TECHNICAL MANUAL ORGANIZATIONAL MAINTENANCE

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP75B00300R000100050001-7
• Release Decision: RIFPUB
• Original Classification: U
• Document Page Count: 209
• Document Creation Date: December 9, 2016
• Document Release Date: December 8, 2000
• Sequence Number: 1
• Publication Date: September 22, 1965
• Content Type: MISC

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TECHNICAL MANUA
ORGANIZATIONAL MAINTE CE
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UNCLASSIFIED
-2-4
VOLIE
COPY NO.
TECHNICAL MANUAL
ORGANIZATIONAL MAINTENANCE
POWERPLANT
(J75P-13)
MODELS U-2C AND U-2F AIRCRAFT
NOTICE
THE DOCUMENT TO WHICH THIS A CHANGE (TMOM -2-4 VOL II)
HAS SUPERSEDED TMOM -2-4 VOL II DATED 1 NOVEMBER 1964,
CHANGED 8 JANUARY 1965; AND ONLY SECTION -2-4 VOL II OF
TMOM -2 ( ) SERIES SUPPLEMENT, DATED 15 FEBRUARY 1965,
CHANGED 21 SEPTEMBER 1965.
DESTROY SUPERSEDED DATA IN ACCORDANCE WITH AFR 205-1.
LATEST CHANGED PAGES SUPERSEDE
THE SAME PAGES OF PREVIOUS DATE
Insert changed pages into basic
publication. Destroy superseded pages.
UNCLASSIFIED
22 SEPTEMBER 1965
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VOL II
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INTRODUCTION
U-2
This is one in a series of manuals which comprise the Technical Manual
of Organizational Maintenance for U-2 aircraft.
Each manual in. the series is a complete and separate book prepared in
support of a particular system. (Some of these manuals, due to simi-
larity or direct interrelation of system functions, contain added sections
wherein the associated systems are covered.)
c
Ma
data or
anual, or each section within it, is broken down into four parts,
cription, Operational Checkout, Trouble Shooting, and
Thus, interested personnel are provided with all necessary
ational Maintenance.
In additi
to include
useful to qua
?
icable, appendices are made part of these manuals,
ling with bench adjustments, and other data
?
NOTES, CAUTIONS, AND WARNINGS
These adjuncts to the text are defined as follows:
Note - An operation, procedure, condition, et cetera, which it
is essential to emphasize.
CAUTION - Operations, procedures, practices, et cetera, which if
not strictly observed, will result in damage to or destruc-
tion of equipment.
LWARNING1
- Operations, procedures, practices, et cetera, which will
result in personal injury or loss of life, if not correctly
followed.
iii
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iv
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LIST OF U-2 TECHNICAL MANUALS
MANUAL NO. TITLE
-2-1
-2-2,
-2-3
-2-4 VOL I
-2-4 VOL II
-2-5
-2-6
-2-7
-2-8
-2-9
General Airplane, Airframe, Landing Gear
Flight Controls and Instruments
Air Conditioning and Pressurization
Powerplant J57
Powerplant J75
Fuel System and Hydraulic System
Oxygen System
Electrical and Electronics
Ground Handling
Special Equipment
-3-1 Structural Repair Instructions
-6-i Inspection Requirements
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VOL II
TABLE OF CONTENTS
Page
Security Classification i/ii
Introduction 111
List of IJ-2 Technical Manuals iv
POWERPLANT (J75) 1-1
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v/vi
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VOL II
POWERPLANT
J75
CONTENTS
DESCRIPTION
Page No.
Engine
1-5
Intake Air System
1-17
Engine Instruments
1-18
Engine Mounting System
1-26
Engine Starting System
1-29
Throttle Control System
1-31
Fuel System
1-33
Ignition System
1-49
Exhaust System
1-58
Lubrication, Scavenge and Breather System
1-59
OPERATIONAL CHECKOUT
Engine
1-75
Engine Instruments
1-82
Fuel System
1-88
Ignition System
1-95
TROUBLE SHOOTING
Engine
1-99
Engine Instruments
1-111
Engine Starting System
1-113
Ignition System
1-115
MAINTENANCE
Engine
1-117
Engine Pressure Ratio System
1-166
Engine Starting System
1-167
Throttle Control System
1-169
Fuel System
1-171
Ignition System
1-176
Exhaust System
1-180
Lubrication, Scavenge and Breather System
1-183
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1-1
Contents
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VOL II
LIST OF ILLUSTRATIONS
Figure Title Page No,
1-1 Cutaway of J75 Engine 1-6
1-2 Major Engine Components (Typical) 1-7
1-3 Dual Rotor Type Engine Reference Stations 1-9
1-4 Engine Airflow (Typical) 1-11
1-5 Accessory Location 1-15
1-6 Engine Tachometer Indicating System 1-19
1-7 Exhaust Gas Temperature Indicating System 1-21
1-8 Engine Oil Pressure Indicating System 1-22
1-9 Engine Oil Temperature Indicating System 1-24
1-10 Engine Pressure Ratio Indicating System 1-25
L-11 Engine Mount Installation (Sheet 1) 1-27
1-11 Engine Mount Installation (Sheet 2) 1-28
1-12 Engine Pneumatic Starter 1-30
.L-13 Throttle Control System 1-32
1-14 Engine Fuel System 1-34
1-15 Engine Fuel Pump 1-37
1-16 Fuel Control (Sheet 1) 1-38
1-16 Fuel Control (Sheet 2) 1-39
1-17 Fuel Control Schematic (Normal System Operation) 1-41
1-18 Fuel Control Schematic (Emergency System Operation) ^ ? ? 1-44
1-19 Fuel Pressurizing and Dump Valve 1-48
1-20 Normal and Continuous Ignition System Block Diagram ? ? 1-49
1-21 Normal and Continuous Ignition System Simplified
Wiring Diagram 1-50
1-22 Normal and Continuous Ignition System Schematic .1-52
1-23 Normal System Sparkigniter 1-54
1-24 Continuous Duty Sparkigniter 1-57
1-25 Lubrication, Scavenge, and Breather System Schematic . ? ? 1-60
1-26 Main Oil Pump 1-63
1-27 Main Oil Strainer 1-65
1-28 Air-Oil Cooler Schematic, 14-Inch
(Left-Hand Installation) 1-66
1-29 Thermostatic Temperature Control Valve 1-67
1-30 Thermostatic Temperature Control Valve Schematic 1-69
1-31 Air-Oil Cooler Schematic, 9-Inch (Right-Hand Installation) 1-70
1-32 Fuel-Oil Cooler 1-71
1-33 Fuel-Oil Cooler Schematic 1-72
1-34 Breather Pressurizing Valve Schematic 1-73
1-35 Temperature Measurement 1-85
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VOL II
LIST OF ILLUSTRATIONS (CONT'D)
Contents
Figure Title Page No.
1-36
Typical Trim Sheet
1-93
1-37
Engine Part Power Trim Curve
1-94
1-38
Removal of Flexible Preservation Container from Engine
While Mounted on Transportation Trailer (Sheet 1)
1-128
1-38
Removal of Flexible Preservation Container from Engine
While Mounted on Transportation Trailer (Sheet 2)
1-129
1-39
Engine Transportation Trailer
1-130
1-40
Engine Buildup (Sheet 1)
1-133
1-40
Engine Buildup (Sheet 2)
1-134
1-40
Engine Buildup (Sheet 3)
1-135
1-40
Engine Buildup (Sheet 4)
1-136
1-41
Transfer of Engine from Transportation Trailer to
Installation Trailer
1-139
1-42
Engine Quick Disconnect Points (Sheet 1)
1-143
1-42
Engine Quick Disconnect Points (Sheet 2)
1-144
1-42
Engine Quick Disconnect Points (Sheet 3)
1-145
1-43
Removal of Engine from Airplane (Sheet 1)
1-146
1-43
Removal of Engine from Airplane (Sheet 2)
1-147
1-43
Removal of Engine from Airplane (Sheet 3)
1-148
1-43
Removal of Engine from Airplane (Sheet 4)
1-149
1-44
Installation of Engine in. Airplane (Sheet 1)
1-153
1-44
Installation of Engine in Airplane (Sheet 2)
1-154
1-44
Installation of Engine in Airplane (Sheet 3)
1-155
1-44
Installation of Engine in Airplane (Sheet 4)
1-156
1-45
Installation of Flexible Preservation Container on Engine
While Mounted on Transportation Trailer (Sheet 1)
1-164
1-45
Installation of Flexible Preservation Container on. Engine
While Mounted on Transportation Trailer (Sheet 2)
1-165
1-46
Inspection of Starter
1-169
1-47
Inspection of Sparkigniter
1-179
1-48
Tailpipe Installation
1-181
1-49
Servicing of Oil Tank
1-185
1-50
Cleaning of Main Oil Strainer
1-189
1-51
Oil Pressure Relief Valve
1-191
LIST OF TABLES
Table
Title
Page No.
1-1
Generator Speed versus Engine Speed
1-16
1-2
Fuel System Components
1-35
1-3
Lubrication and Breather System Components
1-62
1-4
Engine Operational Check Limits
1-80
1-5
Engine Overtemperature and Overspeed Limits
1-81
1-6
Standard Alurn.el Chromel Voltage Scale
1-86
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VOL II
LIST OF TABLES (CONTrD)
Table
Title
Page No.
1-7
Corrections for External Reference Junction Temperature
1-87
1-8
Torque Values for Nuts, Bolts, and Screws
1-119
1-9
Torque Values for Flexible Tube Connections
1-119
L-10
Torque Values for Hose, Tube and Threaded Connections
1-120
1-11
Torque Values for Crush Type Asbestos Filled Gaskets
1-121
:L-12
Torque Values for Steel Pipe Plugs in Aluminum
or Magnesium
1-121
1-13
Special Torques
1-122
1-4
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VOL II
DESCRIPTION
1-1. ENGINE.
1-2. GENERAL.
1-3. The J75P-13 engine (See figures 1-1 and 1-2.) is a continuous flow, turbojet
engine incorporating an axial flow compressor, an eight unit canannular combustion
chamber and a split three stage reaction turbine.
1-4. The multistage axial flow split compressor consists of an eight stage low
pressure unit, which is connected by a through-shaft to the second and third stage
turbine wheels, and a seven stage high pressure, high speed unit, which is con-
nected by a hollow shaft to the first stage turbine wheel.
1-5. The accessory section is located under the "wasp waist" of the compressor
section.
1-6. The tailpipe assembly is attached to the turbine exhaust case and extends
to the augmentor assembly.
1-7. SPLIT COMPRESSOR OPERATION.
1-8. Greater flexibility for starting and part load operation is achieved by
splitting the compressor into two mechanically separated rotors. Each rotor is
driven by separate turbines. The low pressure rotor being free to rotate at its
best speed.
1-9. The starter drives but one section, thus reducing the size and weight of
the starting system.
1-10. The high pressure rotor is geared to the starter drive because it is the
smaller of the two and so requires the lesser torque for starting.
1-11. With the rear or high pressure compressor rotor turning at the governed
speed, the front or low pressure compressor rotor is rotated by its turbine at the
rpm ensuring optimum flow through the compressor.
1-12. JET ENGINE SYMBOLS AND THEIR MEANING.
1-13. The following list of jet engine symbols (See figure 1-3.) represents most of
the more common everyday symbols that you will find used in this text as well as
on the flight or maintenance line.
Fn net jet thrust, lb
N1 low pressure compressor rotational speed, rpm
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1 COMPRESSOR INLET GUIDE VANES AND SHROUD
2 LOW-PRESSURE COMPRESSOR (EIGHT -STAGE)
3 HIGH-PRESSURE COMPRESSOR (SEVEN-STAGE)
4 COMBUSTION CHAMBER (8)
5 FUEL NOZZLE (6 IN EACH COMBUSTION CHAMBER)
6 TURBINE NOZZLE
7 TURBINE WHEEL, FRONT (ONE, DRIVES HIGH PRESSURE COMPRESSOR)
8 TUBRINE WHEEL, REAR (TWO, DRIVE LOW PRESSURE COMPRESSOR)
9 SWIRL STRAIGHTENER VANE (6.)
10 EXHAUST CONE
11 EXHAUST TAILPIPE (NOT PART OF ENGINE)
10 11
23
22
21 20 19 18
17 16 15
12 EXHAUST GAS TEMPERATURE PROBE
13 TURBINE EXHAUST CASE
14 TURBINE NOZZLE CASE
15 COMBUSTION CHAMBER CASE
16 FUEL MANIFOLD AND NOZZLES
17 DIFFUSER CASE
18 COMPRESSOR INTERMEDIATE CASE
19 ENGINE MOUNT (BALL BAT)
20 ACCESSORY CASE (N2)
21 OIL TANK
22 FRONT COMPRESSOR CASE
23 ACCESSORY CASE (N1)
Figure 1-1. Cutaway of J75 Engine
uopdTapsoa
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L-1?00090001?000t100?0089/dC1U-VIO 8Z/80/1.00Z eseeleu JOd peAwddv
1 COMPRESSOR INLET GUIDE VANE AND SHROUD
2 FRONT COMPRESSOR CASE
3 COMPRESSOR INTERMEDIATE CASE
4 DIFFUSER CASE
5 COMBUSTION CHAMBERS
6 SPLIT TYPE FUEL MANIFOLD
7 REAR COMPRESSOR ROTOR (Ne)
8 REAR COMPRESSOR CASE
9 FRONT COMPRESSOR ROTOR (Ni)
10 NO.1 BEARING SUPPORT
11 COMBUSTION CHAMBER OUTER CASE
12 TURBINE NOZZLE CASE
13 TURBINE EXHAUST CASE
14 REAR COMPRESSOR TURBINE ROTOR (Ne)
15 FRONT COMPRESSOR TURBINE ROTOR (N1)
16 COMBUSTION CHAMBER OUTLET DUCT
1
2
A
10
Figure 1-2. Major Engine Components (Typical)
15
14
uo TIAT..1 s aa
Des cription
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VOL II
N2 high pressure compressor rotational speed, rpm
P
am ambient absolute pressure
Ptl
Pt2
Pt3
Pt4
Pt7
total pressure at entrance to inlet duct
total pressure at low pressure inlet
total pressure at low pressure discharge
total pressure at high compressor discharge
total pressure at low pressure turbine outlet
tam ambient temperature, ?F
tt2 temperature at low compressor inlet, ?F
Tt2 total temperature at low compressor inlet
111:7 total temperature at low pressure turbine inlet
tsfc thrust specific fuel consumption, pound of fuel hour per pound of thrust
Wf engine fuel flow, lb/hr
1-14. SYMBOL SUBSCRIPTS.
Example - Pt4:
Means pressure, but what pressure and where?
4
Means total pressure as differentiated from static pressure
Means engine reference station No. 4 (See figure 1-3.)
Thus Pt4 means total pressure existing at the discharge of the high compressor.
The engine reference station numbers will be among the most common subscripts
used. Some more of the common ones are as follows:
1-15. SUBSCRIPTS.
am n ambient (tam)
burner, combustion chamber (Pb)
compressor (77c)
exhaust, exit
fuel (Wf)
static (Ps)
total (pt); turbine (77 t)
1-8
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VOL II
Description
?
I
;
rx1
LOW PRESSURE
COMPRESSOR
BURNER
TAIL-
PIPE
am
2
3
4
5
6
8
REFERENCE STATIONS
It is customary in the jet engine field to utilize engine reference stations
when wishing to indicate the characteristics of the many aerodynamic or
thermodynamic variables at a specific point in the air's progress through
the engine. The standard location of these reference stations, for dual
rotor type axial flow compressor jet engines, are shown above.
Figure 1-3. Dual Rotor Type Engine Reference Stations
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Description
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VOL II
1-16. ENGINE AIRFLOW.
1-17. Air enters the engine (See figure 1-4.) through the compressor inlet guide
vane and shroud assembly (1). Air from this source enters the front compressor (2),
which consists of eight rotor stages and seven vane stages. The gas path of this
compressor has an increasing inside diameter and a decreasing outside diameter.
The front compressor provides initial compression of air.
1-18. The compressor intermediate case (4) separates the front compressor (2)
from the rear compressor (3). Inlet vanes (8th stage) direct compressed air
from the front compressor (2) to the rear compressor(3).
1-19. The rear compressor (3) has seven rotor stages and six vane stages. The
gas path of this compressor has an increasing inside diameter, and a constant
outside diameter. As the air passes through, it increases from low velocity to
high velocity.
1-20. The diffuser case (5) serves to diffuse the air flow discharged by the rear
compressor and adapt it for entry into the combustion chambers. The exit guide
vanes, mounted in the air stream in the forward part of the case, accomplish
diffusion and widening of the air passage formed by the inner inlet duct. The inner
diameter of the case diverts the air to the combustion chambers for burning.
1-21. Most of the highly compressed air discharged from the front and rear com-
pressors passes into the combustion section, there it combines with fuel from the
nozzles to form a combustible fuel-air mixture. When the fuel-air mixture is
ignited, the exhaust-gases are heated and accelerated, then discharged into the
turbine section. As the gases leave the turbine, their pressure forces them at
very high speeds through the Jet nozzles at the rear of the engine producing
reactive thrust.
1-22. From the combustion chambers (7), the gases enter the turbines, producing
power to drive the front and rear compressors, and fuel pump and accessories.
After the gases leave the turbines, their pressure forces them at very high speeds
through the Jet nozzle at the rear of the engine. The engines thrust comes from
taking a large mass of air in at the front end and pushing it out the jet nozzle at a
much higher speed than it had when it entered the front compressor.
1-23. ENGINE BLEED AIR SYSTEM.
1-24. GENERAL.
1-25, Air is bled from the engine primarily for cabin pressurization. Two high
pressure bleed ports on the engine are used for this purpose. A high pressure
line for hydraulic tank pressurization and fuel tank pressurization is taken off the
cabin pressurization line. The maximum pressure available at the high pressure
port is 160 psig.
1-40
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15
A8TH STAGE COOLING AIR
14
13
12
1
COMPRESSOR INLET GUIDE VANE AND SHROUD
9
FRONT COMPRESSOR TURBINE ROTOR
2
FRONT COMPRESSOR ROTOR
10
TURBINE EXHAUST CASE
3
REAR COMPRESSOR ROTOR
11
EXHAUST CONE
4
COMPRESSOR INTERMEDIATE CASE
12
TURBINE NOZZLE CASE
5
DIFFUSER CASE
13
COMBUSTION CHAMBER OUTER CASE
6
FUEL MANIFOLD AND NOZZLES
14
ACCESSORY SECTION ( N Z)
7
COMBUSTION CHAMBERS
15
FRONT COMPRESSOR CASE
8
REAR COMPRESSOR TURBINE ROTOR
Figure 1-4. Engine Airflow (Typical)
uopdTzosa a
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VOL II
1-26. ENGINE BLEED AIR MANIFOLD. An engine bleed air manifold, con-
structed of corrosion resistant steel tubing is bolted to the engine. The manifold
connects to two high pressure bleed air ports on the top right-hand side of the
engine diffuser case. These ports come equipped with Pratt and Whitney short
bolts, covers, and gaskets.
Note
To install the bleed air manifold longer bolts must be used.
(Refer to Engine Buildup.)
1-27. HYDRAULIC TANK PRESSURIZATION. The high pressure bleed air system
of the engine is used to pressurize the hydraulic tank. A quarter inch tee is plumbed
into the high pressure line taken off the cabin pressurization line. From the tee a
corrosion resistant steel tube is plumbed to a pressure regulator in the fuselage.
For additional information, refer to -2-5 Maintenance Manual.
1-28. FUEL TANK PRESSURIZATION. The fuel tanks are pressurized by engine
bleed air through the pressure regulator for level flight and descent conditions, and
by air expansion during a climb. The tank pressurization air is taken off the engine
compressor bleed line above and aft of the sump tank by a one-half inch steel line.
For additional information, refer to -2-5 Maintenance Manual.
1-29. ENGINE COOLING AIR SYSTEM,
1-30. COMPRESSOR AIR. (See figure 1-4.)
1-31. FRONT COMPRESSOR AIR. Cooling air for the rear face of the third stage
turbine disc is furnished by the front compressor. Leakage at the rear of the eighth
stage compressor blade platform enters the front compressor rotor assembly through
holes in the front compressor rear hub. The air then enters the front end of the
front compressor drive turbine shaft and passes rearward through the hollow shaft.
Cooling air leaves the shaft through holes in the front compressor drive turbine hub.
Some of the air is directed past the double air seal on the hub, along the rear face
of the third stage turbine disc and mixes with the exhaust gases. The remaining
air passes rearward through holes in the No. .6 bearing seal housing flange, the
bearing support, and the No. 6 bearing oil suction pump cover flange to the annulus
formed by the cover and the No. 6 bearing rear heatshield. Air leaving the rear
of the heatshield passes outward to enter holes at the inner end of the No. 6 bearing
oil pressure, oil scavenge and breather inner tubes and flows outward along the
tubes, leaving by means of holes in the tube heatshields to mix with the exhaust
gases.
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VOL II
Description
1-32. REAR COMPRESSOR AIR. Rear compressor cooling air passes through
drilled holes in the twelfth stage compressor stator outer shroud and enters the
passage formed by the inner diameter of the necked down portion of the diffuser case
and the rear compressor stator outer shrouds. The air then enters holes in the
outer ends of the diffuser case struts and flows inward to leave by means of holes at
the inner ends of the struts. The cooling air then passes rearward between the
combustion chamber inner case and the turbine shafts heatshield. From there it
passes through holes in the rear flange of the combustion chamber inner case to the
air space between the turbine front bearing seal housing and the turbine seal support.
Cooling air for the front face of the first stage turbine disc leaks past the double
air seal of the seal support and flows outward along the disc face to mix with the
exhaust gases.
1-33. Cooling air for the rear face of the first stage turbine disc and the front
and rear face of the second stage disc and the front face of the third stage disc
passes between the low compressor shaft and the inner diameter of the first and
second stage turbine disc. After the air passes the first stage disc it flows
outward past the single air seal of the second stage turbine nozzle to cool the
rear face of the first stage disc and past the double air seal on the front face of
the second stage disc to cool the front face of the second stage disc. After the
air passes the second stage turbine disc, the air flows through the holes in the
inner seal between the second and third stage disc. After serving to cool the
rear face of the second stage turbine disc and the front face of the third stage
turbine disc the air passes outward to mix with the exhaust gases.
1-34. EXTERNAL COOLING AIR.
1-35. ENGINE COMPARTMENT AND AFT SECTION COOLING. The engine
compartment and aft fuselage section are cooled by means of flush scoops.
a. On U-2C airplanes the flush scoops are located as follows: top engine
hoist cover, and left and right-hand engine mount access covers.
b. On TJ-2F airplanes the flush scoops are located as follows: upper left
and right fairing covers (F.S. 420) and left and right engine mount access covers.
1-36. Additional cooling for the portion aft of the engine burner section is
obtained from the oil cooler air scoop on the left-hand side and the hydraulic and
oil cooler air scoop on the right-hand side.
1-37. All cooling air is exhausted at the aft end of the airplane, between the
augmentor and the tailpipe.
1-38. Heat shields in the upper half of the airplane in the area of the hot section
of the engine protect the structure from radiated heat.
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1-13
Description
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VOL II
1-39. AUGMENTOR ASSEMBLY. The augmentor assembly/Part No. 75P1,
is located in the aft section of the airplane.
1-40. It is fabricated from corrosion resistant steel and is supported on a
bulkhead at the forward end and eight spring clips at the aft end.
1-41. The augmentor is designed to provide pumping of cooling air for aft
section as well as cooling for the air-oil cooler during ground running of the engine.
1-42. ENGINE MOUNTED ACCESSORIES.
1-43. GENERAL.
1-44. Engine mounted accessories (see figure 1-5.) are located on the N1 and
N2 accessory sections.
1-45. AC GENERATOR. The ac generator is a 208 volts, 400 cps, at 6000 rpm,
3 phase, 30 kva unit, Bendix type 28B54-14A. It is a class C, high temperature,
salient pole, brushless generator.
1-46. The ac generator is mounted on an adapter on the centerline of the front
accessory section. It is directly coupled to the N1 rotor and is driven at N1
rotor speed. This engine pad gear ratio is 1:1 (N1).
Note
When installing the generator, no gasket is used between the
generator adapter flange and engine pad.
1-47. The ac generator is air-cooled and receives the air through the left-hand
boundary layer scoop. A two inch diameter aluminum tube ducts the air aft to
the blast cap of the generator. Refer to -2-7 Maintenance Manual for details.
I1-48. STARTER. The engine starter is an air turbine type, Model ATS140-29-1,
Airesearch Part No. 350520 or Model ATS140-16-1, Airesearch Part No. 210250.
1-49. The starter is mounted on six studs provided on the center pad of the oil
pump and accessory drive gearbox. The engine pad gear ratio is 0.823:1 (N2).
1-50. DC GENERATOR. The dc generator installed in J75 equipped airplanes
is a Bendix 30B26-21A. It is a 400 ampere generator and is derated at altitude
to a maximum of 225 amperes.
1-14 Approved For Release 2001/08/28 : CIA-RDP75B00300R000100050001-7
hanged 22 November 1965
Approved For Release 2001/08/28 : GIARDP751300300R000100050001-7Description
VOL II
12
7
8
11
BOTTOM VIEW
10
1. AC GENERATOR
2. STARTER
3. DC GENERATOR
4. GENERATOR ADAPTER
5. ENGINE OIL PUMP ASSEMBLY
6. OIL PUMP AND ACCESSORY
DRIVES GEARBOX (N2)
7. TACHOMETER GENERATOR
8. FUEL CONTROL
9. FUEL PRESSURIZING AND
DUMP VALVE
10. FUEL PUMP
11. HYDRAULIC PUMP
9
12. FRONT ACCESSORY SECTION (N1)
Figure 1-5. Accessory Location
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1-15
Description
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VOL II
1-51. The generator is mounted on a gearbox which is mounted on the left-hand
(forward) side of the oil pump and accessory drive gearbox. The mounted gear-
box, Part No. 75P29, is used to change the engine pad ratio of 0.433:1 (N2) to
0.866:1 (N2). This provides the necessary step-up in rpm to run the generator.
1-52. There are three Allen head plugs in the engine accessory pad, one in the
mounting face, which is the oil supply hole and mates with a hole in the gearbox.
The other two are located in the accessory pad recess and are oil return holes
which drain into the accessory gear case of the engine.
Note
Be sure the Allen plugs in the accessory pad oil holes
have been removed before installing the gearbox.
1-53. As noted on the nameplate of the gearbox, use only the designated pad
gasket between the gearbox and engine accessory pad. This is to ensure that
oil will pass from the engine oil system into the gearbox.
1-54. The dc generator is air-cooled and receives the air through a scoop on the
bottom just forward of the generator. A two inch diameter aluminum tube ducts
the air aft to a flame proof flexible duct which connects to a special blast cap bolted
to the generator.
1-55. The generator speed versus the engine speed is given in table 1-1.
Table 1-1. Generator Speed versus Engine Speed
GENERATOR RPM
ENGINE RPM (N2)
% RPM (APPROX)
3500
4265
49
4500
5485
63
6200
7560
87
1-56. FUEL CONTROL. The fuel control is a hydromechanical unit manufactured
by Hamilton Standard, Model No. JFC-25-15.
1-57. The fuel control is mounted on six studs on the left-hand (aft side) pad of
the oil pump and accessory drive gearbox. The engine pad gear ratio is 0.433:1 (N2).
Refer to Engine Fuel System for details,'
1-16
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Approved For Release 2001/08/28 : Cd94DP751300300R000100050001-7
VOL II
Description
1-58. FUEL PRESSURIZING AND DUMP VALVE. The fuel pressurizing and
dump valve is a pressure operated check valve, manufactured by Pratt and Whitney
Aircraft.
1-59. It is mounted with two bolts on the outside of the engine diffuser case at
its rear bottom center. Three lines connect to the fuel pressurizing and dump
valve and must be disconnected at the time that the valve is being removed.
1-60. ENGINE FUEL PUMP. The fuel pump is a single gear type unit with booster
manufactured by Pesco Products, Inc.
1-61, The fuel pump is mounted on six studs on the right-hand (aft) side of the
oil pump and accessory drive gearbox. The engine pad gear ratio is 0.433:1
(N2).
1-62. HYDRAULIC PUMP. The hydraulic pump is a variable delivery type with
integral flow regulation, controlled by system pressure.
1-63. The hydraulic pump is mounted on six studs on the right-hand forward side
of the oil pump and accessory drive gearbox. The engine pad gear ratio is
0.433:1 (N2). The pressure, suction, and bypass connections are made through
flexible hoses and disconnect at the pump for engine removal.
1-64, TACHOMETER GENERATOR. The tachometer generator is a two pole
alternating current generator, Part No. AN5544-3.
1-65. The tachometer generator is mounted on the left-hand aft side of the oil
pump and accessory drive gearbox. The engine pad gear ratio is 0.481:1 (N2)
and indicates only the speed of the high pressure compressor rotor.
1-66. INTAKE AIR SYSTEM.
1-67. GENERAL.
1-68. The engine intake air system consists of two branches of air ducting ex-
tending from intakes at the sides of the airplane from fuselage station 267 to
fuselage station 365 where they converge to form a cylindrical inlet to the engine.
1-69. The interior of the intake ducts is zinc chromated from the aft end of the
leading edge to the aft end of the duct.
1-70. COMPONENTS.
1-71. The intake air system is composed of the scoop nose, fuselage station 267
to fuselage station 319, and the duct from fuselage station 319 to fuselage station 365.
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1-17
Description
Approved For Release 200401428 : CIA-RDP75600300R000100050001-7
VOL II
1-72. The essential differences between the scoop and the duct is one of location;
the scoop is outside the airplane and the duct is inside.
1-73. Fuselage station 319 bulkhead is the member which effects the scoop to
duct transition. The scoop, being external, has an outside skin which provides
fairing over the structural rings in addition to an inside skin which forms an air
duct wall. The duct is a simple sheet metal part encircled by rings at a six inch
spacing.
1-74. ENGINE INSTRUMENTS.
1-75. GENERAL.
1-76. The engine instruments consist of the tachometer, exhaust gas temperature,
engine oil pressure, engine oil temperature, and engine pressure ratio indicating
systems.
1-77. TACHOMETER INDICATING SYSTEM.
1-78. The tachometer system (See figure 1-6.) provides a visual indication in the
cockpit of the high pressure compressor rotor (N2) rpm in percent of cruise.
1-79. The tachometer generator is engine-driven and supplies the indicator with
a continuous signal during engine operation. The strength of the signal varies
with rpm, enabling the tachometer indicator to indicate rpm to the pilot.
1-80. Refer to -2-7 Maintenance Manual for electrical circuit wiring diagram.
1-81. TACHOMETER INDICATOR. This indicator is located on the upper right-
hand corner of the center instrument panel. It receives a signal from the tach-
ometer generator. The signal strength varies as a function of rpm. Indicator
calibration is based on the tachometer generator turning at 4200 rpm at 100 per-
cent. One hundred (100) percent on the indicator is equivalent to 8730 rpm of (Nz).
1-82. The indicator is composed of a mechanism which gives percentage indi-
cations of high pressure compressor rotor speed. The mechanism is enclosed
in a hermetically sealed case.
1-83. The instrument has a range of 0 to 110 percent and is driven by a three
phase permanent magnet rotor. Electrical connections to the indicator are made
at a single mating plug on the unit.
1-84. TACHOMETER GENERATOR. The tachometer generator, AN5544-3
(MIL-G-6027), is a two pole alternating current generator driven by the engine
at a maximum speed of 4200 rpm during military operation. The generator
supplies a signal to the indicator, enabling the indicator to indicate rpm. It is
mounted on the accessory case pad located on the left-hand aft side of the engine.
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1-18
Approved For Release 2001/08/28 : CIA-RDP75600300R000100050001-7
-2 -4 Description
TACHOMETER
DETAIL A
CENT ER INSTRUMENT PANEL
RIGHT SIDE
FS
211
FS
252
VOL II
i
11
,,.....".?.....
/
, -..,...
,
,
,
r)
,
---?
) .,.
-.....,, N.-..
j---............ .,
-.
, ...2
F:-.'.....%..' .,,''''''''''' ?
FS
365
'..--"
S
319
\
INSTRUMENT
INSTRUMENT
PANEL
PANEL
ml
A
TACHOMETER
GENERATOR
C
TACHOMETER INDICATOR
AppiaVetrFoi-14,1eaFt40.20104/060281PCLA4141BPJZSB00.100R
A
''DI>rYrd
TACHOMETER
GENERATOR
(ENGINE)
dA01150001-7
1 - 1 9
Description
Approved For Release 2001p8428 : CIA-RDP75600300R000100050001-7
VOL II
1-85. The pad gear ratio is 0.481:1 (Nz) high pressure compressor rotor.
1-86. EXHAUST GAS TEMPERATURE INDICATING SYSTEM.
1-87. The thermocouples used in the exhaust gas temperature system (See
figure 1-7.) supply a signal to a temperature indicator on the center instrument
panel.
1-88. THERMOCOUPLE PROBES. Thermocouple probes (6) are located in the
turbine frame just aft of the turbine wheel and are supplied on the engine. A read-
ing is obtained of the average temperature of all the thermocouples. The basic
system is chromel-alurnel and leads are supplied on the engine to a disconnect
point at the lower right-hand side of the main engine access panel, fuselage station
405. From fuselage station 405 disconnect the chromel-alumel leads go forward
to fuselage station 252 disconnect and then to the indicator amplifier in the cockpit.
1-89. Refer to -2-7 Maintenance Manual for electrical circuit wiring diagram.
1-90. EXHAUST GAS TEMPERATURE INDICATOR/AMPLIFIER. The indicator/
amplifier is located in the lower right-hand corner of the center instrument panel.
The unit provides a means of reading exhaust gas temperature and is calibrated
in degrees centigrade times 100 with a range from 0 to 10 . It is a transistorized?
servo driven, hermetically sealed unit. A cannon plug on the back of the instru-
ment provides connections for chromel and alumel leads. It is a Howell indicator,
Part No. BH185R-11B.
1-91.
ENGINE OIL PRESSURE INDICATING SYSTEM.
1-92. The engine oil pressure system (See figure 1-8.) consists of a pressure
transmitter, Edison Part No. 318-100, and an indicator, Edison Part No. 290-100K.
1-93. OIL PRESSURE TRANSMITTER. The transmitter is mounted on the left
side of the engine at approximately fuselage station 419. A flexible line connects the
transmitter with the vent port on the Nz case.
1-94. Identification of the vent port is shown on the engine accessory case.
1-95. Refer to -2-7 Maintenance Manual for electrical circuit wiring diagram.
1-96. OIL PRESSURE INDICATOR. The oil pressure indicator, Edison Part
No. 290-100K, is located on the right-hand side of the center instrument panel.
The range of the indicator is from 0 psi to 100 psi.
1-20
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Approved For Release 2001/08/28 : CA-RpP751300300R000100050001-7
VOL II
DETAR A
CENTER INSTRUMENT PANEL
RIGHT SIDE
FS
211
INSTRUMENT
PANEL
N 3 9
FS
FS
252
FS
365
TO POWER
EXHAUST GAS
TEMPERATURE
FS
508
Des cription
ENGINE
DISCONNECT
TO
B/W SYSTEM
C-26
FS
252
EGT INDICATOR/AMPLIFIER
D-.3
CH
AL
,?
THERMOCOUPLES Z:\
(EXHAUST
GAS TEMP)
INVERTER
CO BUS
1/2 A
C>
F E. G. T.
TEST
(INSTRUMENT
PANEL)
FS
319
F-2
CH
AL
E.G. T.
THERMO-
COUPLES
(ENGINE)
NOTE
E.G. T. THERMOCOUPLES
TYPICAL 6 PLACES
Figure 1-7. Exhaust Gas Temperature Indicating System
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1-.21
Description
Approved For Release 2001)8428 : CIA-RDP75600300R000100050001-7
VOL II
DETAIL
CENTER INSTRUMENT PANEL
RIGHT SIDE
FS
252
FS
365 7
-
FS \\1\
319
il
)
A
INSTRUMENT PANEL ?
INSTRUMENT TRANSFORMER
CIRCUIT BREAKER
(EQUIPMENT BAY)
Z-
2A
CND
1-22
LV
26V AC
OIL PRESSURE TRANSMITTER
OIL PRESSURE
INDICATOR
(COCKPIT)
INSTRUMENT/ ISOLATION
TRANSFORMER
(EQUIPMENT BAY)
1 /2A
0B
Figure 1-8. Engine Oil Pressure Indicating System
OIL PRESSURE
TRANSMITTER
(ON ENGINE)
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Approved For Release 2001/08/28 : CIA-4DP751300300R000100050001-7
VOL II
Description
1-97. The normal operating pressure range (green arc) is 40 to 55 psi and
indicates oil pressure pump discharge pressure. The indicator has red marks
at the 35 psi and 60 psi limits.
1-98. ENGINE OIL TEMPERATURE INDICATING SYSTEM.
1-99. The engine oil temperature system (See figure 1-9.) includes an electrical
resistance temperature bulb, MS28034-1, (MIL-B-7990) and an indicator, Lewis
Engineering Part No. 163B2 (MIL-I-7749).
1-100. OIL TEMPERATURE BULB. The oil temperature bulb is installed in
an adapter, Part No. 75P69, which is installed on the right-hand (forward) side
of the oil pump and accessory drive gearbox adjacent to the starter.
1-101. Refer to -2-7 Maintenance Manual for electrical circuit wiring diagram.
1-102. OIL TEMPERATURE INDICATOR. The oil temperature indicator, Lewis
Engineering Part No. 163B2, is located on the right-hand side of the center instru-
ment panel.
1-103. The range of the indicator is from -70?C to +150?C and registers the
temperature of the oil entering the engine. The maximum oil temperature (red
mark) is 125?C.
1-104. ENGINE PRESSURE RATIO INDICATING SYSTEM.
1-105. The engine pressure ratio system (See figure 1-10.) is designed to give
the pilot an indication of power or thrust for all the throttle settings. The system
consists of a transmitter and an indicator. The transmitter senses a pressure
ratio between engine inlet pressure (Pt2) and exhaust pressure (Pt7) and transmits
the ratio of these pressures to an indicator on the center instrument panel.
1-106. ENGINE INLET PRESSURE SENSING (P)*A pressure sensing probe
(Pt2), Pratt & Whitney Part No. 533039, is installed in the right side of the corn-
pressor inlet case at the 7 o'clock position. A quarter inch steel line is routed
aft to the forward side of the N2 accessory case where it crosses over to the left
side. At this point a short length of hose (1/4") connects to the airplane portion
of the Engine Pressure Ratio System taking it forward to the pressure ratio trans-
mitter located on the upper right side of the main wheel well.
1-107. ENGINE EXHAUST PRESSURE SENSING (Pt7). An exhaust pressure
sensing manifold containing four probes is located on the outer perimeter of the
exhaust case. A five-eighth inch steel line connects to the manifold at the
9 o'clock position and goes forward to the left trunnion (ball bat) position. At this
point a short length of hose (3/8") connects to th9 airplane portion of the Engine
Pressure Ratio System taking it forward to the pressure ratio transmitter located
in the main wheel well.
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1-23
Description Approved For Release/2141/08/28 : CIA-RDP75600300R000100050001-7
VOL II
OIL TEMPERATURE
INDICATOR
DETAIL A
CENTER INSTRUMENT PANEL
RIGHT SIDE
N.ss
INSTRUMENT
PANEL
FS
252
FS
65
FS
319
/
[OIL TEMPERATURE
INDICATOR (COCKPIT)
1-24
FS
211
A
...=?=1????=1M111?Mill
OIL TEMPERATURE
? BULB
?
LOWER
INS T RUMENT
PANEL
5A
28V DC
Figure 1-9. Engine Oil Temperature Indicating System
A
OIL TEMPERATURE
BULB (ON ENGINE)
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Approved For Release 2001/08/28 : Ct1-RIDP75B00300R000100050001-7
VOL II
Description
PRESSURE RATIO
INDICATOR
DETAIL A
CENTER INSTRUMENT PANEL
RIGHT SIDE
"MAIN
I DIAL
SYNCHRO
1
ISDUIABL
ISYNCHRO
A
A
FS
252
FS
FS 365
319
FS
211
PRESSURE RA TIO INDICA TOR
(INS TR TJME NT PANEL)
D31
TRANSMITTER
6"*.o
115 VAC
400., BQ) IA
ENG. PRESS.
RATIO
2 S 252
F 1 0
X
FS 3 9
TO ENG TO ENG
INLET PRESS. EXHAUST
PITOT PITOT
r 41r1
L _I I_ _i
5-
3-
K
4.1?????
L
26V4
--J
TRANSMIT TER (R SIDE
MAIN WHEEL WELL)
ApproveAti7elle;g :Pcinbisf56,6660hafittdowayoern
1-25
Description
Approved For Release 2004198428 : CIA-RDP75600300R000100050001-7
VOL II
1-108. ENGINE PRESSURE RATIO TRANSMITTER. The pressure ratio trans-
mitter, Honey-well Part No. DLG80D1 or FLG80D1, is installed in an insulated
container and is mounted on the upper right-hand side of the main wheel well. The
transmitter consists of a bellows actuated servoed-ratio-computer, a cam and gear
train, an amplifier, a two-phase motor, and a transmitting synchro. The mounting
rack has vibration isolators and pressure and electrical connections.
1-109. ENGINE PRESSURE RATIO INDICATOR. The indicator, Honeywell Part
No. JG151A6, is mounted on the right-hand side of the center instrument panel.
It contains two synchro receivers, a main dial pointer and a subdial pointer. The
subdial increases the readability and accuracy capability of the indicator.
1-110. ENGINE MOUNTING SYSTEM.
1-111. GENERAL.
1-112. The engine has mounting provisions (see figure 1-11.) on the compressor
intermediate case, turbine case, and the exhaust tailpipe, which consist of side
mounts, aft top mount, and tailpipe side mounts.
1-113. COMPONENTS DESCRIPTION.
1-114. SIDE MOUNTS. These are the main load carrying mounts and utilize
the socket-type fittings at the left and right sides of the compressor intermediate
case. A trunnion (ball bat) is inserted in each socket-type fitting when the engine
is mounted in the airplane. The trunnions attach to the airplane at fittings at
fuselage station 425. The attachment of each fitting is by means of two clamp-
type caps which are held closed by eyebolts.
1-115. These side mounts take inertial and side loads. Provisions for engine
expansion (axial load) is also provided in the right side trunnion.
1-116. AFT TOP MOUNT. The aft support consists of a yoke suspended from
a fitting fastened to the fuselage ring at fuselage station 509 (top centerline of
fuselage). The yoke, in turn, is attached by truss type links to the engine.
1-117. This type mount facilitates engine removal and installation. Vertical
adjustment of engine position is made by loosening a locknut and adjusting the
position of a rod end fitting in the yoke. The truss type links are attached to this
rod end fitting.
1-118. The top mount is capable of takihg only side and vertical loads and also
allows for engine expansion rearward.
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Approved For Release 2001/08/28 : ClAfr751300300R0001000500017 Description
VOL II
/
/
Figure 1-11. Engine Mount Installation (Sheet 1)
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1 -27
Description
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VOL II
2
F. S.
SECTION D-D
LH SHOWN - RH OPPOSITE
509
DETAIL A
SIDE MOUNT
LH SHOWN - RH OPPOSITE
EXCEPT AS SHOWNIN DETAIL E
15
17 18
DETAIL E
RH SIDE MOUNT
1-28
19
DETAIL C
TAILPIPE MOUNT
LH SHOWN - RH OPPOSITE
14
16A
15
DETAIL
TOP REAR MOUNT
1 SUPPORT FITTING
2 ENGINE
3 FRONT MOUNT SOCKET
4 3/4-16 BOLT
5 ADJUSTING BOLT
6 EYEBOLT 150 LB IN. TORQUE
7 ENGINE MOUNT CAP
8 LH TRUNNION (BALL BAT)
9 LINK
10 LH REAR ENGINE MOUNT FITTING
11 SUPPORT FITTING
12 REAR ENGINE MOUNT ROD END
13 REAR ENGINE MOUNT YOKE
14 RH REAR ENGINE MOUNT FITTING
15 EYEBOLT (75 LB IN. TORQUE)
16 RH TRUNNION (BALL BAT)
17 TAILPIPE
18 TAILPIPE TRACK
NOTE 19 TAILPIPE MOUNT ROLLER
A TIGHTEN ADJUSTING BOLTS SUFFICIENTLY TO CREATE ENOUGH
FRICTION TO SUPPORT WEIGHT OF TRUNNION AT ANY POSITION
TO WHICH IT IS SWUNG MANUALLY. NO SPECIFIC TORQUE REQUIRED.
A\ AFTER PROPER ADJUSTMENT HAS BEEN OBTAINED, LOCK SOCKET
AND BOLT IN PLACE. (TYPICAL 2 PLACES)
/3\ CENTER TRUNNION BETWEEN ENGINE MOUNT LUGS.
AFTER ENGINE IS INSTALLED, ROTATE LINKS OUTBOARD AND DOWN
TO CLEAR CONTROL CABLES.
Figure 1-11. Engine Mount Installation (Sheet 2)
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Approved For Release 2001/08/284V4-RDP751300300R000100050001-7
VOL II
Description
1-119. TAILPIPE SIDE MOUNTS. The tailpipe is supported by the engine and
by two rollers at approximately fuselage station 636.
1-120. Due to the long overhang of the tailpipe, a flexible joint is built into it
to prevent load pickup from fuselage deflection. As the engine expands due to
temperature rise, the rollers move aft in the aft fuselage section support tracks.
1-121. ENGINE STARTING SYSTEM.
1-122. GENERAL.
1-123. The engine starting system (see figure 1-12.) is manually controlled and
pneumatically operated. The pneumatic starter is installed on the center accessory
pad on the N2 case. An air adapter is attached to the starter inlet port for con-
nection of the ground starting equipment. The air from the ground starting equip-
ment passes through the nozzle causing the turbine wheel assembly to rotate. An
exducer which is a part of the turbine wheel assembly helps exhaust expended air
through the outlet port on the forward side of the starter.
1-124. There is an access to the starter area at the underside of the airplane.
1-125. STARTER.
1-126. The pneumatic starter operates on compressed air from a ground source.
1-127. It is mounted on six studs provided on the center accessory pad on the
Nz case. The engine pad gear ratio is 0.823:1 (Nz).
1-128. Air is introduced at the bottom of the starter and is directed against the
turbine wheel which drives a reduction gear assembly. The starter incorporates
an internal - engaging mechanism. This mechanism is composed of a ratchet
driven by the reduction gear assembly and a set of spring-loaded pawls attached
to the splined starter output shaft. This shaft engages a mating input drive shaft
on the engine and rotates when the engine is turning.
1-129. Leading particulars of Airesearch penumatic starter, Part No. 350520,
Model No. ATS140-29-1, are as follows:
Turbine Type
Reduction Gearing
Output Shaft Assembly Speed
Supply Air Requirements
Air Inlet Total Pressure
Air Inlet Total Temperature
Air Outlet Static Pressure
Inward - radial - flow
Helical and spur mesh
3300 rpm (min) to 3550 rpm (max) at
cutoff speed
45 psi abs (nominal)
288?C (550?F) (nominal)
14.7 psi abs
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1-29
Description
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VOL II
Rated Performance of 2840 rpm
Output Torque (min)
Air Flow (max)
Operating Limitations
Air Inlet Total Pressure
Air Inlet Temperature
System Lubrication
Oil
Capacity
Operating Level
Operating Temperature
Weight
182.5 lb ft.
110.0 lb per min.
60 psi abs (max)
371?C (700?F) (max)
Specification MIL-L-23699A or
MIL-L-7808
600 cc (Max - approximately)
300 cc (min)
177?C (350?F) (max)
27.5 lb (max)
Note
Airesearch starter, Part No. 210250, Model No. ATS140-16-1,
is Interchangeable with Aires earch starter, Part No. 350520.
1
STARTER NAMEPLATE
5
OUTPUT SHAFT ASSY
2
AIR OUTLET PORT
6
OIL FILLER PLUG
3
BALANCE LINE PORT
7
OIL DRAIN PLUG
4
ELECTRICAL RECEPTACLE
(NOT ILLUSTRATED)
(NOT USED ON THIS AIRPLANE)
8
AIR INLET PORT
2
1
8
7
6
Figure 1-12. Engine Pneu.matic Starter
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1-30 Changed 22 November 1965
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VOL II
1-130. OPERATION.
Description
1-131. At the beginning of the starting cycle, the ratchet commences rotation,
engaging the pawls and transmitting torque to the engine. While the starter, is
exerting torque on the engine the pawl and ratchet mechanism will remain in the
engaged position. After the engine reaches ignition speed, the starter continues
to assist the engine to accelerate until the cutoff speed of the starter is reached.
At this speed, the engine overruns the starter and the engaging mechanism ratchets
without transmission of torque. When the output shaft reaches pawl throwout
speed, the engaging mechanism is completely disengaged, the pawls being thrown
outward by centrifugal force so as to clear the ratchet without contact.
Note
Refer to Operational Checkout Section for starter
operational limits.
1-132., THROTTLE CONTROL SYSTEM.
1-133, GENERAL.
1-134. The throttle control system (see figure 1-13.) provides the mechanical
motion necessary to operate the throttle arm on the fuel control. It is cable oper-
ated by means of a drive pulley below the throttle in the left console in the cockpit.
A torque shaft connected to the throttle lever moves the drive pulley. The cables
are routed aft along the left side of the fuselage to the power control lever. A
short adjustable pushrod connects between the cable system termination and the
power control lever. The power control lever operates the throttle arm on the
fuel control.
1-135. COMPONENTS.
1-136. THROTTLE LEVER QUADRANT. The throttle lever quadrant is mounted
at the forward end of the left console. It is attached to the console and airplane
structure by screws and nuts and may be removed and replaced.
1-137. The flap control is also mounted in the quadrant.
1-138. The throttle lever is spring-loaded in the inboard direction so that the
forward motion, from OFF will cause it to drop into IDLE without forcing. Throttle
lever travel from IDLE to FULL is accomplished by straight forward movement.
1-139. A gate type throttle stop is provided to serve as a limit on takeoff power.
This stop will be adjusted to give approximately 93 to 94 percent rpm. In order to
go past this stop, the throttle lever must be moved outboard. The device will
automatically reset when the throttle is retarded.
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1-31
L-1.00090001.000t100?0089/dCIU-VI3 : 8Z/80/1?00Z eseeieu Jod peAwddv
FULL
78?
OFF
FS
252
.0 1.88
@ FULL
POWER OFF
FUEL CONTROL
LEVER ARM
1 5 0
IDLE
10030' --Ai
4) 310 "
9 0 ?
FULL
POWER
22? 30'
VERTICAL
CABLE TRAVEL - 3.82 INCHES
QUICK
DISCONNECTS
NOMINAL CABLE TENSION 60 (j.- 5)LB AT 70?F
COMPENSATION FOR TEMPERATURE CHANGE
a. IF AMBIENT TEMPERATURE IS ABOVE 70?F; ADD 0.35 LB TO
60 LB TENSION FOR EACH DEGREE ABOVE 70.
b. IF AMBIENT TEMPERATURE IS BELOW 70?F, SUBTRACT 0.35
LB FROM 60 LB TENSION FOR EACH DEGREE BELOW 70.
EXAMPLE: AMBIENT TEMPERATURE = 50?F (20?F LESS THAN 70?-w)
20 x 0.35 = 7 LB. CABLE TENSION AT 50?F = 53 LB
Figure 1-13. Throttle Control System
uopd-FaDsou
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VOL II
Description
1-140. THROTTLE LEVER. The throttle lever is mounted in the quadrant on
the left side of the cockpit. It actuates the throttle cable system by moving the
quadrant torque shaft that, in turn, moves a drive pulley.
1-141. A vernier wheel is installed just inboard of the throttle lever and pro-
vides for very small movements of the lever for power adjustment at high altitude.
1-142. Throttle friction may be regulated by a knob located in the center of the
vernier wheel on the left side console.
1-143. A toggle switch installed on top of the grip is used to operate the speed
brakes.
1-144. A pushbutton switch for the microphone is also installed on the grip.
1-145. The wiring for the grip extends out the bottom, joins with other wiring
from the quadrant and connects to a terminal strip at fuselage station 221.
1-146. POWER CONTROL LEVER. The power control lever is installed on the
fuel control and rotates through an arc of 90 degrees. Power lever OFF position
is set 22 degrees 30 minutes forward of vertical .
1-147. FUEL SYSTEM.
1-148. GENERAL.
1-149. The fuel system (see figure 1-14.) consists of a fuel-oil cooler, main
fuel strainer, engine fuel pump, hydromechanical fuel control, fuel flow totalizing
transmitter, fuel pressurizing and dump valve, and the fuel manifolds.
1-150. SYSTEM OPERATION.
1-151. Fuel from the aircrafts boost system is delivered to the fuel-oil cooler
where it passes through cooling coils to reduce the temperature of the oil. From
the fuel-oil cooler the fuel passes through a 60-mesh strainer and into the engine-
driven fuel pump. The two-stage pump delivers fuel at predetermined pressures
and quantities to the hydromechanical control.
1-152. Due to the characteristics of the engine it is necessary that fuel flow be
maintained within certain limits which vary depending upon operating conditions.
The variables sensed by the control are those of burner pressure, engine rpm,
and compressor inlet pressure. Subject to these variables, the control is capable
of accurately maintaining the desired engine rpm during steady state operation by
a governor droop system.
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1-33
Description
\\,..ACCESSORY
...
SEC TION (N1)
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VOL II
? LOW PRESSURE
COMPRESSOR
HIGH PRESSURE
COMPRESSOR
FUEL NOZZLES ? T URBINE
EMERGENCY
ACCESSORY
SECTION (N2)
e# ireff Z., #17 Ael'
RPM
SENSE
PT
Pb
FUEL
CONTROL
ENG
DRIVEN
THROTTLE
FUEL
PUMP
FUEL
FILTER
FUEL PRESS.
TRANSMITTER
DRAIN
INLET
sFUEL
1-34
4111` ENGINE COMPARTMENT
MINI,????
]
PRESSURIZING
AND DUMP VALVE
DRAIN
FUEL FLOW
TOTALIZING
TRANSMIT TER
AP. COCKPIT
FUEL
PRESSURE
INDICATOR
NORMAL H
EMER
EMERGENCY FUEL
CONTROL SWITCH
METERED FUEL PRESSURE MEIN SECONDARY FUEL FLOW
PUMP DISCHARGE PRESSURE MA"' PRESSURE SENSE
(UNMETE RED FUEL)
CZ= FUEL CONTROL BODY
PUMP INT ERSTAGE PRESSURE RE TURN
PRESSURE
UM= BYPASS
BOOST PUMP PRESSURE
="""*...? FUEL DRAIN
PRIMARY FUEL FLOW
0=3:730 DUMP SIGNAL
1:132
GAL FUEL
REMAINING
COUNTER
Figure 1-14. Engine Fuel System
Approved For Release 2001/08/28 : CIA-RDP75600300R000100050001-7
DRAIN
Approved For Release 2001/08/28 :_glAtRDP751300300R000100050001-7 Description
VOL II
1-153. During acceleration and starting, the control senses burner pressure,
and engine rpm, and as a result, schedules fuel flow to permit the maximum rate
of acceleration allowable within the engine temperature limits without compressor
surge while discouraging "rich blow-out".
1-154. During deceleration the fuel control schedules fuel flow as a function of
burner pressure to ensure the maintenance of sufficient fuel flow at the minimum
flow level to support combustion, thus preventing the condition called "lean die-out".
1-155. Metered fuel from the control passes through a fuel flow totalizing trans-
mitter, then into the fuel pressurizing and dump valve. The purpose of this valve
is to provide the division of flow between the primary and secondary nozzle orifices
to ensure proper fuel atomization. Also incorporated in the fuel pressurizing
and dump valve body is a dump valve which drains the fuel manifold at shutdown.
Fuel from the pressurizing and dump valve enters the engine fuel manifolds, which
provide separate paths for primary and secondary fuel flow, and finally into the
48 duel-orifice nozzles where it is atomized for burning in the combustion chambers.
1-156. Fuel System components are listed in table 1-2.
Table 1-2. Fuel System Components
NAME
PART NO.
VENDOR
Fuel-Oil Cooler
87880-3
Airesearch
Main Fuel Strainers
Strainer Assy (200 Mesh)
301385 or 748-1
Airline Welding Prod.
Strainer Assy (60 Mesh)
300720
Airline Welding Prod.
Engine Fuel Pump
023341-030-038P1
Pesco
Fuel Control
597797
Hamilton Standard
Fuel Flow Totalizing Transmitter
H222
Contractor
Fuel Pressurizing and Dump
476828
Pratt & Whitney
Valve
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1-35
Description
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-2-4
VOL II
1-157. COMPONENTS DESCRIPTION.
1-158. FUEL-OIL COOLER. The fuel-oil cooler, by transferring heat from
the engine oil, heats abnormally cool fuel, prevents ice formation in the fuel
strainer and fuel control, and provides additional oil cooling at high altitude.
For further details refer to the Lubrication, Scavenge, and Breather System.
1-159. MAIN FUEL STRAINERS. One fuel strainer (200-mesh) is located
near the outlet of the left-hand sump tank half at approximately fuselage station
389; the other fuel strainer (60-mesh) is located on the right side of the fuselage
structure slightly above the fuel-oil cooler. For further details refer to -2-5
Maintenance Manual.
1-160. ENGINE FUEL PUMP (See figure 1-15.) The function of the engine
fuel pump is to supply fuel under pressure to the engine fuel system.
1-161. The fuel pump is a high-pressure engine-driven pump consisting of one
gear-type pump element and one centrifugal -type booster element combined as
a single unit.
1-162. The booster element is located opposite the drive end of the pump and is
driven through a step-up gear train. A shear section is incorporated in the
centrifugal element drive.
1-163. A No. 40-mesh removable filter is located in the pump body between
the discharge side of the booster stage and the inlet side of the gear stage. The
filter is designed to bypass fuel in the event of clogging.
1-164. A fuel pressure relief valve is contained in the pump body on the dis-
charge side of the gear pump. The pressure relief is adjusted to limit the pres-
sure rise of the pump to a maximum pressure of 835 to 845 psi.
1-165. Operation. Fuel enters the booster stage through the pump inlet on the
end of the impeller casting where the fuel is boosted approximately 20 psi. The
fuel passes through the filter to the inlet side of the main pump and is discharged
through the outlet port to the fuel control. Excess fuel from the fuel control is
returned to the main pump through the return port and is recirculated within the
pump.
1-36
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Approved For Release 2001/08/28 : CIA-RDP75600300R000100050001-7
-2-4
VOL II
FUEL INLET
FUEL
FILTER
Description
FUEL PUMP
DRIVE SHAFT
ADAPTER
FUEL PRESSURE
TRANSMIT TER AND
SNUBBER CONNECTION
(REMOVE PLUG)
FUEL PUMP BODY ?
FUEL BYPASS
(FROM FUEL CONTROL) ?
Figure 1-15. Engine Fuel Pump
1-166. FUEL CONTROL. (JFC-25-15)
PRESSURE
RELIEF
VALVE
FUEL OUTLET
(TO FUEL CONTROL)
1-167. The fuel control (see figure 1-16.) is a hydromechanical unit designed
to meter fuel to the engine. Fuel is metered in the normal system according to
a predetermined flow schedule, which varies as a function of the pilot's throttle
lever position, burner pressure, and engine rpm.
1-168. The hydromechanical control is made up of hydraulic and mechanical
components. The operating forces required within the control are delivered by
servos which operate on approximately a 2:1 pressure ratio.
1-169. The fuel control also incorporates a standby emergency control system
designed to provide emergency control operation with minimum performance.
This system is manually selected by the pilot by means of a toggle switch. When
the system is in operation an amber light on the center instrument panel comes on.
1-170. NORMAL SYSTEM OPERATION. (See figure 1-17.) The normal operating
system of the fuel control consists of a metering system and a computing system.
The metering system alters the fuel supplied to the fuel control by the engine-driven
fuel pump to provide the engine thrust output required, but subject to engine oper-
ating limitations which are sensed and scheduled by the control computing section.
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1-37
Des cription
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VOL II
Pet
0
0
c4 1=1
f=1
Z Z
? fx1 1-1 ?d 0
Z 0 E-1 H
(121 0 r-11
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rxi z 9 Z