
VAPOR EJECTOR JET
RELIEF VALVE
METERING DISC
DIAPHRAGM
MIXTURE
CONTROL
BYPASS VALVE
o ORIFICE
DIAPHRAGM
IDLE CUT-OFF
CHECK VALVE
FUEL-Am
CONTROL
UNIT
TO
-a~~~~~~~
~FUELFLOW---~~~~~~~~
INDICATOR
•
MANIFOLD
VALVE
A
o
SOME ENGINE-DRIVEN FUEL PUMPS
ARE EQUIPPED WITH AN ADJUSTABLE
ORIFICE INSTEAD OF THE FIXED
ORIFICE SHOWN.
~.
LEGEND:
h:::1
INLET PRESSURE
Im:w::]
PUMP PRESSURE
/
INJECTION
MIXTURE
OUTLET
i?13 FUEL METERED BY MIXTURE CONTROL
•
Detail
A
~ FUEL METERED BY THROTTLE CONTROL
Figure 10-7. Fuel Injection Schematic
10-33
10-51. TROUBLE SHOOTING.
TROUBLE
NO FUEL DELIVERED
TO ENGINE.
HIGH FUEL PRESSURE.
ENGINE RUNS ROUGH
AT IDLE.
10-34
PROBABLE CAUSE
REMEDY
Fuel tanks empty.
Check visually. Service with
desired quantity of fuel.
Defective aircraft fuel system.
Refer to Section 11.
Vaporized fuel. (Mose likely
to occur in hot weather with a
hot engine.)
Refer to paragraph 10-103.
Fuel pump not permitting fuel
from electric pump to bypass.
Check fuel-now through pump.
Replace engine-driven fuel
pump if defective.
Defective fuel control unit.
Check fuel flow through unit.
Replace fuel-air control unit
if necessary.
Defective fuel manifold valve,
or clogged screen inside valve.
Check fuel flow through valve.
Remove and clean in accordance with paragraphs 10- 58
and 10-59. Replace if defective.
Clogged fuel injection lines or
discharge nozzles.
Check fuel flow through lines and
nozzles. Clean and replace if
defective.
Restricted discharge nozzles.
Clean or replace plugged nozzle
or nozzles.
Restriction in vapor vent return
line or check valve.
Clean vapor return line. Clean
or replace check valve.
Improper idle mixture adjustment.
Refer to paragraph 10-55.
Restriction in aircraft fuel
system.
Refer to Section 11.
Low unmetered fuel pressure.
Refer to paragraph 10-68.
High unmetered fuel pressure.
Refer to paragraph 10-68.
Worn throtUe plate shaft or
shaft 0- rings.
Replace shaft and/or O-rings.
Intake manifold leaks.
Repair leaks or replace
defective parts.
Leaking intake valves.
Engine repair required.
Discharge nozzle air vent
manifolding restricted or
defective.
Check for bent or loose
connections, restrictions
or defective components.
Tighten loose connections;
replace defective components.
•
•
•
•
10-51. TROUBLE SHOOTING (Cont) •
TROUBLE
FLUCTUATING FUEL
PRESSURE OF FUEL
FLOW.
LOW METERED FUEL
PRESSURE.
•
FUEL DRAINING
FROM MANIFOLD
VALVE VENT.
PROBABLE CAUSE
Defective manifold valve.
Replace manifold valve.
Restriction in engine-driven
fuel pump vapor ejector.
Clean vapor ejector on fuel
pump. Do not use wires to
clean jet.
Defective check valve in
vapor vent return line.
Clean vapor return vent
line and repair or replace
check valve.
Air in line from manifold
valve to gage.
Bleed air from line.
Malfunctioning relief valve
in engine-driven fuel pump.
Clean or replace relief
valve if defective.
Defective gage or restricted
gage line.
Replace gage. Clean
restriction from line.
Plugged main fuel strainer.
Clean strainer.
Air leak on suction side
of engine-driven fuel pump •
Repair leak. Replace
defective parts.
Ruptured diaphragm.
Replace diaphragm or
manifold valve.
..
.,
POOR IDLE CUT-OFF.
Dirt in fuel pump or
defective pump.
Remove pump and flush
out thoroughly. Check
that mixture arm contacts
cut-off stop.
Dirty or defective fuel
manifold valve.
Remove and clean in
accordance with paragraphs
10-58 and 10-59. Replace
if defective.
10-52. FUEL-AIR CONTROL UNIT.
10-53. DESCRIPTION. The fuel-air control unit,
located at the inlet to the intake manifold, contains
the air throttle and fuel metering unit. The function
of the fuel-air control unit is to meter fuel and air
in the proper ratio for engine operation. The throttle
shaft extends into the fuel metering unit, where it
also operates the fuel metering valve. Idle speed
and idle mixture adjustments are provided in the fuelair control unit. The main mixture control unit is
incorporated in the engine-driven fuel pump.
•
REMEDY
10-54. REMOVAL AND INSTALLATION.
a. Remove cowling as required to gain access.
b. Turn fuel selector valves to OFF position.
c. Tag and disconnect hoses at fuel metering unit.
Cap or plug disconnected hoses and fittings.
d. Disconnect manifold pressure line at fuel-air
control unit.
e. Disconnect throttle control at air throttle arm.
Note position of washers.
f. Disconnect induction airbox spring from tab on
mounting bolt.
g. Remove four bolts, washers and nuts attaChing
air inlet duct to throttle body. Lay parts of landing
gear warning switch to one side. Note any other parts
attached by these bolts.
h. Loosen clamps securing throttle body to intake
manifold and slide hoses away from throttle body.
1. Remove bolts, washers and nuts attaching fuelair control unit to bracket onengi.ne and remove unit .
Cover open ends of manifold and air inlet duct.
j. Reverse the preceding steps for reinstallation.
Rig throttle and throttle-operated landing gear warning
switch.
10-35
10-55. ADJUSTMENTS. (Refer to figure 10-8.) The
idle speed adjustment is a conventional spring-loaded
screw located in the air throttle lever. The idle mixture adjustment is the screw at the metering valve end
of the linkage. Turning the screw clockwise (CW)
leans the mixture and counterclockwise (CCW) richens
the mixture. Adjust mixture control to obtain a slight
and momentary gain of 25 rpm maximum at 1000 rpm
engine speed as Dnxture control is moved slowly from
full RICH toward idle cut-off. H mixture is set too
LEAN, engine speed will drop immediately, thus requiring enrichment. If mixture is set too RICH, engine speed will Plcrease above 25 rpm, thus requiring
leaning. Idle speed is 600 ± 25 rpm on the front engine and 650 ± 25 rpm on the rear engine. Return
mixture control to full RICH position as soon as leaning effect is observed, to keep engine running.
NOTE
Engine idle speed may vary among different
engines. An engine should idle smoothly,
without excessive vibration and the idle speed
should be high enough to maintain idling oil
pressure and to preclude any poSsibility of
engine stoppage in flight when the throttle is
closed. When checking or setting idle speed
or idle mixture, "clear" the engine between
adjustments to prevent false indications.
10-56. FUEL MANIFOLD VALVE (FUEL DISTRIBUTOR).
10-57. DESCRIPTION. Metered fuel flows to the
fuel manifold valve, which provides a central point
for distributing fuel to the individual cylinders. An
internal diaphragm, operated by fuel pressure,
raises or lowers a plunger to open and close the individual cylinder supply ports simultaneously. A
needle valve in the plunger ensures that the plunger
fully opens the outlet ports before fuel flow starts
and closes the ports simultaneously for positive
engine sbut-down. A fine-mesh screen is included
in the fuel manifold valve.
NOTE
The fuel manifold valves are supplied in two
flow ranges. When replacing a valve assembly, be sure the replacement valve has the
same suffix letter as the one stamped on the
cover of the valve removed.
10-58. REMOVAL AND INSTALLATION.
NOTE
Cap all disconnected lines, hoses and fittings.
a. Disconnect the fuel lines and the six fuel injection lines at the fuel manifold valve.
b. Remove the two crankcase bolts which secure
the fuel manifold and mounting bracket. After removal, the bracket may be removed from valve, if
desired.
c. Reverse the preceding steps for reinstallation.
10-36
10-59. CLEANING.
a. Remove fuel manifold valve from engine and remove safety wire from cover attaching screws.
b. Hold the top cover down against internal spring
until all four cover attaching screws have been removed, then gently lift off the cover. Use care not
to damage the spring-loaded diaphragm below cover.
c. Remove the upper spring and lift the diaphragm
assembly straight up.
•
NOTE
H the valve attached to the diaphragm is
stuck in the bore of the body, grasp the
center nut, rotate and lift at the same
time to work gently out of the body.
ICAUTION\
Do not attempt to remove needle or spring
from inside plunger valve. Removal of
these items from the valve will disturb the
calibration of the valve.
d. Using clean gasoline, flush out the chamber below the screen.
e. Flush above the screen and inside the center
bore making sure that outlet passages are open.
Use only a gentle stream of compressed air to
remove dust and dirt and to dry.
ICAUTION]
The filter screen is a tight fit in the body and
may be damaged if removal is attempted. It
should be removed only if a new screen is to
be installed.
•
f. Clean diaphragm, valve and top cover in the
same manner. Be sure the vent hole in the top
cover is open and clean.
g. Carefully replace diaphragm and valve. Check
that valve works freely in body bore.
h. Position diaphragm so that horizontal hole in
plunger valve is 90 degrees from the fuel inlet port
in the valve body.
i. Place upper spring in position on diaphragm.
j. Place cover in position so that vent hole in
cover is 90 degrees from inlet port in valve body.
Install cover attaching screws and tighten to 20±1
lb-in. Install safety wire on cover screws.
k. Install fuel manifold valve assembly on engine
and reconnect all lines and hoses to valve.
1. Inspect installation and install cowling.
10-60. FUEL DISCHARGE NOZZLES.
10- 61. DESCRIPTION. Fromtbe fuel manifold valve,
individual, identical size and length fuel lines carry
metered fuel to the fuel discharge nozzles located in
the cylinder heads. The outlet of each nozzle is
directed into the intake port of each cylinder. The
nozzle body contains a drilled central passage with a
counterbore at each end. The lower end is used as a
chamber for fuel-air mixture before the spray leaves
the nozzle. The upper bore contains an orifice for
•
•
FUEL-AIR CONTROL UNIT - - - - - - - - ,......
IDLE MIXTURE ADJUSTMENT
LEFT SIDE
RIGHT SIDE
IDLE SPEED ADJUSTMENT
Refer to ConUnental Motor. Sentce BWleUn
aad
all renlllau tbereto U clrUUIII or flucbaUaa of fuel fiow.,
fuel prea.area or idle mtzturea 0CC1Il'II.
Figure 10-8. Idle Speed and Idle Mixture Adjustments
•
calibrating the nozzles. Near the top, radial holes
connect the upper counter bore with the outside of the
nozzle body for air adm1ssion. These radial holes
enter the counterbore above the orifice and draw outside air through a cylindrical screen fitted over the
nozzle body. This screen prevents dirt and foreign
material from entering the nozzle. A press-fit shield
is mounted on the nozzle body and extends over the
greater part of the Wter screen, leaving a small
opening at the bottom of the shield. This provides an
air bleed into the nozzle which aids in vaporizing the
fuel by breaking the high vacuum in the intake manifold at idle rpm and keeps the fuel lines filled. The
nozzles are calibrated in several ranges. All nozzles furnished for one engine are the same range and
are identified by a number and a suffix letter stamped
on the flat portion of the nozzle body. When replacing
a fuel discharge nozzle be sure it is of the same calibrated range as the rest of the nozzles in the engine.
When a complete set of noZzles is being installed,
the number must be the same as the one removed,
but the suffix letters may be different, as long as
they are the same for all nozzles being installed on a
parUcular engine.
10-62. REMOVAL.
a. Remove engine cowling as required for access.
NOTE
Plug or cap all disconnected lines and fittings.
Use care to prevent damage to fuel injection
lines.
•
b. Disconnect fuel injection line at each discharge
nozzle •
c. Using a standard l/2-inch deep socket, remove
fuel discharge nozzle from cylinder.
10-63. CLEANING AND INSPECTION. To clean
nozzles, immerse in clean solvent and use compressed air to dry them. When cleaning, direct air through
the nozzle in the direction opposite of normal fuel flow.
Do not remove the nozzle shield or distort it in any
way. Do not use a wire or other metal object to clean
the orifice or metering jet. After cleaning, check the
shield height from the hex portion of the nozzle~ The
bottom of the shield should be approximately 1/16
inch above the hex portion of the nozzle.
10-64. INSTALLATION.
a. Using a standard l/2-inch deep socket, install
nozzle body in cylinder and tighten to a torque value
of 60-80 lb-in.
b. Connect the fuell1nes at discharge nozzles.
c. Check installation for crimped lines, loose fittings, etc.
d. Install cowling.
10-65. FUEL INJECTION PUMP.
10-66. DESCRIPTION. The fuel pump is a positivedisplacement, rotating vane type, located opposite
the propeller governor at the propeller end of the engine. Fuel enters the pump at the swirl well of the
pump vapor separator. Here, vapor is separated by
a swirling motion so that only liquid fuel is fed to the
pump. The vapor is drawn from the top center of
the swirl well by a small pressure jet of fuel and is
fed into the vapor return line, where it is returned
to the fuel line manifold. Since the pump is enginedriven, changes in engine speed effect total pump
flow proportionally. A check valve allows the auxiliary fuel pump pressure to bypass the engine-driven
fuel pump for starting, or in the event of enginedriven fuel pump failure in flight. The pump supplies
more fuel than is required by the engine; therefore,
Change 1
10-37
a spring-loaded, diaphragm type relief valve is provided to maintain a constant fuel pump pressure. Refer to paragraph 10-68 for pressure adjustments.
The fuel pump is equipped with a manual mixture control to limit the fuel pump output from full rich to
idle cut-off. Non adjustable mechanical stops are
located at these positions. DUring the 1967 model year
and for :tll service parts, the fixed orifice is replaced
with an adjustable orifice to allow the exact desired
pressure setting at the full-throttle position. Fuel
pumps With the adjustable orifice feature are identified by the presence of a brass plug with a stainless
steel adjusting needle having a screwdriver slot located below the fuel inlet fitUng.
10-67. REMOVAL AND INSTALLATION.
a.- Turn fuel selector valves to the OFF pOsition.
b. Remove cowling, baffles and covers as necessary to gain access.
c. Disconnect miXture control from lever on pump.
Note position of washers.
d. Tag and disconnect fuel hoses and vent line attached to pump. Plug or cap all disconnected hoses
and fittings.
e. Remove mounting nuts and bolts and pull pump
and gasket from engine pad.
IWARNING,
Residual fuel draining from lines and hoses
constitutes a fire hazard. Use caution to
prevent accumulation of fuel wheo lines or
hoses are disconnected.
f. The drive shaft coupling may come off with the
fuel pump, or it may remain in the engine. If it
comes off with the pump, reinstall it io the engine
to prevent dropping or losing it.
g. If a replacement pump is not being installed immediately, a temporary cover should be installed on
the fuel pump mount pad.
h. Reverse the preceding steps for reinstallation.
Using a new gasket, do not force engagement of the
pump drive. Rotate engine crankshaft and pump
drive will engage smoothly when aligned properly.
Check mixture control rigging.
i. Start engine and perform an operational check,
adjust fuel pressure as required in accordance with
paragraph 10-68.
10-68. ADJUSTMENTS. (Refer to figure 10-9.) The
full rich performance of the fuel injection system is
controlled by manual adjustment of the air throttle,
fuel mixture and pump pressure at idle and only by
pump pressure at full throttle. To make full rich
adjustments, proceed as follows:
a. Remove engine cowling in accordance with paragraph 10-3.
NOTE
Inspect the slot- headed adjustable orifice
needle valve (located just below the fuel
pump inlet fitting) to see if it is epoxy
sealed or safety wired to the brass nut.
10-38
Change 1
If the needle valve is epoxy sealed, Con-
tinental Aircraft Engine Service Bulletin
No. 70-10 must be complied with before
calibration of the unit can be performed.
b. Disconnect the engine-driven fuel pump outlet
fitting or the fuel metering unit inlet fitting and "tee"
the test gage into the fuel injection system as illustrated in figure 10-9.
•
NOTE
Cessna Service Kit No. SK320-2J proVldes
a test gage, line and. fittings for connecting
the test gage into the system to perform
accurate calibratioD of the engine-driven
fuel pump.
c. The test gage MUST be vented to atmosphere and
MUST be held as near to the level of the engine-driven
fuel pump as possible.
NOTE
The test gage should be checked for accuracy
at least every 90 days or anytime an error is
suspected. The tachometer accuracy should
also be determined prior to making any adjustments to the pump.
d. Start engine and warm-up thoroughly. Set mixture control to full rich positlon and propeller control full forward (low pitch, high rpm).
e. Adjust engine idle speed to 600 ± 25 rpm (front
engl.ne) or 650 ± 25 rpm (rear engine). Refer to
figure 10-8 for idle speed adjustment. Check fuel
pressure on indicator for 6 to 8 PSI.
•
NOTE
Do not adjust idle mixture untll idle pump
pressure is obtained.
IWARNINGt
DO NOT make fuel pump pressure adjust-
ments while engine is operating.
f. If the pump pressure is DOt 6 to 8 PSI, stop eng1De and turn the fuel pump relief valve adjustment,
on the ceDterline of the fuel pump clockwise (CW) to
increase pressure and countercloc:kwi!le (CCW) to
decrease pressure.
g. Maintaining idle pump pressure and idle RPM,
obtain correct idle miXtUre in accordance with paragraph 10-55.
h. Completion of the preceding steps have provided:
1. Correct idle pump pressure.
2. Correct fuel flow.
3. Correct fuel metering cam to throttle plate
orientation.
1. Advance to full throttle and maximum rated engiDe speed With the mixture control in full rich position and propeller control in full forward (low pitch,
high rpm).
•
•
ENGINE -DRIVEN
FUEL PUMP
FUEL METERING
UNIT
EXISTING FUEL PUMP
OUTLET HOSE
.
• _----1_
I
PRESSURE
INDICATOR
TEE
TEST HOSE
TEST HOSE
•
NOTE
WHEN ADJUSTING UNMETERED FUEL PRESSURE, TEST EQUIPMENT MAY
BE "TEED" INTO THE ENGINE-DRIVEN FUEL PUMP OUTLET HOSE AT THE
FUEL PUMP OR AT THE FUEL METERING UNIT
Figure 10-9. Fuel Injection Pump Adjustment Test Harness
NOTE
Fuel injection pumps with a fixed orifice
cannot be adjusted for full throttle pressure.
IWARNING,
DO NOT make fuel pump pressure adjustments while the engine is operating.
j. Check test gage for pressures specified in paragraph 10-8. If pressure is incorrect, stop engine
and adjust pressure by loosening locknut and turning
the slotheaded needle valve located just below the fuel
pump inlet fitting counterclockwise (CCW) to increase
pressure and clockwise (CW) to decrease pressure .
•
NOTE
U at static run-up, rated RPM cannot be
achieved at full throttle, adjust pump
pressure slightly below limits making
certain the correct pressures are obtained when rated RPM is achieved during take-off roll.
k. After correct pressures are obtained, safety
adjustable orifice and Orifice locknut.
1. Remove test equipment, run engine to check for
leaks and install cowling.
m. Repeat the preceding steps for other engine if adjustment is required.
10-39
10-69. EXHAUST SYSTEMS. (Refer to figure 10-10.)
10-70. DESCRIPTION.
a. FRONT. The front exhaust system incorporates
an individual exhaust stack and cylindrical muffler on
each side of the engine. A shroud surrounds each
muffler and heats air that is routed through a mixing
chamber into the cabin. The exhaust stacks are made
in sections that are clamped together, the tailpipe
routing the exhaust gases overboard. The muffler
and tailpipe is supported by a shock-mounted circular
disc with braces attached to the firewall. Beginning
with aircraft serials 33701317 and F33700025, the
shock-mounts are replaced with a cord reinforced
neoprene rubber strap. To be compatible with the
new shock-mounts, the supporting brackets were also
redesigned.
b. REAR. The rear exhaust system routes the exhaust gases through a flattened tank-type muffler,
then overboard through twin tailpipes. The exhaust
stacks are made in sections that are clamped together.
The muffler contains a vertical passage through which
the rear engine oil may be drained: The muffler is
supported by rigid braces which attach to the engine
mount.
10-71. REMOVAL.
a. Remove engine colwing as necessary for access.
If the rear engine exhaust muffler is to be removed,
it will be necessary to remove the lower tail cap.
b. If installed, remove exhaust gas temperature
probes or disconnect leads.
c. Remove nuts securing stacks to cylinders.
d. On the front engine, disconnect heater hoses
from mufflers, loosen tailpipe clamp, remove bolts
and springs securing mufflers to stacks and slide
mufflers down until stacks and mufflers can be removed.
e. Remove shrouds from mufflers if desired.
f. On the rear engine, remove clamps securing
muffler to stacks, disconnect braces and remove
stacks and muffler.
10-72. INSPECTION. The exhaust systems must be
thoroughly inspected, especially the heat exchanger
section of the mufflers on the front engine. A leak in
the muffler would allow poisonuous gases to enter the
cabin heating system. An inspection of the exhaust
system must be performed every 100 hours of operating time. Any time exhaust fumes are detected in
the cabin, an immediate inspection must be performed. All components that show cracks and general deterioration must be replaced with new parts.
a. Remove engine cowling as necessary for access.
b. Loosen or remove shrouds so that ALL surfaces
of the exhaust system is visible.
c. Check for holes, cracks and burned spots. Especially check the areas adjacent to welds. Look for
exhaust gas deposits in surrounding areas which indicate an exhaust leak.
d. Where a surface is not accessible for a visual
inspection or for a positive test, proceed as follows:
1. Remove exhaust stacks, mufflers and tailpipes.
2. Remove all shrouds.
3. Seal openings with expansion rubber plugs.
4. Using a manometer or gage, apply approximately 1-1/2 psi (3 inches of mercury) air pressure
While the unit being tested is submerged in water.
Any leaks will appear as bubbles and can be readily
detected.
5. It is recommended that any component found
defective be replaced with new parts before the next
flight.
6. If no defects are found, remove plugs and
dry components with compressed air.
e. Install exhaust system and engine cowling.
10-73. INSTALLATION.
a. When installing exhaust stacks, use new gaskets,
regardless of apparent condition of those removed.
b. To prevent pre-stressing of exhaust stack assemblies, place all sections of the assembly in position and join together, with attaching clamps and
braces not tightened. Tighten nuts securing stacks
to cylinders first, then tighten all clamps joining sections, then position and tighten braces.
c. Torque exhaust stack nuts at cylinders to 2002101h-in.
d. After installation of cowling, check for adequate
clearance where tailpipes emerge. They must not
contact cowling or cowl flaps during flight. The minimum clearance between tailpipe and cowling for the
front engine is .75" and for the rear engine is .62".
10-74. OIL SYSTEM.
•
•
(Refer to figure 10-11.)
10-75. DESCRIPTION. A wet-sump, pressurelubricating oil system is employed in the engine.
Oil under pressure from the oil pump is fed through
drilled crankcase passages which supply oil to the
crankshaft main bearings and camshaft bearings.
Connecting rod bearings are pressure-lubricated
through internal passages in the crankshaft. Valve
mechanisms are lubricated through the hollow pushrods, which are supplied with oil from the crankcase
oil passages. The propeller is supplied oil, boosted
by the governor, through the forward end of the crankshaft. Oil is returned by gravity to the engine oil
sump. Cylinder walls and piston pins are spraylubricated by oil escaping from connecting rod bearings. The engine is equipped with an oil cooler and
a thermostat valve to regulate engine oil temperature. A pressure relief valve is installed to maintain
proper oil pressure at higher engine speeds. An external, replaceable element oil filter may be installed.
•
10-40
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
FRONT ENGINE EXHAUST SYSTEM
TYPICAL BOTH SIDES OF ENGINE
Riser
Clamp
Collector
Muffler
Shroud
Spring
Shock-Mount
Bracket
Clamp
TaU Pipe
Exhaust Brace
Support
NOTES
Torque exhaust clamp nuts
to 25 - 30 Ib-in.
Install exhaust flange gaskets
with raised bead toward ex-
haust port on engine .
•
•
9
REAR ENGINE EXHAUST SYSTEM
Figure 10-10. Exhaust Systems
10-41
10-76. TROUBLE SHOOTING.
TROUBLE
NO On.. PRESSURE.
PROBABLE CAUSE
HIGH OIL PRESSURE.
10-42
REMEDY
No oil in sump.
Check oil with dipstick. Fill sump
with proper grade and quantity of
oil. Refer to Section 2.
Oil pressure line broken,
disconnected or pinched.
Check Visually. Replace or
connect.
Oil pump defective.
Remove and inspect. Examine
engine. Metal particles from
damaged pump may have entered
engine oil passages.
Defective oil pressure gage.
Check with a known good gage.
Replace gage if defective.
Oil congealed in gage line.
Disconnect line at engine and gage;
flush with kerosene. Pre-fill with
kerosene and install.
Relief valve defective.
Remove and check for dirty or defective parts. Clean and install;
replace defective parts.
Low oil supply.
Check with dipstick. Replenish
with proper grade and quantity.
Low viscosity oil.
Check visually. Drain sump and
refill with proper grade and
quantity of oil.
Oil pressure relief valve spring
weak or broken.
Remove and inspect. Replace
weak or broken spring.
Defective oil pump.
Remove and inspect. Examine
engine. Metal particles from
damaged pump may have entered
engine oil passages.
Defective oil pressure gage.
Check with a known good gage.
Replace gage if defective.
Secondary result of high oil
temperature.
Observe oil temperature gage for
high indication. Determine and
correct reason for high oil temperature.
High viscosity oil.
Check visually. Drain sump and
refill with proper grade and
quantity of oil.
Relief valve defective.
Remove and check for dirty or
defective parts. Clean ani
install; replace defective parts .
Defective oil pressure gage.
Check with a known good gage.
Replace oil pressure gage.
-.
LOW On.. PRESSURE.
.
•
•
•
•
10-76. TROUBLE SHOOTING (Cont) .
TROUBLE
LOW OIL TEMPERATURE.
HIGH OIL TEMPERATURE.
•
PROBABLE CAUSE
REMEDY
Defective oil temperature gage
or temperature bulb.
Check with another gage. If
reading is normal, aircraft
gage is defective. If reading
is similar temperature bulb
is defective. Replace defective part/or parts.
Oil cooler thermo-bypass
valve defective or stuck
closed.
Remove valve and check for
proper operation. Replace
valve if defective.
Defective Wiring.
Check continuity. Repair Wiring.
Oil cooler air passages clogged.
Check visually. Clean air passage's
Oil cooler oil passages clogged.
Attempt to drain cooler. Inspect
for sludge. Remove cooler and
flush thoroughly.
Low oil supply.
Replenish.
011 viscosity too high.
Drain and fill sump with proper
grade and quantity.
Prolonged high speed operation
on ground .
Hold ground rwming above 1500,
rpm to a minimum.
Defective oil temperature gage.
Check With another gage. If
second reading is normal, aircraft gage is defective. Replace gage.
Defective oil temperature bulb.
Check for correct oil pressure,
oil level and cylinder head temperature. If they are correct,
check oil temperature gage for
being defective; if similar reading is observed, bulb is defective.
Replace bulb.
Oil congealed in cooler.
If congealing is suspected, use
external heater or a heated
hangar to thaw the congealed
oil.
•
Secondary result of low
oil pressure.
Check for low oil pressure
reading. Determine cause
and correct.
Secondary result of high
cylinder head temperature.
Check for high cylinder head
temperature. Determine
cause and correct •
10-43
4
1&
CODE:
_
RETURNAND
SUCTION OIL
IHI
PRESSURE OIL
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
Pressure Gage
Propeller Governor
Sump Drain Plug
Filler Cap
Dipstick
Temperature Transmitter
Oil Cooler
Temperature Gage
Thermostat
Pressure Relief Valve
Oil Dilution Solenoid Fuel Line
Filter Screen (Suction)
Engine Oil Pump
Filter Screen (Pressure)
Filter Bypass Valve
External Filter (Optional)
Figure 10-11. Oil System Schematic
10-44
Detail
A
•
•
10-77. FULL-FLOW On.. Fn..TER. (Refer to figure 10-12.)
10-78. DESCRIPTION. An external full-flow oil
filter may be installed on each engine. The filter and
filter adapter replace the internal 011 pressure
screen. The adapter incorporates a bypass valve
which will open in the event the filter element becomes clogged, allOWing the engine 011 to flow directly to the engine oil passages.
• Before assembly, place a straightedge
across the bottom of filter can (12).
Check for distortion or out-of-flat condition greater than 0.010 inch. Install
a neW can if either of these conditions
exists.
• After installing a new upper gasket (8) on
the lid (9), turn lid over. H gasket falls,
try a different gasket and repeat test. H
this gasket falls off, install a new lid.
10-79. ELEMENT REMOVAL AND INSTALLATION.
a. Remove engine coWling as necessary.
b. Remove both safety wires from filter can and
unscrew stud (16) to detach filter assembly (5) from
adapter (7). Remove assembly and discard gasket
(8). Oil will drain from filter as assembly is removed from adapter.
c. Press downward on stud (16) to remove and discard metal gasket (15).
d. Lift lid (9) from can (12) and discard gasket (10).
e. Pull filter element (11) from can.
NOTE
•
Before discarding the removed filter element,
remove the outer perforated paper cover;
using a sharp knife, cut through the folds of
the element at both ends, close to the metal
caps. Carefully unfold the pleated element
and examine the trapped material for evidence
of internal engine damage such as chips or
particles from bearings. In new or newly
overhauled engines, some small particles or
metallic shavings might be found, these are
generally of no consequence and should not
be confused with particles produced by impacting, abrasion or pressure. Evidence of
internal engine damage found justifies further
examination to determine the cause.
f. Wash lid (9), stud (16) and can in Stoddard solvent or equivalent and dry with compressed air.
NOTE
When installing a new filter element (11),
it is important that all gaskets are clean,
lubricated and positioned properly, and
that the correct amount of torque is applied to the filter stud (16). If the stud is
under-torqued, oil leakage will occur. If
the stud is over-torqued, the filter can (12)
might possibly be deformed, again causing
oil leakage.
• Lubricate the rubber grommets in each end
of the new filter element (11) and gaskets
(8, 10 and 15) with clean engine oil or general purpose grease before installing. Dry
gaskets can cause false torque readings,
again resulting in oil leakage .
•
g. Inspect adapter gasket seat for gouges, deep
scratches, wrench marks and mutilation. H any of
these conditions are found, install a new adapter (7).
h. Place a new filter element (11) in can (12) and
insert stud (16) with a new metal gasket (15) in place,
through the can and filter element.
i. Position a new lower gasket (10) inside flange of
lid (9) and place lid in position on can.
j. Install filter assembly (5) on adapter (7) with a·
new upper gasket (8) in place. While holding can to
prevent turning, tighten stud (16) and torque to 20-25
Ib-ft (240-300 lb-in).
k. Install parts removed for access and service
engine with proper grade and quantity of oil. One
additional quart of oil is required each time filter
element is changed.
1. Start engine and check for proper oil pressure.
Check for leaks after warming up engine.
m. Again check for leakage after engine has been
run at a high power setting (preferably a flight
around the field).
n. Check to make sure that the filter has not been
in contact with adjacent parts due to engine torque.
o. While engine is still warm, recheck torque on
stud (16) then safety stud to tab (14) on can, safety
adapter nut to other tab on filter can.
10-80. ADAPTER REMOVAL.
(Refer to figure
10-12. )
a. Remove filter can as outlined in paragraph
10-79.
b. Remove alternator in accordance with procedures outlined in Section 15.
NOTE
When removing Wter adapter from the
FRONT engine, it is necessary that the
rear of the engine be raised so the adapter
will clear the nose gear tunnel When being
unscrewed. After disconnecting and/or
unclamping items to permit raising the
rear of the engine as required, remove
the rear mount bolts. Attach a suitable
hoist to the hoisting lug and slowly raise
the hoist, watching for any additional items
that may need to be disconnected or unfastened.
10-45
•
NOTE
Do NOT subsitute automotive gaskets
for any gasket used in this assembly.
Use only approved gaskets listed in
Parts Catalog.
A
7
NOTE
•
Use thread lube on threads of nut (3) and
torque nut to 50-60 lb-ft (600-700 lb-in).
Do not allow can (12) to spin when tightening stud (16). Torque stud (16) to 2025 Ib-ft (240-300 lb-in).
1. a-Ring
2. a-Ring
3. Adapter Locknut
4. Bypass Valve
5. Filter Assembly
6. Thread Insert
7. Adapter
8. Upper Gasket
9. Lid
10. Lower Gasket
11. Filter Element
12. Filter Can
13. Upper Safety Wire Tab
14. Lower Safety Wire Tab
15. Metal Gasket
16. Hollow Stud
Detail
Figure 10-12. Full-FlOW Oil Filter
10-46
A
•
•
1-7/8 "R (TYP)
L2.135".-J
MATL : 4130 (Rc.
35-38)
Figure 10-13. Oil Filter Adapter Wrench Fabrication
•
c. Note angular position of the adapter (7), then
remove safety wire and loosen adapter nut (3). Also,
remove screw attaching adapter to bracket.
NOTE
A special wrench adapter (Part No. SE-709)
for the adapter nut is available from the
Cessna. Service Parts Center, or one may
be made as shown in figure 10-13.
d. Unscrew adapter and remove from the engine.
10-81. ADAPTER DISASSEMBLY, INSPECTION
AND REASSEMBLY. (Refer to figure 10-12.) The
relative position of the adapter and associated parts
are shown in figure 10-12 and may be used as a
guide for parts replacement. The bypass valve (4)
is replaced as a unit being staked three places at
installation. Inspect that bypass valve is not held
open by carbon or other foreign material. The helicoil insert (6) in the adapter may be replaced, although special tools are required for installation.
Follow instructions of the tool manufacturer for their
use. Inspect threads on adapter and in engine for
damage. Clean adapter in Stoddard solvent or equivalent and dry with compressed air.
•
10-82. ADAPTER INSTALLATION. (Refer to
figure 10-12.)
a. Assemble nut (3) and new O-rings (1 and 2) on
adapter as illustrated. O-ring (2) must be centered
in groove between threads on adapter. Lubricate
O-rings lightly with clean engine oil.
b. Apply anti-seize compound sparingly to adapter
threads and screw adapter (7) into engine unW O-ring
(2) seats against engine base without turning nut (3).
Rotate adapter to approximate angular pOsition noted
during removal. Do not tighten nut (3) at this time.
c. Carefully lower front engine and install mount
bolts. Connect or fasten all items disconnected or
unfastened when raising engine.
d. Temporarily install filter assembly (5) on adapter and position adapter so adequate clearance with
adjacent parts is attained. Maintaining this position,
tighten nut (3) to 50-60 lb-ft (600-700 lb-in) and safety.
e. Install screw attaching adapter to bracket. Adjust bracket as required.
f. Install alternator in accordance with procedures
outlined in Section 15.
g. Complete filter installation using procedures outlined in paragraph 10-79.
h. Be sure to service the engine oil system, perform the checks and inspections outlined in paragraph
10-79, resafety all items previously safetied and reinstall all items removed for access.
IO-82A.
OIL DILUTION SYSTEM.
IO-82B. DESCRIPTION. (See Figure lO-13A.)
An optional oil dilution system may be installed on
the forward and rear engines. The system consists
of fuel lines from the fuel strainers to solenoid
valves mounted on the firewalls, and fuel lines
from the solenoid valves to the inlet (screen) side of
the respective engines internal mounted oil pump.
Change 1
10-47
The solenoid valves are controlled by a switch
located on the instrument panel, labeled, "OIL
DILUTE". Power to operate the solenoid valves is
supplied from the Rear Engine Gages circuit
breaker through the "OIL DILU' _E" switch to the
solenoid valves. For system operation, refer to the
Pilot's Operating Handbook.
7. Remove bolts (9) and washers (8) and
remove solenoid valve (6) and clamp (5).
10-82D. INSTALLATION OF OIL DILUTION
SYSTEM (See Figure 10-13A.)
a. FORWARD ENGINE.
l.Position solenoid valve (6) on firewall and
install clamp (5) using washers (8) and bolts (8).
Be sure to position solenoid valve so in port will be
connected to the fuel strainer.
2. Remove cap from tee fitting (10) and
connect line (4).
3. Connect line (4) to "IN" port of solenoid
valve (6).
4. Connect hose (3) to "OUT" port of solenoid
valve (6).
5. Remove cap from fitting (2) and connect
hose (3).
6. Connect electrical lead (7) to solenoid val ve
lO-82C. REMOVAL OF OIL DILUTION
SYSTEM. (See Figure 10-13A.)
a. FORWARD ENGINE.
I. Place fuel selectors in t.he "OFF" position.
2. Remove left hand cowling panel.
3. Disconnect line (4) from tee fitting (10) and
cap fitting. If oil dilution system is not to be
reinstalled, replace tee fitting with an elbow
fitting.
4. Disconnect lines (4) and hose (3) from
solenoid valve (6).
(6).
5. Disconnect hose (3) from fitting (2) and cap
7. Install left hand cowling panel.
fitting. If oil dilution system is not to be
b. REAR ENGINE.
reinstalled, replace fitting (2) with plug.
1. Position solenoid val ve (6) on firewall and
6. Disconnect electrical lead (7) from solenoid
install clamp (5) using washers (8) and bolts (9).
valve (6). If oil dilution system is not to be reinstallBe sure to position solenoid valve so "IN" port will
ed, tie up electrical lead (7) and place tape over
"OIL DILUTE" switch and mark inoperative.
be connected to the fuel strainer.
7. Remove bolts (9) and washers (8) and
2. Remove cap from fitting (13) and connect
remove solenoid valve (6) and clamp (5).
line (4) to fuel strainer (12) and solenoid valve (6).
b. REAR ENGINE.
3. Connect. hose (3) to solenoid valve (6).
1. Place fuel selectors in the "OFF" position.
4. Remove cap from fitting (2) and connect
hose (3).
2. Open cowl flaps and disconnect right hand
cowl flap rod.
5. Connect electrical lead (7) to solenoid valve
(6).
3. Remove right hand cowling panel.
4. Disconnect line (4) from fuel strainer (12)
6. Install right hand cowling panel and
and from solenoid valve (6). Cap fitting (13). If oil
connect cowl flap rod.
dilution system is not to be reinstalled, replace
fitting (13) with a plug.
10-83. IGNITION SYSTEM.
5. Disconnect. hose (3) from solenoid valve (6)
10-84. DESCRIPTION. The ignition system for each
and fitting (2). Cap fitting (2). If oil dilution
engine is comprised of two magnetos, two spark plugs
system is not. to be reinstall, replace fitting (2) wit.h
in each cylinder, an ignition wiring harness, an igniplug.
tion SWitch mounted on the instrument panel and re6. Disconnect electrical lead (7) from solenoid quired wiring between the ignition switches and magvalve (6). If oil dilution system is not to be reinstall- netos.
ed, tie up electrical lead (7) and place tape over
"OIL DILUTE" switch and mark inoperative.
•
•
•
10-48
Change 1
•
... ....:...~:-.
......•..•...••.••.••.:.:.:.:.:.:..•...•.•:
':....
:"
. .····iZ· ;.,
.........:::.....
C SEE SHEET 2 OF 4
:~::~...
..:::;::::......
..
'
f-·\·· .
L__ A;;"
..:.Or•• •••••·
7:,1/j
.......
C SEE SHEET 2 OF 4
•
/~..
/£- [
~,'OOOO
o
- 0,.
0 0
--~-----
::::---
0
01
!!!!!HL!!d!!!!!!!!~l!!!!
"
iiilllll,hiliiiiiiiiIJlI,,'
limmmmmmii!!!iii
f'm'llllllllililliili,ii..h
°0
00
0
0
ffIlIl 'fWTmrmmn
ililliWi,iHiim;j:iiJII
!!Hm!!:!lamm~;:;:::l
"'iIJl"","II11IJIIUIII
," .
•
0
0
0
0 '. ~"
""'\
00
::::::::::.....~
.'
1'------'
0
~I
:'-;nmnmlli:]
n:f:llllVlPJll;:n
~~ia]_~=6'~tF~~~:~
0'
-
au========t=.====::=!=1
OIL DILUTION
SWITCH (TYPICAL)
/?- - .. ====----~
:
~lIt-
.!dbl~
-..::-- - _... -
-il~
-----
i
~I
r~J
--
Figure lO-13A. Oil Dilution System Installation. (Sheet 1 of 4)
Change 1
10-48A
•
•
•
OIL PUMP PRESSURE
PORT (REF)
Remove plug and install elbow fitting (2)
then connect hose (3) to this port.
Detail
C
VIEW LOOKING FORWARD ON FORWARD
ENGINE AND AFT ON REAR ENGINE
1. Oil Pump Inlet Port
Figure lO-13A. Oil Dilution System Installation. (Sheet 2 of 4)
10-488
Change 1
•
•
6
5
4
2
TO FUEL PUMP"
•
1.
2.
3.
4.
5.
6.
7.
Oil Pump Inlet Port
Elbow Fitting
Hose - Oil Dilution
Line - Oil Dilution
Clamp
Solenoid Valve
Electrical Lead
8. Washer
9. 80lt
10. Tee Fitting
11. Line
12. Fuel Strainer
~
~
.........
--11
TO ENGINE PRIMERS ""-
~
~.
{f1ffi .
FROM FUEL TANKS / '
Detail
•
A
FORW ARD ENGINE
Figure lO-I3A. Oil Dilution System InstaJJation. (Sheet 3 of 4)
Change 1
1O-4BC
•
4
12
~..,... yFROM FUEL TANKS
Oil Pump Inlet Port
Elbow Fitting
Hose - Oil Dilution
Line - Oil Dilution
Clamp
Solenoid Valve
Electrical Lead
Washer
Bolt
Fuel Strainer
13. Fitting
1.
2.
3.
4.
5.
6.
7.
8.
9.
12.
* SEE DETAIL C SHEET 2 OF 4
•
4
5
B
Detail
REAR ENGINE
Figure 1O-13A. Oil Dilution System Installation. (Sheet 4 of 4)
10-480
Change 1
•
•
10-85. TROUBLE SHOOTING •
TROUBLE
EN(,!NE FAILS TO START.
PROBABLE CAUSE
REMEDY
Defective ignition switch.
Check switch continuity.
if defective.
Spark plugs defective, improperly
gapped or fouled by moisture or
deposits.
Clean. regap and test plugs.
Replace if defective.
Defective ignition harness.
Replace
If no defects are found by a
Visual inspection, check
witb a harness tester. Replace defective parts.
•
ENGINE WILL NOT IDLE
OR RUN PROPERLY.
Magneto lip" lead grounded.
Check continuity. "P" lead
should not be grounded in the
ON position. but should be
grounded in OFF pOSition.
Repair or replace "P" lead.
Failure of impulse couplings.
Impulse coupling pawls should
engage at cranking speeds.
Listen for loud clicks as impulse couplings operate. Remove magnetos and determine
cause. Replace defective parts.
Defective magneto .
Refer to paragraph 10-92.
Broken drive gear.
Remove magneto and check magneto and engine gears. Replace
defective parts. Make sure no
pieces of damaged parts remain
in engine or engine disassembly
will be required.
Spark plugs defective, improperly gapped or fouled
by moisture or deposits.
Clean. regap and test plugs.
Replace if defective.
Defective ignition harness.
If no defects are found by a
visual inspection, check with
a harness tester. Replace
defective parts.
Defective magneto.
Refer to paragraph 10-92.
Impulse coupling pawls
remain engaged.
Pawls should never engage
above 450 rpm. Listen for
loud clicks as 1mpulse
coupling operates. Remove
magneto and determine cause.
Replace defective parts.
Spark plugs 10088.
Check and install properly .
•
Change 1
•
•
Figure 10-14. Magneto Internal Timiru; Template Cut-Outs
Change 1
•
•
10-86. MAGNETOS.
10-87. DESCRIPTION. Bendix-Scintilla S6LN-25
magnetos, equipped with impulse couplings, are used
on both engines. Each magneto fires at 20 before
top center. The right magnetos fire the upper right
and lower left spark plugs and the left magnetos fire
the upper left and lower right spark plugs. Always
use a timing light for accuracy when checking or
setting magneto timing.
0
10-88. REMOVAL AND INSTALLATION. Access to
the breaker compartment is gained by removing the
breaker compartment cover at the back end of the
magneto. To remove the magneto from the engine,
proceed as follows:
a. Remove cowling as necessary for access.
b. Remove high-tension outlet plate and disconnect
magneto .. P" lead.
c. Disconnect any noise filters used with radio installations.
d. If the right magneto is being removed, disconnect the tachometer pick-up coil installed in the
bottom of the magneto.
e. Note the apprOximate angular position at which
the magneto is installed, then remove magneto mounting clamps and remove magneto from engine.
NOTE
•
Never remove the screws fastening the two
halves of the magneto together. Separating
the halves would disengage distributor gears,
causing loss of internal timing and necessitating complete removal and internal retiming.
f. Reverse the preceding steps for reinstallation.
Time magnetos-to-engine in accordance with paragraph 10-90.
NOTE
The No. 1 magneto outlet is identified with
the number "1." The magneto fires at
each successive outlet in direction of rotation. No. 1 magneto outlet routes to No. 1
cylinder, No. 2 magneto outlet to the next
cylinder to fire, etc. Cylinder firing order
is 1-6-3-2-5-4.
•
10-89. INTERNAL TIMING. The following information gives instructions for adjusting breaker contacts
to open at the proper poSition. It is assumed that the
magneto has not been disassembled and that the distributor gear, rotor gear and cam have been assembled for correct meshing of gears and direction of
rotation. Magneto overhaul, including separating the
tHO major sections of the magneto, is not covered in
this manual. Refer to applicable Bendix publications
for disassembly and overhaul.
a. Fabricate a timing template as follows:
1. Cut a paper template from figure 10-14.
2. Cement paper template to a thin piece of
metal for use as a support plate, then trim the plate
to the shape of the paper template.
3. Drill the two mounting holes with a No. 18
drill.
b. Fabricate a timing pointer as shown in figure
10-15 .
c. Remove magneto from engine per paragraph
10-88, remove breaker compartment cover and remove timing inspection plug from top of magneto.
d. Attach timing template ~! breaker compartment
as shown in figure 10-15, using 8-32 screWs 1/4 inch
long.
e. Turn rotating magnet in its direction of rotation
until the painted chamfered tooth on distributor gear
is approximately in center of inspection window, then
turn rotating magnet back until it locates in its magnetic neutral position.
I
NOTE
Impulse coupling pawls must be depressed to
turn rotating magnet in its normal direction
of rotation.
f. Remove cam screw, lockwasher and washer.
Use cam screw to install timing pointer so it indexes
with 0 mark on template, While rotating magnet is
still in its magnetic neutral poSition.
g. Turn rotating magnet in proper direction of rotation until pOinter indexes with 10° mark ("E" gap).
Using 11-9110 timing light or equivalent, adjust
breaker contacts to open at this point.
h. Turn rotating magnet until cam follower is on
high part of cam lobe and measure clearance between breaker contacts. Clearance must be .018 ±
.006 inch. If clearance is not within these limits,
readjust breaker contacts until they are within tolerance, then recheck the 10° ("E" gap) poSition. Tolerance on the "E" gap position is ±4 0 . Replace
breaker assembly if "E" gap and contact clearance
will not both fall within the specified tolerances.
i. Remove timing pointer and timing template and
install cam screw, lockwasher and washer.
j. Install magneto and time to engine in accordance
with paragraph 10-90.
0
10-90. MAGNETO-TO-ENGINE TIMING.
NOTE
In conducting magneto timing checks, use of
a positive piston top dead center (T.D.C.)
locator is of utmost importance. The Universal Engine Timing Indicator, available
from Hanger Service Co., Muskegon County
Airport, Muskegon, Michigan or its equivalent is recommended.
a. Remove all top spark plugs.
b. Install top dead center locator into number 1
cylinder top spark plug hole.
c. Install the timing disc of indicator on the propeller hub.
d. Turn propeller slowly in direction of rotation
until piston lightly touches the T. D. C. locator.
e. Rotate the timing disc on propeller hub until top
center mark is under the pointer .
f. Turn propeller slowly in opposite direction of
rotation until piston lightly touches the T.D.C.
locator.
10-51
•
,
3/4 "
SOLDER
CAM WASHER
TJMlNG POINTER FABRICATION
TEMPLATE AND POINTER ATTACHED
TO BREAKER COMPARTMENT
Figure 10-15. Magneto Internal Timing Pointer
g. Observe the reading on timing disc under pointer
and move the timing disc EXACTLY one-half of the
number of degrees observed toward the top center
mark.
h. Remove the T.D.C. locator from number 1 cylinder and locate the compression stroke. Place
thumb over the spark plug hole and turn propeller
until a positive pressure is felt, continue to turn propeller untlltiming disc is at the T.D.C. position.
Top dead center on the compression stroke has now
been located.
i. To check the magneto-to-engine timing or to
time the magnetos to the engine, move the propeller
in the OPPOsite direction of rotation past 20 ° BTe,
then rotate propeller back in the direction of rotation
until 20 0 BTC is under the pointer. (This step removed the factor of backlash. )
j. The breaker contacts should be just starting to
open after completion of step "i." U not, proceed
to setp "k. "
k. Loosen magneto mounting clamps enough to permit magneto to be rotated.
1. Using a timing light connected across the breaker
contacts, slowly move magneto in its normal direction of cam rotation until the contacts have just closed,
then rotate in the opposite direction until timing light
indicates position at which contacts break. Secure
magneto.
m. Turn the propeller back a few degrees (approximately 5°) to close contacts.
NOTE
Do not turn propeller back far enough to
engage impulse coupling, or propeller
will have to be turned in normal direction
of rotation until impulse coupling releases,
then again backed up to a few degrees before
the firing position.
n. Slowly advance propeller (tap forward with minute movements as firing position is approached) in
normal direction of rotation until timing light indi-
10-52
Attachm~nt
cates position at which contacts break. The contacts
should break at 20° +00 -2°BTC. Rotate magneto to
make contacts break at correct position.
Do not adjust contacts to compensate for incorrect magneto-to-engine timing. Breaker
contact adjustment is for internal timing only,
and any readjustment after internal timing has
been accomplished will result in a weaker
spark, with reduced engine performance.
o. After tightening magneto mounting clamps and
rechecking magneto-to-engine timing, remove timing
equipment. Install and connect all spark plugs that
were removed.
10-91. MAGNETO CHECK. Advanced timing settings in some cases, is the result of the erroneous
practice of bumping magnetos up in timing in order
to reduce RPM drop on Single ignition. NEVER ADVANCE TIMING BEYOND SPECIFICATIONS IN ORDER TO REDUCE RPM DROP. Too much importance
is being attached to RPM drop on single ignition.
RPM drop on single ignition is a natural characteristic of dual ignition design. The purpose of the
following magneto check is to determine that all cylinders are firing. U all cylinders are not firing,
the engine will run extremely rough and cause for
investigation will be quite apparent. The amount of
RPM drop is not necessarily significant and will be
influenced by ambient air temperature, humidity,
airport altitude. etc. In fact, absence of RPM drop
should be cause for suspicion that the magneto timing has been bumped up and Is set in advance of the
settings specified. Magneto checks should be performed on a comparative basis between individual
right and left magneto performance.
a. Start and run engines unW the oil and cylinder
head temperatures are in the normal operating
ranges.
•
•
•
b. Place the propeller control in the full low pitch
(high rpm) position.
c. Advance engine speed to 1800 rpm.
d. Turn the ignition switch to the "R" position and
note the rpm drC'o, then return the sw itch to the
"BOTH" positior: to clear the OPPOSite set of plugs.
e. Turn the switch to the "L" position and note the
rpm drop, then return the switch to the "BOTH"
position.
f. The rpm drop should not exceed 150 rpm on
either magneto or show greater than 50 rpm differential between magnetos. A smooth rpm drop-off
past normal is usually a sign of a too lean or too
rich mixture. A sharp rpm drop-off past normal
is usually a sign of a fouled plug, a defective harness
lead or a magneto out of time. If there is doubt concerning operation of the ignition system, rpm checks
at a leaner mixture setting or at higher engine speeds
will usually confirm whether a deficiency exists.
NOTE
An absence of rpm drop may be an indication
of faulty grounding of one side of the ignition
system, a disconnected ground lead at magneto or possibly the magneto timing is set
too far in advance.
•
10-92. MAINTENANCE. At the first 25-hour inspection and at each 100-hour inspection thereafter,
the breaker compartment should be inspected. Magneto-to-engine timing should be checked at the first
25-hour inspection, first 50-hour inspection, first
100-hour inspection and thereafter at each 100-hour
inspection. If timing is 20° (plus zero, minus 2°),
internal timing need not be checked. If timing is
out of tolerance, remove magneto and set internal
timing, then install and time to the engine.
NOTE
If ignition trouble should develop, spark plugs
and ignition wires should be checked first. If
the trouble appears definitely to be associated
with a magneto, the following may be used to
help disclose the source of trouble without
overhauling the magneto.
a. Moisture Check.
1. Remove the high-tension outlet plate, cables
and grommet, and inspect for moisture.
2. Inspect distributor block high-tension outlet
side for moisture.
3. If any moisture is evident, lightly wipe with
a soft, dry, clean, lint-free cloth.
Do not use gaSOline or other solvents, as
these will remove the wax coating on some
parts and could cause electrical leakage.
•
3. Check breaker contacts for excessive wear,
burning, deep pits and carbon depoSits. Contacts
may be cleaned with a hard-finish paper. Replace
defective breaker assemblies. Make no attempt to
stone or dress contacts. Clean new contacts with
clear, unleaded gasoline and hard-finish paper before installing.
4. Check cam follower oiling felt. If it appears
dry, re-oil with 2 or 3 drops of lubricant (Scintilla
10-86527, or equivalent). Allow about 30 minutes
for the felt to absorb the oil, then blot off excess
with a clean cloth. Too much oil may result in fouling and excessive burning of contacts.
5. Check that the condenser mounting bracket
is not cracked or loose. If equipment is available,
check condenser for a minimum capacitance of . 30
microfarads. If equipment for testing is not available and a defective condenser is suspected, replace
with a new one.
c. If the trouble has not been corrected after ac- .
complishing steps "a" and "b, " check magneto-toengine timing. If timing is not within prescribed
tolerance, remove magneto and set internal timing,
then reinstall and time to the engine.
d. If the trouble has still not been corrected, magneto overhaul or replacement is indicated.
10-93. TACHOMETER BREAKER POINT ADJUSTMENT. (BEGINNING WITH AIRCRAFT SERIALS
337-0526 AND F33700001.) The right magneto of
each engine contains a second set of breaker points
for operation of the tachometer. A tachometer pickup coil is installed in the bottom of the magneto. To
adjust the breaker points, turn rotating magnet until
the tachometer breaker point cam follower is on the
highest part of cam lobe and measure the clearance
between contacts. Adjust clearance to 0.019.:1:0.003
inch.
10-94. SPARK PLUGS. Two spark plugs are installed in each cylinder and screw into helicoil type
thread inserts. The spark plugs are shielded to
prevent spark plug noise in the radios and have an
internal resistor to provide longer terminal life.
Spark plug service life will vary with operating conditions. A spark plug that is kept clean and properly
gapped will give better and longer service than one
that is allowed to collect lead depOSits and is improperly gapped.
NOTE
At each 100-hour inspection, remove, clean,
inspect and regap all spark plugs. Install
lower spark plugs in upper portion of cylinders and install upper spark plugs in lower
portion of cylinders. Since deterioration of
lower spark plugs is usually more rapid
than that of the upper spark plugs, rotating
helps prolong spark plug life.
10-95. STARTING SYSTEM.
b. Breaker Compartment Check .
1. Remove breaker cover.
2. Check all parts of the breaker assembly for
security.
10-96. DESCRIPTION. An electric starter, mounted on a 90-degree starter adapter, is used on each
engine. The starter solenoid for the front engine, or
10-53
the one for the rear engine, is acUvated when the corresponding ignition switch is turned to the "START"
position. Wben the starter solenoid is actuated, its
contacts close and electrical current energizes the
starter motor. Initial rotation of the starter motor
engages the starter through an overrunning clutch in
the starter adapter, which incorporates worm reduction gears. The starter is located just aft of the right
rear cylinder.
•
10-97. TROUBLE SHOOTING
TROUBLE
STARTER Wll.L NOT OPERATE.
STARTER MOTOR RUNS, BUT
DOES NOT TURN CRANKSHAFT.
STARTER MOTOR DRAGS.
,
STARTER EXCESSIVELY
NOISY.
_.
PROBABLE CAUSE
REMEDY
Defective master switch or
circuit.
Check continuity. Install new
switch or wires.
Defective starter SWitch or
switch circuit.
Check continuity. Install new
SWitch or wires.
Defective starter motor.
Check voltage to starter. Repair
or replace starter motor.
Defective overrunning clutch
or drive.
Remove starter and inspect.
Install new starter adapter.
Starter motor shaft broken.
Install new starter motor.
Low battery.
Charge or install new battery.
Starter switch or relay contacts
burned or dirty.
Check continuity. Install
serviceable unit.
Defective starter motor
power cable.
Check visually. Install
new cable.
Loose or dirty connections.
Check visually. Remove, clean
and tighten all terminal connections.
Defective starter motor.
Check starter motor brushes,
brush spring tension, thrown
solder on brush cover. Repair
or install new starter motor.
Dirty or worn commutator.
Check visually. Clean and
turn commutator.
Worn starter pinion.
Remove starter and inspect.
Replace starter drive.
Worn or broken teeth
on crankshaft gears.
Check visually. Replace
crankshaft gear.
•
(
•
10-54
•
10-98. STARTER MOTOR •
10-99. REMOVAL AND INSTALLATION.
a. Remove engine cowling as required for access.
I:SAUTIONI
When disconnecting starter electrical cable,
do not permit terminal bolt to rotate. Rotation of the bolt could break the conductor
between bolt and field coils causing the
starter to be inoperative.
b. Disconnect battery cables and insulate terminals
as a safety precaution.
c. Disconnect electrical cable at starter motor.
d. Remove nuts and washers securing motor to
starter adapter and remove motor. Refer to engine
manufacturer's overhaul manual for adapter removal.
e. Reverse the preceding steps for reinstallation.
Install a new 0- ring seal on motor, then install motor.
Be sure motor drive engages with the adapter drive
when installing.
10-100. PRIMARY MAINTENANCE. The starting
circuit should be inspected at regular intervals, the
frequency of which should be determined by the
amount of service and conditions under which the
equipment is operated. Inspect the battery and wiring. Check battery for fully charged condition, proper electrolyte level with approved water and terminals for cleanliness. Inspect wiring to be sure that
all connections are clean and tight and that the wiring
insulation is sound. Check that the brushes slide
freely in their holders and make full contact on the
commutator. When brushes are worn to one-half of
their orig1nallength, install new brushes (compare
brushes with new brushes). Check the commutator
for uneven wear, excessive glazing or evidence of
excessive arcing. If the commutator is only slightly
dirty, glazed or discolored, it may be cleaned with a
strip of No. 00 or No. 000 sandpaper. If the commutator is rough or worn, it should be turned in a lathe
and the mica undercut. Inspect the armature shaft
for rough bearing surfaces. New brushes should be
properly seated When installing by wrapping a strip
of No. 00 sandpaper around the commutator (with
sanding side out) 1-1/4 to 1-1/2 times maximum.
Drop brushes on sandpaper covered commutator and
turn armature slowly in the direction of normal rotation. Clean sanding dust from motor after sanding
operations.
10-101. EXTREME WEATHER MAINTENANCE.
•
10-102. COLD WEATHER. Cold weather starting
is made easier by using the engine priming system
and the ground service receptacle. The priming
system is manuall y-operated from the cockpit. Fuel
is supplied by a line from the fuel strainer to the
plungers. Operating the primers forces fuel to the
intake manifold of each engine. With the external
power receptacle, an external power source may be
connected to assist in cold weather or low battery
starting. Refer to paragraph 10-106 for use of the
ground service receptacle.
The following may also be used to assist engine starting in extreme cold weather. After the last flight of
the day, drain the engine oil into a clean container so
the oil can be preheated. Cover the engines, including the rear air scoop opening to prevent ice or snow
from collecting inside the cowling. When preparing
the aircraft for flight or engine runup after these conditions have been followed, preheat the drained engine
oU.
IWARNING'
Do not heat the oil above 121 DC (250 DF). A
flash fire may result. Before pulling the
propeller through, ascertain that the ignition
switches are in the OFF positioQ to prevent
accidental firing of the engines.
After preheating the engine oil, gasoline may be mixed with the heated oil in a ratio of 1 part gasoline to
12 parts engine oil before pouring into the engine oil
sumps. If the free air temperature is below minus
29 DC (_20 DF), the engine compartments should be
preheated by a ground heater. After the engine compartments have been preheated, inspect all engine
drain and vent lines for presence of ice. Remove
the protective covers placed on the engines and rear
air scoop opening. After this procedure has been
complied with, pull propellers through several revolutions by hand before attempting to start the engines.
I~~UTION\
Due to the desludging effect of the diluted
oil, engine operation should be observed
closely during the initial warm-up of the
engines. Engines that have considerable
amount of operational hours accumulated
since their last dilution period be seriously
affected by the dilution process. This will
be caused by the diluted oil dislodging sludge
and carbon deposits within the engines.
This residue will collect in the oil sumps
and possibly clog the screened inlets to the
oil sumps. Small depOSits may actually
enter the oil pumps and be trapped by the
main oil filters. Partial or complete loas
of engine lubrication may result from either
condition. If these conditions are anticipated
after oil dilution, the engines should be run
for several minutes at normal operating temperatures and then stopped and inspected for
evidence of sludge and carbon deposita in the
oil sumps and oil filters. Future occurrence
of this condition can be prevented by diluting
the oil prior to each engine oil change. This
will also prevent the accumulation of the sludge
and carbon deposits.
10-103. HOT WEATHER. In hot Weather, With a hot
engine, fuel may vaporize at certain points in the fuel
system. Vaporized fuel may be purged by setting the
mixture control in the "IDLE CUT-OFF" poSition and
operating the auxiliary fuel pump on "HI. "
10-55
Engine mis-starts characterized by weak, intermittent explosions followed by puffs of black smoke from
the exhausts are caused by over-priming or flooding.
This situation is more apt to develop in hot weather
or When the engine is hot. U it occurs, repeat the
starting routine with the throttle approximately onehalf "OPEN, " the mixture control in "IDLE CUTOFF" and the auxiliary fuel pump switch "OFF." As
the engine fires, move the mixture control to full
"RICH" and decrease the throttle to desired idling
speed.
Engine mis-starts characterized by sufficient power
to disengage the starter but dying after 3 to 5 revolutions are the result of an excessively lean mixture
after the start. This can occur in either warm or
cold temperatures. Repeat the starting routine but
allow additional priming time with the auxiliary fuel
pump switch on "LO" before cranking is started, or
place the auxiliary fuel pump switch on "HI" immediately for a richer mixture while cranking.
@AUTION\
If prolonged cranking is necessary, allow the
10-104. SEACOAST AND HUMID AREAS. In salt
water areas, special care should be taken to keep
the engines, accessories and airframe clean to prevent oxidation. In humid areas, fuel and oil should
be checked frequently and drained of condensation to
prevent corrosion.
•
10-105. DUSTY AREAS. Dust induced into the intake
systems of the engines is probably the greatest single
cause of early engine wear. When operating in high
dust conditions, service the induction air filter daily
as outlined in Section 2. Also change engine oil and
lubricate airframe items more often than specified.
10-106. GROUND SERVICE RECEPTACLE. With
the ground service receptacle installed, the use of
an external power source is recommended for cold
weather starting, low battery starting and lengthy
maintenance of the aircraft electrical system. Refer
to Section 15 for additional information.
10-107. HAND-CRANKING. Starting may also be
accomplished by band-cranking the front engine.
After the front engine has been started, use electrical power to start the rear engine.
starter motor to cool at frequent intervals,
since excessive heat may damage the starter.
SHOP NOTES:
•
•
10-56
SECTION lOA
•
ENGINES
(TURBOCHARGED)
(AIRCRAFT SERIALS 337-0526 THRU 3370l.}}8
AND F33700001 THRU F33700045)
TABLE OF CONTENTS
•
•
ENGINE COWLING
Description
Front.
.
Rear
Removal and lnstallation
Cleaning and lnspection
Repair
ENGINES
Description
Engine Qata
Trouble Shooting
Removal
Front.
Rear
Cleaning
Accessories Removal
Inspection .
Build-Up
Installation
Front.
Rear
.
Flexible Fluid Hoses
Pressure Test
Replacement
Engine BaIfles .
Description
Cleaning and lnspection
Removal and lnstallation
Repair
Engine Mount
Description
Removal and lnstallation
Repair
Engine Shock-Mount Pads
COWL FLAPS
Description •
Trouble Shooting.
.
Removal and lnstallation
Rigging
.
Front.
Rear
CONTROL QUADRANT
.
Description
•
Removal and lnstallation
Disassembly and Reassembly
Page
IOA-2
IOA-2
IOA-2
IOA-2
IOA-2
IOA-2
IOA-2
IOA-2
IOA-2
IOA-3
lOA-4
IOA-8
IOA-8
IOA-9
10A-IO
10A-IO
10A-IO
10A-IO
10A-IO
10A-IO
10A-12
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-13
IOA-14
IOA-14
IOA-14
IOA-14
ENGINE CONTROLS
Description
.
Removal and lnstallation
Rigging
.
Throttle-Operated Gear Warning
Switches
Description
Rigging
•
INDUCTION AIR SYSTEM .
Description
.
•
Removal and lnstallation
Front..
.
Rear
.
Cleaning Induction Air Filter
FUEL INJECTION SYSTEM
Description
Trouble Shooting .
Fuel-Air Control Unit
Description
Removal and lnstallation
Adjustments.
.
Fuel Manifold Valve (Fuel
Distributor)
.
Description
.
Removal and lnstallation
Cleaning.
.
Fuel Discharge Nozzles .
Description
Removal
Cleaning and lnspection .
lnstallation
Fuel lnjection Pump
Description
Removal and Installation
Adjustments .
EXHAUST SYSTEMS
Description
Front Engine.
Rear Engine .
Removal
Front Engine
Rear Engine .
lnspection .
lnstallation
.
Front Engine .
IOA-14
lOA.., 14
IOA-14
IOA-14
IOA-14
IOA-14
IOA-14
IOA-14
IOA-14
IOA-14
IOA-14
IOA-16
IOA-16
IOA-16
IOA-16
IOA-18
IOA-19
IOA-19
IOA-19
IOA-19
IOA-19
IOA-20
IOA-20
IOA-20
IOA-20
IOA-20
IOA-20
IOA-20
IOA-20
IOA-20
IOA-20
IOA-20
IOA-21
IOA-21
IOA-21
IOA-21
IOA-21
IOA-21
IOA-21
IOA-24
IOA-24
IOA-24
IOA-25
IOA-l
Rear Engine .
ruRBOCHARGER
.
Description
.
Removal and Installation
Front Engine.
Rear Engine .
Controllers and Waste-Gate
Actuator..
.
Functions
.
Operation. . .
•
Trouble Shooting .
Removal and Installation
Variable Controller.
Rate-of-Change Controller.
Waste-Gate and Actuator
Adjustments •
Variable Controller.
Rate-of-Change Controller
Waste-Gate and Actuator
Operational Flight Check
JIL SYSTEM
Description
Trouble Shooting .
Full-Flow Oil Filter
Description
Element Removal and Installation
Adapter Removal. . • . . .
Adapter Disassembly, Inspection
and Reassembly
.
10A-25
10A-25
10A-25
10A-25
10A-25
10A-27
lOA-27
10A-27
10A-27
10A-3l
10A-33
10A-33
10A-33
10A-33
10A-33
10A-33
lOA-33
10A-36
10A-38
10A-40
10A-40
10A-40
lOA-40
10A-40
10A-40
10A-40
Adapter Installation.
IGNITION SYSTEM .
Description
Trouble Shooting
Magnetos
Description
Removal and Installation
Internal Timing
Magneto-to- Engine Timing
Magneto Check .
Maintenance .
.
Tachometer Breaker Point
Adjustment
Spark Plugs
ST ARTING SYSTEM
Description
Trouble Shooting
Starter Motor
Removal and Installation
Primary Maintenance .
EXTREME WEATHER MAINTENANCE
Cold Weather
Hot Weather .
Seacoast and Humid Areas
Dusty Areas .
.
Ground Service Receptacle
.
•
Hand Cranking.
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
10A-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
•
lOA-40
lOA-40
10A-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
lOA-40
10A-40
IDA-I. ENGINE COWLING.
10A-2. DESCRIPTION.
a. FRONT. The front engine cowling is similar to
that described in Section 10, except it is wider at the
front, with additional ram air openings in the right
and left DOse caps. The opening in the right side
supplies ram air to the turbocharger. The opening
in the left side supplies ram air to the cabin heating
system.
b. REAR. The rear engine cowling is similar to
that described in Section 10, except it is larger at
the tail cap and only one exhaust outlet protrudes
through the lower portion of the cowl. The larger
tall cap permits greater engine cooling and the additional space needed for the installation of the turbocharger system.
stalled on the aircraft. Both engines are located on
the fuselage centerline, one forward and one aft of
the cabin. The engines themselves are similar, although their front (propeller) ends point in opposite
directions. A conventional tractor propeller is required for the front engine and a pusher propeller is
required for the rear engine. Each propeller rotates
in the same direction in relation to its engine, but
rotates in opposite directions in relation to each other.
Cooling for the rear engine is obtained by an overhead
air scoop and laterally mounted cowl flaps. Refer to
paragraph lOA-8 for engine data. For repair and
overhaul of the engines, accessories and propellers,
refer to the appropriate publications issued by their
manufacturer's. These publications are available
from the Cessna Service Parts Center.
•
NOTE
10A-3. REMOVAL AND INSTALLATION. Refer to
paragraph 10-3.
10A-4. CLEANING AND INSPECTION. Refer to
paragraph 10-4.
IOA-5. REPAIR. Refer to paragraph 10-5.
10A-6. ENGINES.
lOA-7. DESCRIPTION. Air cooled, wet sump, six
cylinder, horizontally-opposed, fuel-injected, turbocharged Continental TSIo-360 series engines are in-
lOA-2
Since the installed engines face in opposite
directions, some confusion might arise
from terms such as "right, " "left, " "front"
and "rear". Except where further clarified
in the text, these terms shall be applie~ to
the rear engine as though it were removed
from the aircraft and Viewed from its accessory case end. Rear engine bafnes, cowling
and firewall are not considered part of the
baSic engine and shall be identified as "right, "
"left, " etc., in relation to the aircraft.
•
•
10A-8. ENGINE DATA .
MODEL (Continental)
Aircraft serials 337-0526 thru 337-0755
BegilU'J.ng With aircraft serial 337 -0756
TSIQ-360-A (Front) TSIQ-360-B (Rear)
TSIQ-360-A (Front and Rear)
BHP at RPM
210 at 2800
Limiting Manifold Pressure (Sea Level)
32 Inches Hg.
Number of Cylinders
6-Horizontally-Opposed
Displacement
Bore
Stroke
360 Cubic Inches
4.438 Inches
3.875 Inches
Compression Ratio
7.5:1
Magnetos
Right Magneto
Left Magneto
Bendix-Model S6LN- 25
Fires 20° BTC Upper Right and Lower Left
Fires 20° BTC Upper Left and Lower Right
Firing
1-6-3-2-5-4
Spark Plugs
18MM x .750-20 (Refer to current Continental active
factory approved spark plug chart.)
Torque Value
•
330±30 Lb- In.
Fuel Metering System
Unmetered Fuel Pressure
Continental Fuel Injection
6.5 to 7.5 PSI at 600 RPM FRONT or 650 RPM REAR
29 to 32 PSI at 32.5 Hg. and 2800 RPM
Oil Sump CapaCity
With Filter Element Change
10 U. S. Quarts
11 U.s. Quarts
Tachometer
Electric (Operated by Magneto Pick-Up)
Oil Pressure
Minimum Idling
Normal
Maximum (Cold Oil Starting)
Connection Location
10 PSI
30 to 60 PSI
100 PSI
Between No. 2 and No. 4 Cylinders (Front and Rear)
Oil Temperature
Normal Operation
Maximum Permissible
Within Green Arc
Red Line (240°F)
Cylinder Head Temperature
Probe Location
460°F Maximum
Lower Side No. 2 Cylinder (Rear) Thru aircraft
serial 337-1193
Lower Side No. 5 Cylinder (Front) Thru aircraft
serial 337-1193
Lower Side No. 1 Cylinder (Front and Rear) Beginning with aircraft serials 33701194 and F33700001
Approximate Dry Weight with Standard
Accessories (Excluding Turbocharger
System)
327 Pounds
•
lOA-3
lOA-9. TROUBLE SHOOTING.
TROUBLE
ENGINE FAILS TO START.
ENGINE STARTS BUT DIES, OR
WILL NOT IDLE PROPERLY.
10A-4
PROBABLE CAUSE
REMEDY
--
Engine flooded or improper use
of starting procedure.
Use proper starting procedure.
Refer to Owner's Manual.
Defective aircraft fuel system.
Refer to Section 11.
Fuel tanks empty.
Service fuel tanks.
Spark plugs fouled or defective.
Remove, clean, inspect and regap.
Use new gaskets. Check cables
to persistently fouled plugs. Replace if defective.
Magneto impulse coupling failure.
Repair or install new coupling.
Defective magneto switch or
grounded magneto leads.
Repair or replace switch and leads.
Defective ignition system.
Refer to paragraph 10-92.
Induction air leakage.
Correct cause of air leakage.
Clogged fuel screen in fuel control
unit or defective unit.
Remove and clean. Replace
defective unit.
Clogged fuel screen in fuel
manifold valve or defective
valve.
Remove and clean screen. Replace
defective valve.
Clogged fuel injection lines or
discharge nozzles.
Remove and clean lines and nozzles.
Replace defective units.
Defective auxiliary fuel pump.
Refer to Section 11.
Engine-driven fuel pump not
permitting fuel from auxiliary
pump to bypass.
Install new engine-driven
fuel pump.
Vaporized fuel in system. (Most
likely to occur in hot weather with
a hot engine. )
Refer to paragraph 10-103.
Propeller control in high pitch
(low rpm) position.
Use low pitch (high rpm) poSition
for all ground ope rations.
Improper idle speed or idle
mixture adjustment.
Refer to paragraph 10-55.
Defective aircraft fuel system.
Refer to Section 11.
Spark plugs fouled or defective.
Remove, clean, inspect and regap.
Use new gaskets. Check cables to
persistently fouled plugs. Replace
if defective.
Water in fuel system.
Drain fuel tank sumps, lines
and fuel strainer.
Defective ignition system.
Refer to paragraph 10-92.
•
•
•
•
10A-9. TROUBLE SHOOTING (Cont).
TROUBLE
ENGINE STARTS BUT DIES, OR
WILL NOT IDLE PRO PERL Y
(Cont).
•
ENGINE HAS POOR ACCELERATION, RUNS ROUGHLY
AT SPEEDS ABOVE IDLE OR
LACKS POWER.
•
PROBABLE CAUSE
lnl~'.1ction
air leakage.
REMEDY
Correct cause of air leakage.
Clogged fuel screen in fuel
control unit or defective unit.
Remove and clean. Replace
defective unit.
Clogged fuel screen in fuel manifold valve or defective valve.
Remove and clean. Replace
defective valve.
Restricted fuel injection lines
or discharge nozzles.
Remove, clean lines and nozzles.
Replace defective units.
Defective engine-driven fuel
pump.
lnstaU and calibrate new pump.
Vaporized fuel in system.
(Most likely to occur in hot
weather with a hot engine. )
Refer to paragraph 10-103.
Manual engine primer leaking.
Disconnect primer ouUet line.
If fuel leaks through primer,
repair or replace primer.
Obstructed air intake.
Remove obstruction; service
air filter, if necessary.
One or more cylinder head
drain lines broken or disconnected.
Connect lines; replace if broken.
Discharge nozzle air vent
manifolding restricted or
defective.
Check for bent lines or loose connections. Tighten loose connections. Remove restrictions and
replace defective components.
Defective engU:ie.
Check compression and listen for
unusual engine noises. Check oil
filter for excessive metal. Repair
engine as required.
Idle mixture too lean.
Refer to paragraph 10-55.
Propeller control in high pitch
(low rpm) poSition.
Use low pitch (high rpm) position
for all ground operations.
lncorrect fuel-air mixture,
worn control linkage or
restricted air filter.
Replace worn elements of
control linkage. Service
air filter.
Defective ignition system.
Refer to paragraph 10-92.
Malfunctioning turbocharger.
Check operation, listen for unusual
noise. Check operation of wastegate valve and for exhaust system
defects. Tighten loose connections.
Improper fuel-air mixture.
Check intake manifold connections
for leaks. Tighten loose connections. Check fuel controls and linkage for setting and adjustment.
Check fuel filter screens for dirt.
Check for proper pump pressure.
lOA-5
IOA-9. TROUBLE SHOOTING (Cant).
TROUBLE
ENGINE HAS POOR ACCELERATlON, RUNS ROUGHLY
AT SPEEDS ABOVE IDLE
OR LACKS POWER (Cant).
POOR IDLE CUT-OFF.
ENGINE LACKS POWER, REDUCTION IN MAXIMUM
MANIFOLD PRESSURE OR
CRITICAL ALTITUDE.
IOA-6
PROBABLE CAUSE
REMEDY
Defective fuel injection system.
lkler to paragraph IOA- 51.
Spark plugs fouled or defective.
Remove, clean, inspect and regap.
Use new gaskets. Check cables to
persistently fouled plugs. Replace
of defective.
Engine or engine mount attaching
bolts loose or broken.
Torque as specified. Replace if
defective.
Defective engine shock-mount.
Replace defective parts.
Interference between engine
mount and cowling.
Check for positive clearance.
Edges of cowling stiffener s
and doublers may be ground
for clearance.
Propeller out of balance.
Check and balance propeller.
Defective engine.
Check compression, check oil
filter for excessive metal.
Listen for unusual noises.
Repair engine as required.
Exhaust system leakage.
Refer to paragraph IOA-72.
Turbocharger wheels rubbing.
Replace turbocharger.
Improperly adjusted or defective
variable controller.
Refer to paragraph IOA-82.
Leak in turbocharger discharge
pressure system.
Correct cause of leaks. Repair
or replace damaged parts.
Manifold pressure overshoot.
(Most likely to occur when engine
is accelerated too rapidly. )
Move throttle about two-thirds
open. Let engine accelerate
and peak. Move throttle to
full open.
Engine oil viscosity too high
for ambient air.
Refer to Section 2 for proper
grade of all.
Mixture controlllnkage improperly rigged.
Refer to paragraph 10-41.
Defective or dirty fuel manifold
valve.
Remove and clean manifold
valve.
Fuel contamination.
Drain all fuel and flush out fuel
system. Clean all screens, fuel
strainers, fuel manifold valves,
nozzles and fuel lines.
Defective mixture control
valve in fuel pump.
Replace fuel pump.
Incorrectly adjusted throttle
control, "sticky" linkage or
dirty air filter.
Check movement of linkage by moving control through range of travel.
Make proper adjustments and replace worn components. Service
air filter.
•
•
•
•
10A-9. TROUBLE SHOOTING (Cont) .
TROUBLE
ENGINE LACKS POWER, REDUCTION IN MAXIMUM
MANIFOLD PRESSURE OR
CRITICAL ALTITUDE (Cont).
REMEDY
Defective ignition system.
Inspect spark plugs for fouled
electrodes, heavy carbon deposits, erosion of electrodes,
improperly adjusted electrode
gaps and cracked porcelains.
Test plugs for regular firing
under pressure. Replace damaged or misfiring plugs.
Improperly adjusted waste-gate
valve.
Refer to paragraph lOA-82.
Loose or damaged exhaust
system.
Inspect entire exhaust system to
turbocharger for cracks and
leaking connections. Tighten
connections and replace damaged
parts.
Loose or damaged manifolding.
Inspect entire manifolding system
for possible leakage at connections.
Replace damaged components,
tighten all connections and clamps.
Fuel discharge nozzle defective.
Inspect fuel discharge nozzle vent
manifolding for leaking connec~
tions. Tighten and repair as required . Check for restricted
nozzles and lines and clean and
replace as necessary.
Malfunctioning turbocharger.
Check for unusual noise in turbocharger. If malfunction is suspected, remove exhaust and/or
air inlet connections and check
rotor assembly, for possible
rubbing in housing, damaged
rotor blades or defective bearings.
Replace turbocharger if damage is
noted.
BLACK SMOKE EXHAUST.
Turbo coking, oil forced through
seal of turbine housing.
Clean or change turbocharger.
HIGH CYLINDER HEAD
TEMPERATURE.
Defective cylinder head temperature indicating system.
Refer to Section
Improper use of cowl flaps.
Refer to Owner's Manual.
Defective cowl flap operating
system.
Refer to paragraph 10-3l.
Engine baffles loose, bent or
missing.
Install baffles properly. Repair or
replace if defective.
Dirt accumulated on cylinder
cooling fins.
Clean thoroughly .
Incorrect grade of fuel.
Drain and refill with proper fuel ..
•
•
PROBABLE CAUSE
<0
~4.
IOA-7
10A-9. TROUBLE SHOOTING (Cont).
TROUBLE
PROBABLE CAUSE
HIGH CYLINDER HEAD
TEMPERATURE (Cont).
Incorrect
i~tion
REMEDY
Refer to paragraph 10-90.
timing.
Defective fuel injection system.
Refer to paragraph lOA-51.
Improper use of mixture control.
Refer to Owner's Manual.
Defective engine.
Repair as required.
HIGH OR LOW On..
TEMPERATURE
OR PRESSURE.
•
Refer to paragraph 10-76.
NOTE
Refer to paragraph 10A-aO for trouble shooting of controller
and waste-gate actuator.
10A-10. REMOVAL. If an engine is to be placed in
storage or returned to the manufacturer for overhaul, proper preparatory steps should be taken for
corrosion prevention prior to beginning the removal
procedure. Refer to Section 2 for storage preparation. The routing and location of Wires, cables,
lines, hoses and controls will vary with optional
equipment installed, however, the following general
procedure may be followed.
a. FRONT. The front engine may be removed as a
complete unit with the turbocharger and accessories
installed.
I~AUTIONl
Place suitable padded stands under the tail
boom tie-down rings before removing front
engine. The loss of front engine weight will
cause the aircraft to be tail heavy.
NOTE
Tag each item when disconnected to aid in
idenWying wires, hoses, lines and control
linkages when engine is reinstalled. Likewise, shop notes made during removal will
often clarify reinstallation. Protect openings, exposed as a result of removing or
disconnecting units, against entry of foreign
material by installing covers or sealing With
tape.
1. Place all cabin switches in the OFF position.
2. Place fuel selector valves in the OFF posi-
tion.
3. Remove engine cowling and nose caps in accordance with paragraph 10-3.
4. Disconnect battery cables, remove battery
and battery box for additional clearance, if desired.
5. Drain fuel strainer and lines with strainer
drain control.
lOA-a
NOTE
During the follOWing procedures, remove
any clamps which secure controls, wires,
hoses or lines to the engine, engine mounts
or attached brackets, so they will not interfere With the engine removal. Some of these
items listed can be disconnected at more
than one place. It may be desirable to disconnect some of these items at other than
the placed indicated. The reason for engine
removal should be the governing factor in
deciding at which point to disconnect them.
Omit any of the items which are nor present
on a particular engine installation.
•
6. Remove induction air inlet flexible duct at
right front side of engine for access to engine mount.
7. Disconnect control and remove heater from
left side of engine.
8. Remove vacuum hoses from pump and suction
relief valve and remove de-ice components from firewall.
9. Place propeller control in high-rpm position.
Release unfeathering accumulator pressure through
the filler valve and disconnect hose at accumulator.
10. Drain the engine oil sump and oil cooler.
11. Disconnect magneto primary lead wires at
magnetos.
IWARNING,
The magnetos are in a SWITCH ON condition
when the switch Wires are disconnected.
Ground the magneto points or remove the
high tension wires from the magnetos or
spark plugs to prevent accidental firing.
12. Remove the spinner and propeller in accordance with Section 12. Cover the exposed end of the
crankshaft to prevent entry of foreign material.
•
•
13. Disconnect throttle, mixture and propeller
governor controls. Remove clamps attaching controls to engine and pull controls aft clear of engine.
Use care to avoid bending controls too sharply.
14. Disconnect oil temperature wire at sending
ur.:t.
15. Disconnect tachometer pick-up wire from
' .... m ')i right magneto.
When disconnecting starter cable do not
permit starter terminal bolt to rotate.
Rotation of the bolt could break the conductor between bolt and field coils causing the starter to be inoperative.
•
16. Disconnect starter electrical cable at starter.
17. Disconnect cylinder head temperature wire at
probe.
18. Disconnect electrical wires and wire shielding
ground at alternator.
19. Disconnect electrical wires at throttle-operated switch.
20. Disconnect exhaust gas temperature wires at
probe leads.
21. Disconnect ground strap and any other electrical wiring not previously noted which may be damaged during engine removal.
22. Disconnect fuel strainer drain control wire at
strainer bellcrank and remove control housing lock
nuts securing housing to nose gear tunnel. Pull control and housing from tunnel area.
23. Disconnect vacuum hose at suction relief
valve if not completed during step 8.
24. Disconnect supply and pressure hoses at firewall and hydrauliC filter. Remove hydraulic pump
drain line.
25. Disconnect manifold pressure line at firewall.
26. Disconnect fuel supply hose at nose gear tunnel and vapor return hose at firewall. Remove fuel
pump drain line.
IWARNING'
Residual fuel and oil draining from disconnected lines and hoses constitutes a fire
hazard. Use caution to prevent accumulation of such fuel and oil when lines or hoses
are disconnected.
•
27. Disconnect fuel-flow gage hose at firewall.
28. Disconnect oil pressure hose at firewall.
29. Visconnect cylinder fuel drain line at hose
connection on each side of engine.
30. Disconnect fuel-flow gage vent line at firewall.
31. Disconnect engine primer line at firewall.
32. Disconnect air inlet duct at turbocharger
compressor.
33. Carefully check the engine again to ensure
ALL hoses, lines, wires, cables and clamps are
disconnected or removed which would interfere with
the engine removal. Ensure all wires, cables and
engine controls have been pulled aft to clear the engine.
34. Attach a hoist to the lifting eye at the top center of the engine crankcase. Lift engine just enough
to relieve the weight from the engine mounts.
35. Remove bolts attaching engine to engine
mounts and slowly hoist engine and pull it forward.
Checking for any items which would inter'~re with
the engine removal. Balance the engine l'y hand and
carefully guide the disconnected parts out as the engine is removed.
36. Remove the engine shock-mounts.
b. REAR. The rear engine may be removed as a
complete unit WITH the turbocharger, accessories
and engine mount installed, or WITHOUT the turbocharger and engine mount installed. The following
procedures outline engine removal with the turbocharger system and engine mount left on the aircraft.
NOTE
Tag each item disconnected to aid in identifying Wires, hoses, lines and control linkages
when the engine is reinstalled. Likewise,
shop notes made during removal will often
clarify reinstallation. Protect openings, exposed as a result of removing or disconnecting
units, against entry of foreign material by installing covers or sealing with tape.
1. Place all cabin switc.hes in the OFF position.
2. Place fuel selector valves in the OFF position.
3. Remove ALL engine cowling in accordance
with paragraph 10-3.
4. Remove front engine left upper cowl section,
disconnect battery ground cable and insulate terminal
as a safety precaution.
5. Drain fuel strainer and lines with strainer
drain control.
NOTE
During the following procedures, remove any
clamps or lacings which secure controls,
Wires, hoses or lines to the engine, engine
mount or attached brackets, so they will not
interfere with engine removal. Some of the
items listed can be disconnected at more
than one place. It may be desirable to disconnect some of these items at other than
the places indicated. The reason for engine
removal should be the governing factor in
deCiding at which point to disconnect them.
Omit any of the items which are not present
on a particular engine.
6. Remove induction air filter and adapter fastened to engine baffle. Disconnect compressor inlet
duct and remove duct.
7. Remove de-ice components from firewall.
8. Drain the engine oil sump and oil cooler.
9. Disconnect magneto primary lead wires at
magnetos.
10A-9
IWARNING'
The magnetos are in a SWITCH ON condition
when the sw'tcb wires are disconnected.
Ground the luagneto points or remove the
high tension wires from tbe magnetos or
spark plugs to prevent accidental firing.
10. Remove the spinner and propeller in accordance with Section 12. Cover the exposed end of the
crankshaft to prevent entry of foreign material.
11. Disconnect throttle, mixture and propeller
governor controls. Remove any clamps attaching
controls to engine and pull controls clear of engine.
Use care to avoid bending controls too sharply.
12. Disconnect oil temperature wire at sending
unit.
13. Disconnect tachometer pick-up wire from
bottom of right magneto.
f~AUTIONI
When disconnecting starter cable do not permit starter terminal bolt to rotate. Rotation
of the bolt could break the conductor between
bolt and field coils causing the starter to be
inoperative.
14. Disconnect starter electrical cable at starter.
15. Disconnect cylinder bead temperature wire
at probe.
16. Disconnect electrical wires and wire shielding ground at alternator.
17. Disconnect electrical wire s at tbrotUe-operated switch.
18. Disconnect exhaust gas temperature wires at
probe leads.
19. Disconnect ground strap and any other electrical wiring not previously noted whicb may be damaged during engine removal.
20. Disconnect fuel strainer drain control Wire
at strainer and remove control housing lock nuts securing housing to fuselage structure. Pull control
and housing from structure area.
21. Disconnect vacuum hose at vacuum pump if
not completed during step 7.
22. Disconnect manifold pressure line at firewall.
23. Disconnect fuel supply hose at auxiliary pump,
vapor return hose at firewall and fuel pump drain line.
IWARNING'
Residual fuel and oil draining from disconnected lines and hoses constitutes a fire
hazard. Use caution to prevent accumulation of such fuel and oil wben lines or hoses
are disconnectad.-- - - 24. Disconnect fuel-flow gage bose at firewall.
25. Disconnect oil pressure hose at firewall.
26. Disconnect cylinder fuel drain line at hose
connection at each side of engine.
27. Disconnect fuel-flow gage vent line at firewall.
10A-10
28. Disconnect engine primer line at firewall .
29. Disconnect drain lines protruding through
fuselage skin to prevent damage.
30. Disconnect oil hoses at waste-gate actuator.
Plug or cap hoses and fittingS.
31. Disconnect oil hoses to turbocharger. Plug
or cap hoses and fittings.
32. Disconnect supply and pressure hoses at firewall and hydraulic filter. Remove hydraulic pump
drain line.
33. Disconnect exhaust pipes at collector on eacb
side of tbe engine, so that turbocharger, waste-gate
actuator, waste-gate and exhaust tailpipe may be left
in tbe aircraft. Note that exhaust system braces are
attached to tbe aft engine mount bolts.
34. Remove bolts attacbing turbocharger to support brackets.
35. Remove turbocharger ouUet air duct.
36. Carefully check tbe engine again to ensure
ALL boses, lines, wires, cables and clamps are
disconnected or removed which would interfere with
the engine removal. Ensure all wires, cables and
engine controls have been pulled forward to clear the
engine.
37. Attach a boist to the lifting eye at the top center of tbe engine crankcase. Lift engine just enougb
to relieve the weigbt from the engine mount assembly.
•
/CAUTION\
Be sure there is clearance at tbe top of tbe
tail section, as tbe tall section of the aircraft will rise with the loss of engine weight.
38. Remove bolts attaching engine to engine
mount, slowly hoist the engine and pull it aft.
39. Balance the engine by hand and carefully
work the engine from aircraft, guiding the disconnected parts as the engine is removed.
40. Remove engine shock-mounts.
•
lOA-U. CLEANING. Refer to paragraph 10-1l.
10A-12. ACC ESSORIES REMOVAL. Refer to paragraph 10-12.
10A-13. INSPECTION. Refer to paragraph 10-13.
lOA-H. BUILD-UP. Refer to paragraph 10-14.
lOA-IS. INSTALLATION.
a. FRONT. Before installing the front engine on
the aircraft, install any items Which were removed
from tbe engine or aircraft after the engine was removed.
NOTE
Remove all protective covers,
and identification tags as each
nected or installed. Omit any
present on a particular engine
plugs, caps
item is conitems not
installation.
1. Hoist the engine to a point just above the
nacelle.
•
•
2. Install engine shock-mount pads as illustrated
in figure 10-1.
3. Carefully lower engine slowly into place on
the engine mount pads. Route controls, lines, hoses
and wires in place as the engine is positioned on the
engine mounts.
NOTE
Be sure engine shock-mount pads, spacers
and washers are in place as the engine is
lowered into position.
4. Install engine mount bolts, washers and nuts,
then remove the hoist and tail boom support stands.
Torque bolts to 450-500 lb-in.
5. Connect air inlet duct to turbocharger compressor.
6. Route throttle, mixture and propeller governor controls to their respective units and connect.
Secure controls in position with clamps.
7. Connect engine primer line at firewall.
8. Connect fuel-flow gage vent line at firewall.
9. Connect cylinder fuel drain lines at hose connection on each side of engine.
10. Connect oil pressure hose at firewall.
11. Connect fuel-flow gage hose at firewall.
12. Connect fuel supply hose and vapor return
line at tunnel and firewall. Install fuel pump drain
line.
•
•
NOTE
Throughout the aircraft fuel system, from the
fuel tanks to the engine-driven fuel pump, use
RAS-4 (Snap-On Tools Corp., Kenosha, Wisconsin), MIL- T- 5544 (Thread Compound, Antiseize, Graphite- Petrolatum) or equivalent, as
a thread lubricant or to seal a leaking connection. Apply sparingly to male fittings only,
omitting the first two threads. Always ensure
that a compound. the residue from a previously
used compound or any other foreign material
cannot enter the system. Throughout the fuel
injection system, from the engine-driven pump
through the discharge nozzles, use only a fuel
soluble lubricant, such as engine lubricating
oil, on fitting threads. Do not use any other
form of thread compound on the injection system fittings.
13. Connect manifold pressure line at firewall.
14. Connect vacuum hose at suction relief valve.
15. Connect supply and pressure hoses at firewall and hydraulic filter. Install hydraulic pump
drain line.
16. Install all clamps and lacings securing hoses
and lines to the engine or structure.
17. Connect ground strap to engine mount.
18. Connect exhaust gas temperature wires to
probe leads. Be sure wires are not crossed.
19. Connect electrical Wires to throttle-operated
switch .
20. Connect wires and wire Shielding ground to
alternator.
21. Connect cylinder head temperature wire to
probe .
/CAUTION\
When connecting starter cable, do not permit
starter terminal bolt to rotate. Rotation of
the bolt could break the conductor between
bolt and field coils causing the starter to be
inoperative.
22. Connect starter electrical cable at starter.
23. Connect tachometer pick-up wire at bottom of
right magneto.
24. Connect oil temperature wire at sending unit.
25. Install all clamps and lacings securing \!tires
and cables to the engine or structure.
26. Route the fuel strainer drain control through
the nose gear tunnel structure to the strainer, install
the lock nuts to secure housing and connect control
wire to strainer control bell crank.
27. Install propeller and spinner in accordance
with instructions outlined in Section 12.
28. Complete a magneto switch ground-out and
continuity check, then connect primary lead wires to
the magnetos. Remove the temporary ground or connect spark plug leads, whichever procedure was used
during removal.
IWARNING'
Be sure magneto switch is in OFF position
when connecting switch wires to magnetos.,
29. Connect unfeathering accumulator hose at
accumulator and service accumulator in accordance
with Section 12.
30. Install de-ice components and hoses on firewall.
31. Install heater and connect control.
32. Clean induction air filter and install filter
and induction air inlet duct.
33. Service engine with proper grade and quantiy
of engine oil. Refer to Section 2 if engine is new,
newly overhauled or has been in storage.
34. Check all switches are in the OFF position,
install battery box and battery and connect cables.
35. Rig engine controls in accordance with paragraph 10-41.
36. Inspect engine installation for security, correct routing of controls, lines, hoses and electrical
wiring, proper safetying and tightness of all components.
37. Install engine cowling in accordance with
paragraph 10-3.
38. Check cowl flaps and rig in accordance with
paragraph 10-33, if necessary.
NOTE
When installing a new or newly overhauled
engine and prior to starting the engine,
disconnect the oil inlet line at the controller
and oil outlet line at the controller. Connect
these oil lines to a full flow oil filter, allow-
lOA-ll
ing oil to bypass the controller. With filter
installed, operate engine for approximately
15 minutes to filter out any foreign particles
from the oil. This is done to prevent foreign
material from entering the controller.
S~. Perform an engine run-up and make final adjustments on the engine controls.
b. REAR. Before installing the rear engine on the
aircraft, reinstall any items which were removed
from the engine or aircraft after the engine was removed.
NOTE
Remove all protective covers, plugs, caps
and identification tags as each item is connected or installed. Omit any items not
present on a particular engine installation.
1. Hoist the engine assembly to a point near the
engine mount and route controls, lines, hoses and
wires in place.
2. Install engine shock-mount pads as illustrated
in figure 10-1.
3. Carefully work engine assembly in position
on the engine_D!0unt.
NOTE
Be sure shock-mount pads, spacers and
washers are in place as the engine is
lowered into position.
4. Install engine mount bolts, washers and nuts,
then remove the hoist. Torque bolts to 450-500 lb-in.
NOTE
Exhaust stack braces are secured to the
aft engine mount bolts.
5. Install bolts attaching turbocharger to support
brackets.
6. Connect exhaust pipes to collector on each
side of the engine.
7. Connect supply and pressure hoses at firewall and hydraulic filter. Install hydraulic pump
drain line.
8. Connect oil hoses to turbocharger.
9. Connect oil hoses to waste-gate actuator.
10. Route throttle, mixture and propeller governor controls to their respective units and connect.
Secure controls in position with clamps.
11. Connect drain lines protruding through fuselage skin.
12. Connect engine primer line at firewall.
13. Connect fuel-flow gage vent line at firewall.
14. Connect cylinder fuel drain line at hose connections on each side of engine.
15·. Connect oil pressure hose at firewall.
16. Connect fuel flow gage hose at firewall.
NOTE
Throughout the aircraft fuel system, from
the fuel tanks to the engine- driven fuel
pump, use RAS-4 (Snap-On Tools Corp •.
Kenosha, Wisconsin), MIL-T-5544 (Thro:ad
Compound, Antiseize, Graphite-Petrolatum)
or equivalent, as a thread lubricant or to
seal a leaking connection. Apply sparingly
to male fittings only, omitting the first two
threads. Always be sure that a compound,
the residue from a previously used compound or any other foreign material cannot
enter the system. Throughout the fuel injection system, from the engine-driven fuel
pump through the discharge nozzles, use
only a fuel soluble lubricant, such as engine
lubricating Oil, on the fitting threads. Do
not use any other form of thread compound
on the injection system fittings.
17. Connect fuel supply hose to auxiliary pump,
vapor return hose at firewall and fuel pump drain line.
18. Connect manifold pressure line at firewall.
19. Connect vacuum pump hose at vacuum pump.
20. Install all clamps and lacings securing hoses
and lines to engine, engine mount or structure.
21. Route the strainer drain control through fuselage structure to the strainer, install control housing
lock nuts securing housing to structure and connect
control wire to strainer.
22. Connect ground strap to engine mount.
23. Connect exhaust gas temperature wires at
probe leads. Be sure wires are not crossed.
24. Connect electrical wires at throttle-operated
switch.
25. Connect electrical wires and wire Shielding
ground at alternator.
26. Connect cylinder head temperature wire at
probe.
•
I~AUTIONI
When connecting starter cable, do not permit starter terminal bolt to rotate. Rotation of the bolt could break the conductor
between bolt and field coils causing the
starter to be inoperative.
27. Connect starter electrical cable at starter.
28. Connect tachometer pick-up wire at bottom
of right magneto.
29. Connect oil temperature wire at sending unit.
30. Install all clamps and lacings securing wires
and cables to engine, engine mount or structure.
31. Install propeller and spinner in accordance
with instructions outlined in Section 12.
32. Complete a magneto switch ground-out and
continuity check, then connect primary ground or
connect spark plug leads, whichever procedure was
used dUring removal.
IWARNING'
Be sure magneto switch is OFF when connecting primary leads to magnetos.
10A-12
•
•
•
33. Install de-ice components and hoses on firewall .
34. Install induction air ducts, clean air filter
and install adapter.
35. Service engine with proper grade and quantity of engine oil. Refer to Section 2 if engine is new,
newly overhauled or has been in storage.
36. Check all switches are in the OFF position
and connect battery ground cable.
37. Rig engine controls in accordance with paragraph 10-4l.
3B. Check engine installation for security, correct routing of controls, lines, hoses and electrical
wiring, proper safetying and tightness of all components.
39. Install engine cowling in accordance with
paragraph 10-3.
40. Check cowl flaps and rig in accordance with
paragraph 10-33, if necessary.
NOTE
•
When installing a new or newly overhauled
engine and prior to starting the engine, disconnect the oil inlet line at the controller
and oil outlet line at the controller. Connect these oil lines to a full flow oil filter,
allowing oil to bypass the controller.
With filter installed, operate engine for
approximately 15 minutes to filter out any
foreign particles from the oil. This is
done to prevent foreign material from
entering the controller.
41. Perform an engine run-up and make final
adjustments on the engine controls.
10A-16. FLEXIBLE FLUID HOSES. Refer to paragraph 10-16.
IOA-17.
10-17.
PRESSURE TEST. Refer to paragraph
lOA-lB. REPLACEMENT.
10-IB.
Refer to paragraph
10A-19. ENGINE BAFFLES. Refer to paragraph
10-19.
10A-20. DESCRIPTION. Refer to paragraph 10-20.
IOA-21. CLEANING AND INSPECTION. Refer to
paragraph 10-21.
10A-22. REMOVAL AND INSTALLATION. Refer
to paragraph 10-22.
10A-23. REPAIR.
Refer to paragraph 10-23.
10A- 24. ENGINE MOUNT. Refer to paragraph
10-24.
•
IOA-25. DESCRIPTION. Refer to paragraph 10-25 .
IOA-26. REMOVAL AND INSTALLATION. Refer
to paragraph 10-26.
IOA-27. REPAIR. Refer to paragraph 10- 27.
IOA-2B. ENGINE SHOCK-MOUNT PADS. Refer to
paragraph 10-2B.
IOA-29. COWL FLAPS.
IOA-30. DESCRIPTION. The front and rear cowl
flaps are the same as described in paragraph 10-30
except for the follow-up control attachment location.
10A-3l. TROUBLE SHOOTING. Refer to paragraph
10-31.
IOA-32. REMOVAL AND INSTALLATION. Refer to
paragraph 10- 32.
lOA-33. RIGGING.
a. FRONT. (Refer to paragraph 10-33.) Rigging
of the FRONT cowl flaps on turbocharged aircraft is
the same as outlined in paragraph 1C- 33 except for
the follow-up control attachment location.
b. REAR. (THRU AIRCRAFT SERIAL 337-0755
WHEN NOT MODIFIED IN ACCORDANC E WITH
SK337-B.) (Refer to paragraph 10-33.)
1. Complete steps 1 thru 16 of subparagraph
"d."
2. Operate cowl flaps to the single-engine position and measure travel at trailing edge. The cowl
flap should open 6. 50 +.25 -.00 inches, but still remain open. 50 inch in the CLOSED position. Readjust push-pull rods to bellcranks and cowl flaps and
select a different hole in the bellcranks as required
for proper travel and opening in the closed position.
Check that the stops on the bellcranks just clear the
engine mount tubes. Lower wing flaps cautiously
with cowl flaps full open, and check for at least 1/4
inch clearance in any pOSition.
3. Place rear cowl flap control lever in the
NORMAL OPEN position (twin-engine operation) and
check that the cowl flaps are open 4 ± • 25 inches,
measured at the trailing edges.
4. Check that all rod ends and cleviS ends have
sufficient thread engagement, all jam nuts are tight
and all safeties are installed.
c. REAR. (AIRCRAFT SERIALS 337-0756 THRU
337-097B AND ALL AIRCRAFT MODIFIED IN ACCORDANCE WITH SK337-B. )
1. Complete steps 1 and 2 of paragraph 10-33,
subparagraph "e. "
2. Complete steps 1 thru 16 of paragraph 10-33,
subparagraph "d. "
3. Complete steps 1 thru 3 of paragraph 10A-33,
subparagraph ''b.''
4. Complete steps 4 thru 7 of paragraph 10-33,
subparagraph "e. "
d. REAR. (BEGINNING WITH AIRCRAFT SERIALS
337-0979 AND F33700001.)
1. Complete steps 1 thru 3 of paragraph 10- 33,
subparagraph "c. "
2. Install RIGHT HAND cowl panel, connect
horizontal push-pull rod (49) to torque tube arm (24).
The cowl flap (56) must be open. 50 inch in the
CLOSED POSITION. If not, readjust push-pull rods
as necessary.
10A-13
3. Using jumper wires and external power supply, run the right hand cowl flap to open 6.50+.25-.00
inches, measured from the trailing edge of cowl flap
to aft edge of cowl flap opening.
10A-37. DISASSEMBLY AND REASSEJdBLY. Refer
to paragraph 10- 37.
10A-38. ENGINE CONTROLS. Refer to paragraph
10-38.
(CAUTIONI
lOA-39. DESCRIPTION. Refer to paragraph 10-39.
Do not use master switch before rigging has
been completed. When using jumper wires
connect only one wire to motor lead and
"strike" the other Wire against other motor
lead. If motor does not move in the correct
direction, reverse jumper leads.
10A-40. REMOV AL AND INSTALLATION. Refer to
paragraph 10-40. Omit any references to the intake
heater controls.
10A-41. RIGGING. Refer to paragraph 10-41.
Omit any references to the intake heater controls.
4. Loosen screws on OPEN-LIMIT switch (43)
and adjust switch toward actuating bracket (47) until
switch just de-actuates.
10A-42. THROTTLE-OPERATED GEAR WARNING
SWITCHES. Refer to paragraph 10-42.
NOTE
10A-43. DESCRIPTION. Refer to paragraph 10-43.
Opening of the microawitch may be determined
by listening for a faint "click, " or continuity
may be checked.
5. Complete step 10 of paragraph 10- 33, subparagraph "d. "
6. Install LEFT HAND cowl panel, connect vertical push-pull rod (49) to torque tube arm (24) and
horizontal push-pull rod (53) to cowl flap (56). The
cowl flap should be open 6. 50+.25-.00 inches in the
open position. If not, readjust push-pull rods as
necessary.
7. Place master switch in the ON position and
using cowl flap toggle switch (index 41, sheet 1),
slowly run the cowl flaps to the CLOSED position
and check that the LEFT cowl flap is open . 50 inch
in the CLOSED position. If not, readjust the push-pull
rods as necessary.
NOTE
In all cases, the final result of rigging, is
that the cowl flaps are to be open . 50 inch
in the CLOSED pOSition and are to be open
6.50+.25-.00 inches in the OPEN position.
8. Using toggle switch (index 41, sheet 1), run
cowl flaps through several cycles. Check pOSition
indicating lights for operation. Stop cowl flaps at
intermediate openings to check toggle switch operation.
9. Check that all rod ends and clevis ends have
sufficient thread engagement, all jam nuts are tight,
all safeties are installed and reinstall upper cowling
section.
10A-34. CONTROL QUADRANT. Refer to paragraph
10-34.
lOA-35. DESCRIPTION. Refer to paragraph 10-35.
IOA-36. REMOVAL AND INSTALLATION. Refer to
paragraph 10-36.
lOA-14
•
10A-44. RIGGING. Refer to paragraph 10-44.
10A-45. INDUCTION AIR SYSTEM.
10A-46. DESCRIPTION. Ram air to the front engine
induction system enters an air duct at the right side
of the nose cap cowling. Ram air to the rear engine
induction system enters from the air scoop above the
fuselage. The air is filtered through a dry filter, located in the induction airbox on each engine. From
the filter, the air passes through an air duct to the
inlet of the turbocharger compressor where the air
is compressed. The pressurized induction air is
then routed through an air duct to the fuel-air control
unit mounted on the top side of the engine. From the
fuel-air control unit, the air is supplied to the cylinders through the right and left intake manifolds located on the top side of the cylinders. The fuel-air
control unit is connected to the cylinder intake manifold by hoses and clamps. The intake manifold is attached to each cylinder by two bolts through a welded
flange, which is sealed by a gasket. An alternate air
door, mounted in the air duct between the filter and
the turbocharger compressor is held closed by a
small magnet. If the filter should become clogged,
suction from the turbocharger compressor will cause
the alternate air door to open. This permits the compressor to draw heated unfiltered air from within the
engine compartment. The alternate air door should
be checked periodically for freeness of operation
and complete closing. The induction filters should
be cleaned, inspected and replaced as outlined in
Section 2.
10A-47. REMOVAL AND INSTALLATION.
a. FRONT. (Refer to figure lOA-I.)
1. Remove filter access door on right side of
lower cowl.
2. Pull filter (12) from airbox (6).
3. Remove engine cowling as required for access to the upper duct (4) and airbox (6).
4. Loosen clamps and remove turbocharger compressor outlet duct from compressor outlet and engine
baffle.
•
•
•
TO
1. Clamp
2. Compressor Inlet Duct
3. Seal
4. Duct
5. Access Cover
6. Airbox Assembly
7. Bracket
8. Magnet
9. Alternate Air Door
10. Shim
11. Hinge Pin
12. Air Filter
COMPRESSOR
INLET
l
....
<)
<)
<)
..
6
~
.
l~.{· ,
•
10
Figure lOA-I. Frcnt Engine Induction Air System
IOA-15
5. Working through air filter access door, remove bolt from inboard side of filter cavity.
6. Working through air filter access door, remove screws attaching upper inlet air duct to lower
cowl.
7. Loosen clamp and disconnect upper duct (4)
from nose cap inlet.
8. Work upper inlet air duct (4) aft and out of
aircraft.
9. Remove screws attaching airbox assembly to
nose gear tunnel.
10. Remove screws attaching airbox assembly to
lower cowling.
11. Loosen clamps (1) and remove compressor
inlet air duct (2) from airbox and compressor.
12. Work airbox assembly from aircraft.
13. Reverse the preceding steps for reinstallation.
NOTE
When installing the air filter, ascertain that
the filter fits snugly in airbox. The area
between the upper inlet air duct and airbox
is adjustable by the addition of a shim between the upper air duct and duct mounting
bracket on the lower cowling. Also, the
inboard side of the filter area is adjustable
by loosening the bolt and sliding the duct up
or down.
b. REAR. (Refer to figure 10A-7.)
1. Remove left half of rear cowling.
2. Remove hardware attaching filter (1) to air
inlet duct (2).
3. Remove hardware attaChing air inlet duct to
horizontal baffle and remove duct.
4. Remove bolt attaching airbox to clamp on
engine mount.
5. Loosen and remove clamps securing flexible
duct (6) to compressor and engine mount. Remove
duct and airbox assembly.
6. Remove clamps securing compressor discharge tube (18) to compressor and throttle body
and work tube out of engine compartment.
7. Reverse the preceding steps for reinstallation.
10A-48. CLEANING INDUCTION AIR FILTER.
Refer to Section 2.
10A-49. FUEL INJECTION SYSTEM. (Refer to
figure 10A-2.)
10A-16
lOA-50. DESCRIPTION. The fuel injection system
is a low-pressure system of injecting metered fuel
into the intake valve ports in the cylinders. It is a
multi-nozzle, continuous-flow system which controls
fuel now to match engine airflow. Any change in
throttle position, engine speed or a combination of
both, causes changes in fuel flow in the correct relation to engine airflow. A manual mixture control and
a fuel-flow indicator are provided for leaning at any
combination of altitude and power setting. The four
major components of the system are: the fuel injection pump, fuel-air control unit, fuel manifold (distributor) valve and the fuel discharge nozzles. The
fuel injection pump incorporates an adjustable aneroid sensing unit which is pressurized from the discharge side of the turbocharger compressor. Turbocharger discharge air pressure is also used to vent
the fuel discharge nozzles and the vent port of the fuelflaw indicator. Since the intake manifolds are installed
on the top side of the cylinders, drain lines are installed in the bottom side of the intake ports to drain fuel
which may have accumulated in the intake ports during
engine shut-down.
•
NOTE
Throughout the aircraft fuel system, from the
tanks to the engine-driven fuel pump, use
Never-Seez RAS-4 (Snap-On Tools Corporation,
Kenosha, Wisconsin) or MIL-T-5544 (Thread
Compound, Antiseize, Graphite-Petrolatum) or
eqUivalent, as a thread lubricant or to seal
a leaking connection. Apply sparingly to
male fittings only, omitting the first two
threads on the fitting. Always be sure that
a compound, the residue from a previously
used compound or any other foreign material
cannot enter the system. Throughout the
fuel injection system, from the engine-driven
fuel pump through the discharge nozzles, use
only a fuel soluble lubricant, such as engine
lubricating oil, on the fitting threads. Do
not use any other form of thread compound
on the injection system fittings.
IWARNING,
Residual fuel draining from lines and hoses
constitutes a fire hazard. Use caution to
prevent accumulation of fuel when lines or
hoses are disconnected throughout the fuel
injection system.
•
•
TO
FUEL
~--
VAPOR EJECTOR JET
THROTTLE VALVE
TO
FUEL FLOW
-INDICATOR --g~~m~~iim5
•
AIR FROM
TURBOCHARGER
DISCHARGE
MANIFOLD
VALVE
TURBOCHARGER
DISCHARGE AIR
TO FUEL FLOW
INDICATOR
NOTE
Turbocharger discharge air from the
fuel pump relief valve to the aneroid
chamber is an internal passage in the
fuel pump.
CALIBRATED ORIFICE
LEGEND:
IIIIIIII
•
INLET FUEL
PUMP PRESSURE
petail
\
CD
FUEL METERED BY ANEROID VALVE
[JJ
FUEL METERED BY MIXTURE CONTROL
rmJ
FUEL METERED BY THROTTLE CONTROL
~
FUEL VAPOR RETURNED TO TANK
o
TURBOCHARGER DISCHARGE AIR PRESSURE
Figure 10A-2.
A
<
INJECTION
MIXTURE
OUTLET
Fuel Injection Schematic
lOA-17
lOA-51. TROUBLE SHOOTING.
TROUBLE
NO FUEL DELIVERED
TO ENGINE.
HIGH FUEL PRESSURE.
ENGINE RUNS ROUGH
AT IDLE.
lOA-I8
PROBABLE CAUSE
REMEDY
Fuel tanks empty.
Service with desired quantity of
fuel.
Defective aircraft fuel system.
Refer to Section 11.
Vaporized fuel. (Most likely
to occur in hot weather with a
hot engine. )
Refer to paragraph 10-103.
Fuel pump not permitting fuel
from electric pump to bypass.
Check fuel-flow through pump.
Replace engine-driven fuel pump
if defective.
Defective fuel control unit.
Check fuel flow through unit.
Replace fuel-air control unit
if necessary.
Defective fuel manifold valve,
or clogged screen inside valve.
Check fuel flow through valve.
Remove and clean in accordance with paragraphs 10-58
and 10-59. Replace if defective.
Clogged fuel injection lines or
discharge nozzles.
Check fuel flow through lines and
nozzles. Clean and replace if defective.
Restricted discharge nozzles.
Clean or replace plugged nozzle
or nozzles.
Restriction in vapor vent return
line or check valve.
Clean vapor return line. Clean
or replace check valve.
Improper idle mixture adjustment.
Refer to paragraph 10-55.
Restriction in aircraft fuel
system.
Refer to Section 11.
Low unmetered fuel pressure.
Refer to paragraph 10A-68.
High unmetered fuel pressure.
Refer to paragraph IOA-68.
Worn throttle plate shaft or
shaft O-rings.
Replace shaft and/or O-rings.
Intake manifold leaks.
Repair leaks or replace
defective parts.
Leaking intake valves.
Engine repair required.
Discharge nozzle air vent
manifolding restricted or
defective.
Check for bent or loose
connections, restrictions
or defective components.
Tighten loose connections;
replace defective components.
•
•
•
•
lOA-51. TROUBLE SHOOTING (Cont).
TROUBLE
FLUCTUATING FUEL
PRESSURE OF FUEL
FLOW.
Replace manifold valve.
Restriction in engine-driven
fuel pump vapor ejector.
Clean vapor ejector on fuel
pump. Do not use wires to
clean jet.
Defective check valve in
vapor vent return line.
Clean vapor return vent
line and repair or replace
check valve.
Air in line from manifold
valve to gage.
Bleed air from line.
Malfunctioning relief valve
in engine-driven fuel pump.
Clean or replace relief
valve if defective.
Defective gage or restricted
gage line.
Replace gage. Clean
restriction from line.
Plugged main fuel strainer.
Clean strainer.
Air leak on suction side of
engine-driven fuel pump.
Repair leak. Replace
defective parts.
FUEL DRAINING
FROM MANIFOLD
VALVE VENT.
Ruptured diaphragm.
Replace diaphragm or
manifold valve.
POOR IDLE CUT-OFF.
Dirt 1n fuel pump or
defective pump.
Remove pump and flush
out thoroughly. Check
that mixture arm contacts
cut-off stop.
Dirty or defective fuel
manifold valve.
Remove and clean in
accordance with paragraphs 10- 58 and 10- 59.
IDA-52. FUEL-AIR CONTROJ.. UNIT. Refer to
paragraph 10-52.
IDA-53. DESCRIPTION. Refer to paragraph 10-53.
•
REMEDY
Defective manifold valve.
LOW METERED FUEL
PRESSURE.
•
PROBABLE CAUSE
IDA-54. REMOVAL AND INSTALLATION.
a. Remove cowling as required to gain access.
b. Turn fuel selector valves to OFF position.
c. Tag and disconnect hoses at fuel metering unit.
Cap or plug disconnected hoses and fittings.
d. Disconnect manifold pressure line at fuel-air
control unit.
e. Disconnect throttle control at air throttle arm.
Note position of washers.
f. Disconnect variable controller rod at air throttle
arm. Note position of washers and spacers. Do not
rotate rod end.
,
g. Remove four bolts, washers and nuts attaching
air inlet duct to throttle body. Lay parts of landing
gear warning switch to one side. Note any other parts
attached by these bolts.
h. Loosen clamps securing throttle body to intake
manifold and slide hoses away from throttle body.
1. Remove bolts, washers and nuts attaching fuelair control unit to bracket on engine and remove unit.
Cover open ends of manifold and air inlet duct.
j . Reverse the preceding steps for reinstallation.
Rig throttle, throttle-operated landing gear warning
switch and variable controller.
IDA-55. ADJUSTMENTS. Refer to
para~raph
10-55.
lOA-56. FUEL MANIFOLD VALVE (FUEL DISTRIBUTOR). Refer to paragraph 10-56.
10A-19
lOA-56. DESCRIPTION. Refer to paragraph 10-57.
lOA-58. REMOVAL AND INSTALLATION. Refer to
paragraph 10- 58.
lOA-59. CLEANING. Refer to paragraph 10-59.
10A-60. FUEL DISCHARGE NOZZLES.
lOA-61. DESCRIPTION. From the fuel manifold
valve, individual, identical size and length fuel lines
carry metered fuel to the fuel discharge nozzles located in the cylinder heads. The outlet of each nozzle is directed into the intake port of each cylinder.
The nozzle body contains a drilled central passage
with a counterbore at each end. The lower end is
used as a chamber for fuel-air mixture before the
spray leaves the nozzle. The upper bore contains
an orifice for calibrating the nozzles. Near the top,
radial holes connect the upper counterbore with the
outside of the nozzle body for air admission. These
radial holes enter the counterbore above the orifice
and draw outside air through a cylindrical screen
fitted over the nozzle body. This screen prevents
dirt and foreign material from entering the nozzle.
A press-fit shield is mounted on the nozzle body and
extends over the greater part of the filter screen, .
leaving a small opening at the bottom of the shield.
This provides an air bleed into the nozzle which aids
in vaporizing the fuel by breaking the high vacuum
in the intake manifold at idle rpm and keeps the fuel
lines filled. The nozzles are calibrated in several
ranges. All nozzles furnished for one engine are the
same range and are identified by a number and a suffix letter stamped on the flat portion of the nozzle
body. When replacing a fuel discharge nozzle, be
sure it Is of the same calibrated range as the rest
of the nozzles in the engine. When a complete set of
nozzles is being installed, the number must be the
same as the one removed, but the suffix letters may
be different, as long as they are the same for all
nozzles being installed on a particular engine.
10A-62. REMOVAL.
a. Remove engine cowling as required for access.
NOTE
Plug or cap all disconnected lines and fittings.
Use care to prevent damage to fuel injection
lines.
b. Disconnect nozzle pressurization line at nozzles
and disconnect pressurization line at union fitting so
that pressurization line may be moved away from discharge nozzles.
c. Disconnect fuel injection line at discharge nozzle.
d. Using care to prevent damage or loss of washers
and 0- rings, lift sleeve assembly from discharge
nozzle.
e. Using a standard 1/2-inch deep socket, remove
fuel discharge nozzle from cylinder.
10A-63. CLEANING AND INSPECTION. Refer to
paragraph 10- 63.
10A-20
10A-64. INSTALLATION.
a. Using a standard 1/2-inch deep socket, install
nozzle body in cylinder and tighten to a torque value
of 60-80 lb-in.
b. Install O-rings, sleeve assembly and washers
on nozzle bodies.
c. Align sleeve assembly and connect pressurization lines to nozzles. Connect pressurization line to
union fitting.
d. Install 0- ring and washer at top of discharge
nozzle and connect fuel injection line to nozzle.
e. Inspect installation for crimped lines and loose
fittings.
f. Inspect nozzle pressurization vent system for
leakage. A tight system is required, since turbocharger discharge pressure is applied to various
other components of the injection system.
g. Install cowling.
•
10A-65. FUEL INJECTION PUMP.
lOA-66. DESCRIPTION. The fuel pump is a positivedisplacement, rotating vane type, located OPPOSite the
propeller governor at the propeller end of the engine.
Fuel enters the pump at the swirl well of the pump
vapor separator. Here, vapor is separated by a
swirling motion so that only liquid fuel is fed to the
pump. The vapor is drawn from the top center of
the swirl well by a small pressure jet of fuel and is
fed into the vapor return line, where it is returned
to the fuel line manifold. Since the pump is enginedriven, changes in engine speed effect total pump
flow proportionally. A check valve allows the auxiliary fuel pump pressure to bypass the engine-driven
fuel pump for starting, or in the event of enginedriven fuel pump failure in flight. The pump supplies
more fuel than is required by the engine; therefore,
a spring-loaded, diaphragm type relief valve is provided to maintain a constant fuel pump pressure.
The engine-driven fuel pump is equipped with an
aneroid valve. The aneroid valve and relief valve
are pressurized from the discharge side of the turbocharger compressor to maintain a proper fuel/air
ratio at altitude. The aneroid valve is adjustable for
fuel pump outlet pressure at full throttle and the
relief valve is adjustable for fuel pump outlet pressure in the idle rpm range. Refer to paragraph
10A-68 for pressure adjustments. The fuel pump is
equipped with a manual mixture control to limit the
fuel pump output from full rich to idle cut-off. Nonadjustable mechanical stops are located at these
positions.
10A-67. REMOVAL AND INSTALLATION.
a. Turn fuel selector valves to the OFF pOSition.
b. Remove cowling, baffles and covers as necessary to gain access.
c. Disconnect mixture control from lever on pump.
Note position of washers.
d. Tag and disconnect fuel hoses and vent line
attached to pump. Plug or cap all disconnected hoses
and fittings.
e. Disconnect and plug or cap air vent line at fuel
pump.
f. Remove mounting nuts and bolts and pull pump
and gasket from engine pad.
•
•
•
IWARNINGt
Residual fuel draining from lines and hoses
constitutes a fire hazard. Use caution to
prevent accumulation of fuel when lines or
hoses are disconnected.
g. The drive shaft coupling may come off with the
fuel pump, or it may remain in the engine. If it
comes off with the pump, reinstall it in the engine
to prevent dropping or losing it.
h. If a replacement pump is not being installed immediately, a temporary cover should be installed on
the fuel pump mount pad.
i. Reverse the preceding steps for reinstallat1on.
Using a new gasket, do not force engagement of the
pump drive. Rotate engine crankshaft and pump
drive Will engage smoothly when aligned properly.
Check mixture control rigging.
j. Start engine and perform an operational check,
adjust fuel pressure as required in accordance with
paragraph IOA-68.
IOA-68. ADJUSTMENTS. (Refer to figure IOA-3.)
a. Remove engine cowling as required for access.
b. Remove cap from fuel metering unit. Using test
hose and fittings, connect test gage pressure port
into the fuel injection system as illustrated in figure
IOA-3.
•
NOTE
Cessna Service Kit No. SK320-2J proVides
a test gage, line and fittings for connecting
the test gage into the system to perform
accurate calibration of the engine-driven
fuel pump.
c. Allow engine to warm-up. Set mixture control
full rich and propeller control full forward (low pitch,
high-rpm).
d. Idle engine at 600 :t: 25 rpm (front engine) or 650
:!:: 25 rpm (rear engine) and check for fuel pressure
specified in paragraph 10A-8.
IWARNING'
DO NOT make fuel pressure adjustments
while engine is operating.
e. If pressure is not Within prescribed tolerances,
stop engine and adjust pressure by turning the screw
on the fuel pump relief valve (turn IN to increase
pressure and OUT to decrease pressure) to obtain
correct pressure and repeat steps tIc and d. "
NOTE
After adjusting fuel pressure, idle speed and
idle mixture must be readjusted (refer to
paragraph 10-55).
•
f. Advance throttle to obtain maximum rpm and
check for fuel pressure speCified in paragraph IOA-8.
IWARNINGa
DO NOT make fuel pump pressure adjustments while the engine is operating.
g. U pressure is not Within prescribed tolerances,
stop engine and adjust pressure by loosening locknut
and turning the slot-headed needle valve located just
below the fuel pump inlet fitting counterclockwise
(CCW) to increase pressure and clockwise (CW) to
decrease pressure. Repeat steps "f and g" unW
pressure is witbin prescribed tolerances speCified in
paragraph lOA-So
h. After correct pressure is obtained, safety adjustable orifice locknut and remove test equipment.
i. Install cowling.
j. Repeat preceding steps for other engine if adjustment is required.
IOA-69. EXHAUST SYSTEMS.
10A-70. DESCRIPTION. Each engine exhaust system consists of two exhaust stack assemblies, one
for the left and one for the right bank of cylinders.
The exhaust stack assemblies of each engine are
joined together to route the exhaust from all cVlinders of that engine through the waste-gate or turbine.
a. FRONT ENGINE. The three risers on the left
bank of cylinders are joined together into a common
pipe to form the left stack assembly. The three
risers on the right bank of cylinders are joined together into a common pipe to form the right stack
assembly. The left stack assembly connects to the
right stack assembly at the front of the engine.
Mounting pads for the waste-gate and turbine are
provided at the rear of the right stack assembly.
From the exhaust port of the turblne, a tailpipe
routes the exhaust overboard through the lower
cowling. The exhaust port of the waste-gate is
routed into the tailpipe so the exhaust gases can be
expelled from the system when not needed at the
turbine.
b. REAR tNGINE. The rear exhaust system routes
the exhaust gases into a common turbine inlet assembly, then overboard through a single tailpipe. The
exhaust stacks are made in sections that are clamped
together. The turbine inlet assembly contains an
outlet for the waste-gate valve. Exhaust from the
waste-gate is routed into the tailpipe 10 the exhaust
gases can be expelled from the system when not
needed at the turbine.
IOA-7l. REMOVAL. (Refer to figure IOA-4. )
a. FRONT ENGINE.
1. Remove engine cowling, right and left nose
caps and front engine baffles.
2. Remove nuts attaching each riser assembly
to the cylinders on the left bank. It may be necessary to remove clamp from riser assembly between
number 2 and 4 cylinders.
3. Work left exhaust stack assembly down from
cylinders and out of right exhaust stack assembly at
front of engine .
Change 1
lOA-21
ENGINE -DRIVEN
FUEL PUMP
FUEL METERING
UNIT
'.
EXISTING FUEL PUMP
OUTLET HOSE
•
I
I.
r----L.
TEST HOSE
PRESSURE
INDICATOR
TEE
TEST HOSE
NOTE
WHEN ADJUSTING UNMETERED FUEL PRESSUHE, TEST EQUIPMENT MAY
BE "TEED" INTO THE ENGINE-DRIVEN FUEL PUMP OUTLET HOSE AT THE
FUEL PUMP OR AT THE FUEL METERING UNIT
•
Figure IOA-3. Fuel InjecUon Pump Adjustment Test Harness (Turbocharged Engine)
SHOP NOTES:
•
IOA-22
ChaDge 1
•
r
SLIP JOINT
WASTE-GATE (BY-PASS VALVE)
INSTALLED HERE
NOTE
Minimum gap between tailpipe
and cowling to be .75 inch.
TAILPIPE ATTACHES TO
TURBINE OUTLET
SLIP JOINT
FRONT ENGINE
EXHAUST SYSTEM
- - - - - - - - - - - - - TORQUE ALL EXHAUST CLAMP
NUTS TO 25 - 30 LB-IN.
•
NOTE
Minimum gap between tailpipe
and cowling to be .62 inch.
TURBINE INLET
ATTACHES HERE
Tighten nut until cotter pin
will just fit into hole.
WASTE-GATE (BY-PASS VALVE)
INSTALLED HERE
•
TAILPIPE ATTACHES TO
TURBINE OUTLET
REAR ENGINE
EXHAUST SYSTEM
Figure 10A-4. Exhaust Systems
10A-23
4. Remove bolts, washers and nuts attaChing
waste-gate exhaust tube to waste-gate.
5. Loosen clamp at turbine exhaust outlet and
work tailpipe from turbine and waste-gate exhaust
outlet.
6. If installed, disconnect exhaust gas temperature wires.
7. Loosen clamps and disconnect compressor
air outlet duct at compressor.
8. Loosen clamps and disconnect compressor
air inlet duct at compressor and induction air box.
9. Remove nut and spacer attaching turbocharger
mounting bracket to crankcase and remove bolts attaching bracket to the engine rear mounting leg.
10. Remove bolts, washers and nuts attaching
waste-gate and actuator to exhaust stack assembly.
Tie waste-gate and actuator up to provide clearance
for removal of exhaust stack.
11. Remove bolts, washers and nuts attaChing
turbocharger to exhaust stack assembly. Support
turbocharger as the bolts are removed and lower
turbocharger into cowling.
12. Remove bolts, nuts and clamps attaching
right exhaust stack assembly to riser pipes on right
side of engine. Work exhaust stack from engine.
13. Remove nuts attaching riser pipes to cylinders at right side of engine. Remove riser pipes
and gaskets. Riser pipes should be marked so that
they may be installed on the same cylinder.
b. REAR ENGINE.
1. Remove engine cowling and tail caps as required for access.
2. Remove cotter pins, nuts, washers, bolts
and springs at lower end of collector assembly on
the right side.
3. Remove exhaust gas temperature probe if
installed.
4. Remove two nuts attaChing exhaust pipe
riser to each cylinder on right bank of cylinders and
remove collector assembly and gaskets. The risers
may be removed from collector by removing clamps
attaching riser pipes to collector assembly.
5. Remove clamp attaching right exhaust pipe
to turbine inlet assembly.
6. Remove clamp attaChing waste-gate exhaust
outlet to tailpipe and loosen clamp attaching tailpipe
to turbine exhaust outlet and work tailpipe from turbine.
7. Remove clamp attaching waste-gate exhaust
inlet to turbine inlet assembly.
8. Remove cotter pins, nuts, waShers, bolts
and springs at lower end of collector assembly at left
bank of cylinders.
9. Remove two nuts attaching exhaust pipe riser
to each cylinder on left bank of cylinders and remove
collector assembly and gaskets. The risers may be
removed from collector by removing clamps attaching riser pipes to collector assembly.
10. Remove bolts, washers and nuts attaching
turbine inlet assembly to the turbine.
11. Work turbine inlet assembly from aircraft.
10A-72. INSPECTION. Since exhaust systems of
this type are subject to burning, cracking and general
deterioration from alternate thermal stresses and
IOA-24
vibrations, inspection is important and should be
accomplished every 100 hours of operation. Also, a
thorough inspection of the engine exhaust system
should be made to detect cracks causing leaks which
could result in loss of optimum turbocharger efficiency and engine power. To inspect the engine exhaust system, proceed as follows:
a. Remove engine cowling as required so that ALL
surfaces of the exhaust assemblies can be visually
inspected.
•
NOTE
Especially check the areas adjacent to welds
and slip joints. Look for gas deposits in surrounding areas, indicating that exhaust gases
are escaping through a crack or hole or
around the slip joints.
b. After visual inspection, an air leak check should
be made on the exhaust system as follows:
1. Attach the pre ssure side of an industrial
vacuum cleaner to the tailpipe opening, using a rubber plug to effect a seal as required.
NOTE
The inside of the vacuum cleaner hose should
be free of any contamination that might be
blown into the engine exhaust system.
2. With vacuum cleaner operating, all joints
in the exhaust system may be checked manually by
feel, or by using a soap and water solution and
watching for bubbles. All joints should be free of
air leaks with the exception of the waste gate bearing which will show some bubbling. Also, some
bubbles will appear at the joint of the turbocharger
turbine and compressor bearing housing.
c. Where a surface is not acceSSible for a visual
inspection, or for a more positive test, the following
procedure is recommended:
1. Remove exhaust stack assemblies.
2. Use rubber expansion plugs to seal openings.
3. Using a manometer of gage, apply approximately 1-1/2 psi (3 inches of mercury) air pressure
while each stack assembly is submerged in water.
Any leaks will appear as bubbles and can be really
detected.
4. It is recommended that exhaust stacks found
defective be replaced before the next flight.
d. After installation of exhaust system components,
perform the inspection in step ''btl of this paragraph
to ascertain there are no leaks at the joints of the
system.
•
10A-73. INSTALLATION.
NOTE
Since it is important that the complete exhaust system, including the turbocharger
and waste gate, be installed without preloading any section of the exhaust stack
assembly, follow the sequence outlined for
•
•
•
•
installation on the applicable engine. Use
new gaskets at each end of the waste-gate
and between turbocharger and exhaust
stack assembly. The gasket between each
riser pipe and cylinder may be re-uc;ed as
long as it is not damaged in any way.
a. FRONT ENGINE.
1. Place all sections of the exhaust stack assemblies in position with all clamps loose.
2. Torque nuts attaching riser pipes to the
cylinders to 200-210 Ib-in.
3. Manually check that slip-joints do not bind.
4. Raise turbocharger mounting bracket to
. crankcase. Install and tighten bolts attaching mounting bracket to engine rear mounting leg. Torque
crankcase thru-bolt to 490-510 Ib-in and install
"Palnut". Torque bracket to mounting leg bolts to
160-190Ib-in.
5. Install bolts, washers and nuts attaching
turbocharger to right exhaust stack assembly.
Tighten securely.
6. Install bolts, washers and nuts attaching wastegate to right exhaust stack assembly. Tighten securely.
7. Install tailpipe and tighten clamp securing tailpipe to turbine.
8. Install bolts, washers and nuts attaChing
waste-gate exhaust outlet tube to waste-gate. Tighten
securely.
9. Tighten clamps attaching stack assemblies to
the riser pipes •
10. Install or connect exhaust gas temperature
probe if installed.
11. Connect turbocharger compressor outlet air
duct and tighten clamps.
12. Install turbocharger compressor inlet air
duct. Tighten clamps securely.
13. Be sure all parts are secure and safetied as
required, then perform step ''b'' of paragraph lOA-72
to check for any air leaks. Correct any leaks found
as result of check.
14. Install any parts removed for access, then
install nose caps and cowling.
b. REAR ENGINE.
1. Place all sections of the exhaust stack assemblies in position with all clamps loose.
2. Install bolts, washers and nuts attaching turbine inlet assembly to the turbine outlet. Tighten
securely.
3. Install bolts, waShers and nuts attaChing
waste-gate inlet and outlet tubes to waste-gate.
4. Install tailpipe and tighten clamp securing
tailpipe to turbine. Tighten bolts attaching wastegate exhaust and inlet tubes to tailpipe and turbine
inlet assembly.
5. Torque nuts attaching riser pipes to the
cylinders to 200 to 210 lb-in.
6. Install bolts, springs, washers and nuts at
collector and tube on each side of engine. Tighten
nut until cotter pin will just fit in hole of bolt and
install cotter pin.
7. Tighten clamps attaching collector to risers
on both sides of the engine.
8. Be sure all parts are secure and safetied
as required, then perform step fib" of paragraph
10A-72 to check for any air leaks. Correct any leaks
found as result of check.
9. Install any parts removed for access, then
install tailcaps and cowling.
10A-74. TURBOCHARGER.
10A-75. DESCRIPTION. The turbocharger is an exhaust gas-driven compressor, or air pump, which
provides high velocity air to the engine intake manifold. The turbocharger is comprised of a turbine
wheel, compressor wheel, turbine housing and compressor housing. The turbine Wheel, compressor
wheel and interconnecting drive shaft comprise one
complete assembly and are the only moVing parts in
the turbocharger. Turbocharger bearings are lubricated with filtered oil supplied from the engine lubricating oil system. Engine exhaust gas enters the
turbine housing to drive the turbine wheel. The turbine wheel, in turn, drives the compressor wheel,
producing a high velocity of air entering the engine
inducation intake manifold. Exhaust gas is then
dumped overboard through the exhaust outlet of the
turbine housing and exhaust tailpipe. Air is drawn
into the compressor housing through the induction
air filter and is forced out of the compressor housing
through a tangential outlet to the intake manifold.'
The degree of turbocharging is varied by means of
a waste-gate valve, which varies the amount of exhaust gas allowed to bypass the turbine wheel. The
waste-gate is controlled by the air-oil operated
waste-gate controller.
10A-76. REMOVAL AND INSTALLATION.
a. FRONT ENGINE.
1. (Refer to figure 10A-6.) Remove engine
cowling as required for access to turbocharger components.
2. Remove right cowl flap by disconnecting the
push-pull rod at the cowl flap and at the torque tube.
Remove screws securing cowl flap hinge to lower
fuselage and remove cowl flap.
3. Loosen clamp (13) at turbine exhaust outlet
and work tailpipe (16) from turbine and waste-gate
outlets.
4. (Refer to figure 10A-8, sheet 1.) Remove
the four bolts attaching waste-gate (15) and actuator
(13) to the exhaust stack assembly. Tie waste-gate
and actuator up to provide clearance for removal of
the turbocharger.
5. (Refer to figure 10A-6.) Loosen clamps and
remove compressor air outlet duct and compressor
air inlet duct from compressor (10).
6. Disconnect oil inlet check valve (8) at adapter (15) and oil scavenger line (2) at adapter (11).
Plug or cap disconnected lines and fittings.
7. Remove hardware attaChing front mounting
bracket (7) to engine.
8. Remove bolts, washers and nuts attaching
turbine (14) to the exhaust stack assembly (5). Support turbocharger assembly as the bolts are removed
and work assembly from aircraft through the cowl
flap opening.
10A-25
INDUCTION
SYSTEM
NOTE
Front engine system is shown. Rear engine
is identical except for routing of exhaust
stacks, oil lines and lines which apply turbocharger pressure to fuel discharge nozzles,
fuel pump, controllers and fuel flow gage.
•
TO FUEL
DISCHARGE
NOZZLES
VARIABLE
CONTROLLER
(REGULATES OIL
THRU WASTE-GATE
ACTUATOR
RAM AIR
AUTOMATIC
ALTERNATE
AIR DOOR
(HELD CLOSED
BY MAGNET)
TO ENGINE-
~:r:N FUEL
/-
FUEL FROM FUEL
MANIFOLD VALVE
COMPRESSOR
FUEL FLOW
GAGE
~
PRESSURE RELIEF VALVE
(AIRCRAFT SERIALS 337-0979,
F33700001 AND ALL AIRCRAFT
MODIFIED IN ACCORDANCE
WITH SK337 -14)
~.'--------'~~~
r----:-~~...,
WASTE-GATE
._---t-t-----I ACTUATOR
(S PRING- LOADED
OPEN)
RATE-QF-CHANGE CONTROLLER
(REGULATES TIME REQUIRED FOR
l'vIANIFOLD PRESSURE CHANGE)
(THRU AIRCRAFT SERIAL 337 -0978
AND ALL AIRCRAFT NOT MODIFIED
IN ACCORDANCE WITH SK337-14)
.----------OVERBOARD
THRU
TAILPIPE
OVERBOARD
DRAIN
LEGEND:
E'Mffl
ENGINE OIL
·"f~"
COMPRESSED AIR
+
EXHAUST AIR
¢
RAM AIR
WASTE GATE
CONTROLS VOLUME
THRU TURBINE
OIL RETURN
TO ENGINE
MECHANICAL LINKAGE
Figure lOA-5. Turbocharger System Schematic
10A-26
•
•
•
•
NOTE
Wben installing a NEW turbocharger on the
FRONT engine, it will be necessary to remove the six bolts attaching the exhaust turbine housing to the center section of the unit.
Rotate the exhaust portion of housing 180 degrees. This is done so that oil outlet (11) in
the center section will be pointed downward
When installed in the aircraft. Also loosen
the band on the compressor portion of turbocharger and rotate housing so that the outlet
can be connected to the duct going to the
throttle body. Refer to figure 10A-6 for
torque value of bolts attaching center section
to exhaust turbine housing.
9. Reverse the preceding steps for reinstallation. Install neW gaskets between turbocharger and
exhaust manifold and between waste-gate and exhaust
manifold. Reinstall all safety wire where removed.
Refer to figure lOA-6 for torque values of the attaching bolts.
b. REAR ENGINE.
1. (Refer to figure 10A-7.) Remove engine
cowling and tail caps as required for access to the
turbocharger components.
2. Remove clamp attaChing waste-gate exhaust
to tailpipe (12).
3. Loosen clamp at turbine exhaust outlet and
work tailpipe (12) from turbine (9) and waste-gate
exhaust .
4. Loosen clamps at compressor and slide
coupler securing discharge tube (18) to compressor
(7) upward.
5. Loosen clamps (5) and disconnect air inlet
duct (6) from compressor (7).
6. Disconnect oil inlet line (17) from check
valve (19) and scavenger line (8) from check valve
(10). Plug or cap disconnected lines and fittings.
7. Remove bolts, washers and nuts attaching
turbine to exhaust assembly.
8. Remove bolts (14) securing turbocharger
to support assembly (16) and bolts securing turbocharger to bracket (13). Support turbocharger as
these bolts are removed.
9. Work turbocharger from aircraft.
10. Reverse the preceding steps for reinstallation. Install a new gasket between the exhaust stack
and turbine and reinstall all safety wire where removed.
10A-77. CONTROLLERS AND WASTE-GATE
ACTUATOR.
•
10A-78. FUNCTIONS. The waste-gate actuator,
variable controller and rate-of-change controller
use engine oil for supply power to control the turbocharger. The waste-gate is used to control engine
exhaust flow through the turbine and regulate its
speed. Since the exhaust energy Is the force that
drives the turbocharger unit, the output of the compressor is controlled by bleeding or dumping of
excess exhaust energy as needed. The waste-gate
actuator, which is physically connected to the waste-
gate by mechanical linkage, controls the position of
the waste-gate butterfly valve. The butterfly valve
position is controlled by the variable controller.
Engine oil is supplied to the waste-gate actuator
through the capillary tube Where the pressure of oil
determines the position of the valve. The variable
controller cam arm is connected to the throWe linkage and controls the output of the compressor discharge pressure. Thru aircraft serial 337-0978,
the rate-of-change controller regulates time required
for manifold pressure change. Beginning with aircraft serials 337-0979 AND F33700001, the rate-ofchange controller is deleted and a pressure relief
valve is installed in the induction air inlet. This
pressure relief valve bleeds off compressor discharge pressure that is in excess of maximum manifold pressure. This helps control overboosting of the
engine in cold temperatures.
lOA-79. OPERATION. The waste-gate actuator is
spring-loaded to position the waste-gate butterfly
valve to the open position when there is no oil pressure. When the engine starts, oil pressure is admitted into the actuator through the capillary tube.
This automatically fills the cylinder and lines leading to the controller metering valves. At engine idle
the turbocharger runs slowly with low compressor
output and the metering valve in the variable controller remains open. As the throttle is advanced, the
cam of the variable controller is rotated, calling for
an increase in compressor output by closing Its
metering valve, resulting in a build up of oil pressure in the waste-gate actuator cylinder. The oil
pressure overcomes the spring force in the actuator
cylinder, causing the waste-gate butterfly valve to
close, which causes the engine exhaust gases to pass
through the turbine. As the engine increases in
power and speed, the increase in temperature and
pressure of the exhaust gas causes the turbocharger
to spin faster, raising the compressor and outlet pressure. The variable controller senses the compressor
outlet pressure on an aneroid bellows. As engine output increases, the proper absolute pressure is reached and the force on the aneroid bellows opens the
metering valve. This lowers the oil pressure in the
waste-gate actuator cylinder. When this oil pressure
is lowered suffiCiently, the spring force causes the
waste-gate butterfly valve to partially open. A portion of the engine exhaust gases then bypasses the
turbocharger turbine, thus preventing further increase
of turbocharger speed and holding the compressor discharge pressure to the preselected manifold pressure
as determined by the throttle control. The waste-gate
will modulate toward the closed position or open position to maintain the selected manifold pressure during
changes of altitude, airspeed or engine speed. Above
20,000 feet the variable controller will continue to
maintain 32 inches of mercury manifold pressure at
full throttle. It is necessary to reduce manifold pressure with the throttle to follow the maximum pressure
versus altitude schedule shown on instrument panel
placard. The rate-of- change controller is connected
in parallel with the variable controller and regulates
the rate of change in compressor discharge pressure
and prevents engine overboost. This controller sen-
10A-27
•
ATTACHES TO ENGINE
REAR MOUNTING FOOT
(TORQUE BOLT TO 160190 LB-IN)
WASTE-GATE
ATTACHES TO
EXHAUST STACK
\
•
COMPRESSOR DISCHARGE
TUBE ATTACHED HERE
AND TO THROTTLE BODY
1. Line (To Oil Pressure Gage)
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
Line (Oil Return From Turbine)
Check Valve
Rear Mounting Bracket
Exhaust Stack Assembly
Line (Pressure Oil To Turbine)
Front Mounting Bracket
Check Valve
Stud
Compressor
Adapter (Oil Out)
Cover
Clamp
Turbine
Adapter (Oil In)
Tailpipe
ATTACHES TO ENGINE
THRU-BOLT (TORQUE
TO 490-510 LB-IN)
TORQUE ATTACHING BOLTS
TO 100 + 10 - 00 LB-IN
• Safety wire these items.
* 00046
Beginning with air·craft serials 33701362 and F337and on and all service parts, a new turbine
oil inlet adapter and check valve is used. When an
oil inlet adapter or check valve is to be replaced
on aircraft prior to the above serials, it will be
necessary to install BOTH the check valve and oil
inlet adapter.
Figure lOA-6. Front Engine Turbocharger Installation
IOA-28
•
•
NOTE
Turbine support assembly (16) is
attached with engine thru-bolts.
Torque thru-bolts to 490-510 lb-in.
2---1
• Remove bolts (14) when removing
turbocharger.
F
~,
e
• Safety wire these items.
21
•
Induction Air Filter
Air Inlet Duct
Airbox Assembly
Alternate Air Door
Clamp
Flexible Duct
Compressor
Lin~ (Oil Return From Turbine)
Turbine Assembly (Bolts to Ex.'1aust Assembly)
Check Valve (Oil Out)
Adapter (Oil Out)
Tailpipe
Turbine Bracket
Bolt
Turbine Brace
Turbine Support Assembly
Line (Oil Pressure to Turbine)
Compressor Discharge Tube (Connects
to Throttle Body)
19. Check Valve (Oil In)
20. Adap~er (Oil In)
21. Engine Mount
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
•
I"igure lOA -7. Rear
E!'.g:~e
10
'*
•
Beginning with aircraft serials 33701362
and F33700046 and on and all service
parts, a new turbine oil inlet adapter and
check valve is used. When an oil inlet
adapter or check valve is to be replaced
on aircraft prior to the above serials, it
will be necessary to install BOTH the
check valve and oil inlet adapter •
Turbocharge!" and Induction Ai!" Ir!stallation
10A-29
ses the compressor outlet pressure in the upper
chamber through an internal capillary tube in the
lower chamber. When compressor discharge pressure increases more rapidly than approximately 6.5
inches of mercury per second, a pressure diHerentia1.
eXists between the upper and lower chambers of the
diaphragm. As the pressure in the upper chamber
becomes greater than that of the lower chamber, the
diaphragm between the upper and lower chamber is
forced downward, causing the metering valve to open
and lower the oil pressure in the waste-gate actuator
power cylinder, causing the waste-gate butterfly
valve to open. This prevents the turbocharger compressor discharge pressure from increasing at too
fast a rate and prevents overboosting the engine. The
pressure relief valve is installed in the induction air
duct ahead of the throttle control unit. This valve
senses the compressor outlet pressure and bleeds off
the pressure that is in excess of maximum manifold
pressures.
[~Au!(o~l
The turbocharged engines are equipped with
controller systems which automatically control the engine power within prescribed manifold pressure limits. Although these automatic controller systems are very reliable
and eliminate the need for manual control
through constant throttle manipulation, they
are not infallible. For instance, such things
SHOP NOTES:
lOA-30
as rapid throttle manipulation (especially with
cold oil), momentary waste-gate sticking,
air in the oil system of the controller, etc.,
can cause overboosting. Consequently, it is
still necessary that the pilot observe and be
prepared to control manifold pressure, particularly during take-off and power changes
in flight. Slight overboosting of manifold
pressure beyond established maximums,
which is occasionally experienced during
initial take-off roll or during a change to
full throttle operation in flight, Is not considered detrimental to the engines as long
as it Is momentary. Momentary overboost
is generally in the area of 2 to 3 inches and
can usually be controlled by slower throttle
movement. No corrective action is required
where momentary overboosting corrects
itself and Is followed by normal engine operation. However, if overboosting of this
nature perSists, or if the amount of overboost goes as high as 6 inches, the controllers
and pressure relief valve should be checked
for necessary adjustment or replacement.
overboost exceeding 6 inches beyond established maximums is excessive and can result
in engine damage. It is recommended that
overboosting of this nature be reported to
your Cessna Dealer, who will be glad to
determine what, if any, corrective action
needs to be taken.
•
•
•
lOA-80. TROUBLE SHOOTING.
TROUBLE
REMEDY
Controller not getting enough oil
pressure to close the waste-gate.
Check oil pump outlet pressure,
oil filter and external lines for
obstructions. Clean lines and
replace if defective. Replace oil
filter.
Controllers out of adjustment
or defective.
Refer to paragraph lOA-82.
Replace controllers if defective.
Defective actuator.
Refer to paragraph lOA-82. Replace actuator if defective.
Leak in exhaust system.
Check for cracks and other obvious defects. Replace defective
components. Tighten clamps and
connections.
Leak in intake system.
Check for cracks and loose connections. Replace defective
components. Tighten all clamps
and connections.
Defective controllers.
Refer to paragraph lOA-S2.
Replace if not adjustable.
Waste-gate actuator linkage
binding.
Refer to paragraph lOA-82.
Waste-gate actuator leaking
oil.
Replace actuator.
Turbocharger overspeeding from
defective or improperly adjusted
controllers.
Refer to paragraph lOA-82.
Replace if defective.
Waste-gate sticking closed.
Correct cause of sticking. Refer
to paragraph lOA-82. Replace
defective parts.
Controller drain line (oil return
to engine sump) obstructed.
Clean line. Replace if defective.
ENGINE POWER INCREASES
SLOWL Y OR SEVERE MANIFOLD PRESSURE FLUCTUATIONS WHEN THROTTLE
ADVANCED RAPIDLY.
Waste-gate operation is
sluggish.
Refer to paragraph lOA-82.
Replace if defective. Correct
cause of sluggish operation.
ENGINE POWER INCREASES
RAPIDLY AND MANlFOLD
PRESSURE OVERBOOST
WHEN THROTTLE ADVANCED
RAPIDLY.
Rate-of- change controller/
overboost control valve out
of adjustment or defective.
Refer to paragraph lOA-82.
Replace if defective.
Waste-gate operation is
sluggish.
Refer to paragraph lOA-82.
Replace if defec'tive. Correct
cause of sluggish operation .
UNABLE TO GET RATED
POWER BECAUSE MANIFOLD PRESSURE IS LOW.
ENGINE SURGES OR
SMOKES.
•
TURBOCHARGER NOISY
WITH PLENTY OF POWER.
•
PROBABLE CAUSE
10.\-31
lOA-BOo TROUBLE SHOOTING (Cont).
TROUBLE
FUEL PRESSURE DECREASES
DURING CLIMB, WHILEMANIFOLD PRESSURE REMAINS
CONSTANT.
PROBABLE CAUSE
REMEDY
Compressor discharge pressure
line to fuel pump aneroid
restricted.
Check and clean out restrictions.
Leaking or otherwise defective
engine-driven fuel pump
aneroid.
Replace engine-driven
fuel pump.
Leak in intake system.
Check for cracks and other
obvious defects. Tighten all
hose clamps and fittings.
Replace defective components.
Leak in compressor discharge
pressure line to controller.
Check for cracks and other
obvious defects. Tighten all
clamps and fittings. Replace
defective components.
Controller seal leaking.
Replace controller.
Waste-gate actuator leaking oil.
Replace actuator.
Waste-gate butterfly - closed
gap is excessive.
Refer to paragraph IOA-82.
Intake air filter obstructed.
Service air filter. Refer to
Section 2 for serviCing inStructions.
FUEL FLOW DOES NOT DECREASE AS MANIFOLD
PRESSURE DECREASES AT
Defective engine- driven fuel
pump aneroid mechanism.
Replace engine-driven fuel
pump.
PART-THROTTLEC~CAL
Obstruction or leak in compressor
discharge pressure line to enginedriven fuel pump.
Check for leaks or obstruction.
Clean out lines and tighten
all connections.
FUEL FLOW INDICATOR
DOES NOT REGISTER
CHANGE IN POWER
SETTINGS AT HIGH
ALTITUDES.
Moisture freezing in indicator
line.
Disconnect lines, thaw ice and
clean out lines.
SUDDEN POWER DECREASE
ACCOMPANIED BY LOUD
NOISE OR RUSHING AIR.
Intake system air leak from
hose becoming detached.
Check hose condition. Install
hose and hose clamp securely.
MANIFOLD PRESSURE GAGE
INDIC' ATION WILL NOT REMAIN STEADY AT CONSTANT
POWER SETTINGS.
Defective variable controller.
Replace controller.
Waste-gate operation is
sluggish.
Refer to paragraph lOA-82.
Replace if defective. Correct
cause of sluggish operation.
MANIFOLD PRESSURE DECREASES DURING CLIMB
AT ALTITUDES BELOW
NORMAL PART THROTTLE
CRITICAL ALTITUDE, OR
POOR TURBOCHARGER
PERFORMANCE INDICATED
BY CRUISE RPM FOR CLOSED
WASTE-GATE. (Refer to
paragraph lOA-82.)
ALTITUDE.
lOA-32
•
•
•
•
•
•
10A-81. REMOVAL AND INSTALLATION.
a. VARIABLE CONTROLLER. (Refer to figure
10A-8. )
1. Remove engine cowling as required for access.
2. Disconnect and tag oil lines (2 and 9) at controller (1) and plug or cap open lines and fittings.
3. Disconnect compressor outlet pressure sensing line (8) from controller and plug or cap open
line and fitting.
4. Disconnect control rod (7) from controller.
Note position and size of washers and spacers. Do
not disturb control rod length.
5. Remove screws securing controller to bracket
on top of engine.
6. Remove bolts, washers and nuts securing aft
end of controller to bracket on top of engine.
7. Remove controller from engine, reinstall
screws removed in step 5.
8. Reverse the preceding steps for reinstallation. Tighten firward mounting screws to 20-30
lb-in. Adjust controller in accordance with paragraph 10A-82.
9. The rear engine controller may be removed
in a similar manner using figure lOA-8 as a guide.
b. RATE-OF-CHANGE CONTROLLER. (Thru aircraft serail 337-0978.) (Refer to figure 10A-8.)
1. Remove engine cowling as required for access.
2. Disconnect and tag olllines (2 and 9) at controller (10) and plug or cap open lines and fittings.
3. Disconnect compressor outlet pressure sensing line (11) from controller and plug or cap open
line and fitting.
4. Remove controller mounting bolts.
5. Remove controller from engine.
6. Reverse the preceding steps for reinstallation. Adjust controller in accordance with paragraph
10A-82.
7. The rear engine controller may be removed
in a similar manner using figure lOA-8 as a guide.
c. WASTE-GATE AND ACTUATOR. (Refer to
figure 10A-8. )
1. Remove cowling as required for access.
2. Disconnect and tag oil lines (9 and 12) from
actuator (13) and plug or cap open lines and fittings.
On the rear engine remove clamp securing turbocharger oil inlet line to bracket on waste-gate.
3. Remove bolts, washers and nuts attaching
waste-gate and actuator assembly to tailpipe.
4. Loosen clamp attaching tailpipe to turbine exhaust outlet and work tailpipe from aircraft.
5. Remove bolts, washers and nuts attaChing
waste-gate and actuator. assembly to the exhaust
stack.
6. Carefully work assembly from aircraft.
7. Reverse the preceding steps for reinstallation using new gaskets. Adjust waste-gate in accordance With paragraph 10A-82.
8. The rear engine assembly may be removed
in a similar manner using figure lOA-8 as a guide .
10A-82. ADJUSTMENTS.
a. VARlABLE CONTROLLER. (Refer to figure
10A-8. )
1. Place throttle in full OPEN pOSition and
check that throttle arm (6) and controller arm contact their stops at the same time. If not, adjust
control rod (7) until the stops are contacted at the
same time.
2. With engine running and oil temperature at
middle of green arc, slowly open throttle and note
maximum manifold pressure obtainable. Do not exceed 32±. 5 in Hg.
3. Loosen the high manifold pressure adjustment
screw locknut and adjust screw (4) counterclockwise
(CCW) to increase or clockwise (CW) to decrease
manifold pressure. Tighten locknut after adjustment.
NOTE
Approximately one turn of the high setting
screw will change the manifold pressure
reading about one inch Hg.
4. Operate engine as in step 2 to check that adjustment has not caused a radical change in manifold
pressure.
NOTE
When making adjustments on the ground, the
hotter the engine gets, the lower the manifold pressure will be.
5. Flight test the aircraft after each adjustment
to check results until desired results are obtained.
6. The rear engine controller is adjusted in a
similar manner using figure 10A-8 as a guide.
b. RATE-OF-CHANGE CONTROLLER. (Thru aircraft serial 337-0978.) (Refer to figure 10A-9.)
1. Remove controller as outlined in paragraph
10A-81.
2. Remove fitting from drain port of controller.
3. Remove ambient (low pressure) plug from controller.
4. Insert tool (Part No. 5090002-1) into drain
port. Insert small bladed screwdriver into low pressure port. Rotate poppet assembly until screwdriver
blade engages slot provided in bellows assembly boot.
5. Holding bellows assembly boot, rotate poppet
assembly clockwise (CW) to Increase, counterclockwise (CCW) to decrease. Lightly tap the unit after
each adjustment to seat internal parts.
NOTE
When adjusting, rotate in VERY small increments as this is an extremely sensitive adjustment.
6. Reinstall plug and fitting. Reinstall controller as outlined in paragraph 10A-81.
10A-33
2
FRONT ENGINE
I
I
•
7
• Torque to 8 - 10 Ib-in.
*,Safety wire these items.
o BEGINNING WITH AIRCRAFT SERIALS 337-0979,
F33700001 AND ALL AIRCRAFT MODIFIED IN
ACCORDANCE WITH SK337-14, THE RATE-OFCHANGE CONTROLLER (10) IS DELETED. THE
VARIABLE CONTROLLER (1) THEN CONNECTS
DIRECTLY TO THE WASTE-GATE ACTUATOR (13) .
~l
•
13
1&
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
Variable Controller
Line (Oil Return to Engine Sump)
Low Manifold Pressure Adjustment Screw
High Manifold Pressure Adjustment Screw
Rod End
Throttle Control Arm
Control Rod
Line (Connects to Compressor Discharge
Tube at Throttle Body)
Line (Controller to Waste-Gate Actuator)
Rate-of-Change Controller
Line (To Turbocharger Compressor Discharge Tube)
Line (Pressure Oil From Engine Pump)
Waste-Gate Actuator
Overboard Drain Line
Waste-Gate
Tailpipe
Figure lOA-So Controllers and Waste-Gate Installation (Sheet I of 2)
IOA-34
•
REAR ENGINE
_ _ _ _ _- - -
/
11
_179'ai:"S:::!!..
/
100
,
I
•
/
_3*
•
4
/
I
,yo
I
!
8
,-
STA~K"~
~,
14
FROM EXHAUST
(AHEAD OF TURBINE
~ET)
•
TO TAILPIPE
Figure lOA-B. Controllers and Waste-Gate Installation (Sheet 2 of 2)
IOA-35
o RATE-OF-CHANGE
CONTROLLER
o THRU AmCRAFT SERIAL
337 -0978 WHEN NOT MODIFIED IN ACCORDANCE
WITH SK337-l4
•
SCREWDRIVER
INLET PORT
DRAIN PORT
TOOL (PART NO. 5090002-1) IS AVAILABLE FROM
THE CESSNA SERVICE PARTS CENTER
•
Figure IOA-9. Rate-Of-Change Controller Adjustment
7. Fllght test aircraft after each adjustment to
check results until desired results are obtained as
outlined in step 4 of paragraph lOA -83.
c. WASTE-GATE AND ACTUATOR. (Refer to
figure IOA-IO.)
1. Remove waste-gate and actuator in accordance with paragraph lOA-8l.
2. Install a plug in the actuator outlet port and
apply a 50-60 psig air pressure to the inlet port of
the actuator.
3. Check for. 005 to .015 inch gap between
butterfly and waste-gate body as illustrated.
4. If adjustment is required, release the air
pressure and remove the pin from the actuator shaft.
5. Hold clevis end and turn shaft clockwise (CW)
to increase gap or counterclockwise (CCW) to decrease gap of butterfly. Install pin through clevis
and shaft, securing pin with washer and cotter pin.
6. After adjusting closed position of waste-gate
and with zero pressure in cylinder, check butterfly
for a clearance of .700 to .800 inch in the full-open
position as illustrated.
10A-36
7. If adjustment is required, loosen locknut and
turn screw clockWise to decrease or counterclockwise
to increase opening of butterfly.
8. Recheck butterfly in the closed position to ascertain that gap tolerance has been maintained.
NOTE
To assure correct spring loads, actuate
butterfly with air pressure. Actuator
and butterfly should move freely. Actuator should start to move at 15 ± 2
psig and fully extend at 35 ± 2 psig.
Two to four psig hysteresiS is normal
due to friction of O-rings against cylinder wall.
9. Remove air pressure line and plug from
actuator.
10. Install waste gate and actuator in accordance
with paragraph 10A-8t.
•
•
*
•
•
.005 INCH MINIMUM
• 015 INCH MAXIMUM
• 700 INCH MINIMUM
• 800 INCH MAXIMUM
*_.i1-
LOCKNUT
INLET
CLEVIS END
Figure 10A-IO. Waste-Gate Adjustments
•
20,000 FT
PRESSURE
ALTITUDE
----- ,...••..•.•........
20, 000 FT
TO
15,000 FT
PRESSURE
ALTITUDE
.;._
~
NOTE
2,000 FT
ABOVE
GROUND
Figure lOA-H.
Circled numbers refer to corresponding
flight checks required in paragraph lOA-83.
Operational Flight Check
10A-37
10A-83. CONTROLLER AND TURBOCHARGER OPERATIONAL FLIGHT CHECK. The following procedure
details the method of checking the operation of the variable reference and rate-of-change controller and a
performance check of the turbocharger.
CD a.TAKEOFF
- VARIABLE REFERENCE CONTROLLER CHECK.
Cowl Flaps - Open
b.
c.
d.
e.
f.
Airspeed - 100 MPH lAS
Middle of green arc
Engine Speed - 2800 ± 25 RPM
Fuel Flow - 21 to 22 GPH (Full Rich Mixture)
Full Throttle M.P. - Variable reference controller should maintain 32 ± .5 in Hg (stabilized).
au Temperature -
•
Climb 2000 feet after takeoff to be sure manifold pressure has stabilized. It is normal on the first takeoff of
the day for full throttle manifold pressure to decrease 1/2 to 1.0 inch of mercury within one minute after the
initial application of full power. Refer to paragraph 10A-82 for variable reference controller adjustment.
CD a.CLIMB
- VARIABLE REFERENCE CONTROLLER AND TURBOCHARGER PERFORMANCE CHECK.
Cowl Flaps - Open
b.
c.
d.
e.
f.
Engine Speed - 120 MPH lAS
Engine Speed - 2600 RPM
Fuel Flow - Adjust mixture for 14.5 GPH
Part-Throttle M. P. - 28 in. Hg.
Climb to 20,000 feet· Check manifold pressure stability during climb.
Once the climb power setting is established after take-off, the controller should maintain a steady manifold
pressure up to 24,000 feet which is the maximum operating altitude for 28 inches Hg.
CD a.CRUlSE
- TURBOCHARGER PERFORMANCE CHECK.
Cowl flaps - closed
b.
c.
d.
e.
f.
g.
Airspeed - Level flight
Pressure Altitude - 20,000 feet
Engine Speed - 2800 RPM
Part - Throttle M. P. - 28 in. Hg.
Fuel Flow - Lean to 15 GPH
Propeller Control (1) Slowly decrease engine speed to 2200 RPM or until manifold pressure starts to drop, indicating
waste-gate is closed.
NOTE
•
If the waste-gate closes at engine speeds lower than 2200 RPM, the turbocharger performance is normal. If the waste-gate closes at engine speeds higher than 2200 RPM,
refer to the trouble shooting chart in paragraph IOA-80.
(2) Note outside air temperature and RPM as manifold pressure starts to drop, which should be in
accordance with the follOwing chart.
(3) After noting temperature and RPM, increase engine speed 50 RPM to stabilize manifold pressure,
with the waste-gate modulating exhaust flow to control compressor output.
Outside Air Temperature
RPM where M. P. Starts to Decrease
40 0 F Above Standard
Standard Temperature
40°F Below Standard
2400
2300
2200
CD a.DESCENT
- RATE-OF-CHANGE CONTROLLER.
Cowl Flaps - Closed
b.
c.
d.
e.
f.
Airspeed - 100 MPH lAS
Pressure Altitude - 20,000 to 15,000 feet
Propeller - High RPM
Mixture - Full Rich
Throttle - Idle, until M. P. stabilizes
(1) Rapidly advance throttle to full power.
(2) Note time required for M. P. to increase from 20 to 30 in. Hg.
Time required should be 1.8 to 2.9 seconds (3.5 to 5.5 in. Hg. per second). Refer to paragraph IOA-82 for
rate-of-change controller adjustment.
IOA-38
•
•
TO PROPELLER
ENGINE AND
ACCESSORY
BEARINGS
3
(
,/
)
\oJ
•
7
8
"* THRU Am CRAFT SERIAL
9
337 -0978 WHEN NOT MODIFIED IN ACCORDANCE
WITH SK337 -14
18
CODE:
............ PRESSURE OIL
;#&-, RETURN OIL AND
SUCTION OIL
•
1.
2.
3.
4.
5.
6.
7.
8.
Pressure Gage
Propeller Governor
Oil Sump Drain Plug
Filler Cap
Dipstick
Oil Temperature Gage
Oil Coo~er
Check Valve
9. Turbocharger
10. Check Valve
11. Waste-Gate (Bypass
Valve) Actuator
12. Variable Controller
13. Fuel Line From Oil
Dilution Solenoid
14. Thermostat
15.
16.
17.
18.
19.
20.
21.
Temperature Transmitter
Pressure Relief Valve
Oil Pump
Filter Bypass Valve
External Filter
Scavenger Pump
Rate-Of-Change
Controller
Figure 10A-12. Oil System Schematic
10A-39
10A-S4. ENGINE OIL SYSTEM.
10A-S5. DESCRIPTION. The engine 011 system is
the same as described in paragr~ph 10-75 except the
external 011 filter is standard equipment on turbocharged engines. Also, the engine oil is used to control the waste-gate and lubricate the turbocharger
bearings. Engine 011 is returned from the turbocharger sump by a scavenger pump, which is an integral part of the engine. Refer to figure 10A-12 for a
schematic diagram of the oil system.
10A-S6. TROUBLE SHOOTING. Refer to paragraph
10-76.
10A-S7. FULL-FLOW OIL FILTER. Refer to paragraph 10-77.
10A-IOl. MAGNETO CHECK. Refer to paragraph
10-91.
10A-102. MAINTENANCE. Refer to paragraph
10-92.
10A-103. TACHOMETER BREAKER POINT ADJUSTMENT. Refer to paragraph 10-93.
10A-l04. SPARK PLUGS. Refer to paragraph 10-94.
10A-105. STARTING SYSTEM. Refer to paragraph
10-95.
10A-106. DESCRIPTION. Refer to paragraph
10-96.
10A-SS. DESCRIPTION. Refer to paragraph 10-7S.
10A-I07. TROUBLE SHOOTING. Refer to paragraph
10-97.
10A-S9. ELEMENT REMOVAL AND INSTALLATION.
Refer to paragraph 10-79.
10A-108 . STARTER MOTOR. Refer to paragraph
10-98.
10A-90. ADAPTER REMOVAL. Refer to paragraph
10-SO.
to paragraph 10-99.
10A-9l. ADAPTER DISASSEMBLY, INSPECTION
AND REASSEMBLY. Refer to paragraph 10-81.
10A-UO. PRIMARY MAINTENANCE. Refer to
paragraph 10-100.
10A-92. ADAPTER INSTALLATION. Refer to
paragraph 10-82.
10A-U1. EXTREME WEATHER MAINTENANCE.
Refer to paragraph 10-101.
10A-93. IGNITION SYSTEM. Refer to paragraph
10-S3.
10A-U2. COLD WEATHER. Refer to paragraph
10-102.
10A-94. DESCRIPTION. Refer to paragraph 10-84.
10A-U3. HOT WEATHER. Refer to paragraph
10-103.
10A-95. TROUBLE SHOOTING. Refer to paragraph
10-85.
10A-109. REMOVAL AND INSTALLATION. Refer
•
10A-U4. SEACOAST AND HUMID AREAS. Refer
to paragraph 10-104.
10A-96. MAGNETOS. Refer to paragraph 10-86.
10A-97. DESCRIPTION. Refer to paragraph 10-87.
10A-U5. DUSTY AREAS. Refer to paragraph
10-105.
10A-98. REMOVAL AND INSTALLATION. Refer to
paragraph 10-88.
10A-U6. GROUND SERVICE RECEPTACLE.
Refer to paragraph 10-106.
10A-99. INTERNAL TIMING. Refer to paragraph
10-89.
.
10A-117. HAND CRANKING. Refer to paragraph
10-107.
10A-IOO. MAGNETQ-TQ-ENGINE TIMING. Refer
to paragraph 10-90.
•
10A-40
SECTION 11
•
•
FUEL SYSTEM
TABLE OF CONTENTS
Page
FUEL SYSTEM
11-1
Description (Thru 337F) .
11-1
Description (337G)
11-2
Description (Thru T337F) .
11-2
Description (337G Long-Range)
11-2
Precautions
.
...
11-2
Trouble Shooting.
•
11-3
Main Fuel Tanks .
11-10
11-10
Description (thru 337F)
Description (337G)
11-10
11-10
Removal (Thru 337F) .
Installation (Thru 337F) .
11-10
11-10
Removal of Outboard Tanks (337G)
Installation
.
11-10
Removal of Inboard Tank
11-10
Installation .
11-10
Fuel Quantity Transmitters
11-10
Fuel Quantity Sending Units
11-10
Fuel Sump Tanks .
11-15
Description
11-15
Removal.
11-15
Installation
11-15
Fuel Vents.
11-15
Description (Except T337 thru 337F) 11-15
Removal
11-15
Checking
11-15
Description (T337-Series and 337G). 11-15
Removal
11-18
Checking
11-18
Fuel Line Manifolds
11-18
Removal and Installation
11-18
Removal and Installation of Fuel Lines
11-18
Auxiliary Fuel Pumps
11-18
Description
11-18
Removal and Installation
11-18
Pump Circuit Adjustment
11-19
Fuel Selector Valves
11-19
Description (Thru 337F) .
11-19
Description (337G)
11-20
Removal and Installation (Thru
33701316 and F33700024)
11-20
11-1. FUEL SYSTEM (EXCEPT T337-Series).
•
11-2. DESCRIPTION. (Thru 337F) The main fuel
supply is contained in fuel tanks in the Wings. Two
interconnected metal tanks are located in each wing,
just outboard of the booms. Fuel flows from the
main tanks to sump tanks, one in each Wing. Fuel
from the main tanks will drain completely into the
sump tanks. From the sump tanks, fuel flows
through a bypass in electrical fuel pumps to both
(front engine and rear engine) fuel selector valves
in the wing roots. By using the selector valves,
fuel can be selected from either the right or left
main tank for either engine. This arrangement
permits both engines to operate from either tank.
Either electric pump will sustain both engines in the
highly unlikely circumstance that the two enginedriven pumps and one electric pump should become
inoperative. Fuel flows from each selector valve,
Removal and Installation (33701317
and F33700025 thru 33701462 and
F33700055)
.
.
Removal and Installation (Beginning
with 33701463 and F33700056)
Removal and Installation of Selector
Gearbox (Thru 33701316 and
F33700024)
.
.
.
.
Removal and Installation of Selector
Gearbox (33701317 and F33700025
thru 33701462 and F33700055)
Removal and Installation of Selector
Gearbox (Beginning with 33701463
and F33700056) . .
Installing New Selector Valve Handle
Rigging (Thru 33701316 and F33700024).
Rigging (33701317 and F33700025
thru 33701462 and F33700055)
Rigging (Beginning With 33701463
and F33700056)
Fuel Strainers .
.
Description
Removal and Installation
Disassembly
Primer System
Description
Removal and Installation
Auxiliary Fuel System
Removal of Aux. Tank
Installation of Aux. Tank
Fuel Quantity Transmitter or
Sending Unit .
Removal and Installation
Fuel Vent
Removal
Checking
Installation
Drain Valve
Fuel Line
Fuel Quantity Indication
11-22
11-22
11-22
11-22
11-22
11-2.4
11-24
11-24
11-26
11-26
11-26
11-26
11-26
11-26
11-26
11-26
11-26
11-26
11-28
11-28
11-31
11-31
11-31
11-31
11-21
11-31
11-31
11-31
through a fuel strainer at each engine, into the
engine-driven fuel pump of each engine. Each iuel
strainer contains a remotely controlled drain valve.
Each engine primer receives its fuel supply from
the front strainer. The optional oil dilution fuel line
connects at each fuel strainer. Fuel vapor return
lines return vapor and unused fuel from the front
engine-driven fuel pump to the left fuel tanks, and
from the rear engine-driven fuel pump to the right
fuel tanks, regardless of selector valve position.
Auxiliary fuel tanks are available as optional equipment, and are installed, one in each wing, between
the cabin and the boom. The left auxiliary fuel tank
feeds direcUy into the front engine selector valve
only, and the right auxiliary tank feeds directly into
the rear engine selector valve only. On models prior
to the 337C-Series, fuel level is indicated on electrically-operated fuel quantity gages. Each gage is
operated by two interconnected fuel quantity transmitters, one in each main tank. Each auxiliary tank has
11-1
a fuel quantity transmitter which operates its individOn Models 337C thru 337F, only two fuel
quantity indicators are provided in the instrument
cluster on the panel. The indicators are for left and
right fuel tanks and indicate both main and auxiliary
fuel tank levels.
ual gage.
11-3. FUEL SYSTEM (EXCEPT T337-SERIES).
11-4. DESCRIPTION. (Beginning with 337G) This
fuel system is basiCally similar to the system de
scribed in paragraph 1-2, except that the fuel selector valves in the wing roots have been changed from
a five-port valve to a three-port valve, and the fuel
selector control device above the cabin console has
been changed from a four-position gearbox configuration to a three-position caromed bellcrank design.
The vapor and fuel return lines return unused fuel
and vapor to the sump tanks.
11-5. FUEL SYSTEM (T337-SERIES).
11-6. DESCRIPTION. (Thru 337F). The main fuel
supply is contained in fuel tanks in the Wings. Two
interconnected metal tanks are located in each wing,
just outboard of the booms. Fuel flows from the
main tanks to sump tanks, one in each wing. Fuel
fl'Qm the main tanks will drain completely into the
sump tanks. From the sump tanks, fuel flows directly to both (front engine and rear engine) fuel selector
valves in the Wing roots. By using the selector valves,
fuel can be selected from either the right or left main
tank for either engine. This arrangement permits
both engines to operate from either tank. Fuel flows
from each selector valve into its fuel line manifold,
through a fuel strainer at each engine, through a bypass in an electric fuel pump for each engine, into
the engine-driven fuel pump of each engine. Each
fuel strainer contains a remotely controlled drain
valve. Each engine primer receives its fuel supply
from the front strainer. The optional oil dilution
fuel line connects at each fuel strainer. The front
engine electric fuel pump will sustain the front engine
if its engine-driven fuel pump should become inoperative, and the rear engine electric fuel pump will
sustain the rear engine if its engine-driven fuel pump
should become inoperative. Fuel vapor return lines
return vapor and unused fuel from the front enginedriven fuel pump into the front fuel line manifold,
where the fuel is recirculated and the vapor is returned to the left fuel tanks. Fuel vapor return lines
return vapor and unused fuel from the rear enginedriven fuel pump into the rear fuel line manifold,
where the fuel is recirculated and the vapor is returned to the right fuel tanks. This arrangement is
always true, regardless of selector valve position.
11-7. FUEL SYSTEM (LONG-RANGE 337G).
11-8. DESCRIPTION. (Beginning with 337G). The
fuel supply is contained in three metal fuel tanks
located in each wing. Two interconnected tanks are
located just outboard of the booms. An additional
fuel tank is installed in each wing, between the cabin
and the boom. This tank is interconnected with the
outboard tanks. A fuel quantity sendin~ unit is located in all three fuel tanks in each Wing. The units
11-2
transmit fuel tank quantities to indicators located
in a cluster on the instrument panel. Fuel flows
from the main tanks to a sump tank, located in each
boom, immediately beneath the wing. From the
sump tanks, fuel flows direcUy to both (front engine
and rear engine) fuel selector valves, located in each
wing root area. These valves are mechanically connected to selector handles located in the pilot's overhead console in the cabin. By using the selector
valves, fuel can be routed from either the right or
left main tanks for either engine. This arrangement
permits both engines to operate from either set of
tanks. Fuel flows from each selector valve through
each fuel strainer and a bypass in each auxiliary fuel
pump, into an engine-driven fuel pump for each engine. Each fuel strainer contains a remotely ~on­
trolled drain valve. Each engine primer receives its
fuel supply from the front strainer. The optional oil
dilution fuel line connects at each fuel strainer. The
front engine electric fuel pump will sustain the front
engine if its engine-driven fuel pump should become
inoperative and the rear engine fuel pump will sustain
the rear engine if its engine-driven fuel pump should
become inoperative. Fuel vapor return tines return
vapor and unused fuel from the front and rear enginedriven pumps into the respective fuel line tees 10cated between the sump tank and inboard fuel tank in
each Wing. This arrangement is always true, regardless of selector valve position.
11-9. PRECAUTIONS.
NOTE
There are certain general precautions and
rules concerning the fuel system which
should be observed when performing the
operations and procedures in this Section.
These are as follows:
•
•
a. During all fueling, defueling, tank purging, and
tank repairing or disassembly, ground the airplane
to a suitable ground stake.
b. Residual fuel draining from lines and hoses constitutes a fire hazard. Use caution to prevent the
accumulation of fuel when lines or hoses are disconnected.
NOTE
Throughout the aircraft fuel system, from the
fuel tanks to the engine-driven fuel pump, use
RAS-4 (Snap-On Tools Corp., Kenosha, Wisconsin), MIL-T-5544 (Thread Compound,
Antiseize, Graphite-Petrolatum) or equivalent,
as a thread lubricant or to seal a leaking connection. Apply sparingly to male threads only,
omitting the first two threads. Always ensure
that a compound, the residue from a previously
used compound, or any other foreign material
cannot enter the system. Throughout the fuel
injection system, from the engine-driven fuel
pump through the discharge nozzles, use only
a fuel soluble lubricant, such as engine lubricating oil, on fitting threads. Do not use any
other form of thread compound on the injection
system.
•
•
11-10. TROUBLE SHOOTING.
NOTE
Use this trouble shooting chart in conjunction with
the engine trouble shooting charts in Sections 10
or lOA.
TROUBLE
NO FUEL FLOW TO
ENGINE-DRIVEN PUMP.
•
FUEL STARVATION AFTER
STARTING.
NO FUEL FLOW WHEN
ELECTRIC PUMPS
OPERATED.
PROBABLE CAUSE
REMEDY
Selector valve not turned on.
Turn selector valve on.
Fuel tanks empty.
Service with proper grade and
amount of fuel.
Fuel line disconnected or broken.
Connect or repair fuel lines.
Defective selector valve.
Repair or replace selector valve.
Selector valve not rigged properly.
Re-rig selector valve.
Sump tank strainer or auxiliary
strainer plugged.
Clean screens and flush out tanks.
Plugged fuel strainer.
Clean strainer and screen.
Defective bypass valve in electric
fuel pump.
Repair pump. Replace bypass
valve.
Fuel line plugged.
Disconnect lines as necessary
to locate obstructions, then
clean.
Partial fuel flow from the preceding
causes.
Use the preceding remedies.
Malfunction of engine-driven fuel
pump or fuel injection system.
Refer to Sections 10 or lOA.
Fuel vents plugged.
See paragraphs 11-26 and 11-66.
Water in fuel.
Drain fuel tank sumps, fuel lines
and fuel strainer.
Defective auxiliary pump switch.
Replace defective switch.
Open or defective circuit breaker.
Reset. Replace if defective.
Loose connections or open circuit.
Tighten connections; repairor replace
wiring.
Defective electriC fuel pump.
Replace defective pump.
Defective engine-driven fuel pump bypass Refer to Sections 10 or lOA.
or defective fuel injection system.
•
NO FUE L QUANTITY
INDICA TlON.
Fuel tanks empty.
Service with proper grade and
amount of fuel.
Defective indicator, transmitter, sending Refer to Section 14.
unit or electrical circuit.
11-3
•
··fUEL flOW
INDICA Tal
THIOTTLE
~IITUIE
CONTIOl
CMECIC
~----~II:;::..::II"11
AND
AIR
~ ____ _
VALYE.
fUEL QUANTITY
YENT
WITH
CHICK
VALVE
VALVE
DIAIN
...... LvE
DRAIN
VALVE
WITH
IY.'ASS VALVE
CHECl
VAL VE
AUX.
fUEl PUMP
SWITCH
•
AUX. fUEL PUMP
SWlfeH
DIAIN KNOI
FUEL 'UM'
AND
_ _ _ _~ MllTUlt CONTlOl
II I....._ , . .
~ '"Ionu
____
CODE
lllllrnnrnnWIIWIlDUDmrnUDll1
U
n
...............
fUEL flOM LEFT MAIN '''NIS TO
flONT ENGINE AND CIOSSfUO TO
lEA. lNGINE. fUEL flOM UfT
··fUEl flOW
INDICA'OI
AUJULlAIY TANK TO nONT
ENGINE ONLY.
~~~~!m 'UEL nOM IIG"T MAIN JANKS TO
1'''1 f:HGINE AND CIOSS'EfD TO
flONT lNGIHE. 'UU flOM liGHT
Ii
11:,,1 ENGINE fUEL NOZZLES
AUXIUAU TANK TO IU,' ENGINE
ONLY.
.... :~.:~ . :.:I :~~p :~OD y;~~U~:f~~~~ ~60~:I~lL
r:::
fUEl TANI{S
-
-
-
-
~
NON-TURBOCHARGED 337-SERIES
(THRU 337B)
• nONT ENGINE fUEL 5(:UC1'OI HANDLE
SHOWN IN "LEfT MAIN TANK." 'OSITION.
RofAR !NGINE fUEL SElECTOR HANDLE
SHOWN IN ··RIGHT AUllliARY TANI(
'OSITION
--A SINGLE DUAL. INDICATING fUEl flOw
INDICATOR IS USED fOI 10TH ENGINES.
MI!CHANICAl LINKAGE
llECUI(Al CONNECflON
Figure 11-1. Fuel System Schematic (Sheet 1 of 6)
11-4
•
•
'ION1 INGIHI 'Ull HOUUS
5"AI"'1I DIAl"" IiNOI
\lJ
~'
~~
'ut. PUMP
awne"
AUI.
AUK"
YI;J""~.:';U~:fL.:~~'1S 'UII
VALYI
~::::::::~===:::::::::;,
Qu ....." ' ' '
INOIC.TOI
CHICI
VAI"'I
VINI
vi HI
•
CODE
IIIIIIIIID
fUll flO,. UfT ..... IN
no".'
''''N''~
10
INGlHf "''''0 (tQU'UO
10 ItA. IN(;.fIooII 'Utl '10_ LUI
AUIILlA.' 'AHI 10
INGrHI
no ... ,
aNI'
lUll '10_ llOM1' .... ~ 'ANI:5 10
ltoU lNG-IN! AHO 'IOSSJlUD TO
.tONI INCo""I. 'UII .. 0. hOItl
.t .... INGINf 'UIl NOlllfS
&U'II./A" U.N. 10 ...... tNO'"
0"'1'
.u.•
lUll AND "A'OI IIIUI ...
'1.1'1
"0.
AND .'Uull UNIIS
10 'UII .M .. "NrfOID
c=:::::J
TURBOCHARGED 337-SERIES
(THRU 337B)
• '10"" INC...... 'un slLterol ""NOLI
SHOw ... IN un .... IN 'ANIl POS.'IO'"
1.,,1 ING''''I lUll tUlero. "ANCtil
,"OWN ''''
IIGHI AUIIU...... , .......
P051110""
ytNI IIHIS
'''Gil DUA' "'Of("'ING 'ull 'LOw
""01(""01 I! UUD 'Ot 10'''1 INC, .... U
• • .It.
""ICMAHICA, II ....... GI
•
~
(IICI'(AI CONH(CIION
Figure 11-1.
Fuel SY"3tem Sr.h~matic (Sheet 2 of 6)
11-5
•
""tI
'UIL Qu ..
INDiCAfOI
001(_
VAl"l
Aua.
'ulL PUM'
,WIf(H
NON-TURBOCHARGED 337 -SERIES
(MODEL 337C THRU 337D)
•
CODE
IIDDDD
.... 11 MOM un .....IN ' .. Nr.S TO
.tONI ING'" "NO CIOUfllD
10 IIA' I~I. fUll '10M
AU'''1A1Y
' 'MI:
un
10 NONI
INGINION'"
'UII .IOM 'GMt ......... ~
-,
TURBOCHARGED 337-SERlES
(MODEL 337C THRU 337D)
to
...... INCMII AND aosII'tlO TO
.ItOMI INGlNI, fUU .IOM IIOWf
AU...... n .'AN" 10 lUliNG'"
flUl' ANO .....PO. ItlVlN .~
NIL .".., AND . . .",1, UNtf1,
ro 'Ull 1M .......H)ID
VU" lINIS
NOTE
• "0'" INGINI fUll IIUC'1OI .........
...0-... ... "un ...... '''NI ~1I'tON.
I'A' .........,,' HUCIO. "ANOlI
1N0"",, ... "ltGtIT .. u........ ,
Remainder of the fuel system scbematic is
the same as that shown on sheets 1 and 2.
'Os_no ...
' 'NI.'
'un
• . . . PtOU: OUA,.NDtC"'MG
"OW
NDlCAJOe ISo USID PO' '0'" I..o ... n
_,(MANICA, "MIAG!
••• laO_'''o
____
IUCIlICA' CONNI(110N
Figure 11-1. Fuel System Schematic (Sheet 3 of 6)
11-6
U'
.110
.oe.AID
•
•
MODEL 337E THRU 337F
AUXILIARY FUEL
INDICATOR LIGHT (OFF)
s::
•
.*
FUEL QUANTITY INDICATOR
(SHOWING LEFT MAIN FUEL)
•
RIGHT
FUEL TANK
RIGHT MAIN
~~l~~~~it§FUEL
TANKS
FUEL SENSOR (TYPICAL)
FUEL SUMP
(TYPICAL)
AUX. FUEL
PUMP
(TYPICAL)
_*-H~--a:t;-
FUEL SELECTOR HANDLES
VALVE
•
THRU 1970-SERIES
:
BEGINNING WITH 1971-SERIES
NOTE
Remainder of the fuel system schematic is
the same as that shown on sheets 1 and 2.
*
•
A SINGLE CONTROL MONITOR ELECTRICALLY
SERVES BOTH RIGHT AND LEFT MAIN TANKS.
A SEPARATE MONITOR SERVES BOTH AUXILIARY TANKS .
Figure 11-1. Fuel System Schematic (Sheet 4 of 6)
11-7
•
FRONT ENGINE FUEL NOnUS
INDICATOR-
THROTTUO- - - - FUEL
INDICATOR
~
INDICATOR
."'''~~.~~*MONITOR
t~O
TANKS
VENT WITH
CHECK VALVE
•
- - - :,:f'MIXTURE CONTROL
CONTROL
CODE
mm
~
~
O.
.
o
..
FUEL FROM
UFT TANKS TO FRONT
ENGINE AND CROSSFUD
TO REAR ENGINE.
REAR ENGINE FUEL NOZZLES
SERVES
MAIN TANKS.
_ A SINGLE DUAL-INDICATING FUEL FLOW
INDICATOR IS USED FOR BOTH ENGINES.
FUEL FROM
RIGHT TANKS TO REAl
ENGINE AND CR055FEED
TO FRONT ENGINE.
FUn AND VAPOR
RnURN fROM Fun PUMP
AND MIXTURE UNITS TO
MAIN AND SUMP TANKS.
BEGINNING WITH 1973 MODEL 337G
STANDARD-RANGE TANK INSTALLATION
VENT
- - - - MECHANICAL
LINKAGE
~ __ ELECTRICAL
-,.-- CONNECTION
-
FLOW
Figure 11-1. Fuel System Schematic (Sheet 5 of 6)
11-8
•
•
*A SINGLE CONTROL MONITOR
ELECTRICALLY SERVES BOTH
RIGHT AND LEFT MAIN TANKS.
FRONT ENGINE FUEL NOZZLES
FUEL
1::JD~11 JlimIlnl DIS TRIB U TOR
FUEL FLOW
INDICATOR ():Ul2tl~
THROTTLE Vooi::!~-"v=:FUEL PUMP
AND MIXTURE UNIT
FUEL OUANTITY INDICATOR
------t
FRONT
FUEL STRAINER
FUEL O\JANTITY
INDICATOR
ONITOR*
FILLER
CAP
CHECK
VALVE
VE T
LECTOR
LVE
QUICK
•
DRAIN
CODE
IDIFUEL FROM
LH MAIN TANKS
TO FRONT ENGINE
_FUEL FROM
RH MAIN TANKS
TO REAR ENGINE
~
FRONT
L~
FUEL
SELECTOR
HANDLES
c=>
VENT LINES
-'
SELECTOR VALVE
!m)"'O-;+-REAR FUEL STRAINER
CHECK VALVE
AUX. FUEL PUMP
WITH BY-PASS VALVE
DDCROSSFEED
.. FROM LH
MAIN TANKS TO
RH FUEL SELECTOR
c::J
-
REAR
_CROSSFEED
FROM RH
MAIN TANKS TO
- ___ LH FUEL SELECTOR
r>.'] FUEL AND VAPOR
RETURN FROM
FUEL PUMPS
AND MIXTURE
UNITS
-
MIXTURE
---~ CONTROL
FUEL AND AIR
THROTTLE UNIT
FUEL
DISTRIBUTOR
----- MECHANICAL
LINKAGE
•
7. ELECTRICAL
REAR ENGINE FUEL NOZZLES
FUEL FLOW
INDICATOR
BEGINNING WITH 1973 MODEL 337G
LONG-RANGE INSTALLATION
CONNECTION
Figure 11-1. Fuel System Schematic (Sheet 6 of 6)
11-9
11-11. MAIN FUEL TANKS.
11-12. DESCRIPTION (Tbru 337F). The main tank in
each wing consists of two interconnected metal
tanks. The tanks are connected by two hoses, one
at the forward bottom edge and one at the aft bottom
edge. The outboard tank in each wing has a vent
line which extends outboard from the fuel tank to the
wing tip and then aft to the wing trailing edge. A
check valve is installed in each vent line at the wing
tip. The inboard tank is vented to the outboard tank
by an interconnecting hose at the top forward outboard corner. Both tanks are serviced through a
single filler neck in the outboard tank. The inboard tank has two lines, one at the forward inboard corner and one at the aft inboard corner,
through which fuel flows from both tanks to the fuel
sump tank. Fuel flow from the tanks to the sump
tank is complete, eliminating unusable fuel in the
tanks and the need for drains in the tanks. All fuel
draining is done through the quick-drain valve or
strainer in the bottom of the sump tank. Through
the Model 337D-Series, each fuel tank has a fuel
transmitter mounted in the top of the tank. These
transmitters are wired in parallel in each wing to
give only one reading for each set of tanks. Beginning
with the Model 337E-Serles, each tank has a sending
unit installed on a bracket inside the tank. These
Rending units are wired in parallel in each wing.
11-13. DESCRIPTION. (Beginning with 337G). The
standard-range aircraft is the same as described in
the preceding paragraph. The long-range aircraft is
equipped with three interconnected metal tanks in each
wing. The two outboard tanks are connected by three
hoses, one at the aft bottom edge, and one each at the
top and bottom forward edge. The inboard tank is connected to the center tank by a large and a small hose
and a metalllne. A sump tank, located in each boom
between the center and inboard tanks, is fed fuel by
two lines from the center and inboard tanks. The
sump tanks are connected to the fuel selector valves
located in each wing root. Fuel flow from the main
tanks to the sump is complete, eliminating need for
drains in the tanks. All fuel draining is accomplished
through the quick-drain valve or strainer in the bottom of the sump tanks. The outboard tank in each
wing has a vent line which extends outboard from the
fuel tank to the wing tip. A check valve is installed
in each vent line at the wing tip. Each main tank has
a sending unit installed on a bracket inside the tank.
11-14. REMOVAL OF MAIN TANKS. (Tbru 337F).
Each tank is retained by two metal. straps and may
be removed as an individual unit.
a. Place fuel selector valves in OFF position.
b. Remove sump tank access cover and drain all
fuel from tanks by removing quick-drain valve.
Str.ainer-can be removed to expedite fuel draining.
NOTE
Support outer wing panel and tail boom with
cradle supports, before removing fuel tank
covers, to prevent wing and boom deflection.
c. Remove tank cover from top of wing by remOVing
11-10
screws around outer edge of cover and around filler
opening. After screws are removed, the forward
edge of the cover must be pulled aft from under the
leading edge skin. Retain gaskets between filler neck
and top wing cover.
d. Remove bolts from retaining straps securing
tank to be'removed.
e. Disconnect electric wire from fuel quantity transmitter or sending unit at each tank to be removed.
f. Remove two access plates from bottom of wing
between fuel tanks to gain access to two lower interconnect hoses. Remove hose clamps and lower hoses.
Remove clamps and upper interconnect hose through
top of Wing.
g. If outboard tank is being removed, disconnect
vent line at ta~, and lift tank from wing.
h. If inboard tank is being removed, disconnect fuel
lines from inboard side of tank and lift tank from
wing.
•
11-15. INSTALLATION. (Thru 337F). Installation
of the main fuel tanks may be accomplished by reversing the steps of paragraph 11-14. A cradle to support the outer wing panel should be provided to pre-
vent wing deflection. Wing deflection can cause misalignment of holes in wing and fuel tank caver, making
installation of the cover extremely difficult. When installing fuel tank cover, make sure that forward edge
of cover is under wing leading edge skin. Be sure
that gaskets are placed between scupper and fuel tank
cover. A maximum of three gaskets may be used to
maintain wing contour and prevent canning of the cover.
11-16. REMOVAL OF OUTBOARD TANKS. (Beginning with 337G) •. Each tank is retained by two metal
straps and may be removed as an individual unit.
NOTE
•
Remove outboard tanks tn accordance with
procedures outlined in paragraph 11-14.
11-17 • INSTALLATION (Beginning with 337G). Install outboard tanks in accordance with procedures
outlined tn paragraph 11-15.
11-18. REMOVAL OF INBOARD TANK (Beginning
with 337G). Removal of either inboard fuel tank
is accomplished through the top of the wing.
NOTE
Remove inboard tanks in accordance with
procedures outlined in paragraph 11-62.
11-19. INSTALLATION (Beginning with 337G). Install inboard fuel tanks in accordance with procedures
outlined in paragraph 11-63.
11-20. FUEL QUANTITY TRANSMITTERS. (Thru
3370). Fuel quantity transmitters are installed in
the top of fuel tanks. A complete description, along
with procedures for removal, installation and adjustment are contained in Section 15.
11-21. FUEL QUANTITY SENDING UNITS (Beginning with 337E). A fuel quantity sending unit is
•
•
EXCEPT T337-SERIES
THRU 337F
Detail
A-
5
Detail
•
B
13
SEE FIGURE 11-4
FOR VENT VALVE
INSTALLATION _ _...I
_BEGINNING WITH 1971 337-SERIES
•
1.
2.
3.
4.
5.
Auxiliary Fuel Pump
Fuel Selector Valve
Fuel Strainer
Strainer Drain Control
Rear Priming System
6.
O-Ring
7.
8.
9.
10.
11.
12.
13.
Quick-Drain Valve
Sump Tank Strainer
Outboard Fuel Tank
Vent Line
Interconnect Hose
Inboard Fuel Tank
Sump Tank
14.
15.
16.
17.
18.
19.
20.
Auxiliary Fuel Tank
Check Valve
Left Selector Control
Selector Gear Box
Right Selector Control
Engine Primers
Front Priming System
Figure 11-2. Fuel System (Sheet 1 of 4)
11-11
•
T337 -SERIES
THRU 337F
Detail
A·
11
Detail
B
15
•
SEE FIGURE 11-4
FOR DETAIL OF
INSTALLA TION _ _--.I
-BEGINNING WITH 1971 T337-SERIES
1.
2.
3.
4.
5.
6.
7.
Outboard Fuel Tank
Interconnect Hose
Inboard Fuel Tank
Sump Tank
Auxiliary Fuel Tank
Fuel Selector Valve
Strainer Drain Control
B.
9.
10.
11.
12.
13.
14.
Auxiliary Fuel Pump
Check Valve
Fuel Strainer
Rear Priming System
O-Ring
Quick-Drain Valve
Sump Tank Strainer
Figure 11-2. Fuel System (Sheet 2 of 4)
11-12
15.
16.
17.
lB.
19.
20.
21.
Vent Line
Fuel Line Manifold
Front Priming System
Engine Primers
Left Selector Control
Selector Gear Box
Right Selector Control
•
•
"
.......:..............
~
/
7
......
...............
~~~:~~...:~~...: ......:..
9
.. .'.'.'
4
'
'.
18
14
•
13
1&
17
BEGINNING WITH 1973 MODEL 337G
STANDARD FUEL TANK INSTALLATION
1.
2.
3.
4.
5.
6.
•
Selector Gearbox
Right Selector Control
Check Valve
RH Auxiliary Fuel Pump
Right Fuel Sump Tank
RB Inboard Fuel Tank
7.
8.
9.
10.
11.
12.
RH Fuel Selector Valve
Fuel Strainer
LH Fuel Selector Valve
Left Fuel Sump Tank
LH Inboard Fuel Tank
LH Outboard Fuel Tank
13.
14.
15.
16.
17.
18.
Vent Line
LH Auxiliary Fuel Pump
Check Valve
Left Selector Control
Strainer Drain Control
Fue 1 Strainer
Figure 11-2. Fuel System (Sheet 3 of 4)
11-13
•
•
19
BEGINNING WITH 1973 MODEL 337G
LONG-RANGE FUEL TANK INSTALLATION
1. Selector Gearbox
2. Right Selector Control
3. Check Valve
4. RH Inboard Tank
5. RH Center Tank
6. RH Outboard Tank
7. Vent Line
8.
9.
10.
11.
12.
13.
14.
15.
RH Sump Tank
Selector Valve
Rear Auxiliary Fuel Pump
Strainer Drain Control
Rear Fuel Strainer
LH Inboard Tank
LH Sump Tank
LH Center Tank
16.
17.
18.
19.
20.
21.
22.
Figure 11-2. Fuel System (Sheet 4 of 4)
11-14
LH Outboard Tank
Front Fuel Strainer
Strainer Drain Line
Pump Drain Line
Front Auxiliary Fuel Pump
Strainer Drain Control
Right Selector Control
•
•
•
located in each tank. A complete description, along
with procedures for removal, installation' and calibration are contained in Section 15.
11-22. FUEL SUMP TANKS.
11-23. DESCRIPTION. A fuel sump tank is installed
in the forward part of the boom in each wing. Each
sump tank has a qulck-drain valve and strainer installed in the bottom of the tank. The quick-drain
valve is used for draining water or sediment which
may have collected in main tanks or sump tanks. The
qulck-drain valve may be removed to drain fuel from
the main tanks. The strainer can be removed to expedite fuel draining.
11-24. REMOVAL.
a. Place fuel selector valves in OFF position.
b. Support outer wing panels and tail boom with
cradle supports before removing tank covers, to
prevent wing and boom deflection.
c. Remove access cover beneath sump tank in boom
and remove inboard fuel tank cover from top of wing
between boom and cabin.
d. Completely drain all fuel from main and sump
tanks by removing quick-drain valve in sump tank.
Strainer can be removed to expedite fuel draining.
e. Remove inboard tank as outlined in paragraph
11-18. This is necessary for access to fuel line
connections on top of sump tank.
f. Disconnect all fuel lines at sump tank.
g. Loosen bolts and remove two retaining straps;
remove sump tank.
NOTE
Quick-drain valve or strainer in bottom of
sump tank may be removed for replacement
or cleaning.
11-25. INSTALLATION. Install sump tank by reversing procedures outlined in the preceding paragraph.
11-26. FUEL VENTS.
•
11-27. DESCRIPTION. (Except T-337-Series thru
337F). The main tank vent line extends outboard from
the upper forward corner of the outboard fuel tank to
the wing tip. This vent line contains a swing check
valve to prevent fuel drainage through the vent line,
but still allows the positive pressure from expanding
fuel to escape from the tanks. The inboard tank is
vented to the outboard tank through a hose which connects the two tanks at the forward top adjacent corners. The fuel vent line on each auxiliary tank runs
from the forward outboard corner of the tank to the
flap gap panel at the trailing edge of the wing. The
vent line on some aircraft contains a reStrictor,
shown in figure 11-3. The main fuel tank vent outlet at the trailing edge of the Wing and the auxiliary
fuel tank vent outlet should be checked dally for evidence of foreign matter. Check all fittings and
clamps for tightness and all tubes or lines for clearance to prevent chafing against inner wing structure.
11-28. REMOVAL. Figure 11-2 illustrates the various vent lines and components, and may be used as a
guide during removal. Drain fuel from tanks if line
to be removed is below fuel level. Remove wing tips,
access covers, fairings, upholstery and trim as required for access to fittings and clamps along the vent
line routing. When necessary to remove main or
auxiliary fuel tank covers for access, support outer
wing panel and tail boom with cradle supports before
removing the covers, to prevent Wing and boom deflection.
11-29. CHECKING FUEL VENTS. Field experience
has demonstrated that fuel vent lines can become
plugged, with possible fuel starvation of the engine
or collapse of fuel tanks. Also, the bleed hole in
the vent valve assembly could possibly become
plugged, allowing pressure from expanding fuel to
pressurize the tank.
.
NOTE
Remember that a plugged vent line or bleed
hole can cause either fuel starvation or collapse of fuel tanks, or pressurization of
tanks by fuel expansion.
NON-TURBOCHARGED AIRCRAFT:
a. Attach a rubber tube to the end of vent line at
trailing edge of wing tip.
b. Blow into tube to pressurize tank. If air can be
blown into tank, vent line is open.
c. After tank is slightly pressurized, insert end of
tube into a container full of water and watch for continuous stream of bubbles which indicate bleed hole
in valve assembly is open and relieving pressure.
d. Any vents found plugged or restricted shall be
corrected prior to returning airplane to service.
e. Check auxiliary fuel tank by remOving fuel filler
cap and blOwing through vent line with the rubber tube
attached to vent line at flap gap panel. A restrictor
is used in this line instead of a check valve with
bleed hole.
TURBOCHARGED AIRCRAFT:
a. Remove wing tip.
b. Disconnect fuel line manifold vapor return line
from tee and plug the tee.
c. Disconnect auxiliary fuel tank vent line, if installed, at tee and plug the tee.
d. Check the main vent in the Same manner as nonturbocharged aircraft.
e. To check the auxiliary fuel tank vent, disconnect
and plug tees for main vent line and fuel line manifold
vapor return line, then check in the same manner as
the main vent.
f. Any vents found plugged or restricted shall be
corrected prior to returning airplane to service.
g. Reconnect all lines and reinstall wing tip.
11-30. DESCRIPTION (Turbocharged aircraft and
337G). These fuel vent systems are the same as
those described in paragraph 11-27, except that the
optional auxiliary fuel tank vents and the fuel line
manifold vapor return lines are also connected to
the main fuel tank vent lines at the wing tips. The
auxiliary tank vent lines do not contain restrictors.
On the Model 337G, only one vent line extends from
the outboard fuel tanks to the wing tips.
11-15
•
3
4
I
8
11
31
----
::::::;;:1
EXCEPT
MODEL T337
14
28~
%1
22
19
1. Fuel Filler Door
2. Fuel Tank Cap
3. Scupper Drain Hose
4. Scupper Drain Line
5. Block
6. Retaining Strap
7. Restrictor
S. Union
9. Main Fuel Cells
10. Interconnect Vent Hose
11.
12.
13.
14.
15.
16.
17.
IS.
19.
20.
~
17
12
18 ...._ - - ~:!!!!!=======
TO
WING
TIP
Auxiliary Fuel Tank
21. Bracket
Auxiliary Tank Vent Line 22. Clip
Sump Tank
23. Sensor Unit
Positioning Block
24. Access Plate
Interconnect Hose
25. Wire Assembly
Finger Strainer
26. Gasket
Retaining Strap
27. Fuel Quantity Transmitter
Retaining Strap
2S. Ground Strap
Spacer
29. Cork Washer
Barrel Nut
30. Washer
31. Main Tank Vent Line
MODEL T337
Figure 11-3. Fuel Tanks and Sump Tanks Installation (Sheet 1 of 2)
11-16
•
•
•
E
Detail
7
--Detail
C
\
12
LONG-RANGE
TANK SYSTEM
Detail
\
Qii&L,
~
~
A
DetailB
c
,~
4
Detail
F
I I
'I.... /
A
•
STANDARD
TANK SYSTEM
G
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
19
Sump Tank
Inboard Fuel Tank (L-R InatI)
Center Fuel Tank (L-R Instl)
Strainer
O-Ring
Drain Valve
Retaining Strap
Barrel Nut
Spacer
Filler Cap
16
11.
12.
13.
14.
15.
16.
17.
18.
19.
Vent Line
Vent Valve
Gasket
Fuel Sensor
Sta-Strap
Outboard Fuel Tank
Positioning Block
Interconnect Hose
Inboard Fuel Tank
I
Detail
G
BEGINNING WITH 33701463
AND F33700056
Figure 11-3. Fuel Tanks and Sump Tanks Installation (Sheet 2 of 2)
11-17
EXCEPT T337 SERIES
4
j
~====:;~z.Q
T337 SERIES
6
4
•
3
NOTE
Hinge for vent valve must be at top and vent
valve installed with arrow in direction shown.
1. Wing Tip Rib
2. Main Tank Vent Line
3. Vent Valve
4. Vent Line
5. Vapor Return Line
6. Auxiliary Tank Vent Line Tee
7. Main Tank Vent Line Tee
Figure 11-4. Fuel Tank Vent Valve
11-31. REMOVAL. Refer to figure 11-2 and paragraph 11-28 for routing of vent lines and information
regarding vent line removal.
11-32. CHECKING.
moving the covers, to prevent wing and boom deflection. When installing fuel lines, check connections
for fuel leaks before reinstalling parts removed for
access.
11-36. AUXILIARY FUEL PUMPS.
NOTE
Check vents in accordance with procedures
outlined in paragraph 11-29. Eliminate
steps which do not pertain to a certain aircraft.
11-33. FUEL LINE MANIFOLDS (Thru T337F).
The front fuel line manifold is located on the left side
of the cabin, just below and aft of the pilot's window.
The rear fuel line manifold is located on the righthand side of the aft wheel well, just below the horizontal fir~wall.
11-34. REMOVAL AND INSTALLATION (Refer to
figure 11-2.) Turn off fuel selector valves before
disconnecting fuel lines. The rear manifold is accessible and may be removed by disconnecting all
lines attached to it. The left side panel must be removed to gain access to the front manifold. Remove
the manifold by disconnecting all lines attached to it.
When installing manifolds, check connections for fuel
leaks before reinstalling parts removed for access.
11-35. REMOVAL AND INSTALLATION OF FUEL
LINES. The various fuel lines are shown in figure
11-2, which may be used as a guide during removal
and installation. Turn off selector valves, drain
fuel strainers, or drain fuel from tanks as required
for lines being removed. Remove access covers,
fairings, upholstery or other components as necessary for access to fittings and clamps along fuel line
routing. When necessary to remove main or auxiliary fuel tank covers for access, support outer wing
panel and tail boom with cradle supports before re11-18
11-37. DESCRIPTION. On non -turbocharged aircraft,
an electric auxiliary fuel pump is installed in the leading edge of each wing between the boom and the cabin.
On turbocharged aircraft and aircraft with long-range
installations beginning with 1973 Model 337G, the
electric auxiliary fuel pump for the front engine is located in the nose wheel well and the electric auxiliary
fuel pump for the rear engine is located on the upper
right side of the rear cabin bulkhead. The p'lmps are
operated by a split rocker switch arrangement, one for
each pump, located on the switch panel. Tbeyare
powered by the airplane electrical system. The
switch pOSitions are labeled HI, OFF, and LO. The
pumps are used in starting and, in the event of an
engine-driven fuel pump malfunction, SuPply pressure to operate the engine. An integral bYpass and
check valve permits fuel flow through the pump even
when the pump is inoperative, but prevents reverse
flow. A separate overboard pump drain line prevents entry of fuel into the electric motor, in the
event of pump internal leakage.
11-38. REMOVAL AND INSTALLATION.
a. Place selector valves in OFF position.
b. On non-turbocharged atrplanes, drain all fuel
from main tanks on side from which pump is being removed, and remove auxiliary fuel pump access cover
in bottom of leading edge skin.
c. On turbocharged airplanes, open the landing gear
doors for access to the front pump, and remove the
right engine cowling for access to the rear pump after
disconnecting the right cowl flap.
d. Disconnect the two fuel lines and electrical leads.
•
•
•
EXCEPr MODEL T337 SERIES
(Left hand installation shown)
~
I!J
7
1111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111111
FRONT-
MODEL T337-SERIES
THRU 337F
9
11
REAR
•
1. Auxiliary Fuel Pump
2. Pump Bracket
3. Fuel Line
4. Wing Rib
5. Front Wing Spar Web
6. Overboard Drain Line
7. Vapor Return Line
8. L.H. Wheel Well Tunnel Wall
9. Horizontal Firewall
10. Aft Cabin Bulkhead
11. Fuel Hose
3
NOTE
An adjustable resistor is included in each fuel'
pump electrical circuit. Refer to Section 15 .
for adjustment of the resistors.
Figure 11-5. Auxiliary Fuel Pump Installation (Sheet 1 of 3)
e.
and
f.
g.
Remove two bolts from pump retaining straps
remove pump.
Remove pump drain line and fitting.
Reverse the preceding steps to install the pump.
11-39. AUXILIARY FUEL PUMP CmCUIT ADJUSTMENT. Each auxiliary fuel pump is adjusted in
the low output poSition. This adjustment is made by
sliding a tap on a variable resistor in each circuit.
The reSistors are mounted on the left side structure
of the control console as viewed from the pilots seat.
The adjustment may be made by the following procedure.
a. Engines oft, aircraft outdoors.
b. Throttle and mixture control full on.
Pump switcl. in LO pOSition.
d. Adjust resistor for 5 GHP reading on instrument panel fuel flow indicator.
e. Repeat this procedure to adjust other pump.
c:-
•
IWARNING'
Operation of the fuel pumps with the mixture
and throttle controls full on will allow fuel to
overflow and spill on the ground from each
engine, thus causing a dangerous fire hazard.
Starting the engines should not be attempted
for at least five minutes in order to allow
drainage of excess fuel from the engines.
11-40. FUEL SELECTOR VALVES.
11-41. DESCRIPTION (thru 337F-Series).
Fuel selector valves are divided into two baSic parts:
the selector valve, located in the wing at the wing
root, and the selector gearbox and handle, located on
the centerline of the cabin top. Through the 1970 337Series, and F337 -Series, the selector gearbox handle
is connected to the selector valve by a control wire,
routed through a steel casing. Beginning with the
1971 Models, connection is made with a control cable
with adjustable cleViS terminals at each end. Figure
11-6 illustrates installation of the controls. Each
selector valve has four positions: LEFT MAIN. RIGHT
11-19
•
FRONT INSTALLATION
3~,,--
__
1. Fuel Strainer
2.
3.
4.
5.
6.
7.
8.
9.
10.
Clamp Bolt
Tee
Plate
Drain Control
Spacer
Auxiliary Fuel Pump
Bracket
Pump Drain Tube
Strainer Drain Tube
10
BEGINNING WITH 1973 MODEL 337G
LONG-RANGE FUEL TANK INSTALLATION
•
Figure 11-5. Auxiliary Fuel Pump Installation (Sheet 2 of 3)
MAIN. AUXILIARY and OFF. The forward selector
in the gearbox controls the fuel selector valve in the
left wing and fuel flow to the front engine, while the
aft selector in the gearbox controls the fuel selector
in the right Wing and fuel flow to the rear engine.
11-42. DEScmPTlON (337G). Fuel selector
valves are divided into two basic parts: the selector
valve, located in the wing root and the selector gearbox, located on the centerline of the cabin top, above
the pilot. The selector gearbox handles are connected to the wing root valves by control cables With adjustable clevis terminals at each end. Figure 11-9
illustrates the fuel selector gearbox installation, and
figure 11-6 illustrates the wing root valve installation.
The fuel selector gearbox glass assembly has three
posltlons: LEFT, OFF and RIGHT. for each selector
handle. The forward selector handle controls the
fuel selector in the left wing and fuel flow to the
front engine. The aft selector handle controls the
fuel selector in the right wing and fuel flow to the
rear engine.
11-20
11-43. REMOVAL AND INSTALLATION OF FUEL
SELECTOR VALVE. (Thru 33701316 and F33700024:
Refer to figure 11-6.) Remove either fuel selector
valve as follows:
a. Remove sump tank access covers and drain all
fuel from tanks by removing quick-drain valve in
bottom of sump tank. Strainer can be removed to
expedite fuel draining.
b. If auxiliary tanks are installed, completely drain
fuel from tanks by removing quick-drain valve in
bottom of tank.
c. Drain all fuel lines by draining each fuel strainer with the fuel selector valves placed in the various
poSitions, then place selector valves in the OFF
position.
d. Remove forward wing-to-fuselage fairing and
fuel selector valve access door from bottom of wing
at wing root.
e. Disconnect all fuel lines at selector valve.
f. Loosen setscrew holding control wire on arm of
selector valve, and loosen clamps holding control
casing on bracket attached to selector valve.
•
•
REAR INSTALLATION
3
•
BEGINNING WITH 1973 MODEL 337G
LONG-RANGE FUEL TANK INSTALLATION
1. Strainer Drain Tube
2. Fuel Strainer
3. Horizontal Bulkhead
•
4. Strainer Drain Control Knob
5. Drain Control
6. Auxiliary Electric Fuel Pump
7. Pump Bracket
8. Aft Firewall
9. Pump Drain Line
Figure 11-5. Auxiliary Fuel Pump Installati on (Sheet 3 of 3)
11-21
g. Remove two bolts securing selector valve to
wing rib, unbend end of control wire, and pull coritrol forward out of clamps.
h. Control may be removed after disconnecting
inboard end of control as outlined in paragraph 11-46.
i. Install fuel selector valve by reversing preceding
steps, rigging controls in accordance with paragraph
11-50.
11-44. REMOVAL AND INSTALLATION OF FUEL
SELECTOR VALVE. (33701317 and F33700025 thru
33701462 and F33700055). (Refer to figure 11-6.)
Remove either fuel selector valve as follows:
a. Remove sump tank access covers and drain all
fuel from tanks by removing quick-drain valve in
bottom of sump tank. Strainer can be removed to
expedite fuel draining.
b. If auxiliary tanks are installed, completely drain
fuel from tanks.by removing quick-drain valve in
bottom of tank.
c. Drain all fuel lines by draining each fuel strainer with the fuel selector valves placed in the various
poSitions, then place selector valves in the OFF pOSition.
d. Remove forward wing-to-fuselage fairing and
fuel selector valve access cover from bottom of wing
at wing root.
e. Disconnect all fuel lines at selector valve.
f. Remove cotter pin and clevis pin from arm of
selector valve and remove clevis.
g. Remove two bolts securing selector valve to
wing rib, and remove selector valve.
h. Control may be removed after disconnecting inboard end of control.
i. Install fuel selector Valve by reversing preceding steps, rigging controls in accordance with paragraph 5-51.
11-45. REMOVAL AND INSTALLATION OF FUEL
SELECTOR VALVE. (Beginning with 3370146 and
F33700056). (Refer to figure 11-6.) Remove either
fuel selectof valve as follows:
a. Remove sump tank access covers and drain all
fuel from tanks by removing quick-drain valve in
bottom of sump tank. Strainer can be removed to
expedi te draining.
b. Drain all fuel lines by draining each fuel strainer
with the fuel selector valves placed in the various
pOSitions, then place selector valves in the OFF position.
c. Remove forward wing-to-fuselage fairings and
fuel selector valve access cover from bottom of
wing at wing root.
d. Disconnect all fuel lines at selector valve.
e. Remove cotter pin and clevis pin from arm of
selector valve and remove clevis.
r. Remove bolts securing selector valve bracket to
wing rib, and remove selector valve and bracket.
g. Reverse the preceding steps to install the fuel
selector valve. Rig controls as outlined in figure
11-8.
11-46. REMOVAL AND INSTALLATION OF FUEL
SELECTOR GEAR BOX. (Thru 33701316 and
F33700024: Refer to figure 11-6.) Remove fuel
selector gear box as follows:
11-22
a. Remove fuel selector handles from overhead
console.
b. Remove the four screws attaching console to
ceiling.
c. If oxygen system is installed, remove oxygen
selector handle knob.
d. Partially pull console down until oxygen cylinder
pressure gage can be held securely while unscrewing bezel attaching gage to console.
•
Use care in removing oxygen cylinder pressure gage to avoid damaging pressure line.
e. Disconnect console light wires at quick-disconnects and remove the console.
f. Loosen set screws holding control wires in
swivels of selector gear box, loosen clamps holding
casings on gear box, unbend end of control wires,
and pull controls outward out of clamps.
g. Remove the two screws attaching gear box to
bracket on ceiling, and remove gear box.
h. Install the fuel selector gear box by reversing
the preceding steps, rigging the controls in accordance with paragraph II-50.
11-47 . REMOVAL AND INSTALLATION OF FUEL
SELECTOR GEARBOX. (33701317 and F33700025
thru 33701462 and F33700055). (Refer to figure
11-6.) Remove fuel selector gearbox as follows:
a. Remove fuel selector handles from overhead
console.
b. Remove screws attaching console to ceiling.
c. If oxygen system is installed, remove oxygen
selector handle knob.
d. Partially pull console down until oxygen cylinder
pressure gage can be held securely while unscrewing
bezel attaching gage to console.
•
I~AUTION\
Use care in remOving oxygen cylinder pressure gage to avoid damaging pressure line.
e. Disconnect console light wires at quick-disconnects and remove console.
f. Remove cotter pin and clevis pin from shaft in
gear box, and remove clevis.
g. Remove screws attaching gear box to bracket
on ceiling, and remove gear box.
h. Install fuel selector gear box by reversing preceding steps, rigging controls in accordance with
paragraph 5-51.
11-48. REMOVAL AND INSTALLATION OF FUEL
SELECTOR GEARBOX. (Beginning with 33701463
and F33700056). (Refer to figure 11-6.)
a. Remove overhead console in accordance with applicable procedures outlined in Section 3.
b. Remove cotter pins and clevis pins from shafts
in gearbox, and remove clevises.
c. Remove screws attaching gearbox to bracket on
ceiling and remove gearbox.
d. Reverse preceding steps to install fuel selector
gearbox. Rig controls in accordance with applicable
paragraph in this Section.
•
•
~~~})K~.~
\-t\
ARM
9
8
ol
.
1
i ,
io
~
~
-..
DetailC
.
10
1971 THRU 1972 -BERIES
•
9
1...:r--i-l0
&
THRU 1970-SERIES
Detail
•
1.
2.
3.
4.
A
Fuel Selector Valve
Arm
Swivel
Control Wire
Detail
5. Control Casing
6. Bracket
7. Washer
8. Adjustable Clevis
9.
10.
11.
12.
Selector Gear Box
Shaft
Handle
Roll Pin.
B
• Roll pins used on
serials prior to
337-0044
Figure 11-6. Fuel Selector Valve 9.nd Fuel Selector Gearbox Installation (Sheet 1 of 2)
11-23
NOTE
Spray fuel selector valve assembly ports
and mating tube assembly "B'! nuts with
MS-122 FLUOROCARBON (Release agent
dry lubricant) before installing fuel line
to valve assembly. AVOID SPRAYING
INTO FUEL VALVE PORTS.
•
RIGHT-HAND WING SHOWN
BEGINNING WIT-H 1973 MODEL 337G
1.
2.
3.
4.
Fuel Selector Valve
Bracket
Root Rib
Detent Plate
5. Control Arm
6. Control
7. Stop Bolt
Figure 11-6. Fuel Selector Valve and Fuel Selector Gearbox Installation (Sheet 2 of 2)
11-49. INSTALLING NEW FUEL SELECTOR
VALVE HANDLE. On serial number 337-0044 and
on, the handles and selector valve shafts are fabricated so they can only be assembled in the correctly
indexed poSition. Prtor to serial nwnber 337-0044,
the handles and shafts were indexed by drilling a hole
part way through them and installing a roll pin. The
roll pins were not installed in any particular position.
Since a replacement handle for these serial numbers
is not drilled to accommodate the roll pin, it is necessary to modify the handle to match the position of the
roll pin on a particular shaft. Proceed as follows:
a. (FRONT gearbox shaft.) Rotate front shaft
clockwise (looking up at the gearbox) as far as it will
go. With the roll pin removed, place new handle on
the shaft with the handle pointing to the RIGHT side
of the airplane. Drill or file new handle to match
existing roll pin hole, then install handle and roll pin.
b. (REAR gearbox shaft.) Rotate rear shaft clockwise (looking up at the gearbox) as far as it will go.
With the roll pin removed, place new handle on the
shaft with the handle pointing FORWARD. Drill or
file new handle to match existing roll pin hole, then
install handle and roll pin.
selector valve arms.
c. Using one of the selector valve handles to turn
the gearbox shaft, rotate shaft of front gearbox clockwise (looking up at the gearbox) until FFONT gearbox
lever is moved to the LEFT as far as it ~ill go, then
turn back very slightly so the handle points straight
toward the right side of the airplane.
d. Rotate shaft of rear gearbox clockwise (looking
up at the gearbox) until REAR gearbox lever is moved
to the RIGHT as far as it will go, then turn back very
slightly so the handle points straight forward.
e. With gearbox levers in these positions and arms
of selector valves in wings in AFT detents, secure
controls in clamps on gearboxes, insert wires
through holes in swivels, tighten set screws, and
bend remaining wire around swivels.
1. Reinstall parts removed for access, then install
fuel selector handles. The handles and shafts are
indexed so the handles cannot be installed incorrectly.
Prior to serial number 337 -0044, a roll pin indexes
the handles, and on all other serials the parts are
fabricated so they can only be assembled in the correct pOSition.
11-50. RIGGING FUEL SELECTOR VALVES. (Thru
33701316 and F33700024: Refer to figure 11-8.) If
fuel selector valves and fuel selector gear box are
already installed, the following rigging procedure
may be accomplished without draining the fuel system. Remove overhead console and wing access
plates as necessary for access.
a. If controls are being installed, position controls
in clamps on brackets attached to selector valves,
allowing enough control wire to protrude through
holes in selector valve swivels to bend around swivels. Tighten set screws and bend the wire around
the swivels.
b. Position control arms of selector valves in
wings in AFT detents, and maintain this position of
11-51. RIGGING FUEL SELECTOR VALVES.
(33701317 and F33700025 thru 33701462 and F33700055.) (Refer to figure 11-6.) U fuel selector valves
and fuel selector gear box are already installed, the
following rigging procedure may be accomplished without draining the fuel system. Remove overhead console and wing access plates as-necessary for access.
a. If controls are being installed, position controls
in brackets and clamps along routing.
b. Position control arms of selector valves in wing
roots in AFT detents, and maintain this position of
selector valve arms.
c. Using one of the selector valve handles to turn
the gear box shaft, rotate shaft of front gear box
clockwise (looking up at the gear box) until FRONT
gear box lever is moved to the LEFT as far as it
11-24
•
•
•
FUEL
SELECTOR
ASSEMBLY
AIRCRAFT
CENTERLINE
•
FWD
FUEL SELECTOR VALVE (LH WING)
FUEL SELECTOR VALVE (RH WING)
FUEL SELECTOR RIGGING INSTRUCTION SCHEMATIC
(CONTROL ARMS IN SELEC TOR ASSEMBLY AND CONTROL
ARM FOR LEFT-HAND AND RIGHT-HAND WING VALVES
SHOWN IN OFF POSITION)
FUEL SELECTOR RIGGING INSTRUCTIONS
1. Position fuel selector control arms parallel to centerline of aircraft as shown.
2. Position left-hand and right-hand fuel valve control arms in the center detent
with the control arm extended inboard as shown.
3. Attach control cables with the control arms in the OFF position as shown •
•
Figure 11-7. Fuel Selector Valve Rigging (1973 Model 337G)
11-25
will go, then turn back very slightly so the handle
points straight toward the right side of the aircraft.
d. Rotate shaft of rear gear box clockwise (looking
up at the gear box) until REAR gear box lever is
moved to the RIGHT as far as it will go, then turn
back very slightly so the handle points straight forward.
e. With gear box levers in these positions, and
arms of selector valves in wings in AFT detents,
attach control terminal clevises to gear box levers.
NOTE
Terminals may be rotated to align with gear
box levers. Loosen lock nut to rotate terminal. Tighten locknut after terminal is secured to gear box lever.
f. Attach terminal clevises to arms of fuel selector valves in wings.
g. Install clevis pins, cotter pins, and safety wire
controls in brackets as shown in figure 11-8.
h. Reinstall parts removed for access, then install
fuel selector handles. The handles and shafts are
indexed so the handle cannot be installed incorrectly.
11-52. RIGGING FUEL SELECTOR VALVES.
(Beginning with 33701463 and F33700056.) Refer
to figure 11-7 for procedures to be followed during
selector valve rigging.
11-53. FUEL STRAINERS. (Refer to figure 11-8.)
11-54. DESCRIPTION. The fuel strainer for the
front engine on either turbocharged or non-turbocharged aircraft is located in the nose wheel well.
The fuel strainer for the rear engine on non-turbocharged aircraft is located on the upper right-hand
side of the rear cabin bulkhead, and on the firewall
in the aft wheel well on turbocharged aircraft. Each
fuel strainer is equipped with a drain valve control
which affords control of the strainers through access
doors in the upper cowling of both engines. Strainer
screens, gaskets and bowls may be removed and
cleaned with the strainer installed in the aircraft.
11-55. REMOVAL AND INSTALLATION.
a. Turn off fuel selector valves and drain each
strainer.
b. Open landing gear doors to gain access to fuel
strainers mounted in wheel wells.
c. Remove rear engine right cowling after disconnecting cowl flap for access to rear strainer on nonturbocharged airplanes.
d. Disconnect all lines and controls attached to
strainers.
e. Remove strainer mounting bolts.
f. Reverse the preceding steps to install fuel
strainers. Check for fuel leaks.
11-56. DISASSEMBLY. (Refer to figure 11-8.)
a. Turn off applicable fuel selector valve and
drain strainer.
b. Remove safety wire, nut, and washer at bottom
of filter bowl and remove bowl.
c. Carefully unscrew standpipe and remove.
d. Remove filter screen and gasket. Wash filter
11-26
screen and bowl with solvent (Federal Specification
P-S-661, or equivalent) and dry with compressed
air.
e. Using a new gasket between filter screen and
top assembly, install screen and standpipe. Tighten
standpipe only finger tight.
f. Using all new O-rings, install bowl. Note that
step-washer at bottom of bowl is installed so that
step seats against O-ring.
g. Turn on fuel selector valve, close strainer
drain, and check for leaks. Check for proper operation.
h. Safety wire bottom nut to top assembly. Wire
must have ri!?;ht hand wrap, at least 45 degrees.
•
11-57. PRIMER SYSTEM.
11-58. DESCRIPTION. The primer system is a manually operated type. Fuel is supplied by a line from the
front fuel strainer to plunger-type primers. Two primer handles, one for each engine, are located on the
control quadrant. Operating the primers force fuel to
the engines. Fuel is delivered to the propeller end
of each intake manifold. This primes the entire length
of the intake manifold for each bank of cylinders. Primer lines should be replaced when crushed or broken,
and should be properly clamped to prevent vibration
and chafing.
11-59. REMOVAL AND INSTALLATION.
a. Remove console cover.
b. Disconnect primer lines at primer bodies.
c. Remove screws from brackets and remove
each primer body and bracket as a unit.
d. Reverse the preceding steps to install the primers, checking for correct pumping action and positive fuel shut-off in the locked position.
•
11-60. AUXILIARY FUEL SYSTEM.
11-61. DESCRIPTION. The system is described in
paragraph 11-2.
11-62. REMOVAL OF AUXILIARY FUEL TANK.
Removal of either auxiliary fuel tank is accomplished through the top of the wing.
a. Place fuel selector valve in the OFF position.
b. Completely drain the auxiliary fuel tank to be
removed by removing the quick-drain valve in the
bottom of the tank.
NOTE
Support outer wing panel and tail boom with
cradle supports, before removing fuel tank
covers, to prevent wing and boom deflection.
c. Remove auxiliary fuel tank cover from top of
wing by removing screws around outer edge of cover
and around fuel filler opening.
d. After screws are removed, the forward edge of
the cover must be pulled aft from under the wing
leading edge skin. Retain gaskets between filler
neck and top wing cover.
e. Remove bolts from retaining straps securing
tank to wing structure.
f. Disconnect wire from fuel quantity transmitter
•
THRU 337E
21
21
19
rrJP
11
-~.
3~
17
./
15
I
-..r\
I
1&
•
I
I
15
i
14
J
v·
13
//
I
J
I
EXCEPT T337 SERIES
NOTE
I
Torque nut (10) to 25-30 Ib in.
2
l.
FRONT INSTALLATION
2.
3.
4.
5.
6.
7.
337F
B.
9.
9
10.
11.
12.
13.
14.
15.
16.
17.
•
lB.
19.
20.
2l.
22.
23.
T337 SERIES
24.
2
T337 SERIES
Figure ll-B. Fuel Strainers Installation
11-27
•
•
2------------~~~
l. Knob
THRU MODEL 337C-SERIES
2.
3.
4.
5.
6.
7.
8.
9.
10.
Cap Assembly
Bulb
Button
Spacer
Shell Assembly
Shim
Stop
Glass Assembly
Placard
11.
12.
13.
14.
15.
16.
17.
18.
19.
Console
Ceiling Bracket
Fuel Selector Gear Box
Base Plate
Microswitch
Switch Bracket
Lever
Washer
Plate
Figure 11-9. Fuel Quantity Indication (Sheet 1 of 3)
or sending unit.
g. Disconnect fuel outlet line at tank.
h. Disconnect fuel vent line at tank and remove
line by pulling up and forward to remove line from
grommets in wing structure on non-turbocharged
airplanes. On turbocharged airplanes, disconnect
the vent line at the hose connection near the tank,
and remove the short section of line outboard of the
hose connection.
i. Disconnect scupper drain line by loosening
clamp on hose just outboard of tank and pull hose
free from tank and lift tank from wing.
may be accomplished by reversing the steps of paragraph 11-62. Cradles to support the outer wing
panel and tail boom should be provided to prevent
wing and boom deflection. Wing and boom deflection
can cause misalignment of holes in wing and fuel
tank cover, making installation of the cover difficult.
When installing fuel tank cover, make sure that forward edge of cover is under wing leading edge skin.
Be sure that gaskets are placed between scupper and
fuel tank cover. A maximum of three gaskets may
be used to maintain wing contour and prevent canning
of the cover.
11-63. INSTALLATION OF AUXILIARY FUEL
TANKS. Installation of either auxiliary fuel tank
11-64. FUEL QUANTITY TRANSMITTER OR SENDING UNIT. Prior to the Model 337E-Series, a trans-
11-28
•
•
"2
3
20
19
10
•
18
MODEL 337D THRU 337F
*BEGINNING WITH MODEL 337E-SERIES
1. Fuel Selector Gear Box
2. Base Plate
3. Nutplate
4. Microswitch
5. Plate Assembly
6. Straight Pin
7. Lever
•
8. Washer
9. Plate Assembly
10. Placard
11. Shell Assembly
12. Bulb
13. Cap Assembly
14. Knob
15.
16.
17.
18.
19.
20.
21 •
Cover
Button
Shim
Glass Assembly
Pin
Lever
Cotter Pin
Figure 11-9. Fuel Quantity Indication (Sheet 2 of 3)
11-29
•
11
10----J1
9
9----+......,~~
7
12
BEGINNING WITH 1973
MODEL 337G
7
6
5
4
•
•
13
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
3
11.
12.
13.
14.
15.
16.
Screw
Cover
Knob
Glass Assembly
Placard Assembly
Plate Assembly
Lower Lever
Pin
Upper Lever
Bracket
Bracket
Washers
Washer
Shell Assembly
Bulb
Cap Assembly
Figure 11-9. Fuel Quantity Indication (Sheet 3 of 3)
11-30
•
•
mUter is installed in each auxiliary tank. ~eginning
with the Model 337E -Series, a fuel quantity sending
unit is installed. Transmitters are described in
paragraph 11-8. Sending units are described in Section 14. Prior to the Model 337C-Series, the transmitter in each auxiliary tank is connected to a separate indicator on the instrument panel. Beginning
with the Model 337C-Series, main and auxiliary tank
readings are registered on a common indicator for
left and right tanks.
11-65. REMOVAL AND INSTALLATION OF FUEL
QUANTITY TRANSMITTER. Removal and installation of the transmitter in the auxiliary fuel tanks is
similar to procedures used for main tanks. Refer
to Section 15.
11-66. FUEL VENT. Auxiliary fuel tank vents are
described in paragraph 11-27 which discusses venting
of the complete fuel system.
11-67. REMOVAL OF FUEL VENT. Refer to paragraph 11-28.
11-68. CHECKING FUEL VENTS. Refer to paragraphs 11-29 and 11-32.
11-69. INSTALLATION. Refer to paragraph 11-28.
•
11-70. DRAIN VALVE. A quick-drain valve is in-
stalled in the bottom of each auxiliary fuel tank.
This valve is used to sample fuel for water and
sediment. The valve is removed or installed simply by screwing it in or out. Be sure to safety valve
after installation.
11-71. AUXILIARY FUEL LINE. The only fuel line
in the auxiliary system is a short outlet line from the
tank to the fuel selector valve. The line can be removed after disconnecting it at the tank and selector
valve.
11-72. FUEL QUANTITY INDICATION. (Refer to figure 11-9.) Beginning with the 1968 Models 337C and
T337C, o~ly two fuel quantity indicators are provided
in the instrument cluster on the panel. The indicators
are for left and right fuel tanks, and show both main
and auxiliary fuel tank levels. A PUSH -TO-GAGE
button on each fuel selector handle in the overhead console is depressed when either handle is turned to the
AUX position. The button mechanically operates microswitches which cause the indicator to register
fuel level in the auxiliary tanks instead of the main
tanks when it is depressed. Either button may be depressed manually to obtain a temporary reading of
fuel level in the corresponding auxiliary tank. Beginning with 1973 Model 337G aircraft, the four-position
gearbox configuration has been replaced with a threeposition cammed belle rank design. The figure may be
used as a guide for replacement of components.
SHOP NOTES:
•
11-31/(11-32 blank)
•
•
SECTION 12
PROPELLERS AND PROPELLER GOVERNORS
TABLE OF CONTENTS
Page
PROPELLERS .
Description
Repair
Trouble Shooting .
Removal
Installation
PROPELLER GOVERNORS
Description
Trouble Shooting .
Removal
Installation
High-RPM Stop Adjustment
Overhaul
Propeller Feathering Controls
Feathering Lift Rod Adjustment
12-1
12-1
12-1
12-2
12-3
12-4
12-4
12-4
12-4
12-4
12-4
12-6
12-6
12-6
12-6
12-1. PROPELLERS. (Refer to figure 12-1.)
•
12-2. DESCRIPTION. The aircraft is equipped with
McCauley all-metal, constant-speed. full-feathering,
governor-regulated. two-bladed propellers employing a six bolt flange mount hub. The front propeller
is a tractor-type and the rear propeller is a pushertype. The front propeller rotates clockwise as viewed from the rear of the aircraft, while the rear
propeller, equipped with left hand blades, rotates
counterclockwise as viewed from the rear of the
aircraft. Both propellers operate in the same manner. Each propeller is Single-acting in which oil
pressure from its engine, boosted and regulated by
a governor, is used to decrease blade pitch while
the forces produced by external counterweights and
internal springs are used to increase blade pitch and
to feather. An internal. pressure-operated latching
mechanism prevents feathering during engine shutdown. Beginning with aircraft serials 33701195 and
F33700001, a new centrifugal feathering latch is installed in the propellers. The function of the new
latch depends on spring tenSion and centrifugal force
thereby eliminating the variables of differential oi.l
pressure used by the propeller on the earlier m01elyears. Beginning with aircraft serials 33701317 and
UNFEA THERING SYSTEMS
Description
Maintenance .
Accumulator Overhaul
PROPELLER SYNCHRONIZER SYSTEM
Description
Controller Removal and Installation
Actuator Removal, Installation and
Rigging
Adjustable Rod End Removal and
Installation
Flexible Shaft and/or Gl.Lide Tube
Removal and Installation
Magnetic Pick-Up Removal, Installation
and Adjustment
Synchronizer Functional Test
12-7
12-7
12-7
12-7
12-7
12-7
12-7
12-7
12-7
12-11
12-11
12~11
F33700025, a new threadless blade propeller is installed. With this deSign, the blades use split retaining rings which are assembled around the blade base
after the blade is assembled into the propeller hub.
Unfeathering the propeller is accomplished by placing the propeller control lever forward of the feathered position and rotating the blades to low pitch position, or by starting the engine with the propeller
control lever forward of the feathered position. An
optional unfeathering system, discussed later, may
be installed. Also, an optional automatic propeller
synchronizing system, discussed later, may be installed. Refer to Section 13 for the propeller antiice system which may be installed as optional eqUipment.
12-3. REPAIR. Metal propeller repair first involves evaluating the damage and determining
whether the repair will be a major or minor one.
Federal Aviation Regulations, Part 43 (FAR 43). and
Federal Aviation Agency. Advisory Circular No.
43.13 (FAA AC No. 43.13). define major and minor
repairs. alterations and who may accomplish them.
When making repairs or alterations to a propeller
FAR 43. FAA AC No. 43. 13 and the propeller manufacturer's instructions must be observed.
12-1
12-4. TROUBLE SHOOTING.
TROUBLE
FAILURE TO CHANGE PITCH.
REMEDY
PROBABLE CAUSE
Governor control disconnected
broken.
Ill'
C 1I11J1l'ct or replacE' (·ont rol.
Governor not correct for
propeller. (Sensing wrong. )
Rt'place
Defective governor.
Refer to llaragraph 12-9.
Defective pitch changing mE'chanism inside propeller 1..'1" excessivE'
propeller blade friction.
Check propeller manually.
repair ur rel>lace as required.
Improper rigging of governor
control.
Check that governor control arm
and control have full ~ravel.
Rig control and arm as required.
Defective governor.
Refer to paragraph 12-9.
SLUGGISH RESPONSE TO
PROPELLER CONTROL.
Excessive friction in pitch
changing mechanism inside
propeller or excessive blade
friction.
C heck propeller manually, repair
or replace as required.
STATIC RPM TOO HIGH.
Governor high-rpm stop set
too high.
Refer to paragraph 12-12.
Defective governor.
Refer to paragraph 12-9.
Incorrect propeller or incorrect
low pitch blade angle.
Check aircraft specification and
install correct propeller with
correct blade angle.
Governor high-rpm stop set
too low.
Refer to paragraph 12-12.
Defective governor.
Refer to paragraph 12-9.
Incorrect propeller or incorrect
low pitch blade angle.
Check aircraft specification and
install correct propeller with
correct blade angle.
Sludge In governor.
Refer to paragraph 12-9.
Air trapped in propeller
actuating cylinder.
Trapped air should be purged
by exercising the propeller
several times prior to take-off,
after propeller has been reinstalled or has been idle for
an extended period.
Excessive friction in pitch
changing mechanism inside
propeller or excessive blade
friction.
Check propeller manually,
repair or replace as requi red.
Defective governor.
Refer to paragraph 12-9.
FAILURE TO CHANGE PITCH
FULLY.
STATIC RPM TOO LOW.
ENGINE SPEED WILL NOT
STABILIZE.
12-2
•
~overnor.
•
•
•
TROUBLE
OIL LEAKAGE A T PROPELLER
MOUNTING FLANGE.
•
•
PROBABLE CAUSE
REMEDY
Damaged O-rirlg seal between
engine crankshaft flange and
propeller.
Remove propeller and install new
O-ring seal.
Foreign material between
engine crankshaft flange and
propeller mating surfaces or
mounting nuts not tight.
Remove propeller and clean
mating surfaces; install new
O-ring and tighten mounting
nuts evenly to torque.yalue shown
in figure 12-1.
OIL LEAKAGE AT ANY
OTHER PLACE.
Defective seals, gaskets,
threads, etc., or incorrect
assembly.
Propeller repair or replacement
is required.
FAILURE TO FEATHER
OR UNFEA THER.
Defective governor.
Refer to paragraph 12-9.
Defective pitch changing
mechanism or excessive
blade friction.
Check propeller manually,
repair or replace as
required.
Incorrect rigging of governor
control.
Check that arm on governor
has full travel. Rig in accordance with Section 10•
Defective latching mechanism inside propeller.
Propeller repair or replacement is required.
Latching mechanism does
not engage.
A propeller may occasionally
feather during shut-down. If
this occurs repeatedly, the
latching mechanism is defective. Repair or replace as
required.
PROPELLER FEATHERS
DURING ENGINE SHUTDOWN.
12-5. REMOVAL. (Refer to figure 12-1.)
a. Start engines, feather propellers and shut down
engines. Propellers should be removed in the
"FEATHERED" positions.
b. If optional unfeathering systems are installed,
dissipate system pressure as follows:
1. After the front propeller has been feathered
and the front. engine shut down, move front propeller
control out of "FEATHER" position until blades start
to unfeather, then quickly pull the control back into
"FEATHER. "
2. Continue to "milk" pressure out of the system with the propeller control until the propeller
blades will no longer move. This may require from
15 to 20 movements of the propeller control.
3. Do not allow propeller blades to rotate far
enough to let high pitch latches engage, or engine
will have to be restarted, propeller feathered again
and the procedure repeated.
,
•
.
4. After the front propeller has been feathered
and system pressure dissipated, repeat the procedure to place the rear propeller in the feathered position with system pressure dissipated.
NOTE
Either the front or rear engine propeller and
propeller spinner may be removed as a complete unit.
c. If spinner is to be removed, remove attaching
screws and remove spinner, spinner support and
spacers. Retain any spacers behind spinner support.
d. (Front propeller.) Remove cowling and nose
cap as necessary to gain access to propeller attaching nuts. Either the right or left nose cap may be
removed.
12-3
e. (Rear propeller.) Remove cowl side panels and
tall cap as necessary to gain access to propeller
attaching nuts.
f. Loosen ]lI'Opeller mounting nats untll they
contact the crankcase, then pull propeller away
from crankcase until halted by mouT,ting nuts.
I
NOTE
As the propeller is separated from the
engine, oU w1ll drain from the propeller
and crankshaft cavities.
g. Remove propeller mounting nuts and washers
and pull propeller forward to remove from engine
crankshaft.
h. If desired the propeller spinner bulkhead may
be removed from the propeller by removing the
attaching bolts.
12-9. TROUBLE SHOOTING. When trouble shooting
a propeller-governor combination, it is recommended
that a governor known to be in good condition be installed to check whether the propeller or the governor is at fault. Removal and replacement, highspeed stop adjustment, desludging and replacement
of the mounting gasket are not major repairs and
may be accomplished in the field. Repairs to governors are classed as propeller major repairs in
Federal Avtation Regulations, which also define who
may accomplish such repairs.
•
12-10. REMOVAL.
a. Remove cowling and baffles as required for
access.
b. If an optional unfeather1ng system is not installed, place propeller control in high rpm-position.
c. Disconnect propeller control from governor.
NOTE
12-6. INSTALLATION. (Refer to figure 12-1.)
a. If removed, install spinner bulkhead on propeller hub. AUgn blade cutouts in bulkhead fillet with
propeller blades.
b. Clean propeller hub and engine crankshaft
cavities and mating surfaces.
c. LlghUy lub\1cate a new O-ring and engine crankshaft pUot with clean engine 011 and install O-ring in
propeller hub.
d. Align propeller mounting studs and dowel pins
with correct holes in engine crankshaft flange and
slide propeller over crankshaft pUot untll hub nange
is approximately 1/4 inch from crankshaft nange.
e. Install propeller attaching washers and nuts
and work propeller aft as far as possible, then
tighten nuts evenly and torque to 55-65 Ib ft.
f. Install spacers and spinner support. The spacers
are used as required (maximum of 4) to cause a snug
fit between the support and the spinner.
g. Install spinner and cowling removed for access.
12-7. PROPELLER GOVERNORS.
12-8. DESCRIPTION. The propeller governor is a
single-acting, centrifugal type, which boosts 011 pressure from the engine and directs it to the propeller
where the 011 is used to increase blade pitch. A
single-acting governor uses 011 pressure to effect a
pitch change in one direction only; a pitch change in
the opposite direction results from a combination of
centrifugal twisting moment of rotating blades and
compressed springs. 011 pressure is boosted in the
governor by a gear type all pump. A pilot valve, fly
weights and speeder· spring act together to open and
close governor 011 passages as required to maintain
a constant engine speed.
NOTE
Whether 011 pressure is used to increase or
decrease blade pitch cannot be determined
by the outward physical appearance of the
governors. Always be sure the correct governors, with correct part numbers, are used.
12-4
Note position of all washers so that washers
may be installed in the same posttion on reinstallation.
d. If an optional unfeathering system is installed,
release accumulator pressure, then disconnect accumulator hose from governor fIttlng.
Always release accumulator pressure through
filler valve, before disconnecting hose between accumulator and governor or removing
accumulator.
e. If an optional propeller synchronizing system
Is installed, remove the magnetic pick-up from
governor.
f. Remove nuts and washers securing governor
and pull governor from mounting studs.
g. Remove gasket between governor and engine
mounting pad.
12-11. INSTALLATION.
a. Wipe governor and engine mounting pad clean.
b. Install a new gasket, with the raised surface of
the screen away from the engine pad.
c. Position governor on mounting studs, aligning
governor splines with splines in engine and install
mounting washers and nuts. Do not force spline engagement. Rotate engine craakshaft and spUnes will
engage smoothly when aligned.
d. If an optional unfeathering system -is installed,
connect accumulator hose to governor and recharge
the accumulator.
e. If an optional propeller synchronizing system
is installed, connect the magnetic pick-up to governor.
f. Connect governor arm. If rod-end adjustment
was not disturbed, it should not be necessary to rig
the control. Check rigging and adjust as required.
Refer to Section 10.
g. Reinstall baffles and cowUng removed for
access.
•
•
eo
eThru aircraft serial 337 -1012
bolt (8) is reversed.
7
••
REAR PROPELLER AND
SPINNER INSTALLATION
SHOWN
10
e
FRONT SPINNER
BULKHEAD
13
• TORQUE TO 660 - 780 LB-IN. (55 - 65 LB-FT. )
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
Spinner
Spinner Support
Spacer
Propeller
Stud
Mounting Nut
Nutplates
Bolt
Bulkhead Assembly
Doubler
O-Ring
Dowel Pin
Screw
NOTE
Use spacers (3) as required (maximum of 4)
to cause a snug fit between the spinner (1)
and the spinner support (2).
The front propeller and spinner installation is
the same as the rear, except that right hand
instead of left hand blades are used, the counterweights are opposite and the hub is shorter.
e
Figure 12-l.
Propeller Installation
12-5
•
FEA THERING '
LlFTROD~
HIGH-SPEED
STOP SCREW
"
LOCK NUT ----~!'b~J
Figure 12-2. High-Rpm Stop Adjustment
12-12. HIGli-RPM STOP ADJUSTMENT. (Refer
to figure 12-2.)
a. Remove engine cowling and baffles as necessary
for access.
.
b. Loosen lock nut on high-speed stop screw.
c. Turn the screw IN to decrease maximum rpm
and OUT to increase maximum rpm. One full turn of
the stop screw causes a change of approximately 25
rpm.
d. Make propeller control adjustments as required
for full travel and proper cushion at the control quadrant. Refer to Section 10.
e. Tighten the lock nut on the high-speed stop screw.
L Reinstall baffles and COWling removed for access.
g. Test operate the propellers and governors.
NOTE
It is possible for either the propeller low
pitch (high-rpm) stop or the governor highrpm stop to be the high-rpm limiting factor.
It is desirable for the governor stop to
limit the high-rpm at the maximum rated
rpm for a particular aircraft. Due to climatic condittons, field evaluation, lowpitch blade angle and other considerations,
an engine may not reach rated rpm on the
ground. n may be necessary to readjust
the governor stop after test flying to obtain
maximum rated rpm when airborne.
12-13. OVERHAUL. The propeller governor should
be overhauled at each recommended engine overhaul
period. If an engine is required to be overhauled
prematurely, and it Is suspected the governor has
been affected also (011 contaminatton, etc.), then
the governor should be overhauled as well. This is
strictly a matter of judgement. The governor overhaul manual is available from the Cessna Service
Parts Center.
12-6
12-14. PROPELLER FEATHERING CONTROLS.
Each propeller feathering control mechanism is
housed in the handle of the I)ropeller control lever:
By lifting the handle (pulling it out) and moving the
control aft, an additional 15° of travel pulls the governor arm into the feathering position. The handle
may be disassembled by removing the knob and carefully lifting the outer sleeve. As sleeve is raised,
the spring and link wtll fall free. Note position of
components for reassembly.
12-15. FEATHERING LIFT ROD ADJUSTMENT.
(Refer to figure 12-2.) Minor adjustment of the
feathering lift rod may be necessary to obtain
proper feathering action and rpm stabilization.
While holding feathering lift rod, loosen jam nut and
then tum feathering lift rod clockwise to increase
stabilization rpm with corresponding increased time
to feather or counterclockwise to decrease rpm and
time.
a. Start and run engine at 1000 rpm until oil and
cylinder head temperature is in normal operating
range.
b. With propeller control lever in full increase
positton, set throttle to obtain 1800 rpm. Retard
propeller control lever to the safety step at the full
decrease position while monitoring the tachometer.
There should be no change in rpm. Retard propeller
control lever over the step to the full feather position.
Rpm should drop to 1200 rpm within 3 seconds.
Promptly recover rpm by moving propeller control
lever to the full increase position.
c. Advance throttle to 2400 rpm. Retard propeller
control lever to the safety step at the full decrease
position. Rpm should stabilize at 2100 plus or minus
100 rpm.
d. Adjust feathering 11ft rod if not within the preceding prescribed limitS. One-half revolution of the
lift rod clockwise will lower the feathering rpm approximately 100 revolutions.
•
•
•
12-16. UNFEA THERING SYSTEMS. (Refer to figure 12-3.)
12-17. DESCmPTION. Each optional unfeathering
system consists of a nitrogen-charged accumulator,
a special governor and a hose running between the
governor and the accumulator. The governor contains a spring-loaded check valve which is unseated
while the propeller control is in any position except
"FEA THER", thus permitting governor-pressurized
oil to flow to and from the accumulator. When the
propeller control is moved to "FEATHER" position,
the check valve is seated and oil under governor
pressure is trapped in the accumulator and hose. As
the propeller control is moved from the "FEATHER"
position, the trapped pressurized oil flows back
through the governor to the propeller to unfeather it.
12-18. MAINTENANCE.
ItAUTION\
Always release system pressure by placing
propeller control in high-rpm position and
release accumulator pressure through the
fUler valve, before disconnecting hose between accumulator and governor or removing accumulator.
•
a. Place propeller control in the high-rpm position
before charging the accumulator to prevent the possibility of oil under pressure being trapped in the accumulator.
b. Although the accumulator will function properly
when charged with air, nitrogen gas is recommended
to minimize corrosion.
c. Either too much pressure or not enough pressure in the accumulator will reduce efficiency of the
unfeathering system. With a normal amount of frictton within the propeller, optimum pressure is the
approximate mid-range of the pressures speCified
in figure 12-3.
d. AIW'ci.ys check that the filler valve does not leak
after charging an accumulator.
12-19. ACCUMULATOR OVERHAUL. The propeller unfeathering accumulator should be overhauled
at each recommended engine overhaul period. If an
engine or governor is required to be overhauled prematurely and it is suspected the accumulator has
been affected also (oil contamination etc), then the
accumulator should be overhauled as well. This
is strictly a matter of judgement. The propeller
unfeathering accumulator overhaul manual is available
from the Cessna Service Parts Center.
12-20. PROPELLER SYNCHRONIZER SYSTEM.
(Refer to figure 12-4.)
•
12-21. DESCRIPTION. The propeller synchronizing system is comprised of a controller mounted in
the cabin, an actuator attached to the rear engine
firewall/mount, special governors with magnetic impulse pick-ups, a control switch mounted on the engine control pedestal, a flexible control shaft from
the actuator to the rear engine governor and electrical wiring. Witb the engines operating within apprOXimately 3C~rpm of each other, placiDg the control
switch to the C'~ position wtll cause the- rear engine
rpm to be aut~ttcally adjusted to ~ same rpm as
that of the front engine. The rear engine rpm may be
manually changed by the governor control lever at
any time. The control range that the front engine and
controller has over the rear engine, when the control
switch is ON, is apprOXimately 60 rpm; therefore,
the propeller should be manually synchronized within
this controlling range before placing the control
switch to the ON position. When the control switch
is in the OFF position, the controller automatically
adjusts the rear engine adjustable rod end to the
center of Its range. The rear engine is then controlled manually by the propeller control lever.
12-22. CONTROLLER REMOVAL AND INSTALLATION.
a. Disconnect electrical plug and remove control
switch from control pedestal.
b. Disconnect indicator light electrical leads from
control switch.
c. Remove four screws, washers and nuts attaching controller to bottom of glove box.
d. Reverse the preceding steps for reinstallation.
12-23. ACTUATOR REMOVAL, INSTALLATION
AND RIGGING.
a. Remove rear engine cowling as necessary for
access.
b. Cut safety wire and disconnect electrical plug
from actuator.
c. Disconnect flexible shaft from actuator.
d. Remove four bolts, washers and nuts attach-.
ing actuator to brackets on engine mount/firewall,.
e. Install actuator by installtng attaching bolts,
washers and nuts and connecting electrical plug to
actuator.
f. With flexible shaft disconnect and control
switch OFF, place master switch ON. This will
cause the actuator to be centered.
g. Rotate flexible shaft to place adjustable rod end
in the center of its travel range.
h. Connect flexible shaft to the actuator and safety
electrical plug.
i. Install engine cowling and perform functional
test.
12-24. ADJUSTABLE ROD END REMOVAL AND
INSTALLATION.
a. Remove engine cowling as necessary to gain
access to propeller governor.
b. Cut safety wire and disconnect flexible shaft
from rod end.
c. Disconnect rod end from governor control arm
and remove rod end from governor control.
d. Install rod end on governor control.
e. With adjustable rod end set at its mid-point of
travel, rig governor as outlined in Section 10.
C. Rotate the splined shaft in rod end assembly to
one end of its travel. Move the propeller control
lever through its entire range of travel and observe
the governor control arm to be certain it hits both
the maximum and minimum rpm stops.
12-7
•
FRONT ENGINE
•
5
AIRCRAFT SERIAIS 337-0181 AND
337-0183 THRU 337-0211
~
/
---
...
.
/
•
/
AIRCRAFT SERIAIS 337-0182 AND
337 -0212 THRU 33701398
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Governor
Elbow
Hose Assembly
Baffle
Grommet
Bracket
Accumulator Assembly
Clamp
Spacer
Fire Sleeve
11. Engine Mount
12. Line Assembly
13. Union
BEGINNING WITH AIRCRAFT
33701399 AND F33700046
NOTE
Beginning with aircraft serials 337-0467 and F33700001,
a Woodward or a McCauley unfeathering accumulator may
be Installed. Removal and installation procedures are
Similar, however, charge pressure for the Woodward
accumulator is 100-125 PSI and for the McCauley Is 90100 PSI.
FIgure 12-3. Unfeathering Systems (Sheet 1 of 2)
12-8
SERIA~
•
•
REAR ENGINE
NOTE
Position accumulator (7) on inboard side
of engine mount (11). Position accumulator forward or aft to provide the best
hose routtng and access to filler.
I
AIRCRAFT SERIALS 337-0181 AND
337-0183 THRU 337-0211
11
•
AIRCRAFT SERIALS 337-0182 AND
33700212 THRU 33701229 AND F33700001 THRU F33700009
11
12
•
BEGINNING WITH AIRCRAFT SERIALS
33701230 AND F33700010
•
Figure 12-3. Unfeathering Systems (Sheet 2 of 2)
12-9
1
•
A
--x------------..
I
1·- ., .-.
7_1- -
~,/ - -'~~.,
~_/
3
..:..,
...•.....
.,~
·.~I
10
*
B
• Beginning with aircraft serial 337-0410 and all service parts
4
15 18 _
~
~
1.
2.
3.
4.
5.
Indicator Ught (7) is used thru aircraft serial 337 -0369
Detail
B.
12
Rear Propeller Governor
Actuator
Wiring
Rear Propeller Control
Controller
6.
7.
8.
9.
10.
11.
Control Switch
Indicator Light
Circuit Breaker
Front Propeller Governor
Electrical Plug
Mounting Plate
Figure 12-4. Synchronizer System
12-10
12.
13.
14.
15.
16.
Flexible Shaft
Guide Tube
Magnetic Pick-Up
Adjustable Rod-End
Lock Tab
•
i.
•
!!. M:\llUalIy J·ulal(· ~plined shaft in rod l'ulJ, dSS,'milly 10 III(> opposite end of its travel and repeat' check
in step "f." This assures that pI'opeller control rig(ting allows stop-to-stop travel with any possible rod
end selling.
h. r. -nnrtt fl£'xible shaft to rod end assembly and
;.;al:-tv.
.
I.
Disculluel't flexible shaft from actuator and with
(:ontrol switch OFF, place master switch ON.
This will allow actuator te, run to the center of its
ran!!l!.
j. Cunnect fl('xible shaft to actuator and safety.
k. Install engine cowling and perform functional
test.
12-25. FLEXIBLE SHAFT AND/OR GUIDE TUBE
REMOVAL AND INSTALLATION.
a. Disconnect flexible shaft from actuator and rod
end assembly.
b. Remove clamps attaching guide tube to engine.
c. At rod end of flexible shaft, remove lock ring
and hex nut and pull flexible shaft from guide tUbe.
d. Secure guide tube to engine using clamps removed in step "b. "
e. Remove lock ring and hex nut from flexible
shaft.
f. Lubricate flexible shaft housing (MIL-G-21164),
where it will slide in the guide tube.
g. Insert flexible shaft through the guide tube so
that lock ring end of the flexible shaft will mate with
adjustable rod end and install hex nut and lock ring .
h. Cunnecl flexible shaft to rod end assembly and
rotate shaft to obtain center of rod end travel range.
i. With control SWitch OFF, place master switch
ON. This will allow actuator to run to the center of
its range.
. j. Connect flexible- shan tu actuator and sarety.
k. Install c'n-::i'1(: cowling.
12<;C. :VIAGNETIC PICK-UP REMOVAL INSTALLATION AND ADJUSTMENT. Refer to Woodward Governor Bulletin No. 33049A for replacement or adjustment.
12-27. SYNCHRONIZER FUNCTIONAL TEST. To
make a functional test of the synchronizer system in
flight, first determine the limited rpm range through
which the rear engine will remain synchronized with
the front engine. To do this, manually synchronize
the propellers and then turn on the control switch.
Slowly move the front engine propeller control lever
to increase and decrease rpnl, noting the range of
rpm through which the rear engine will remain synchronized. This is the limited operating range of
the synchronizer. With the control switch turned
on, move the front engine propeller control lever
close to either end of this limited range. Turn off
the control switch to develop an unsynchronized
condition as the actuator returns to its mid-position.
Turn on the control switch and check that automatic
synchronization occurs.
NOTE
The flexible shaft must be free to slide in the
guide tube when the governor control is operated•
•
12-11/(12-12 blank)
•
SECTION 13
UTILITY SYSTEMS
... TABLE OF CONTENTS
•
•
Page
.13-28
HEATING, VENTILATING AND DEFROSTING
SYSTEM (Non-Turbocharged Aircraft)
13-1
..
Trouble Shooting •
13-3
Removal and Installation
•.
13-3
HEATING, VENTILATING AND DEFROSING
SYSTEM (Turbocharged Aircraft)
.13-10
Trouble Shooting .
.13-10
Removal and Installation
.13-10
DE-ICE SYSTEM (Thru 33701462)
.13-13
System Operation
.13-13
Removal and Installation
.13-13
Trouble Shooting.
"
.13-13
Operational Check . . . . . . . . .
.13-20
DE-ICE SYSTEM (Beginning with 33701463)
.13-20
Description
•
•
.13-20
Component Description
.13-20
System Operation
.13-21
System Removal
.13-21
De-Ice Boot Repair .
• .13-21
Description
.
.13-21
Repair
.
..
.13-21
Replacement of De-Ice Boots
.13-24
PROPELLER DE-ICE SYSTEM
.13-24
Slip Ring Alignment •
• 13-27
Trouble Shooting .
.13-27
Timer Test
••
Installation and Alignment of
Brush Block Assembly
Replacement of De-Icer Boots
OXYGEN SYSTEM
Description
•
Maintenance Precautions
Replacement of Components
Oxygen Cylinder General
Information .
•
Oxygen Cylinder Service
Requirements
Oxygen Cylinder Inspection
Requirements
.
Oxygen System Component
Service Requirements
Oxygen System Component
Inspection Requirements
Masks and Hose.
.
Maintenance and Cleaning
System Purglng •
Functional Testing
System Leak Test
System Charging .
13-1. HEATING, VENTILATING, AND DEFRalTING SYSTEM (NON -TURBOCHARGED AIRCRAFT).
defroster outlets. In addition to fresh air suppUed
through the heating and ventilating system, individual fresh air control valves are provided for the
occupant of each seat, including the optional fifth
and sixth seats, when installed. The air control
valves for the pilot and copilot are located in a plenum box mounted immediately forward of the overhead console. This plenum box receives fresh air
from ducts routed from an inlet in the leading edge
of the wing root fairing of each wing. The four
rear seat air control valves, mounted above the
side windows at each seat, receive fresh air from
ducts routed from an inlet in the leading edge of each
wing. Rotating air control valves counterclockwise
gradually increases air flow through each valve. A
cabin air exhaust vent is installed In the rear firewall to route stale air overboard through an outlet
duct. The exhaust vent also provides better circulation of incoming cabin air. Beginning with the
1967 model, the cabin air ventilation system was
redesigned to incorporate a plenum chamber with
a valve which meters the incoming cabin ventilation air. This provides a chamber for the expansion of cabin air which greatly reduces inlet air
noise. An additional floor air distribution box is
added to each duct across the aft side of the firewall.
13-2. Ram air, routed through ducts connected to
the horizonW baffles of the front engine, is ducted
through the heat exchange section of the engine exhaust mufflers to mixing airboxes on the forward
side of the firewall. Unheated ram air, routed
through ducts connected to the vertical baffles of
the front engine, is also ducted to these airboxes.
The position of a valve on the forward end of each
airbox controls the temperature of air entering the
cabin. Air is distributed Into the forward cabin area
through holes in two ducts, one located behind each
airbox, on the aft side of the firewall. The rear
cabin area receives air which flows around and between the front seats. Additional air for the rear
cabin is routed through ducts attached to the firewall
ducts. Openings for these ducts are provided at the
forward door posts In the side panels. Two temperature control knobs on the instrument panel individually control the position of a valve in each airbox.
Rotating the knob clockwise gradually opens the
heated air passage and simultaneously closes the
unheated air passage. Intermediate settings blend
heated and unheated air. The DEFROST knob operates a damper valve at the defroster supply duct.
Pulling the knob gradually increases airflow to the
.13-28
.13-29
.13-29
.13-30
.13-30
.13-30
.13-34
.13-34
.13-34
.13-34
.13-35
.13-35
.13-35
.13-35
.13-35
.13-35
.13-36
13-1
DetauD
•
'".
1&
•
Detail
tv;l '-~L·
,,'
DetauA
~
. ... \ ,.
, .......................
...
. .- -.'
.
..
PRIOR TO 1967
~
DetauH
Figure 13-1. Heating • Ventilating • and Defrosting Systems (Sheet 1 c:4 7)
13-2
.
!!'
:
(
\
3O~}'J.,-
45
F
,
I
37
DetallG
•
•
•
13-3. Beginning With the 1969 Models 337D and
T337D, the aircraft are equipped With quadr~Lnt-type
heater controls. The standard aircraft has three
vertical operating controls; left and right cabin air
controls and the defroster control. The "OFF' position for all three controls is at the top of the panel.
The cabin air control levers increase the amount of
fresh air entering the cabin as the control is moved
downward. The maximum fresh air position is just
below the center of the panel. As the controls are
moved downward from this position, the fresh air is
slowly closed off and heat is added to the system.
Maximum heat is attained when the control is in the
full down positiOn. The heater controls on turbocharged aircraft have the cabin air volume control on
the left and the defroster control in the center with
the "OFF" position of both controls at the top of the
panel. The thermostat control is located on the right
with the "LOW" position at the top of the panel. Refer
to figure 13-1.for heater controls beginning with
the 337D Series.
13-4. TROUBLE SHOOTING. Most of the operational troubles in the heating, ventilating, and defrosting system are caused by sticking or binding
air valves and their controls, damaged air ducting,
or defects in the exhaust muffler. In most cases,
lubrication will free sticking or binding parts.
Damaged or broken parts should be repaired or
replaced. Check that flexible hoses are properly
secured and replace hoses that are crushed, frayed,
burned, or otherwise damaged. Check that valves
respond freely when operated by their controls,
that they move in the correct direction, and that
they move through "their full range of travel and
seal properly. U fumes are detected in the cabin,
a very thorough inspection of the exhaust muffler
should be conducted. Refer to applicable paragraphs
in Section 10 for this inspection. Since any holes or
cracks may permit engine exhaust fumes to enter the
cabin, replacement of defective parts is imperative
because exhaust fumes in the cabin are extremely
dangerous.
13-5. REMOVAL AND INSTALLATION OF COMPONENTS. Figure 13-1 shows the various parts 01
the heating, ventilating, and. defrosting systems, and
may'be used as a guide for replacement of parts.
When ass"embling components shown on sheet 6 of figure 13-1, apply LocUte, Grade A-A to all contacting
parts of valve plate (3), star washer (13), plate (14),
shaft (18), and associated nuts at final assembly.
Seal between plate assembly (5) and lower housing (8)
with Presstite 579. 6 sealer, or equivalent, as required to prevent air leaks. Use tire talc (powdered
soapstone) between plate (14) and seal (15). At both
ends of springs, use Dow Corning Silicone grease #33
medium, or equivalent. When assembling plenum
chamber and silencer assembly, torque nuts to 40
pound-inches. When installing a new flexible hose,
cut to length and install in the original routing. Trim
the hose winding shorter than the hose to allow hose
clamps to be fitted securely.
References for Figure 13-1 (Sheet 1)
1.
2.
3.
4.
5.
6.
•
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
Cover
Valve
Gasket
Air Outlet
Air Duct
Air Duct
Air Scoop
Noise Filter
Air Duct
Clamp
Adapter
Seal
Wing Leading Edge
Air Duct
Clamp
Aircraft Structure
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
Plenum Chamber
Plate
Valve Assembly
Defrost Valve Outlet
Right Hand Heat Control
Defroster Control
Left Hand Heat Control
Defrost Outlet
Air Outlet Duct
Fresh Air Inlet Duct
Air Duct
Heater Duct Assembly
Valve Assembly
Clamp Bolt
Spacer
32.
33.
34.
35.
36.
37.
38.
39.
40.
41.
42.
43.
44.
45.
46.
Body Valve
Clamp
Air Box Assembly
Clamp
Air Duct
Baffle
Spring
Arm
Air Outlet Duct
Clamp
Adapter
Screen
Retainer
Air Outlet
Exhaust Vent Adapter
13-3
•
3
Detail
B
&
\'\
~~Ji~
27
~
/--
-
TBRU 337 -0978
Deta.1IC
•
",--
10
7
337-0979 AND ON
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
Valve Assembly
Right Band Heat Control
Spacer
Valve Body
Air Box Assembly
Baffle
Defroster Control
Left Band Heat Control
Defroster Outlet
Left Hand Air Mixture Control
Fresh Air Inlet Duct
~tgure 13-1.
13-4
NoN-TURBOCHARGED AIRCRAFT
HEATING SYSTEM( (1967 THRU 1970)
12. Air Duct
13. Beater Duct Assembly
14. Beater Control Assembly
15. RIght Hand Air Mixture Control
16. Spring
17. Washer
18 •. Clamp
19. Knob
20. Lever and Cam Assembly
21. Stiffener
22. Spring
Heating, Ventilating, and Defrosting Systems (Sheet 2 of 7)
•
•
13
DetailB
Detail
A
•
15
c
DetauD
NON-TURBOCHARGED AmCRAFT
HEATING SYSTEM (BEGINNING WITH 1971)
•
1.
2.
3.
4.
5.
6.
Clamp
Control Box
Knob
Lever and Cam Assembly
Stiffener
Spring
7. LH Air Control
8. RH Air Control
9. LH Heat Control
10. RH Heat Control
11. LH Defrost Control
12. RH Defrost Control.
13. Defrost Valve Outlet
Detaile
14.
15.
16.
17.
18.
19.
Control
Arm Assembly
Valve Assembly
Plenum Chamber
Link
Floor Heater Vent
Figure 13-1. Heating, Ventilating, and Defrosting Systems (Sheet 3 of 7)
13-5
13
Detail B
•
11
14
DetauD
337-0979
ANn ON
•
THRU
337-0978
Detail
C
•
Detail
Detail
A
20
E
9
TURBOCHARGED AIRCRAFT
HEATING SYSTEM
25
.1.
4..
3.
4.
5.
6.
7.
8.
9.
10.
Knob
Lever and Cam Assembly
Stiffener
Spring
Left Band Air Mixture Control
Right Band Air Mixture Control
Left Hand Defroster Control
Clamp
Temperature Control
Right Band Defroster Control
11 • Bracket
12.
13.
14.
15.
16.
17.
18.
19.
20.
Fuel Pump
Electrical Lead
Fuel Outlet Line
Fuel Regulator and
Shut-off Valve
Fuel Inlet Line
Heater SWitch Assembly
SWitch
Actuator
Valve Body
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
Reinforcement
Shim
Valve Plate
Spring
Clamp Bolt
Control Arm
Valve Seat Assembly
Air Mixture Control
Washer
Spring
Beater Start SWitch
Awe Cabin Beat Control
Figure 13-1. Heating, Ventilating, and Defrosting Systems (Sheet 4 of 7)
13-6
•
•
5
18
2
~10
Detail
1{
A
11
! f'..12
•
23
TURBOCHARGED AIRCRAFT HEATER DETAILS
Clamp
Support Assembly
Outlet Duct
Cabin Heat Control
Combustion Air Blower
Combustion Air Blower Outlet
7. Combustion Air Blower Inlet
8. Combustion Air Pressure Switch
1.
2.
3.
4.
5.
6.
•
9.
10.
11.
12.
13.
14.
15.
Fuel Inlet Line
Solenoid Valve
Nipple
Drain Line
Spark Plug Lead
Radio Noise Filte r
Ignition Assembly
16.
17.
18.
19.
20.
21.
22.
23.
Bracket
Combustion Air Blower Inlet
Fresh Air Adapter
Clamp
Inlet Duct
Exhaust Stack Extension
Shroud
Bracket
Figure 13-1. Heating, Ventilating, and Defrosting Systems (Sheet 5 of 7)
13-7
•
BEGINNING Wl'm
33701317 AND
F33700025
Detail
B
A
REFER TO SHEET 1 FOR
DETA.D..S OF SYSTEM
NOT SHOWN
NOTE
Beginning with Serials 33701317 and
F33700025, beads on inlets of upper
and lower housings (2) and (8) are deleted to facilitate Cycolac or RoyaUte
ducts and clamps. Seal around inlets
where hose and clamps attach with
579.6 sealer (Presstite Engineering
Co., St. Louis, Missouri), or equivalent sealer.
Detail
•
B
VENTILATING SYSTEM
11 18
1.
2.
3.
4.
5.
6.
7.
8.
9.
Bracket
Upper Housing
Valve Plate
Sealer
Plate Assembly
Bracket
Muffler Ring
Lower Housing
Escutcheon
NOTE
Tighten nut between dome
(12) and star washer (13)
securely, and cement to
plate (14) with an epoxy
base adhesive. Dome (12)
is sealed to body (23) at
final assembly with an
epoxy base adhesive.
10
10.
11.
12.
13.
14.
15.
16.
17.
18.
Knob
Spring
Dome
Star Washer
Plate
Seal
Cap Assembly
Washer
Shaft
19. Spacer
20. Insulator
21. Hose
22. Clamp
23. Body
24. Escutcheon
25. Setscrew
Figure 13-1. Heating, Ventilating and Defrosting Systems (Sheet 6 of 7.)
13-8
•
....
•
.-....................
......
......-
............. ...-. .....
....
......
••.••. 3
••••
. ..•••
Position tube (3) With
". ".
•••••••••• •••••••••••
drain on bottom.
". . •• ::.....
••••
.••••
••••••
•••••
............
.......
'
.....-
. ::::::::......
..•..
.
...........
....
. ........
.
::::............
.. ...........
...
,I'~~.---
. -..
'.
4
5
--,~-------,
,
.......... .
I
\
\
,
I
,
/'
,
I
I'
,I
"
I
-:. . . .
.......
'
2---~N':;;;
A
Detail
., .................,
.......; .....
....
·
·
•
:
.......
........
. , ....,
:r.....
.....
"
................
.'
." .
8
Detail
\
\ iJ'.··· .. ••••""
:/
.'
~
",
.......
.:
.'..... .'
.'.'
.'
.'
.'.'.'
'\'\"
:
B
.....
. . ~. -.. ··<'~L{Jb··········
•.•.•
A
THIS SYSTEM USED WITHOUT
OXYGEN SYSTEM INSTALLATION
l//:j
STANDARD SYSTEM INSTALLATION
'.Onu/OOOOI
OPTIONAL 5TH AND 6TH SEAT INSTALLATION
------'::---,
\
BEGINNING WITH 33701463 AND F33700056
\
\
I ,
I'
,,
(
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
•
Adapter
Scoop Assembly
Tube
Distributor Assembly
Adapter
Seal
Air Outlet Duct
Exhaust Valve Adapter
Air Outlet
Grommet
2-.....;..~('
........ -:
.....
7
THIS SYSTEM USED WITH
OXYGEN SYSTEM INSTALLATION
10
B
Figure 13-1. Heating, Ve!ltilating and Defrosting Systems (Sheet 7 of 7)
13-9
13-6. HEATING, VENTILATING AND DEFROSTING SYSTEM (TURBOCHARGED AffiCRAFT).
I
13-7. Ram air, routed through an elbow duct, connected to an opening in the left side of the forward
engine nose cap, is ducted through the heat exchange
section of a gasoline heater mounted in the left side
of the forward engine cowling, to the aircraft heating,
ventilating, and defrosting systems. A portion of
this ram air is routed through a small opening in the
aft part of the elbow duct, to a combustion air blower
mounted immediately above the heater. The combustion air blower supplies air to the combustion chamber of the heater in such a way that a whirling motion
is created. Fuel is routed from the fuel strainer in
the forward wheel well, through an electric fuel pump
on the forward side of the front firewall to a fuel solenoid regulator which regulates fuel pressure to 7 pSi.
Fuel from the regulator is routed to a spray nozzle
in the combustion chamber of the heater where the
fuel-air mixture is ignited by a spark plug. Electric
current for ignition is supplied by an ignition unit that
converts 24-volt current to a high-voltage, oscillating current, which provides a continuous spark.
Electric current is supplied when the heater switch
is turned from OFF to the START position momentarily, then allowed to return to the RUN position.
I~AUTIONI
Do not operate heater switch unless front engine is running. The heater is dependent upon
front-engine propeller slipstream pressure for
heater airflow during ground operation. Heater
operation in flight is independent of engine operation.
The stable, whirling flame sustains combustion under
the most adverse conditions because it is whirled
around itself many times. This type of flame is selfpiloting, and ignition is continuous. The burning
gases travel the full length of the combustion chamber, flow around the outside of the chamber, pass
through cross-over passages into an outer radiating
area, then travel the length of the assembly and out
the exhaust. This causes the ventilating air passing
through the heater to come in contact with two or
more heated cylindrical surfaces. An auxiliary heat
knob operates a butterfly valve in the combustion air
blower outlet. Pulling out the control knob partially
closes the butterfly valve and decreases the flow of
combustion air into the heater. As combustion airflow increases, a combustion air pressure switch
mounted on the combustion air inlet tube of the heater
closes, and actuates the ignition unit and solenoid
regulator valve. Fuel then flows through the regulator valve into the spray nozzle, which injects a conical spray of fuel into the combustion chamber where
the spark plug is already sparking; thus combustion
occurs. The temperature control rheostat knob actuates a duct switch mounted on the left heater duct,
on the aft side of the front firewall. The duct switch
acts as a thermostat which senses the heater air outlet temperature. As the heated air exceeds the thermostat setting, the thermostat automatically closes
the solenoid in the regulator valve, stopping fuel flow
into the heater. As the heater cools, the thermostat
opens the solenOid, allowing fuel to flow. Combustion takes place since the spark plug is continuously
sparking whenever the heater switch is turned from
the OFF position. By cycling, on and off, the heater
maintains an even air temperature in the cabin. The
heater is protected by an overheat switch mounted on
the heater jacket to sense the outlet temperature of
the ventilating airstream. Should this temperature
become too high, the overheat switch will automatically shut off the flow of fuel to the heater. When the
heater is turned off, unheated ram air passes through
the heater to the aircraft ventilating and defrosting
systems, described in paragraph 13-2.
•
13-8. TROUBLE SHOOTING. For trouble shooting of the heating, ventilating, and defrosting distribution system refer to paragraph 13-4. For
trouble shooting, maintenance, and overhaul of the
gasoline heater, refer to "Cessna Heater and Components Service/Parts Manual. "
13-9. REMOVAL AND INSTALLATION OF
COMPONENTS. Figure 13-1 illustrates the
parts of the heating, ventilating, and defrosting system' and may be used as a guide for replacement of
parts. Also refer to paragraph 13-5.
a. Remove the left engine cowling and nose cap.
b. Disconnect:
1. Electrical wires at terminal block.
2. Auxiliary air control (10) at combustion air
blower.
3. Drain line (11) at front of heater.
4. Exhaust tube (12) at lower aft end of heater.
5. Fuel line (24) at heater nozzle inlet.
c. Remove:
1. Inlet air hose (17) from combustion air
blower.
2. Outlet air hose (22) from combustion air
blower and heater inlet.
3. Fresh air adapter (18) from front of heater.
4. Four bolts securing combustion air blower
to support brackets and remove combustion air
blower.
5. Outlet duct (13) at aft end of heater.
6. Two clamps securing heater to support
assemblies.
d. Remove heater from airplane. Reverse the
preceding steps to install the heater.
•
•
13-10
•
FRONT ENOINE
VACUUM PUMP
OIL SEPARATOR
TO
...
SHUTTLE
DE-ICE
W
I
:.!i~~~
L
I
N
T
G
E
rEl B
RIGHT WING
•
WING ICE
G
DETECTOR
LIGHT
SHUTTLE
VACUUM
RELIEF
VALVE
OIL SEPARATOR
FROM VACUUM
INSTRUMENTS
REAR ENGINE
VACUUM PUMP
...
OIL RETURN
TO REAR ENGINE
ST ABILIZER DE-ICE BOOT
--
====CODE====
•
lIDO
PRESSURE LINES
THRU 33701462
VACUUM LINES
ALTERNATING PRESSURE
AND VACUUM LINES
Figure 13-2. De-Ice Schematic
(~t
1 ri 2)
13-11
PUMP
RH DE-ICE
DE-ICE BOOT
ACUUM
PRESSURE
•
RELIEF VALVE
OVERBOARD AIR
~a:t!ID.I\I~r:a:laza;{J •
OV ER BO AR D
EXHAUST VALVE
-..-+-----!-H+--FLOW CH ECK V ALV E
'-I-_ _ _-Ht~PRESSURE CONTROL VALVE
VACUUM
•
HORIZONTAL STABILIZER
BOOT
LINE CODE
NOTE
OPERA TING PRESSURE
IS 18 PSIG (NOMINAL)
PRESSURE
111111,
VACUUM
CD Pressure Control Valve set at 18= PSIG (Nom).
@
CD
BEGINNING WITH 33701463
Vacuum Relief Valve to be set at 5" HG.
Momentary actuation of control switch wtll
provide one 6-second de-icing cycle.
Figure 13-2. De-Ice Schematic (Sheet 2 of 2)
13-12
•
engine operation, if the vacuum relief valve
to the gyros is set too low, suction to the
gyros will "drop momentarily during the boot
inflation cycle. This suction variation can
be corrected with proper vacuum relief valve
adjustment. Check valves are included in the
standard vacuum system, so that the front
and rear systems will operate independently.
13-10. DE-ICE SYSTEM (Thru 33701462).
•
13-11. An optionalUght weight de-ice system may be
installed on the Models 337 and T337. De-icing of the
wing and horizontal stabilizer leading edge is accompUshed by inflation and deflation of rubber boots attached to these surfaces. The duration of each inflation and deflation cycle is controlled by valves
which in turn are controlled by an electronic timer.
13-12. DE-ICE SYSTEM OPERATION. An enginedriven vacuum pump is mounted on the top center of
each engine accessory housing and provides both
pressure and vacuum for the inflation and deflation
of the de-ice boots. Air from the outlet (pressure)
Side of the pump passes through an oil separator,
across the pressure relief valve, and overboard
when "the system is not operating. When the de-ice
switch is turned on, the timer closes the pressure
relief valve overboard line and directs the air from
the pressure side of the vacuum pump through a filter, shuttle valve, and into the de-ice boots for the
inflation cycle. Inflation time of the boots is approximately six seconds and the de-ice light on the
switch panel should be illuminated during the inflation cycle. At the completion of the inflation cycle,
the timer opens the pressure relief valve, returning
vacuum pump pressure overboard. Pressure in the
boots is returned through the system and overboard
through the pressure relief valve. When the shuttle
valve has less than one psi against it, it closes and
the vacuum side of the vacuum pumps holds the boots
in a deflated pOSition. The timer automatically repeats the cycle after a pause of apprOXimately 3
minutes to allow sufficient ice build-up for efficient
de-icing.
[~~UTION\
Always allow sufficient ice build-up for efficient ice removal before actuating the de-ice
system. H de-ice system is actuated continuously or before ice has reached sufficient
thickness, the ice will build up over the boots
instead of cracking off.
•
The de-ice system consists of two engine-driven
vacuum pumps with an oil separator, pressure relief
valve, air filter, and shuttle valve for each engine.
A pressure switch, timer, two boots on the leading
edge of each wing, and a boot on the leading edge of
the horizontal stabilizer complete the system. The
standard vacuum system components also serve the
de-ice vacuum system and the vacuum relief valve
adjustment should be maintained in the manner outlined in the Relief Valve Adjustment paragraph in
Section 14. The standard dry-type vacuum pumps
are replaced with oil-lubricated pumps. An ice detector light is incorporated in the left side of the
fuselage at the wing leading edge to aid checking for
ice formations during night operation.
NOTE
13-13. REMOVAL AND INSTALLATION OF DE-ICE
SYSTEM. For removal and installation of de-ice
system components refer to figures 13-3 through
13-6. Refer to figure 13-7 for ice, detector light.
The de-ice system will operate satisfactorily
on either or both engines. During single13-14. TROUBLE SHOOTING.
TROUBLE
PROBABLE CAUSE
DE-ICE BOOTS 00 NOT
INFLATE OR INFLATE
SLOWLY
•
REMEDY
Loose or faulty wiring.
Repair or replace wiring.
Loose or damaged hose.
Tighten or replace hose.
Loose or missing gasket.
Tighten fitting and/or replace
gasket.
Shuttle valve malfunction.
Replace shuttle valve.
Pressure relief valve set too low .
Reset or replace valve.
Pressure relief valve malfunction.
Replace pressure relief valve.
Defective timer.
Replace timer.
NOTE
With both vacuum pumps inoperative, this system will not operate.
13-13
•
SEE FIGURE 13-2 FOR
THE DE-ICE SYSTEM
SCHEMATIC
FIGURE 13-5
FOR EQUIPMENT
ON FIREWALL
SEE FIGURE 13-4
FOR EQUIPMENT
ON FIREWALL
'. ".
SEE FIGURE 13-7 FOR
THE ICE DETECTION
LIGHT
",
1. Wing De-Ice Boots
2. Stabilizer De-Icer Boot
3. Pressure Switch
4. Circuit Breaker Panel
5. De- Ice Switch
THRU 33'101462
Figure 13-3. De-Ice System (Sheet 1 of 2)
13-14
•
6. Pressure Indicator Light
'1. Timer
•
•
STABILIZER
DE-ICE BOOT
RH WING
DE-ICE BOOT
FRONT FIREWALL
DE-ICE EQUIPMENT
•
";'-A
{ /Ei!
~/
STALL STRIP
CIRCUIT BREAKER
PANEL
PRESSURE INDIcAToR LIGHT
•
DE-ICE
SWITCH
BEGINNING WITH 33701463
Figure 13-3. De-Ice System (Sheet 2 of 2)
13-15
---~
/
/
I
/
'" '"
/
5
I
/
4
4
I
I
I
\
8
~ssUlm~
TOOVERBO
•
2
9
10
1.
2.
Rubber Mount
er
~~parator
3. O-Ring
4.
5 Shuttie Valve
6· Bracket
7· Air FilterRelief Valve
S• Pressure
Pump
•
ngine Driven
L_l~o~._,:Tim=e=r
9. E
13-16
-:==-::--::n.t
___
THRU 33701462
Fi
Components (Sheet 1 of 2)
Figure 13 -4 . Front Firewall De-Ice
•
•
FRONT INSTALLATION
10
11
2
•
BEGINNING WITH 33701463
1. Vacuum Relief Valve
2. Elbow
3. Bushing
4. Union
•
5.
6.
7.
8.
Tee
Reducer
Check Valve
Timer
9, Control Valve
10. Bracket
11. Exhaust Valve
12. Dry Air Pump
Figure 13-4. Front Firewall De-Ice Components (Sheet 2 of 2)
13-17
•
~
I
\
----
-----
....
... """"--,
I
\~
'.............
~
-----..,
OIL RETURN
TOENGINE
•
4
3
,
,/
' ...........
/'
2
~
TO VAC. SYST.
REAR RELIEF
VALVE
I
~~
(//
.
~~
~VAC. ~
ALT.
TO BOOTS
~
ALT. PRES. & VAC.
FROM FRONT
SHUTTLE VALVE
PRESSURE LINE TO
OVERBOARD VENT
THRU 33701462
1. Pressure Relief Valve
2. O-Riltg
3. Air Filter
4. Bracket
5. Oll Separator
6. Engine Driven Pump
7. Spacer
8. Shuttle Valve
Figure 13-5. Rear Firewall De-Ice Components (Sheet 1 of 2)
13-18
•
•
7
REAR INSTALLATION
7
5~~---7"1
9
3
•
2
7
BEGINNING WI'MI 33701463
1. Pressure Control Valve
2. Bracket
3. Reducer
•
4. Tee
5. Check Valve
6. Tee Assembly
7. Elbow
8. Bracket
9. Pump
Figure 13-5. Rear Firewall De-Ice Components (Sheet 2 of 2)
13-19
13-14. TROUBLE SHOOTING (Cont).
TROUBLE
DE-ICE BOOTS DO NOT
DEFLATE OR DEFLATE
SLOWLY.
Pressure relief valve malfunction.
Replace pressure relief valve.
Shuttle valve malfunction.
Replace shuttle valve.
Defective timer.
Replace timer.
13-15. DE-ICE SYSTEM OPERATIONAL CHECK.
a. Electrical Test:
1. Turn WING DE-ICE Switch to off position.
2. Place master switch in on position.
3. Press WING DE -ICE indicator light to check
light circuit and bulb. Make sure dimming lens on
indicator is open.
4. Turn WING DE-ICE switch on and repeat
step 3.
5. If indicator light does not function in steps 3
and 4, the circuit breaker may have opened. Check
for short in the system. Reset circuit breaker and
repeat step 3.
b. Air Leakage Test:
1. This test can be performed in either the front
or rear engine compartments.
2. Disconnect pressure hose from pressure relief valve inlet port.
3. Disconnect vent tube from overboard port,
and cap port.
4. Connect a source of clean air to the pressure
relief valve inlet port. It is necessary that the inlet
pressure be a minimum of 18-20 psi to perform this
test. Include a pressure gage in the air line to observe the system pressures.
5. Apply 18 psi pressure to the system and, by
means of a hand-operated valve, trap the pressure in
the de-ice system. Observe the system for leakage.
The leakage rate should not exceed a pressure arop
of 4.0 psi per minute.
6. If the leakage exceeds 4.0 psi per minute,
use a soap and water solution to locate leaks. Tighten
connections as required.
7. To check the pressure switch, place master
switch on while de-ice system is pressurized. The
indicator light should illuminate.
8. Remove test equipment, lubricate all threads
and connect all system components disconnected.
c. Vacuum Relief Valve Adjustment and System Tes
1. Adjust vacuum relief valve as outlined in
paragraph 14-21.
2. With vacuum relief valve adjusted and one
engine operating at 2400 rpm, place WING DE-ICE
switch to on pOSition and observe de-ice system op_
eration. System is functioning satisfactorily if the
Wn.lG DE-ICE indicator light illuminates within 4.0
seconds after turning WING DE-ICE switch on.
3. Repeat the above procedure for the other
engine.
d. Timer Cycle Check:
1. With engines operating at 2100 rpm, place
WING DE-ICE switch to on position. As soon as deice boots inflate, reduce engine speed to normal idle
for approximately 2 1/2 minutes. This permits timer
13-20
REMEDY
PROBABLE CAUSE
•
to complete its cycle. At the end of the 2 1/2 minute
idle period, increase engine speed to 2100 rpm and
observe de-ice boots for inflation. Elapsed time
from inflation to inflation should be approximately
3 minutes.
2. If it appears that the timer is defective,
apply 28 vdc to pins tI 1 and #2 and listen for action
of stepping switch.
I~AUTIONl
The negative ground must be applied to pin
*1; pin #2 is positive. A reverse voltage
will ruin the timer diode. The 28 vdc must
be filtered if it is rectified from ac; a battery should be used.
13-16. DE-ICE SYSTEM (Beginning with 337-1463).
13-17. DESCRIPTION. Air pressure and vaccum required for operation of the p~eumatic de-icing system
are provided by engine-driven pumps. Va~uum fro:n
the pu:nps is routed to a va~uum manifold which supplies the instruments and through the exhaust valve to
the de-icers .. Pressure from the pumps is routed to
flow check valves, then through the pressure manifold
to the de-icers. A pressure control valve, located
on a tap between the pump and the flow check valve in
each engine co:npartment, regulates the pump output
pressure. Control of the operation of the pressure
co~trol valves and exhaust valve is provided through a
time-delay relay. A pressure switch, located on a
tap off the pressure line between the flow check valves
and the de-icers, is used in conjunction with a light on
'he instrument panel to indicate that all de-icers are
being inflated. During non-operating periods, vacuum
is applied to the de -icers through the exhaust valve
while the pressure control valves relieve the pressure
produced by the pumps. Operation of the de-icers is
initiated through a control switch which activates the
time-delay relay. The time-delay relay provides
power to the solenoids of the pressure control valves
and the exhaust valve. When energized, the pressure
control valves regulate the pump output air to de-icer
system vacuum. After six seconds, the time-delay
relay shuts off the power to the solenoids of the pressure con trol valves and the exhaust valve. Then,
pressure in the de-icers is released through an integral pressure relief section of the exhaust valve and
vacuum is reapplied.
13-18. COMPONENT DESCRIPTION.
a. Pneumatic De-Ice Boot:
The de-icer consists of a smooth rubber and fabric
•
•
•
T337 -0750 THRU 33701462
NOTE
Inspect screen (12) and
washer (11) in fitting
(13) at each 50- hour
inspection.
A
•
•
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
Oil Separator
Spacer
Rubber Mount
Plate
Bolt
Firewall Bracket
Extension Bracket
Washer
Nut
Cotter Pin
Washer
Screen
Fitting
Figure 13-6. Oil Separator Installation
blanket containing small spanwise de-icing tubes.
All tubes in each de-Icer are simultaneously inflated
through a Single air connection. The de-icer is cemented to the airfoil leading edge. When the system
is "OFF", vacuum is applied to the de-icer tubes.
This is necessary to resist negative aerodynamic
pressures and to maintain the tubes in a flat or deflated condition. When icing conditions are encountered, it is recommended that at least 1/4" of ice be
allowed to accumulate before the de-ice system is
operated; however, the de-icer will effectively remove both thicker and thinner ice accumulations.
b. Dry Air Pump:
An air pump, mounted on the accessory pad of each
eng{ ne, provides positive pressure and vacuum for
the de-icing system.
c. Flow Check Valve:
This valve controls the flow of operating air to the
de-icers:- The valve will open at a predetermined
pressure and remain open during the time the deicers are being pressurized. At the end of the deicing cycle, when the pressure is relieved, the valve
will close automatically to function as a vacuum
check valve.
d. Pressure Control Valve:
This valve, located on a tap between the pump and
flow check valve, regulates the pump output pressure.
e. Exhaust Valve:
This valve is located on a tap off the vacuum manifold
It provides the vacuum necessary to maintain the de-
icing tubes in a denated condition, resisting negative
aerodynamiC pressures. When the de-icer system is
"ON", the exhaust valve solenoid is energized, cloSing the vacuum port. After the de-icing cycle, pressurized air within the de-icers is released through
an integral pressure relief section of the exhaust
valve and vacuum is reapplied.
13-19. SYSTEM OPERATION. Refer to paragraph
13-17.
13-20. REMOVAL AND INSTALLATION OF DE-ICE
SYSTEM. For removal and installation of de-ice system components, refer to figures 13 -3 through 13 -6.
13-21. DE-ICE BOOT REPAm (COLD PATCH).
13-22. DESCRIPTION. There are four types of damage that are most common to the de-icer boots. The
following procedures describe the damage and outline
techniques for the repair.
13-23. REPAm.
13-21
•
ICE DETECTOR LIGHT
CIRCUIT BREAKER PANEL
•
Figure 13-7. Ice Detector Light
NOTE
When repairing the de-ice boots and replacement layers are being installed, exercise care to prevent trapping air beneath
the replacement layers. If air blisters appear after material is applied, remove them
wi tit a hypoder rnic needle.
Scuffed or Iamaged Surface:
This type of damage is the most commonly encountered and is usually caused by scuffing the outer surface of the de-ice boots while using scaffolds, refueling hose, ladders, etc. Repair is generally not
necessary because the thick outer veneer provides
protection to the natural rubber underneath. If the
damage is severe and bas caused removal of the
entire thickness of veneer (exposing the brown
natural rubber underneath), the damage should be
repaired as follows:
•
a. Select a patch (B. F. Goodrich Part Number
3306-1, 3306-2, or 3306-3) large enough to cover
the damaged area.
b. Using a clean cloth dampened with solvent,
thoroughly clean the damaged area.
c. Buff the area around the damage with steel
wool so that the area is moderately but completely
roughened.
d. Wipe the buffed area clean with a cloth slightly
13-22
dampened with solvent to remove all loose particles.
e. Apply one even thorough coat of EC-1403 (Minnesota Mining and Manufacturing Co. ) cement to the
patch and corresponding damaged area of the de-ice
boot and allow cement to dry completely.
f. Reactivate cemented surfaces with solvent.
Apply patch to the de-ice boot with an edge or the
center adhering first, and work the remainder of the
patch down, being careful to avoid air pockets between patch and boot.
g. Roll the patch thoroughly with a stitcher-roller
(Part Number 3306-10) and allow to set for 10 to 15
minutes.
h. Wipe the patch and surrounding area, from the
center of the patch outward, with a cloth slightly
dampened with solvent.
1. Apply one light coat of A-56-B conductive
cement (Part Number 3306-13) to the patched area
to restore conductivity.
NOTE
Satisfactory adhesion should be obtained in
four hours; however, if the patch is allowed
to cure for a minimum of 20 minutes, the deice boots may be inflated to check the repair.
Damage to Tube Area:
This type of damage consists of cuts, tears, or
•
•
ruptures to the inflatable tube area and a fapric reinforced patch must be used for this repair. Damage
to the tube area should be repaired as follows:
a. Select a patch (B. F. Goodrich Part Number
3306-4, 3306-5, or 3306-6) of ample size to extend
at least 5/8-inch beyond the damage area.
NOTE
If none of these patches are of proper Size,
one may be cut to the size desired from one
of the larger patches. If this is done, the
edge should be beveled by cutting with the
shears at an angle. These patches are
manufactured so they will stretch in one
direction only. Be sure to cut patch selected
so that the stretch is in the widthwise direction of the inflatable tubes.
•
b. Using a clean cloth dampened with solvent,
thoroughly clean the area to be repaired.
c. Buff the area around the damage with steel
wool so that the area is moderately but completely
roughened.
d. Wipe the buffed area clean with a cloth slightly
dampened with solvent to remove all loose particles.
e. Apply one even thorough coat of EC-1403 (Minnesota Mining and Manufacturing Co.) cement to the
patch and the corresponding damaged area of the
de-ice boot. Allow cement to dry completely.
f. Reactivate cemented surfaces with solvent.
Apply patch to de-ice boot with the stretch in the
widthwise direction of the inflatable tubes, sticking
edge of patch in place first and working remainder
down with a very slight pulling action so the injury
is closed. Use care to avoid air pockets between
patch and de-ice boot surface.
g. Roll the patch thoroughly with a stitcher-roller
(Part Number 3306-10) and allow to set for 10 to 15
minutes.
h. Wipe the patch and surrounding area, from the
center of the patch outward, with a cloth slightly
dampened with solvent.
i. Apply one light coat of A-56-B conductive
cement (Part Number 3306-13) to restore conductivity.
NOTE
Satisfactory adheSion of patch to de-ice boot
should be reached in four hours; however,
if the patch is allowed to cure for a minimum
of 20 minutes, the de-ice boots may be inflated to check the repair.
Damage to Fillet Area:
•
This includes any tears or cuts to the tapered area
aft of the inflatable tubes. Damage to the fillet area
should be repaired as followS:
a. Trim damaged area square and remove excess
material. Cut must be sharp and clean to permit a
good butt joint of the inlay.
b. Cut an inlay from tapered fillet (B. F. Goodrich
Part Number 3306-7) to match cutout area.
c. Using solvent, loosen edges of de-ice boot
around cutout area approximately 1 1/2 inches from
all edges.
d. Thoroughly "clean the area to be repaired, using
a cloth dampened with solvent.
e. Lift edges of loosened boot around cutout, and
apply one coat of EC-1403 (Minnesota Mining and
Manufacturing Co.) cement to underneath side of
boot.
f. Apply one coat of EC-1403 cement to the wing
skin underneath the loosened edges of de-ice boot,
allowing cement to extend 1-1/2 inches beyond edges
of boot into cutout area.
g. Apply a second coat of EC-1403 cement to underneath side of boot as outlined in step "e. "
h. Apply one coat of EC-1403 cement to one side of
a 2-inch Wide, neoprene-coated fabric tape (Part
Number 3306-8) and allow cement to dry. Trim the
tape to size of cutout. This tape is necessary to reinforce splice.
1. Reactivate cemented surface of tape and wing
skin with solvent and apply tape to wing skin. Use
care to center tape under all edges of cutout.
j. Roll down tape on wing skin with stitcher-roller
(Part Number 3306-10) to assure good adheSion,
being careful to avoid air pockets between tape and
wing skin.
k. Apply one coat of EC-1403 cement to top surface
of tape and allow cement to dry approximately 5 to 10
minutes.
1. Reactivate cemented surfaces of boot wing skin
and tape with solvent. Working toward the cutout,
roll down carefully the edges of the loosened boot
to prevent trapping air. The boot edges should overlap the tape approximately 1 inch.
m. Roughen back surface of inlay repair material
(Part Number 3306-7, previously cut to size) with
steel wool. Thoroughly clean with solvent and apply
one coat of EC-1403 cement.
n. Apply one coat of EC-1403 cement to wing skin
inside cutout area and allow to dry.
o. Apply the second coat of EC-1403 cement to inlay repair mater"ial and allow to dry.
p. Reactivate cemented surfaces with solvent and
carefully insert inlay material with feathered edge
of inlay aft. Working from forward edge aft, carefully roll down the inlay to avoid trapping air.
q. Rooghen area on outer surface of de-ice boot
and inlay with steel wool 1-1/2 inch on either side
of splice. Clean with solvent and apply one coat
of EC-1403 cement.
r. Apply one coat of EC-1403 cement to one side
of 2-inch wide, neoprene-coated fabric tape (Part
Number 3306-8), trim to Size, and center tape
over splice on three sides.
s. Roll down tape on de-ice boot and inlay with
stitcher-roller (Part Number 3306-10) to assure
good adhesion, being careful to avoid trapping air.
t. Apply one light coat of A-56-B conductive cement (Part Number 3306-13) to restore conductivity.
Veneer Loose From De-Ice:
If the veneer should become loose from the de-ice
boot, repair should be made as follows:
a. Peel and trim the loose veneer to a point where
the adheSion of veneer to the de-ice boot is good.
b. Roughen area in which veneer is removed with
13-23
steel wool. Motion must be paralled to cut edge of
veneer ply, to prevent loosening it.
c. Taper edges of veneer down to the tan rubber
ply by rubbing parallel to cut edge of veneer with
steel wool and solvent.
d. Cut a piece of veneer material (Part Number
3306-9) large enough to cover the damaged area and
extend at least 1 inch beyond in all directions.
e. Mask off the damaged boot area 1/2-inch larger
in width and length than the patch.
f. Apply one coat of EC-1403 cement to damaged
boot area and allow to dry.
g. Apply second coat of EC-1403 cement to damaged boot area and allow to dry.
h. Reactivate cement surface with solvent. Peel
the backing from the veneer, and for 6 inches of its
length, and roll the veneer to the boot with a 2-inch
roller. Roll edges with stitcher-roller (Part Number
3306-10).
i. Continue stripping the backing from the veneer
as the rolling progresses, applying a slight tenSion
on the veneer ply to prevent wrinkling.
j. Be careful to prevent trapping air. Hair blisters appear after veneer is applied, remove them
with a hypodermic needle.
k. Wipe the patch and surrounding area, from the
center of the patch outward, with a cloth slightly
dampened with solvent.
1. Apply one light coat of A-56-B conductive cement (Part Number 3306-13) to restore conductivity.
NOTE
B. F. Goodrich "cold patch" Repair Kit No.
74-451C for surface ply de-ice boot repair
is available from the Cessna Service Parts
Center.
13-24. REPLACEMENT OF DE-ICE BOOTS. To
remove or loosen installed de-ice boots, use toluol
or toluene 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. Removal
should be slow enough to allow the solvent to undercut the cement so that parts will not be damaged. To
install a wing de-tcer boot, proceed as follows:
a. Clean the metal surfaces and the bottom Side of
the de-icer thoroughly with Methyl Ethyl Ketone or
Methyl Isobutal Ketone. This shall be done by wiping the surfaces with a clean, lint-free rag soaked
with the solvent and then wiping dry with a clean,
dry, lint-free rag before the solvent has time to dry.
b. Place one inch masking tape on wing to mask
off boot area allowing 1/2 inch margin. Take care
to mask accurately so that clean-up time will be
reduced.
c. Stir EC-1300L (EC-1403) cement thoroughly
before using. Apply one even brush coat to the metal
and to the rough side of the boot, brushing well into
the rubber. Allow cement to air dry for a minimum
of 30 minutes and then apply a second coat to each of
the surfaces. Allow at least 30 minutes, preferably
one hour, for drying.
d. Snap a chalk line along the leading edge line of
the wing and a corresponding line on the inside of the
de-icer if it dues not have a centerline. Securely
attach hoses to the deicer connections. Position tne
13-24
centerline of the boot with the leading edge of the
wing, and using a clean, 'lint-free cloth heavily
moistened with toluol, reactivate the surface of the
cement on the wing and the boot in small spanwise
areas about six inches wide. Avoid excess rubbing
of the cement, which would remove it from the surface. Have enough help to hold boot in a vertical
plane. Place the chalk lines in alignment, and
starting at one end of the boot, tack it to the wing
along the leading edge line. Hold the rest of the
boot clear of the wing. Roll along the leading edge
line with a rubber roller, and an inch or two on
either side. Taking approximately six inches of
chord at a time, roll from the leading edge aft in
firm, overlapping, chordwise strokes of the rubber
roller until the entire boot is in contact with the
airfoil. It is important that all air be removed from
between the rubber and the metal, and that the boots
be distorted to a minimum amount. H any air is
trapped between the rubber and the metal, it may be
removed by the careful use of a small hypodermic
needle, except in the tube area. Use the metal
stitcher roller around the edges of the boot and
connections. Fill any gaps between adjoining boots
with EC -539 sealer. Apply a coat of EC -539 sealer
along the traillng edges of the boot to the surface of
the skin to form a neat straight fillet.
Remove masking tape and clean surfaces with toluol.
e. When installing the large inboard boot, it will be
helpful to place a clean, lint-free, folded liner of
canvas on the top of the wing, back of the leading
edge with the fold forward. The boot can be laid on
top of the liner and the liner pulled back about six
inches at a time as the rolling progresses aft. The
bottom portion of the boot will, of course, hang free
of the wing, preventing premature contact. This
should be done in a manner to align the rear edges
of adjoining boots and the carpenter's chalk line
should be used for this purpose. Trim butting edges
of adjoining boots to keep gaps to a minimum. H
gaps result, they may be filled with EC-539 sealer.
Apply a coat of EC-539 sealer along the trailing edges of the boot to the surface of the skin to form a
neat straight fillet.
f. Remove masking tapes and clean edge surfaces
with toluol.
13-25. PROPELLER DE-ICE SYSTEM. The system
is of an electrothermal type, consisting of electrically heated de-icers bonded to each propeller blade, a
slip ring assembly for power distribution to the propeller de-icers, a brush block assembly to transfer
electrical power to the rotating Slip ring, a timer to
cycle electric power to the de-icers in proper sequence, an ammeter, mounted in the instrument
panel, a switch and a circuit breaker. The de-ice
system applies heat to the surfaces of the propeller
blades where ice normally would adhere. This heat,
plus centrifugal force and the blast from the airstream, removes accumulated ice. Each de-icer
has two separate electrothermal heating elements,
an inboard and an outboard section. When the switch
is turned on, the timer provides power through the
brush block and slip ring to outboard elements for
approximately 34 seconds, reducing ice adhesion in
these areas. Then, the timer switches power to inboard heating elements for apprOximately 34 seconds.
•
•
•
•
Lockwashers (17) and nat washers (16) are
used as required to align plane of slip ring
perpendicular to engine crankshaft Within a
total devlation of .005 inches, with .002 inch
deviation within any four inches at circumference of slip ring. Check with diallndicator.
6
II
•
15
1. Switch
2. Circuit Breaker Panel
•
3.
4.
5.
6.
7.
B.
9.
10.
11.
12.
Brush Block Assembly 13.
Timer
14.
Ammeter
15.
Bolt
16.
Engine Crankshaft
17.
Slip Ring
lB.
Spinner Bulkhead
19.
Clip Assembly
Terminal Block
Boot Strap
Boot
Propeller
Spinner
Washer
Lockwasher
Spacer
Engine Crankcase
13
NOTE
Torque bolts (6) 90 to 110 lb-in.
Figure 13-B. Propeller Anti-Ice System (Sheet 1 of 2)
13-25
*
De-icer boot lead strap must be positioned above the
safety wire. Do not sandwich between safety wire
and balance weights or hub.
•
Install restrainer strap (3) in same manner
as boot installation outlined in steps "a" thru
"n" of paragraph 13-30. Start restrainer
strap approximately 90° from de-icer boot
lead (point a). Wrap around propeller
blade so that a double thickness will cover
de-ice boot lead strap. Trim restrainer
strap so it will end a approximately pOint{b) .
.----.50, +.06 -.06
View
•
AA
I
EC-539
SEALANT
FAmED IN
J
~L....-..---~-.JA
...,<
~t~.
View
BB
1.
2.
3.
4.
Boot
Propeller
Restrainer Strap
Boot Lead Strap
BEGINNING WITH 1971 \
THREADLESS RETEN'"
TION PROPELLER.
4
(REMAINDER OF COMPONENTS REMAIN THE
SAME.)
Figure 13-8. Propeller Anti-Ice System (Sheet 2 of 2)
13-26
•
•
cated on the gauge.
c. Check that total rWl-out does not exceed 0.005
inch (±0.0025 inch) for the Model 337, or 0.008 inch
(±0.004 inch) for the Model T337. Also check that
run-out does not exceed 0.002 inch within any 4
inches of slip ring travel for either type of engine.
It then returns to the outer elements and continues
cycling action. This outboard-inboard sequence is
important since the loosened ice teods to move outboard. Heating may begin at any phase in the cycle,
depending on the timer position when the switch was
turned off from previous use. Ground checkout of
the system is permitted with the engine not running.
System components may be removed and replaced,
using figure 13-8 as a guide. Propeller removal
is necessary before de-ice system comPonents, except brush block assembly, can be installed or removed.
@~~r~~~\
Due to the loose fit of some propeller bearings, a considerable error may be indicated
in the readings by pushing in or pulling out
on the propeller while rotating it. Care must
be taken to exert a uniform push or pull on
the propeller to hold this error to a minimum.
13-26. SLIP RING ALIGNMENT. After installation,
the Slip ring assembly must be checked for run-out,
and adjustments made, if necessary.
d. U slip ring run-out is within the limits specified, no corrective action is required. A small
amount of run-out may be corrected by varying the
torque of the attachment bolts within the limits specified by the propeller manufacturer.
e. U the procedure outlined in step "d" does not
produce acceptable run-out, fabricate small washel"shaped shims (apprOximately .010 inch), and place
on attachment bolts, limit one washer per bolt, between slip ring and spinner bulkhead or mounting
plate.
f. Recheck run-out. Adjust shim thickness and
vary torque of attachment bolts unW Slip ring runs
true Within the prescribed tolerance.
NOTE
Excessive slip ring run-out will result in
severe arcing between the Slip ring and
brushes, and cause rapid brush wear. U
allowed to persist, this condition Will result in rapid deterioration of the Slip ring
and brush contact surfaces, and lead to
the eventual failure of the De-Icing System.
•
a. Securely attach dial indicator gauge to the engine, and place the pointer on the slip ring.
b. Rotate propeller slowly by band, noting the
deviation of the slip ring from a true plane as indi13-27. TROUBLE SHOOTING.
NOTE
The propeller anti-ice ammeter may be used while trouble shooting the system. The ammeter needle
should rest within the shaded band except for "flickers" approximately 34 seconds apart, as the step
switch of the timer operates. The ammeter will also reflect a bad connection or open circuit by reading below normal or zero. A high reading indicates a short circuit.
TROUBLE
ELEMENTS DO NOT HEAT.
SOME ELEMENTS DO NOT
HEAT.
•
PROBABLE CAUSE
REMEDY
Circuit breaker out or defective.
Reset circuit breaker. If it pops out
again, determine cause and correct.
Replace defective parts.
Defective wiring.
Repair or replace Wiring.
Defective switch.
Replace switch.
Defective timer.
Replace timer.
Defective brush-to-slip ring
connection.
Check alignment. Replace defective
parts.
Incorrect wiring.
Correct wiring.
Defective wiring.
Repair or replace Wiring.
Defective timer.
Replace timer •
Defective brush-to-slip ring
connection.
Check alignment. Replace defective
parts.
13-27
•
13-27. TROUBLE SHOOTING (Cont).
PROBABLE CAUSE
TROUBLE
REMEDY
CYCLING SEQUENCE NOT
CORRECT OR NO CYCLING.
Crossed connections.
Correct wiring.
Defective timer.
Replace timer.
RAPID BRUSH WEAR,
FREQUENT BREAKAGE,
SCREECHING OR
CHA TTERING.
Brush block or slip ring out of
alignment.
Align properly.
input pins. (Refer to chart follOWing this step for pin
identification. )
13-28. TIMER TEST.
a. Remove connector plug of wire harness from timer
and jump power input socket of wire harness to timer
Timer PIN
Power Input
Pin & Socket
Ground Pin
Output Sequence,
Time, Voltage
3E1540-1
B (14 VDC)
A (14 VDC)
C, D - 34 sec.
each, then
repeats (14 VDC)
b. Jump bmer ground pm to ground.
c. Turn on De-Icing System.
d. Check timer operation per the chart preceding
step "b. " (Use a voltmeter.)
.
e. Check volts to ground in each case. If engine is
not running, and auxiliary power is not used, voltage
will be battery voltage and cycle time may be slightly
longer than indicated.
f. Hold voltmeter probe on the pin until the voltage
drops to O. Move the probe to the next pin in the sequence shown in the chart. Check voltage at each pin
in sequence. When correctness of the cycling sequence
is established, turn propeller De-ICing switch off at
the beginning of one of the on-time periOds, and record the letter of the pin at which the voltage supply
is present.
NOTE
Timers do not home to pin "C" when turned off.
13-29. INSTALLATION AND ALIGNMENT OF
BRUSH BLOCK ASSEMBLY. (Refer to Figure 13-9.)
NOTE
Installation of the brush block should be deferred, when pOSSible, until after the Slip
ring, propeller, and related components
are installed. However, the brush block
assembly may be replaced without removing the propeller. To avoid breakage when
installing the brush block assembly, keep
brushes retracted in brush block until Slip
ring and propeller assemblies have been
installed.
Total Repeat
Cycle Time
(minutes)
1.1
[~AUTIO~I
Make sure that slip ring run-out has been
corrected before attempting to align brushes
on slip ring.·
a. In order to get smooth, efficient and quiet transfer of electric power from the brushes to the slip
ring, brush alignment must be checked and adjusted,
if necessary to meet the following requirements.
1. Projection must be such that the distance
between the brush block and the Slip ring is • 06" :!:
.03" •
2. The brushes must be lined up with the Slip
ring so that the entire face of each brush is in conD
tact With the slip ring throughout the full 360 of
slip ring rotation.
3. The brushes must contact the Slip ring at an
angle of apprOximately 2 D from perpendicular to the
slip ring surface, measured toward the direction of
rotation of the slip ring.
b. Brush projection can normally be adjusted by
loosening hardware attaching the brush block and
holding the brushes in the desired location while retightening the hardware. Slotted holes are provided.
c. One method for face alignment is described in
step "b". Another is to use shims between brush block
and bracket. Laminated metal shims are generally
provided. Layers of metal • 003" are used to make
up shims which are approximately O. 20" thick overall.
Shims may be fabricated locally.
d. Loosen mounting bolts and twist block while
tightening to attain proper angular adjustment.
I$AUTIONI
Use care not to disturb other adjustments
when adjusting angular alignment.
13-28
•
•
•
SLIP RING ASSEMBLY
I
2°~~
1/16
:I:
I~
1/32-,
f
'\
PROPELLER ROTATION
~o
II
BRUSH BLOCK A$EMBLY
PROJECTION AND ANGULAR BRUSH ALIGNMENT
INCORRECT
CORRECT
G:II"'UIIIIIIIH~
INCORRECT
BRUSH FACE ALIGNMENT
Figure 13-9. Brush Face Alignment and Projection and Angular Brush Alignment
•
•
13-30. REPLACEMENT OF DE-ICE BOOTS. To
remove or loosen installed de-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. Removal should be
slow enough to allow the solvent to undercut the cement so that parts will not be damaged. To install a
propeller anti-ice boot, proceed as follows:
a. Clean the metal to be bonded with Methyl Ethyl
Ketone, (MEK). For final cleaning, wipe the solvent
film off quickly with a clean, dry cloth before it has
time to dry.
b. Prepare a pattern the size of the boot, including
three inches of the boot strap. Draw a centerline
(lengthwise) through the pattern.
c. Draw a line on the centerline of the leading edge
of the blade. Position the pattern centerline over the
leading edge centerline. Position pattern so bottom
of boot is 1/2" below spinner cutout. Draw a line on
the propeller hub on each side of the pattern boot
strap where it crosses the hub. Check boot strap
poaition by fitting restraining strap on the hub and
comparing its position with the marked position of the
strap.
d. Mask off an area 1/2" from each Side and outer
end of the pattern, and remove the pattern.
e. Mix EC-1300L cement (Minnesota Mining & Mfg.
Co.) thoroughly and apply one even coat to the cleaned metal surface. Allow to dry for a minimum of
one hour, then apply a second coat of the cement.
f. Moisten a clean cloth with Methyl Ethyl Ketone
and clean the unglazed back surface of the boot,
changing cloths frequently to avoid contamination of
the cleaned area.
~. Apply one even coat of EC-1300L cement to back
surface of boot. It is not necessary to cement more
than 1/2" of the boot strap.
h. Using a silver-colored pencil, mark a centerline
along the leading edge of the propeller blade and a
corresponding centerline on the cemented side of the
boot.
i. Reactivate the surface of the cement using a
clean, link-free cloth, heavily mOistened with toluol.
AVOid excessive tubbing of cement, which would remove the cement.
j. Position the boot centerline on the propeller
leading edge, starting at the hub end at the position
marked. Make sure that boot strap will fall in the
position marked. Tack the boot centerline to the
leading edge of the propeller blade. If the boot is
allowed to get off -center, pull up with a quick motion
and replace properly. Roll firmly along centerline
with a rubQer roller.
k. Gradually tilting the roller, work the boot carefully over either side of the blade contour to avoid
trapping air in pockets.
1. Roll outwardly from the centerline to the edges.
If excess material at the edges tends to form wrinkles, work them out smoothly and carefully with
fingers.
m. Apply one even coat of EC-539 (Minnesota Mining & Mfg. Co.), mixed per manufacturer's instructions, around the edges of the installed boot.
n. Remove masking tape from the propeller and
clean the surface of the propeller by wiping with a
clean cloth dampened with toluol.
o. Install restraining strap in accordance with figure
13 -8, and secure with screws, washers and sleeves.
13-31. OXYGEN SYSTEM.
13-29
'WARNING'
Under NO circumstances should the ON-OFF
control on the oxygen regulator be turned to
the "ON" position with the outlet (low pressure) ports open to atmosphere. Operation
of these units in this manner will induce serious damage to the regulators and have the
following results:
1. Loss of outlet set pressure.
2. Loss of oxygen now through the regulator which will result in inadequate oxygen being fed
through the aircraft system.
3. Internal leakage of oxygen through regulator.
Opening of the control lever with the outlet ports
open to atmosphere, results in an "overshoot" of
the regulator metering device due to the extreme
flow demand through the regulator. After overshooting, the metering poppet device goes into oscillation,
creating serious damage to the poppet seat and diaphragm metering probe. This condition can occur
even by turning the control lever on and then turning
it quickly off.
A potential hazard exists to aircraft in the field
where inexperienced personnel might remove the
cylinder and regulator assembly from the aircraft
and for some reason, attempt to turn the regulator
to the "ON" position With the outlet ports open. Unfortunately, after the units have been improperly
operated as noted, there is no outward appearance
indicating that damage has occUrred.
Testing these regulators should be accomplished only
after installation in the aircraft, with the "downstream" low pressure line attached.
13-32. DESCRIPTION. The system is comprised of
two oxygen cylinders, a pressure regulator, filler
valve, pressure gage, pressure lines, outlet and
mask assemblies. The oxygen cylinders are mounted
in the cabin top area. Locations of system components
are shown in figure 13-10. The pilot's supply line is
designed to receive a greater flow of oxygen than the
passengers. The pilot's mask is equipped With a microphone, keyed by a switch button on the pilot's control wheel. The filler valve is located in the leading
edge of the right Wing.
,--W~A'="R:-:N~I::":'N::-:::G::"'II'
011, grease or other lubricants in contact
with high-pressure oxygen, create a serious fire hazard and such contact should be
avoided. Do not permit smoking or open
flame in or near aircraft while work is performed on oxygen systems.
13-33. MAINTENANCE PRECAUTIONS.
a. Working area, tools and hands must be clean.
b. Keep oil, grease, water, dirt, dust and all
other foreign matter from system.
13-30
c. Keep all lines dry and capped unW installed.
d. Use only MIL-T-5M2 thread compound or teflon
lubricating tape on threads of oxygen valves, tubing
connectors, fittings, parts of assemblies which might
under any conditiOns, come in contact with oxygen.
The thread compound must be applied sparingly and
carefully to only the first three threads of the male
fitting. No compound shall be used on alumirmm
flared fittings or on the coupling sleeves or on the
outside of the tube flares. The teflon tape shall be
used in accordance with the instructions listed following this step. Extreme care must be exercised
to prevent the contamination of the thread compound
or teflon tape with oil, grease or other lubricant.
1. Lay tape on threads close to end of fitting.
Clockwise on standard threads, opposite on
left hand threads.
2. Apply enough tension while winding so tape
forms into thread grooves.
3. After wrap is complete, maintain tension
and tear tape by pulling apart in direction
it was applied. Resulting ragged end is the
key to the tape staying in place. (If sheared
or cut, tape may unwind. )
4. Press tape well into threads.
5. Make connections.
e. Fabrication of oxygen pressure lines is not
recommended. Lines should be replaced by part
numbers called out in the aircraft Parts Catalog.
f. Lines and fittings must be clean and dry. One
of the following methods may be used.
1. Clean by degreasing with stabllized trichlorethylene, conforming to Federal Specifications
0-T-634 or MIL-T-27602. These items can be obtained from American Mineral Spirits of Houston,
Texas.
•
•
NOTE
Most air compressors are 011 lubricated
and a minute amount of oil may be carri;d
by the airstream. If only an oU lubricated
air compressor is available, drying must
be accomplished by heating at a tempera-
ture of 250 0 to 300°F for a suitable period.
NOTE
Cap lines at both ends immediately after
drying to prevent contamination.
13-34. REPLACEMENT OF COMPONENTS. Removal, disassembly, assembly and installation of
system components may be accomplished while using
figure 13-10 as a guide.
The pressure regulator, pressure gage and
line and filler valve should be removed and
replaced only by personnel famlUar with
high-pressure fittings. Observe the maintenance precautions listed in the preceding
paragraph.
•
•
7
QUICK- DISCONNECT. VALVE
Detail
A
10
3337-0614 THRU 33701193
17
"~ I.
1
10
•
B
'~--/lVfr
MODEL337~-~00~0~1~~
_
THRU___
337-0039
-
35 34 33 32 31
THRU 337-0978
Detail
*
•
25
10
.~n
.~.
r---..
10
~
~i
I:"
Detail
Cradle Support
Strap
Cradle
4.
Strap
Interconnec t Line
l.
C
** 337-0040 T HRU 33701193
'11
337-0756
T HRU 33701193
~.5.•
'.
8.
9,
Cylinder ure Relief Valve
High-Press e Line
Pressure Gag
Outlet Line
ON-OFF Control
10.
11.
12.
Regulator
Low-Pressure Re lief Valve
13.
14,
15.
16.
17.
18.
19.
20.
2l.
22,
23.
24.
21
l
-7"
32
2.
3.
2&*
2&"
t>:);J
~J;d.
D
-~A
'fPSj 337 -0979
. AND ON
31"P
DetaUB'
L'ne
Filler Valve 1 Relief Valve
High-Pressure
Spacer
O-Ring
Base
Jamb Nut
SprIng
Poppet
Core
Escutcheon
Cover
Lock Ring
25.
26.
27.
28.
29.
30.
3l.
32.
33.
34.
35.
36.
ti::::!:ar- ---30
337-0001
755
THRU 337-0
Control Arm
Bracket
Pressure Gage
Console
Knob
Bezel
Cap
Valve
O-Ring
Piston
Seat
Bracket
Figure 13-10. Oxygen System (Sheet 1 of 3)
13-31
•
Detail
A
........... ........
................ ........ B
........
....;;.....
.
.. .::::.::::..
........ ~:......: ........................'
.......,
'
........
-
13
......
"'\:..'/
....,:-:.........
0"
../......
~'"
-,......-
.'_
..::'/"
.. ::,(('"
....
12
'
.......... "·t-
L§.l_-17
ir·.."..
.......:.:.~: : :~ i~ ...:~:· · · · · ·..
..........
...:..
.\.'
DetanB
•
13
'
,
..
'
C
5
33701194 THRU 33701462
20
12
Cap
Valve
Cover
Filler Valve Line
5. RH Low Pressure Line
6. Filler Line
7. Support
1.
2.
3.
4.
8.
9.
10.
11.
12.
13.
14.
DetailC
Interconnect Line
Cradle Support
Oxygen Cylinders
Strap
LH Low Pressure Line
Capillary Line
Regulator
Figure 13-10. Oxygen System (Sheet 2 of 3)
13-32
15.
16.
17.
18.
19.
20.
21.
Valve
Control
Control Arm
Console
Pressure Gage
Knob
Outlet
•
•
6
.'
.'
..
,
.... .;~:...
'
.
..........
•
BEGINNING WITH 33701463
14
1. Outlet
2. Arm
3. Control
4. Support
5. Cylinder and Valve Assembly
6. Stiffener
7. Cradle Support
•
8.
9.
10.
11.
12.
13.
14.
Cradle
Cyl1ner and Regulator Assembly
Strap
Cable Assembly
Spacer
Gage
Knob
Figure 13-10. Oxygen System (Sheet 3 of 3)
13-33
NOTE
Oxygen cylinder and regulator assemblies
may not always be installed in the field
exactly as illustrated in figure 13-10, which
shows factory installation. Important
points to remember are as follows.
a. Before removing cylinder, release low-pressure line by opening cabin ouUets. Disconnect pushpull control cable, filler line, pressure gage line
and outlet line from regulator. CAP ALL LINES
IMMEDIATELY.
b. U it is necessary to replace filler valve O-rhlgs,
remove parts necessary for access to filler valve.
Remove line from quick-discormect valve at the
regulator, then discormect chain, but do not remove
cap from filler valve. Remove screws securing
valve and disconnect pressure line. Referring to
applicable figure, cap pressure ltne and seat. Disassemble valve, replace O-rings and reassemble
valve. Install filler valve by reversing procedures
outlined in this step.
c. A cabin outlet is tllustrated in figure 13-10. Repair kit, (part no. C166006-0108), available from
the Cessna Service Parts Center, may be used for
replacement of components of the outlet assembly.
d. To remove entire oxygen system, headliner
must be lowered and soundproofing removed to expose lines. Refer to Section 3 for headliner removal.
13-35. OXYGEN CYLINDER GENERAL INFORMATION. The following information is permanently
steel stamped on the shoulder, top head or neck of
each oxygen cylinder:
a. Cylinder specification, followed by service
pressure (e. g. '1CC-3AA1800" and "ICC-3HT1850"
for standard and light weight cylinders respectively).
NOTE
Effective 1 January 19'70, all newly-malUlfactured cylinders will be stamped "DOT"
(Department of Transportation), rather
than '1CC" (Interstate Commerce Commission). An example of the new deSignation
would be: ''DOT-3HT1850''.
b. 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.
c. Inspector's official mark near serial number.
d. Date of manufacture: This is the date of the
first hydrostatic test (such as 4-69 for April 1969).
The dash between the month and the year figures
may be replaced with the mark of the testing or inspection agency (e. g. 4L69).
e. Hydrostatic test date: The dates of subsequent
hydrostatic tests shall be steel stamped (month and
year) directly below the original manufacture date.
The dash between the month and year figures can be
replaced with the mark of the testing agency.
13-34
f. A Cessna identiftca~ion placard is located near
the center of the cylinder body.
g. Halogen test stamp: "Halogen Tested", date of
test (month, day and year) and inspector's mark
appears directly underneath the Cessna identification
placard.
•
13-36. OXYGEN CYLINDER SERVICE REQumEMENTS.
a. Hydrostatic test requirements:
1. standard weight (ICC or DOT-3AAl800)
cylinders must be hydrostatically tested to 5/3 their
working pressure every five years commenCing with
the date of the last hydrostatic test.
2. Light weight (ICC or DOT-3HT1850) cylinders must be hydrostatically tested to 5/3 their
working pressure every three years commencing
with the date of the last hydrostatic test.
b. Service life requirements:
1. Standard weight (ICC or DOT-3AA1800)
cylinders have no age life limitations and may continue to be used until they fail hydrostatiC test.
2. Light weight (ICC or DOT-3HT1850) cylinders must be retired from service after 12 years
or 4,380 filling cycles after date of manufacture,
whichever occurs first.
NOTE
These test periods and life limitations are
established by the Interstate Commerce
Commission Code of Federal Regulations,
Title 49, Chapter 1, Para. '73. 34.
13 -3'7 • OXYGEN CYLINDER INSPECTION REQumEMENTS.
a. Inspect the entire exterior surface of the cylinder for indication of abuse, dents, bulges and strap
chafing.
b. Examine the neck of cylinder for cracks, distortion or damaged threads.
c. Check the cylinders to determine if markings
are legible.
d. Check date of last hydrostatic test. U the periodic retest date is past, do not return the cylinder
to service until the test has been accomplished.
e. Inspect the cylinder mounting bracket, bracket
hold-down bolts and cylinder holding straps for
cracks, deformation, cleanliness, and security of
attachment.
f. In the immediate area where the cylinder is
stored or secured, check for evidence of any types
of interference, chafing, deformation or deterioration.
13-38. OXYGEN SYSTEM COMPONENT SERVICE
REQumEMENTS.
a. PRESSURE REGULA TOR. The regulator shall
be functionally tested every two years or I, 000 hours
for aircraft operating under 15, 000 ft. and one year
for aircraft operating over 15,000 ft. The regulator
shall be overhauled every five years ao at time of
hydrostatiC test.
b. FILLER VALVE. The valve shall be functionally tested every two years and overhauled every five
years or at time of hydrostatic test.
•
•
•
c. QUICK-RELEASE COUPLING. The coupUng
shall be functionally tested every two years and
overhauled every five years or at time of hydrostatic
test.
d. PRESSURE GAGE. The gage shall be checked
for accuracy and overhauled by an FAA approved
factuty every five years.
e. OUTLETS. The outlets shall be disassembled
and inspected and the sealing core replaced, regardless of condition, every five years.
13-39. OXYGEN SYSTEM COMPONENT INSPECTION REQUIREMENTS.
a. Examine all parts for cracks, nicks, damaged
threads or other apparent damage.
b. Actuate regulator controls and valve to check
for ease of operation.
c. Determine if the gage is functiOning properly
by observing the pressure build-up and the return to
zero when the system oxygen is bled off.
d. Replace any oxygen line that is chafed, rusted,
corroded, dented, cracked or kinked.
e. Check fittings for corrosion around the threaded area where lines are joined together. Pressurize the system and check for leaks.
•
13-40. MASKS AND HOSE.
a. Check oxygen masks for fabric cracks and rough
face seals. If the mask is a full-faced model, inspect glass or plastic for cleanliness and state of
repair.
b. Flex the mask hose gently over its entirety and
check for evidence of deterioration or dirt.
c. Examine mask and hose storage compartment
for cleanliness and general condition.
13-41. MAINTENANCE AND CLEANING.
a. Clean and disinfect mask assemblies after use,
as appropriate.
NOTE
Use care to avoid damaging microphone
assembly while cleaning and sterilizing.
b. Wash mask with a mild soap solUtion and rinse
it with clear water.
•
c. To steril1ze, swab mask thoroughly with a
gauze or sponge soaked in a water/merthiolate solution. This solution should contain 1/5 teaspoon of
merthiolate per one quart of water. Wipe the mask
with a clean cloth and let air dry.
d. 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.
e. 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.
f. Replace hose if it shows evidence of deterioration.
g. Hose may be cleaned in the same manner as the
mask •
13-42. SYSTEM PURGING. Whenever components:
have been removed and reinstalled or replaced, it is
advisable to purge the system. Charge oxygen sys-
tem in accordance with procedures outlined in paragraph 13-45. Plug masks into all outlets and turn
the pilot's control to ON position and purge system
by allowing oxygen to flow for at least 10 minutes.
Smell oxygen flowing from outlets and continue to
purge until system is odorless. Refill cylinders as
required during and after purging.
13-43. FUNCTIONAL TESTING. Whenever the regulator and cylinder assembly has been replaced or
overhauled, perform the following flow and internal
leakage tests to check that the system functions properly.
a. Fully charge oxygen system in accordance with
procedures outlined in paragraph 13-45.
b. Disconnect line and fitting assembly from pilot's mask and line assembly. Insert outlet end of
line and fitting assembly into cabin outlet and attach
opposite end of line to a pressure gage (gage should
be calibrated in one-pound increments from 0 to 100
PSI). Place control lever in ON position. Gage
pressure should read 75±10 PSI.
c. Insert mask and line assemblies into all remaining cabin outlets. With oxygen flOwing from all
outlets, test gage pressure should still be 75±10 PSI.
d. Place oxygen control lever in OFF position and
allow test gage pressure to fall to 0 PSI. Remove
all adapter assembl1es except the one with the pressure gage. The pressure must not rise above 0 PSI
when observed for one minute. Remove pressure
gage and adapter from oxygen outlet.
NOTE
If pressures specified in the foregOing pro-
cedures are not obtained, the oxygen regulator is not operating properly. Remove
and replace cylinder-regulator assembly
with another unit and repeat test procedure.
e. Connect mask and line assemblies to each cabin
outlet and check each mask for proper operation.
f. Check pilot's mask microphone and control
wheel switch for proper operation. After checking,
return all masks to mask case.
g. Recharge oxygen system in accordance with
procedures outlined in paragraph 13-45.
13 -44. SYSTEM LEAK TEST. When oxygen is being
lost from a system through leakage, a sequence of
steps may be necessary to locate the opening. Leakage may often be detected by listening for the distinct hissing of escaping gas. If this check proves
negative, it will be necessary to soap-test all lines
and connections with a castile soap and water solution or specially compounded leak-test material.
Make the solution thick enough to adhere to the contours of the fittings. At the completion of the leakage test, remove all traces of the leak detector or
soap and water solution.
/CAUTION\
Do not attempt to tighten any connections
while the system is charged.
13-35
•
NOTE
Each interconnected series of oxygen cylinders is
equipped with a single gage. The trailer type
cascade may also be equipped with a nitrogen cylinder (shown reversed) for filling landing gear
struts, accumulators, etc. Cylinders are not
aVailable for direct purchase, but are usually
leased and refilled by a local compressed gas
supplier.
PRESSURE GAGE
OXYGEN PURIFIER
W/REPLACEABLE
CARTRIDGE
Figure 13-11. Typical Portable Oxygen Cascades
13-45. SYSTEM CHARGING.
IWARNING'
BE SURE TO GROUND AIRCRAFT AND
GROUND SERVICING EQUIPMENT BEFORE CHARGING OXYGEN SYSTEM.
a. Do not attempt to charge oxygen cylinders 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.
b. If cylinder is completely empty, do not charge,
as the cylinder must then be removed, inspected
and cleaned.
@AUTION\
A cylinder which is completely empty may
well be contaminated. The regulator and
cylinder assembly must then be disassembled, inspected and cleaned by an FAA
approved faCility, before filling. Contamination, as used here, means dirt, dust
or any other foreign material, as well as
ordinary air in large quantities. If a gage
line or filler line is disconnected and the
fittings capped immediately, the cylinder
will not become contaminated unless temperature variation has created a suction
within the cylinder. Ordinary air contains
water vapor which could condense and
freeze. Since there are very small orifices
in the system, it is very important that
this condition not be allowed to occur.
•
•
13-36
•
c. Connect cylinder valve outlet or outside filler
valve to manifold or portable oxygen cascade.
d. Slowly open valve on cascade cylinder or manifold with lowest pressure, as noted on pressure gage,
allow pressure to equalize, then close cascade cylinder valve.
e. Repeat this procedure, using a progressively
higher pressure cascade cylinder, until system has
been charged to the pressure indicated in the chart
immediately following step "r' of this paragraph.
f. Ambient temperature listed in the chart is the
air temperature in the area where the system is to
be charged. Filling pressure refers to the pressure to which aircraft cylinders should be filled.
This table gives approximations only and assumes
a rise in temperature of approximately 25°F. due
to heat of compression. This table also assumes
the aircraft cylinders will be filled as quickly as possible and that they will only be cooled by ambient
air; no water bath or other means of cooling be used.
Example: If ambient temperature is 70°F., fill
aircraft cylinders to apprOXimately 1,975 psi or as
close to this pressure as the gage may read. Upon
cooling, cylinders should have apprOXimately 1, 850
psi pressure.
TABLE OF FILLING PRESSURES
Ambient
Temp.
OF
0
10
20
30
40
Filling
Press.
psig
1650
1700
1725
1775
1825
Ambient
Temp.
OF
Filling
Press.
psig
50
60
70
80
90
1875
1925
1975
2000
2050
•
SHOP NOTES:
•
13-37/(13-38 blank)
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
•
SECTION 14
INSTRUMENTS AND INSTRUMENT SYSTEMS
TABLE OF CONTENTS
Page
•
•
INSTRUMENTS AND INSTRUMENT
SySTEMS .................................................... 14-2
General ................................................... 14-2
Instrument Panel ..................................... 14-2
Description ......................................... 14-2
Removal and Installation .................... 14-2
Shock Mounts ......................................... 14-2
Description ......................................... 14-2
Instruments ............................................. 14-2
Removal and Installation .................... 14-2
Pitot and Static Systems ......................... 14-3
Description ......................................... 14-3
Maintenance ...................................... 14-3
Static Pressure System Inspection
and Leakage Test.. ............................ 14-3
Pitot Static System Inspection
and Leakage Test.. ............................ 14-10
Blowing Out Lines .............................. 14-14
Removal and Installation .................... 14-14
Troubleshooting ................................. 14-14
True Airspeed Indicator .......................... 14-14
Description ......................................... 14-14
Trouble Shooting ................................ 14-16
Trouble Shooting - Altimeter .............. 14-16
Trouble Shooting - Vertical Speed
Indicator ............................................. 14-16
Trouble Shooting - Pitot Tube Heater 14-17
Vacuum System ...................................... 14-18
Description ......................................... 14-18
Trouble Shooting ............................... 14-16
Trouble Shooting - Gyros ................... 14-20
Trouble Shooting - Vacuum Pump .... 14-21
Removal and Installation .................... 14-21
Cleaning ............................................. 14-21
Relief Valve Adjustment.. ................... 14-21
Engine Indicators .................................... 14-21
Fuel Quantity Indicating System ........ 14-21
Indicators
Description .................................... 14-21
Transmitters ....................................... 14-21
Description .................................... 14-21
Removal ........................................ 14-22
Installation ..................................... 14-22
Sending Units ..................................... 14-22
Description .................................... 14-22
Removal and Installation ............... 14-22
Change 2
Jan 5/2004
Page
Control Monitors ...................................... 14-22
Description ......................................... 14-22
Removal and Installation .................... 14-22
Calibration ............................................... 14-22
Trouble Shooting ................................ 14-23
Fuel Quantity Indicating System
Operational Test ..................................... 14-23
Dual Tachometer .................................... 14-25
Description ......................................... 14-25
Trouble Shooting ................................ 14-25
Dual Manifold Pressure Gage ................. 14-26
Description ......................................... 14-26
Trouble Shooting ................................ 14-26
Dual Fuel Flow Indicator ......................... 14-26
Description ......................................... 14-26
Trouble Shooting ................................ 14-27
Instrument Cluster .................................. 14-27
Description ......................................... 14-27
Cylinder Head Temperature Gages ........ 14-28
Description ......................................... 14-28
Trouble Shooting ................................ 14-28
Oil Pressure Gages ................................. 14-28
Description
14-28
Trouble Shooting
14-28A
Oil Temperature Gages .......................... 14-28A
Description ......................................... 14-28A
Hourmeter ............................................... 14-28B
Description ......................................... 14-28B
Trouble Shooting ................................ 14-28B
Synchroscope ......................................... 14-29
Description ......................................... 14-29
Trouble Shooting ................................ 14-29
Dual Economy Mixture Indicator ............. 14-30
Description ......................................... 14-30
Trouble Shooting ................................ 14-30
Calibration .......................................... 14-30
Removal and Installation .................... 14-30
Miscellaneous Instruments .......................... 14-30
Magnetic Compass ................................. 14-30
Description ......................................... 14-30
Turn-and-Slip Indicator ........................... 14-31
Description ......................................... 14-31
Trouble Shooting ................................ 14-31
Outside Air Temperature Gage ............... 14-32
Description ......................................... 14-32
Trouble Shooting ................................ 14-32
© Cessna Aircraft Company
14-1
CESSNA AIRCRAFT COMPANY
MODEL 337
SERVICE MANUAL
14-1.
INSTRUMENTS AND INSTRUMENT SYSTEMS.
14-2.
GENERAL. This section describes typical instrument installations and their respective operating systems.
Emphasis is placed on troubleshooting and corrective measures only. It does NOT deal with specific
instrument repairs since this usually requires special equipment and data and must be handled by
instrument specialists. Federal Aviation Regulations require malfunctioning instruments be sent to an
approved instrument overhaul and repair station or returned to manufacturer for servicing. This Service
Manual provides preventive maintenance information on various instrument systems and repair of
systems that do not operate. The descriptive material, maintenance and trouble shooting information in
this section is intended to help the mechanic determine why an instrument system does not operate in a
satisfactory manner. Some instruments, such as fuel quantity and oil pressure gages, are so simple and
inexpensive; repairs usually will be more costly than a new instrument. Aneroid and gyro instruments are
usually worth repairing. The words "replace instrument" in the text, therefore, must be taken only in the
sense of physical replacement in the aircraft. If replacement is to be with a new instrument, an exchange
one, or the original instrument is to be repaired must be decided on basis of individual circumstances.
14-3.
INSTRUMENT PANEL.
14-4.
DESCRIPTION. The instrument panel assembly consists of a stationary panel, a removable flight
instrument panel and a shock-mounted panel. The stationary panel is part of the fuselage structure and
not ordinarily removable. The stationary panel contains fuel and engine instruments. The removable panel
contains flight instruments such as airspeed, vertical speed and altimeter. The shock-mounted panel
contains major flight instruments such as the horizontal and directional gyros. Decorative covers are
installed on the panel with screw on buttons thru 1971 models, beginning with 1972 models Velcro
fasteners are used.
14-5,
REMOVAL AND INSTALLATION. Two methods can be used to remove the shock-mounted or removable
panel. Disconnect wiring and plumbing as necessary, tag wiring and cap plumbing. Instruments can be
removed from the panel singular or the panel can be removed as an assembly. A removable door/two
removable doors beginning with 1973 models, forward of the windshield provide access to the area behind
the instrument panel.
14-6.
SHOCK-MOUNTS.
14-7.
DESCRIPTION. Service life of instruments is directly related to proper shock mounting of the panel. Thru
1972 models the shock mounted panel is secured to the stationary panel with seven shock mounts, two
non-adjacent mounts could possibly have been cut through the middle, at the factory, to lessen vibration.
If additional instruments are installed in the field, these two cut shock mounts must be replaced with
standard shock mounts if excessive vibration occurs due to the increase weight. Conversely, if the weight
of the panel is decreased by permanent removal of equipment, two non-adjacent shock mounts can be
cut through the middle to lessen vibration caused by the decreased weight of the panel. Beginning with
1973 models the shock-mounted panel contains only two instruments and is installed with four shock
mounts. Shock mounts must be checked periodically for deterioration and replaced as necessary.
14-8.
INSTRUMENTS.
•
•
REMOVAL AND INSTALLATION. Thru 1972 models most instruments are secured to the panel with
screws inserted through the panel face under the decorative cover. To remove an instrument, remove
decorative cover, disconnect wiring or plumbing