A-12 UTILITY FLIGHT MANUAL

• Document Type: FOIA
• Collection: FOIA Collection
• Document Number (FOIA) /ESDN (CREST): 06535936
• Release Decision: RIFPUB
• Original Classification: U
• Document Page Count: 104
• Document Creation Date: December 28, 2022
• Document Release Date: August 10, 2017
• Case Number: F-2014-00925
• Publication Date: September 15, 1965

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the current change. Insert latest changed and/or added pages;
destroy superseded pages.
NOTE: The portion of text affected by the change is indicated
by a vertical line in the outer margins of the page. 4-Indi-
cates deletion of text.
Issue Code C-2 Changed 15 June 1968
A Approved for Release: 2017/07/25 C06535936
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COPY NO.
Page 1 of 1
TDC No. 13
15 June 1968
A-12 FLIGHT MANUAL
TECHNICAL DATA CHANGE
This TDC transmits revised pages which supersede previously furnished
pages for the Flight Manual dated 15 October 1967. All previously issued
TDC's are incorporated.
In addition, this TDC includes':
a. Rapid Deployment to ARCP data.
b. Revised presentation of normal climb performance
c. Revised presentation of cruise performance for long range and
high altitude cruise (1956 ARDC and "MEAN TROPIC"
atmospheres)
d. Revised single engine descent data for various speeds, powers,
and for both 1956 ARDC and "Mean Tropic" atmospheres.
e. Minor descriptive material.
Previously issued checklist changes conform with procedures supplied
in this manual.
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the current change. Insert latest changed and/or added pages;
destroy superseded pages.
NOTE: The portion of text affected by the change is indicated
by a vertical line in the outer margins of the page.
cotes deletion of text.
Changed 15 June 1968
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NDFX
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INDEX-in ORIGINAL
Page No. Issue
Page No. Issue
Page No. Issue
* The asterisk indicates pages changed, added, or deleted by
the current change. Insert latest changed and/or added pages;
destroy superseded pages.
NOTE: The portion of text affected by the change is indicated
by a vertical line in the outer margins of the page. -I-Indi-
cates deletion of text.
Issue Code C-2
Changed 15 June 1968
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A-12.
TECHNICAL DATA CHANGE SUMMARY
TDC Date Purpose
Status
No. 1
10-16-67
Est. Tropical Atmosphere Climb Performance
Superseded
by TDC 3
No. 2
1-26-68
Emergency Forward Transfer
Inc.
No. 3
2-05-68
Revised Climb and Cruise Performance
Inc.
(1956 ARDC Atmosphere & "MEAN
TROPIC" Atmosphere)
No. 4A
3-04-68
Time Limits ,& EGT Limits
Inc.
No. 5
3-01-68
Supersonic Cruise Flight Characteristics
Inc.
No. 6
3-05-68
Tire Limits
Inc.
No. 7
3-06-68
Climb and Cruise Performance
Inc.
No. 8
3-15-68
Increase Chute Deploy Limit 210 KIAS
Inc.
No. 9
3-16-68
Transmit Printed Change Dated 3-15-68
Inc.
No. 10
5-7-68
Rapid Deployment to ARCP
Inc.
No. 11
5-10-68
Normal Operation for Descent & Engine
Shutdown
Inc.
No. 12
5-16-68
Normal Climb Performance Revised
Inc.
No. 13
6-15-68
Transmit Printed Change Dated 6-15-68
Inc.
Changed 15 June 1968
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A-12
SECURITY INFORMATION
SPECIFIC INSTRUCTIONS FOR SAFEGUARDING THIS INFORMATION
1. This document contains information affecting the national defense
of the United States within the meaning of the Espionage Laws
Title 18, USC Section 793 and 794. The transmission or the
revelation of its contents in any manner to unauthorized persons
is prohibited by law.
The nature of this document is such that dissemination and
handling will be carried out with strict adherence to the following
policies:
a. Distribution will be controlled on a strict, officially
established "need-to-know" basis.
b. Strict accountability of each document will be maintained.
c. This document will be controlled in such a fashion
to prevent its loss, destruction, or falling into the
hands of unauthorized persons.
2. In the event this document is lost or is subject to unauthorized
disclosure or other possible subjection to compromise of
classified information, such fact will be promptly reported to the
authority responsible for the custody of the material for
appropriate action.
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A - 12
FOREWORD
The primary purpose of this manual is to provide systems descriptions,
operating procedures, limitations, and other information necessary for
operation of these aircraft. Experience and basic aircraft familiarity
gained to date has been recognized in the preparation of this manual.
This manual will be reviewed and changed periodically by the
manufacturers test organization to reflect information gained from
further tests and operating experience. Comments, corrections, and/or
questions regarding this manual are welcome. They may be forwarded
through the manufacturer's senior engineer at the test site or directly
to the flight manuals representatives.
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Al2
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SECTION
PAGE
� I Description 1-1
II Normal Procedures 2-1
DI Emergency Procedures 3-1
IV Auxiliary Equipment 4-i
� V Operating Limitations 5-1
VI Flight Characteristics 6-1
VIE Systems Operation 7 -1
IX All Weather Operation 9-1
Appendix: Performance Data A-1
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A-12
4-26-66
F200-30
Iv
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Al2
Section I
TABLE OF CONTENTS
Page
Page
The Aircraft
1-1
Emergency Equipment
1-67
Engine And Afterburner
1-7
Landing Gear System
1-67
Air Inlet System
1-19
Nosewheel Steering System
1-69
Fuel Supply System
1-26
Wheel Brake System
1-71
Air Refueling System
1-35
Drag Chute System
1-71
Electrical Power Supply System
1-39
Air Conditioning and
Hydraulic Power Supply System
1-44
Pressurization System
1-73
I Flight Control System
1-47
Oxygen Systems and Personal
Automatic Flight Control System
1-54
Equipment
1-81
Stability Augmentation System
1-55
Windshield
1-85
Pitot Static System
1-59
Canopy
1-87
Air Data Computer
1-61
Ejection Seat
1-89
Instruments
1-63
THE AIRCRAFT
AIRCRAFT DIMENSIONS
The A-12 is a delta wing, single place air-
craft powered by two axial flow bleed bypass
turbojet engines with afterburners. The
aircraft is built by the Lockheed-California
Company and is designed to operate at very
high altitudes and at high supersonic speeds.
Some notable features of the aircraft are
very thin delta wings, twin canted rudders
mounted on the top of the engine nacelles,
and a pronounced fuselage chine extending
from the nose to the leading edge of the wing.
The propulsion system uses movable spikes
to vary inlet geometry. The surface controls
are elevons and rudders, operated by irre-
versible actuators with artificial pilot con-
trol feel. A single-point pressure refueling
system is installed for ground and in-flight
refueling. A drag chute is provided to re-
duce landing roll.
Changed 15 March 1968
The overall aircraft dimensions are as
follows:
Wing Span
Length (overall)
Height (to top of
vertical stabilizer)
Tread (MLG center
wheels)
AIRCRAFT GROSS WEIGHT
55.62 ft.
101.6 ft.
18.45 ft.
16.67 ft.
The ramp gross weights of these aircraft
may vary from approximately 122,900 lb.
to 124,600 lb. with 10,590 gallons of fuel.
This is based on zero fuel weights between
54,600 lb. and 56,300 lb. , fuel density of
6.45 lb. per gallon, and varying equipment
loading configurations.
1-1
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�
OUTBOARD ELEVON
SPIKE BLEED AIR OUTLET�
FWD BYPASS �
�SPIKE
PITOT AND HF ANTENNA
� ADF LOOP ANTENNA
�TRANSLATOR-ARC-50 UHF ( R.N. SIDE)
� RCVR-XMTR-ARC-50 UHF ( L. H. SIDE)
�RETRACTABLE UHF ANTENNA
�R TACAN ANTENNA
�AR RECEPTACLE DOORS
� EX.PWR.RECEPT.
RUDDER
� ADF SENSE ANTENNA
�PITCH AND YAW GYRO
CHINE EQUIPMENT BOX ( L. H. SIDE)
ROLL RATE GYRO AND LATERAL ACCELEROMETER
BATTERIES
LEFT TACAN ANTENNA
LANDING AND TAXI LIGHTS
NITROGEN TANKS
AIR CONDITIONING BAY AND
INERTIAL NAVIGATION COMPONENTS
�LIQUID OXYGEN TANKS
�Q-BAY
E-BAY
�EJECTION SEAT
� COMPUTER AIR INLET CONTROL
UHF-ADF ANTENNA
� HF TRANSCIEVER
�ANTENNA TUNING UNIT HF
�FRS COMPASS TRANSMITTER
�HF ANTENNA COIL`
�
rn
rn
�INBOARD ELEVON
�ROLL AND PITCH MIXER
� YAW SERVOS RUDDER TRIM >
�EJECTOR FLAPS PO
C)
3
rn
�4
TERTIARY DOORS�
ELEVON ACTUATORS �
E-BAY CONTAINS THE FOLLOWING ITEMS:
a. Air data computer
b. Air data transducer
c. Tacan RCVR-XMTR
d. Inverter (UHF power)
e. Auto pilot
f. Stability augmentation sys.
g. IFF
h. ADF
i. Birdwatcher
j. Temperature control
k. Flight reference gyros
I. Air refuel signal amplifier
m. Rate gyro
n. Back up pitch gyro
�
�
�
NI 01,1 D S
I-1
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A-12
SECTION I
INSTRUMENT PANEL
75
74
73
72
71
11 13 15
0 12 14 16 1718
34
6 8 �
7 9
62
68 65 63 61
70 17 69 67 6664
1 AIR CONDITIONING CONTROL PANEL
2 AIRSPEED-MACH METER
3 BEARING DISTANCE HEADING
INDICATOR (BDHI)
4 ANIARC-50RANGE INDICATOR
5 INS DISTANCE TO GO-GROUND
SPEED INDICATOR
6 WINDSHIELD DEICER SWITCH
7 RAIN REMOVAL SPRAY BUTTON
8 DRAG CHUTE HANDLE
9 AIR REFUEL READY-DISC
LIGHT AND SWITCH
10 ATTITUDE INDICATOR
11 DE-ICING WARNING LIGHT
12 MASTER CAUTION LIGHT
13 ALTIMETER
14 PERISCOPE VIEWING SCREEN27
15 EWS LIGHTS 28
16 COMPRESSOR INLET STATIC 29
PRESSURE GAGE 30
17 FUEL DERICHMENT WARNING 31
LIGHTS (2) 32
18 VERTICAL SPEED INDICATOR 33
19
20
21
22 TRIPLE DISPLAY INDICATOR 38
23 IGNITER PURGE SWITCH
24
25
26
COMPRESSOR INLET
34
TEMPERATURE GAGE
35
ELAPSED TIME CLOCK
36
FIRE WARNING LIGHTS
37
60
59
58
57
56
55
54
TACHOMETERS
EXHAUST GAS
TEMPERATURE INDICATORS
EXHAUST NOZZLE
POSITION INDICATORS
FUEL TANK SWITCHES
FUEL FORWARD TRANSFER SWITCH 44
QUAD HYDRAULIC QUANTITY
AIR REFUEL SWITCH' 45
LIQUID NITROGEN QTY INDICATOR
FUEL TANK PRESURE, INDICATOR 46
RIGHT FORWARD PANEL
FUEL DUMP SVVITCH 47
PUMP RELEASE SWITCH , 48
FUEL QUANTITY INDICATOR 49
ILS PANEL
TEST N AND TANK LIGHT SWITCH 51
52
39
40
41
42
43
20 21 2223
24 25
26 27
44
45
46
47 37
48 42 40 38 35
43 41 39 36
49
50
53
54
55
51 56
57
58
59
60
53 61
62
63
52
FUEL FLOW INDICATORS 64
FWD BYPASS POSITION INDICATOR
OIL PRESSURE INDICATORS 65
SPIKE POSITION INDICATOR 66
HYDRAULIC SYSTEM (A AND B) 67
PRESSURE GAGE 68
HYDRAULIC SYSTEM (L AND R) 69
PRESSURE GAGE
COCKPIT PRESSURE SCHEDULE 70
SWITCH
EMERGENCY FUEL SHUTOFF 71
SWITCHES 72
BACKUP PITCH DAMPER SWITCH 73
A-13A CLOCK 74
ANNUNCIATOR PANELS
PITCH LOGIC OVERRIDE SWITCH 75
YAW LOGIC OVERRIDE SWITCH
LANDING GEAR RELEASE HANDLE
Figure 1-2
28 29
30
33
31
32
LOWER CIRCUIT BREAKER PANEL
RUDDER PEDAL ADJUST HANDLE
NOSE AIR OFF HANDLE
TRIM POWER SWITCH
HYDRAULIC RESERVE OIL SWITCH
PITOT HEAT SWITCH
SURFACE LIMITER HANDLE
INS DEST AND SELECT PANEL
COURSE INDICATOR
EMER SPIKE SWITCH
TURN AND SLIP INDICATOR
SPIKE AND BYPASS CONTROL
PANEL .
STANDBY ATTITUDE INDICATOR
RESTART SWITCHES �
FUEL DERICHMENT ARMING SWITCH
PERISCOPE CONTROL PANEL '
EXHAUST GAS TEMPERATURE,
TRIM SWITCHES., �
LANDING GEAR DOWN
INDICATOR LIGHTS �
LEFT FORWARD. PANEL
LANDING AND TAXI LIGHT SWITCH
ALT STEER AND BRAKE SWITCH
LANDING GEAR WARNING
CUTOUT BUTTON �
P I TCH-ROLL-YAW TRIM. INDICATORS
FZ00-14(i)
Changed 15 March 1968
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1-3
SECTION I
COCKPIT LEFT SIDE
2Th
4%0
20
13
21
20
22
0
STRY All
EAST LRECT
UNMIIIF
16
15
14
18
17
DEFOG
INCREASE
HOLD
OFF
RCN LTS
9 OH
US ITS
FE BRIGHT
INSTR ITS
OFF BRIGIII
PANI I ITS
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A-12
� RH RUDDER
SYNCHRONIZER
OXY
SYSI
H INT
F oVOL 0 MR
0
�� ROLL TRIM
00
s/YS TEN.
ON
OXY
SYS 2
0 TONE
�IF MAIN
80111
ADE
�RCA:
INT
0
CONI
10
0-BAY SPECIAL PACKAGES PANEL
Vol
Figure 1-3
1 AFT BYPASS INDICATOR LIGHTS
2 AFT BYPASS SWITCHES
3 RUDDER SYNCHRONIZER SWITCH
4 ROLL TRIM SWITCH
5 THROTTLE QUADRANT
6 OXYGEN PANEL
7 CANOPY JETTISON HANDLE
8 UHF COMMAND RADIO TRANSLATOR
CONTROL PANEL
9 UHF COMMAND RECEIVER TRANSMITTER
CONTROL PANEL
10 Q-BAY EQUIPMENT PANEL (NOT SHOWN)
11 SUIT VENTILATION BOOST LEVER
12 HF RADIO CONTROL PANEL
13 IFFIS IF CONTROL PANEL
14 PANEL LIGHTS SWITCH
15 INSTRUMENT LIGHTS SWITCH
16 IFR VOLUME CONTROL
17 COMMUNICATION SELECTOR SWITCH
18 BEACON-FUSELAGE LIGHTS SELECTOR SWITCH
19 DE-FOG SWITCH
20 HF MUTE-UNMUTE SWITCH AND LIGHT
21 STANDBY ATTITUDE INDICATOR FAST
ERECT SWITCH
22 RADIO BEACON-SELECTOR SWITCH
rzoo-17(f)
1 -4
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A- 1 2
SECTION I
COCKPIT RIGHT SIDE
1. SAS CONTROL PANEL
2. PITOT PRESSURE SELECTOR LEVER
3. NOSE HATCH SEAL LEVER
4. CANOPY SEAL LEVER
5. SEAT AND CANOPY SAFETY PIN STOWAGE
6. AUTO PILOT SELECTOR SWITCH
7. BDHI NO. 1 NEEDLE SELECTOR SWITCH
8. FLIGHT RECORDER SWITCH
9. FLOOD LIGHT SWITCH
10. FACE PLATE HEAT SWITCH
11. B-W AND SIP CONTROL PANEL
12. FRS CONTROL PANEL
13. ADF RECEIVER CONTROL PANEL
14. TACAN CONTROL PANEL
15. INS CONTROL PANEL
16. AUTO PILOT CONTROL PANEL
16
15
14
13
12
11
F' ITCH''TOIL PAW
ON ON; .
I A, ON
OFCYCIE
_
, LIT TEST
STAB 0 AUG
MACH.
�/11/* , ON
- , KEAN. AUTO HEADING
' HOLD . . NAv ' HOLD ' ,
AUTOPILOT. _. ,. ,
: TRIM - ROI.L
TURN ON
PUSF)S
F ITEINTI.
Lc;
N
t("1 HST ILAT ION /Fly L:(1-7L)
pos
J�
� LOT- LONG� -
.,f;Ts;, , � ABLE':
N LOT ' LONG
VOI
OP, �
R '
0" FRFOU
Of Ar,11
l00
Gikirq
TM OFF OFF
0 0 0 ,
COT A 1CTIIii TA: 'CODE
Figure 1-4
FLIGHT
RECORDER
OS
).)
OFF
noob unrirs
5
----------- 6
10
F200-18(f)
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1-5
s- I a-n2T1
�
�
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18 17 16
15
14
1 AIR INTAKE
2 BLEED BYPASS VALVES
3 9-STAGE COMPRESSOR SECTION
4 STARTING BLEED VALVES
5 CHEMICAL IGNITION (TEB) RESERVOIR
6 BLEED BYPASS TUBES (6)
7 BURNER CANS (8)
, 8 2-STAGE TURBINE
9 SPRAY BARS (4)
13
12
10 AFTERBURNER LINER
11 VARIABLE AREA EXHAUST NOZZLE
12 EXHAUST NOZZLE ACTUATORS (4)
13 FLAME HOLDERS (4)
14 MAIN ENGINE GEARBOX
15 ENGINE FUEL CONTROL
16 REMOTE GEARBOX SHAFT FITTING
17 AFTERBURNER FUEL CONTROL (RIGHT SIDE HIDDEN)
18 BLEED BYPASS VALVES ACTUATOR CYLINDERS (4)
� 0 � �
�
�
I NOLL Das
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SECTION I
A-12
' NOTE
See the weight and balance hand-
book, T. 0. 1-1B-40 for information
regarding specific aircraft and
equipment configurations.
ENGINE AND AFTERBURNER
Thrust is supplied by two Pratt and Whitney
JT11D-20A bleed bypass turbojet engines
with afterburners. The interim maximum
afterburning static thrust rating of each
engine is 31,500 pounds at sea level with
standard day conditions. The engines are
designed for continuous maximum thrust op-
eration at the high compressor inlet tem-
peratures associated with high Mach number
and high altitude operation. There is no
time limit on maximum thrust operation.
The engine has a single rotor, nine stage,
8:1 pressure ratio compressor utilizing a
compressor bleed bypass cycle for high
Mach number operation. The bypass sys-
tem bleeds air from the fourth stage of the
compressor, and six external tubes duct
the air around the rear stages of the com-
bustion section and the turbine. The air
reenters the turbine exhaust ahead of the
afterburner and is used for increased
thrust augmentation. When the engine goes
into bypass operation, the afterburner fuel
control resets to furnish additional fuel to
the afterburner. The transition to bypass
operation is scheduled by the main fuel
control as a function of compressor inlet
temperature (CIT) and engine speed. The
transition normally occurs at a CIT of ap-
proximately 150o to 190 C, corresponding
to a Mach number range of 2.2 to 2.3.
Engine speed on the ground, or at low Mach
numbers, varies with throttle movement
from IDLE to a position slightly below
MILITARY thrust. Between this throttle
position and the maximum afterburning
thrust position the main fuel control sched-
ules engine speed as a function of CIT and
modulates the variable area exhaust nozzle
to maintain approximately constant rpm.
Throttle movement in the afterburning
range varies the afterburner fuel flow, noz-
zle position and thrust. At high Mach num-
ber and constant inlet conditions, the engine
speed is essentially constant for all throttle
positions down to and including IDLE. At a
fixed throttle position, the engine speed will
vary according to schedule when Mach num-
ber and CIT change.
The engine has a two stage turbine. Com-
pressor discharge air cools the first stage
and is then returned to the exhaust gas
stream. Turbine discharge temperatures
are monitored by indications of exhaust gas
temperatures. A chemical ignition system
is used to ignite the low vapor pressure
fuel. .A separate engine driven hydraulic
system, using fuel as hydraulic fluid, op-
erates the exhaust nozzle, chemical ignition
system dump, compressor bypass and
staking bleed systems. The main fuel
pump, engine hydraulic pump and tach-
ometer are driven by the main engine gear-
box. The afterburner fuel pump is powered
by an air turbine driven by compressor dis-
charge air.. The ac generator, aircraft
hydraulic pumps and fuel circulating pump
are located on a remote gearbox driven by
the engine power takeoff pad through a re-
duction gearbox.
ENGINE THRUST RATINGS
The engine thrust ratings are defined by the
power lever position at the main fuel control.
The power lever is mechanically linked to
the throttle, providing a relationship be-
tween throttle position and thrust ratings.
1-7
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A-12
ENGINE AND NB FUEL SYSTEM
COMPRESSOR
INLET TEMP
INDICATOR
CLOSE
STATIC PRESS
INDICATOR
II ni �
COMPRESSOR
BLEED ACTUATOR
STARTING
BLEED ACTUATOR
i�no
CLOSE
COMPR
BLEED
PILOT
VALVE
ENGINE
DRIVEN
MAIN
FUEL PUMP
STARTING
BLEED
PILOT VALVE
BOOST
PUMP
OIL
PRESSURE
INDICATOR
MAIN
GEARBOX
_Jil It
FUEL/OIL
COOLER
MAIN FUEL
CONTROL
FROM
SMART
vALVE
WINDMILL
BYPASS AND'
DUMP VALVE
C)I FILTER
A/B DETENT
THROTTLE
OFF
HYDRAULIC
PUMP
TO
SMART
VALVE
SOLENOID
VALVE
FUEL PRESS LOW_1
EGT
INDICATOR
FUEL
DER I CH
ARM
EXHAUST
NOZZLE
ACTUATOR
EXHAUST
NOZZLE
POSITION
INDICATOR
immo
CLOSELN-
EXHAUST NOZZLE
CONTROL VALVE
DERICH TO SMART VALVE
WARNING WHEN A/B IS OFF
LIGHTS (2)
ENGINE
PRESSURE
REGULATOR
A/B FUEL
CONTROL
AIR IN
'
4:FROM
ENGINE
MAIN FUEL CONTROL COMPONENTS
PRESSURE REGULATOR VALVE
FUEL DENSITY SELECTOR
FROM THROTTLE VALVE
FUEL , PILOT VALVE
SYSTEM BURNER CAN LIMIT VALVE
FLOWMETER
FUEL FLOW
INDICATOR
� CODE
1515EZZLI
FUEL FLOW BURNERS
FUEL HYDRAULIC PRESS
FUEL DERI CH SYSTEM
ELECTR ICAL
Figure 1-6
AFTERBURNER FUEL CONTROL COMPONENTS
THROTTLE VALVE
PUMP REGULATOR
RECIRCULATING BYPASS VALVE
PRESSURE REGULATOR VALVE
PEAK THROTTLE VALVE
Fzoo-lo(e)
1 -8
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SECTION I
A-12
Maximum Rated Thrust
Maximum rated thrust is obtained in after-
burning by placing the throttle against the
quadrant forward stop.
Minimum Afterburning Thrust
MINIMUM afterburning thrust is ob-
tained with the throttle just forward of the
quadrant MILITARY thrust detent. After-
burner ignition is automatically actuated
when the throttle is advanced past the detent
and afterburner fuel flow is terminated when
the throttle is retarded aft of the detent.
Afterburning fuel flow and thrust are mod-
ulated by moving the throttle between the
detent and the quadrant forward stop. Mini-
mum afterburning is approximately 85% of
maximum afterburning thrust at sea level
and approximately 55% at high altitude. The
basic engine operates at MILITARY rated
thrust during all afterburning operation.
Military Rated Thrust
MILITARY rated thrust is the maximum
non-afterburning thrust and is obtained by
placing the throttle at the MILITARY
thrust detent on the quadrant.
Idle
IDLE is a throttle position for minimum
�thrust operation. It is not an engine rating.
Minimum thrust is always obtained when
the throttle is at the IDLE stop on the quad-
rant.
Start
There is no distinct throttle position for
starting. Starting is accomplished by mov-
ing the throttle from OFF to the IDLE posi-
tion as the proper engine speed is reached.
This directs fuel to the engine burners by
actuation of the windmill bypass valve and
actuates the chemical ignition system.
Off
The aft stop on the quadrant is the engine
OFF throttle position. In this position, the
windmill bypass valve cuts off fuel to the
burners and bypasses it back to the aircraft
system. This provides engine oil, fuel pump
and fuel hydraulic pump cooling when an
engine is windrnilling at high Mach number.
ENGINE FUEL SYSTEM
Engine fuel system components include the
engine driven fuel pump, main fuel control,
windmill bypass valve and variable area fuel
nozzles in the main burner section.
Main Fuel Pump
The engine driven main fuel pump is a two
stage unit. The first stage consists of a
single centrifugal pump acting as a boost
stage. The second stage consists of two
parallel gear type pumps with discharge
check valves. The parallel pump and check
valve arrangement permits one pump to op-
erate in the event the other fails. The pump
discharge pressure is determined by the re-
gulating and metering function of the main
fuel control. The maximum discharge pres-
sure is approximately 900 psi. A relief
valve is provided in the second stage dis-
charge to prevent excessive fuel system
pressure.
Main Fuel Control
The main fuel control meters main burner
fuel flow, controls the bleed bypass and
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SECTION I
A-12
start bleed valves and controls exhaust
nozzle modulation. Thrust is regulated as
a function of throttle position, compressor
inlet air temperature, main burner pressure
and engine speed. The bypass and start
bleed valve positions are controlled as a
function of engine speed biased by CIT. For
steady state inlet conditions at high Mach
number, the control provides essentially a
constant engine speed at all throttle posi-
tions down to and including IDLE. On the
ground and at lower Mach numbers, engine
speed varies with throttle position from
slightly below MILITARY down to IDLE.
Afterburner operation is always at MILI-
TARY rated engine speed and EGT. The
fuel control is provided with a pilot op-
erated trimmer for EGT regulation. There
is no emergency fuel control system.
Windmill Bypass and Dump Valve
The windmill bypass and dump valve directs
fuel to the engine burners for normal oper-
ation or bypasses fuel to the recirculation
system for accessory, engine component
and engine oil cooling during windmilling
operation. The valve is actuated by sig-
nals from the main fuel control. The valve
also opens to drain the fuel manifold when
the engine is shut down.
Fuel Nozzles
The engine has eight can-annular type com-
bustion chambers with forty-eight variable
area dual orifice fuel nozzles in clusters of
six nozzles per burner. The nozzles have
a fixed area primary metering orifice and
a variable area secondary metering orifice,
discharging through a common opening. The
secondary orifice opens as a function of pri-
mary orifice pressure drop.
ENGINE FUEL DERICHMENT SYSTEM
The derichment system provides protection
against severe turbine over-temperature
during high altitude operation. When EGT
indicates 860oC or more with the system
armed, the fuel:air ratio in the engine
burner cans is reduced, or deriched, below
normal values. This is accomplished by a
solenoid operated valve and orifice which
bypasses metered engine fuel from the fuel
oil cooler to the afterburner fuel pump inlet.
The solenoid valve is actuated by a signal
from the EGT gage when 860 C is reached.
Once actuated, it remains open until the
system is turned off. Two warning lights
are provided to indicate when the valve is
open. Power for the derich circuits is pro-
vided from the essential dc bus.
Fuel Derichment Arming Switch
A two position fuel derichment arming switch
is located on the left side of the instrument
panel. In the ARM (up) position the derich-
ment circuits are armed and the respective
derichment solenoid valve will open auto-
matically and remain open if the EGT
reaches 860 C. In the OFF position the
derichment solenoid valve is closed and
the system can not provide derichment flow.
Power is furnished from the essential d. c.
bus.
Fuel Derichment Warning Lights
The fuel derichment warning lights, located
on the left and upper center of the instru-
ment panel, illuminate and remain on while
the derichment solenoid valve is open. The
lights will be extinguished when the arming
switch is placed in the OFF position and
will remain extinguished when the arming
switch is reset to ARM if both EGTs are
below 860�C.
1-10
Changed 15 March 1968
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SECTION I
A-12
WARNING
In the event of derichment the arm-
ing switch must be placed in the
OFF position prior to relighting
the afterburner to prevent engine
speed suppression and subsequent
inlet unstart. If engine flameout is
experienced with inlet unstart the
arming switch should also be placed
to OFF prior to relighting the engine.
. Derichment at sea level will cause
a thrust loss of approximately 5%
if in maximum afterburning or 7%
if at Military. Approximately 45%
loss in thrust and 600 rpm speed
suppression will occur during cruise
with maximum afterburning.
AFTERBURNER FUEL SYSTEM
Afterburner fuel system components include
the centrifugal afterburner fuel pump, after-
burner fuel control and spray bars.
Afterburner Fuel Pump
The afterburner fuel pump is a high speed,
single stage centrifugal pump. The pump
is driven by an air turbine which is op-
erated by engine compressor discharge air.
The compressor discharge air supply is re-
gulated by a butterfly valve in response to
the demand of the afterburner fuel control.
The turbine is protected from over speed by
an aero-dynamic speed limiting air dis-
charge venturi.
Afterburner Fuel Control
The afterburner fuel control is a hydro-
mechanical fuel control which schedules
metered fuel flow as a function of throttle
position, main burner pressure and com-
pressor inlet temperature. Fuel flow is
metered on a predetermined schedule to
provide fuel into both zones of the after-
burner spray bars simultaneously. The
control incorporates a reset mechanism
which increases the afterburner fuel flow
when the bypass valves open and decreases
the fuel flow when the valves close.
ENGINE FUEL HYDRAULIC SYSTEM
Each engine is provided with a fuel hy-
draulic system for actuation of the after-
burner exhaust nozzle and the start and by-
pass bleed valves. Engine hydraulic sys-
tem pressure is also required to dump the
unused chemical ignition fluid. Pressure
is supplied by a high temperature, engine
driven, variable delivery, piston type
pump. The pump maintains system pres-
sures up to 2500 psi with a maximum flow
of 50 gpm for transient requirements.
Engine fuel is supplied to the pump from
the main fuel pump boost stage. Some high
pressure fuel is diverted from the hydraulic
system to cool the non-afterburning recir-
culation line and the windmill bypass valve
discharge line. This fuel is returned to the
aircraft system. Low pressure fuel from
the hydraulic pump case is returned to the
main fuel pump boost stage. Hydraulic
system loop cooling is provided by the
compensating fuel supplied from the main
fuel pump.
Exhaust Nozzle Actuation
The exhaust nozzle control and actuation
system is composed of four actuators to
move the exhaust nozzle, and an exhaust
nozzle control modulating the hydraulic
pressure to the actuators in response to
engine speed signals from the main fuel
control. The exhaust nozzle control is
mounted on the aft portion of the engine.
A pressure regulator is contained in a
separate unit located near the exhaust
nozzle control.
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SECTION I A-12
START BLEED AND BYPASS VALVE ACTUATION
Engine Speed-RPM
7000
6000
5000
4000
Ground
Idle
100
Start Bleeds
Bypass Bleed
Military Speed Schedule .
..����
--
.*
/ Bypass And Start
/ Bleeds Open
Compressor inlet Temperature �C
100 200
Windmill Band
300
Start Bleeds Exhaust To Nacelle Secondary Air Flow
Compressor Bleed Air Bypass
400
,Burner Afterburner
Compressor Section Turbine Section
Section Section
Figure 1-7
F200-96
1-12
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SECTION I
A-12
Start and Bypass Bleed Valve Actuation
The bypass bleed control and actuation
system consists of four two-position ac-
tuators to move the bleed valves, and a
pilot valve to establish the pressure to the
actuators. The pilot valve controls the
bleed valve position in response to a me-
chanical signal from the main fuel control.
Bleed valve position is scheduled within the
main fuel control as a function of engine
speed and compressor inlet temperature.
The starting bleed control and actuation sys-
tem is similar to the bypass bleed system
except that three actuators are used and the
pilot valve controls starting bleed valve
position in response to the main fuel pump
boost stage pressure rise.
EXHAUST NOZZLE AND EJECTOR SYSTEM
The variable area, iris type, exhaust nozzle
is comprised of segments operated by a cam
and roller mechanism and four hydraulic
actuators. The actuators are operated by
fuel hydraulic system pressure. The ex-
haust nozzle is enclosed by a fixed contour,
convergent-divergent ejector nozzle to which
free floating trailing edge flaps are attached.
In flight, the inlet cowl bleed supplies sec-
ondary airflow between the engine and na-
celle for cooling. During ground operation,
suck in doors in the aft nacelle area provide
cooling air. Free floating doors around the
nacelle, just forward of the ejector, supply
tertiary air to the ejector nozzle at subsonic
Mach numbers. The tertiary doors and
trailing edge flaps open and close with vary-
ing internal nozzle pressure, which is a
function of Mach number and engine thrust.
Exhaust Nozzle Position Indicator
Each engine is provided with a nozzle posi-
tion indicator located on the right side of
the instrument panel. The indicators are
marked from 0 to 10 and indicate the ap-
proximate percentage of open position. Ad-
ditional dot markings above and below the
0 and 10 position marks are for calibration
purposes. The indicators are remotely op-
erated by electrical transducers located
near the exhaust nozzles. Each transducer
is cooled by fuel and is operated by the
afterburner nozzle feedback link. Power
for the indicators is supplied by the No. 1
inverter.
OIL SUPPLY SYSTEM
The engine and reduction gear box are lu-
bricated by an engine contained, "hot tank",
closed system. The oil is cooled by cir-
culation through an engine fuel-oil cooler.
The oil tank is mounted on the lower right
side of the engine compressor case and has
a usable capacity of 4.5 gals. Total tank
capacity is 6.7 gals. The oil is gravity fed
to the main oil pump which forces the oil
through a filter and the fuel-oil cooler.
The filter is equipped with a bypass in case
of clogging. From the fuel-oil cooler the
oil is distributed to the engine bearings and
gears. Oil screens are installed at the lu-
bricating jets for additional protection.
Scavenge pumps return the oil to the tank
where it is deaerated. The main oil pump
normally maintains an oil pressure of 40
to 55 psi. A pressure regulating valve keeps
flow and pressure relatively constant at all
flight conditions. Because of the high vis-
cosity of the oil, it is diluted with trichloro-
ethlene at lower temperatures and special
cold weather shut down procedures may be
required.
Main Fuel-Oil Cooler
This unit provides oil cooling by using
engine fuel to absorb the heat. The oil
temperature is controlled by fuel flow
through the cooler. A bypass valve is in-
corporated to bypass fuel around the cooler
when the fuel flow is greater than the cooler
flow capacity of approximately 12,000
pounds per hour.
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SECTION I
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A-12
CHEMICAL IGNITION SYSTEM
�;TO THROTTLE VALVE S
CIS DUMP
SOLENOID
FROM
VALVE
FUEL
HYD
PUMP
ON
IGNITOR
PURGE SWITCH OFF
1.1.1
16
CODE
MAIN BURNER IGNITION 1�11�IMMI
MAIN IGNITION SIGNAL mosormi
DUMP SIGNAL
DUMP SIGNAL DRAIN izszastazazo
FUEL COOLING IN Emma
TO A/B FUEL PUMP 0
CHEMICAL IGNITION SYSTEM
COMPRESSOR
DISCHARGE PRESSURE
FUEL COOLING OUT
A/B IGNITION LINE
TURBINE DISCH PRESS
IGNITION SIGNAL
ELECTR ICAL
Figure 1-8
=193
F200 -11(b)
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SECTION I
A-12
Oil Quantity Low Lights
An indicator light for each engines oil sys-
tem is located on the lower instrument an-
nunciator panel. The lights are labeled L
and R OIL QTY LOW and illuminate when
the respective engine oil quantity is reduced
to 2.25 gals. Power is furnished by the es-
sential dc bus.
Engine Oil Temperature Light
_
L and R OIL TEMP lights are installed on
the annunciator panel. These lights will
illuminate when respective engine oil inlet
temperature is less than +15.6 + 3 C or
greater than 282�C + 11�C.
Remote Gear Box Oil System
The remote gear box contains an indepen-
dent, wet sump lubricating system with its
own oil supply and pressure pump. There
is no scavenge pump. It is vented to the
engine breather system through the remote
gear box drive shaft. The oil is cooled by
circulation through the remote gear box
fuel-oil heat exchanger.
CHEMICAL IGNITION SYSTEM
Triethylborane (TEB) is used for ignition of
main burner and afterburner fuel. Special
handling procedures are required because
TEB above 0 F will burn spontaneously
upon exposure to air above -4 F. The TEB
is contained in a 600 cc (1-1/4 pint) storage
tank pressurized with nitrogen. The nitro-
gen provides inerting and operating pres-
sure to supply a metered quantity of TEB to
either the main burner or afterburner
section. Operation is in response to a fuel
pressure signal from the appropriate sys-
tem. Actuation is automatic with throttle
movement. A mechanical counter for each
engine, located aft of the throttles, indicate
TEB shots remaining. A minimum of 16
injections can be made with one full tank of
TEB. The TEB tank is engine mounted and
is cooled by main burner fuel to maintain
the TEB temperature within safe limits. If
the TEB vapor pressure exceeds a safe
level, a rupture disc is provided to dis-
charge the vaporized TEB and tank nitrogen
through the afterburner section. No pilot
indication of TEB tank discharge is pro-
vided. The engine is also equipped with
catalytic igniters installed on the afterburner
flameholders to provide improved after-
burner ignition system reliability and re-
light capability. Turbine exhaust temper-
ature must be above approximately 730 C
to obtain a satisfactory afterburner "light"
by the catalytic igniter s.
Igniter Purge Switch
A lift-lock toggle switch labeled IGNITER
PURGE is installed on the upper right side
of the instrument panel. When the switch
is pulled out and held in the up position a
solenoid operated valve supplies fuel hy-
draulic system pressure to the chemical
ignition system dump valve. This allows
the remaining TEB to be dumped into the
afterburner section; while the engine is
running. Approximately 40 seconds is re-
quired. Electrical power is furnished by
the essential dc bus.
NOTE
Both electrical power and engine
fuel hydraulic pressure are
necessary to purge the TEB sys-
tem. If the engine is not rotating
the system will not normally dump.
Do not actuate the Igniter Purge
switch unless the engine is ro-
tating in the 5000-6000 rpm
range to prevent damage to the
afterburner flame holders.
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SECTION I A-12
THROTTLE QUADRANT
1 THROTTLES
2 TRANSMIT BUTTON
3 MILITARY DETENT
4 THROTTLE FRICTION LEVER
5 MAX AFTERBURNER STOP
6 TEB SHOT COUNTERS
FMA12,13-(a)
Figure 1-9
1-16
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SECTION I
A-12
STARTER SYSTEM
A starter cart is provided for ground starts.
This may be either a self-contained gas
engine cart or multiple air turbine cart.
The output drive gear of either cart connects
to a starter gear on the main gear box at the
bottom of the engine. There are no aircraft
controls for this system. It is turned on and
off by the ground crew in response to signals
from the pilot. Air starts are made by
windmilling the engine.
THROTTLES
Two throttle levers, one for each engine,
are located in a quadrant on the left forward
console. The right throttle is mechanically
linked to the right engine main fuel control
and the left throttle to the left engine after-
burner fuel control with parallelogram type
linkages designed to compensate for heat
expansion. The afterburner and main fuel
controls are interconnected by a closed
loop cable. The throttle quadrant is labeled
OFF, IDLE and AFTERBURNER. When the
throttles are moved forward from OFF to
IDLE, they drop over a hidden ledge to the
IDLE position. This ledge prevents inad-
vertent engine cutoff when the throttles are
retarded to IDLE. When retarding the
throttles from IDLE to OFF they must be
lifted in order to clear the IDLE stop ledge.
Forward throttle movement from IDLE to
a MILITARY stop controls the non-after-
burning thrust range of the engine. The
throttles must be slightly raised to clear
the stop before additional forward move-
ment of the throttle can actuate the after-
burner ignition. The AFTERBURNER
range extends from the Military stop to the
quadrant forward stop. The right throttle
knob incorporates a radio transmission push-
button switch. Throttle quadrants are
marked to indicate 82o power lever angle
(PLA) for assistance in determining the
cruise power position.
Throttle Friction Lever
The throttles are prevented from creeping
by a friction lever located on the inboard
side of the throttle quadrant. When the
lever is full aft, the throttles are free to
move. Moving the lever forward as the
INCREASE FRICTION label indicates, pro-
gressively increases the amount of friction
to hold the throttles in the desired position.
TEB Remaining Counters
A mechanical TEB remaining counter for
each engine is located aft of each throttle.
The counters are spring wound and set to
12 prior to engine start. Each time a
throttle is moved forward from OFF to IDLE
or MILITARY to A/B the counter will reduce
one number.
Exhaust Gas Temperature Trim Switches
Individual exhaust gas temperature trim
switches for each engine are located on the
lower left side of the instrument panel. The
switches are spring loaded, momentary
contact, three position switches with a
center OFF position. When held in the IN-
CREASE (up) position, a remote trim elec-
tric motor on the engine fuel control is ac-
tuated to slightly increase main burner fuel
flow and turbine inlet temperature. The
trim motors have a fuel flow or EGT travel
raw of about 150�C and a rate of change
of 8 C per second. When held in the DE-
CREASE (down) position, the trim motor
reduces the fuel flow and decreases tur-
bine inlet temperature. An increase or
decrease in turbine inlet temperature will
be reflected on the respective exhaust gas
temperature gage. These switches are the
only provision for main engine control when
the throttles are in the afterburning range.
They have no effect on rpm when the nozzle
is modulating to provide the scheduled en-
gine speed. Power for the trim motors is
furnished by the respective ac generator
bus.
Changed 15 March 1968
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SECTION I
A-12
ENGINE INSTRUMENTS
Exhaust Gas Temperature Gages
Two exhaust gas temperature gages, one for
each engine, are mounted on the right side
of the instrument panel. Theci are caliborated
in degrees centigrade from 0 C to 1200 C
and indicate the temperature sensed by tur-
bine discharge thermocouples. The four
digit windows at the top of the gages indicate
the exhaust gas temperature to the nearest
degree. An OFF window at the bottom of
each dial when visible indicates instrument
power failure. A small red light on the dial
will light when EGT reaches 860�C. This
will activate the respective derichment sys-
tem if armed. The indicating system re-
ceives power from the No. 1 inverter.
Fuel Flow Indicators
A fuel flow indicator for each engine is
mounted on the instrument panel and dis-
plays total fuel flow (engine and afterburner)
in pounds per hour. The dial is calibrated
in 2000 pound per hour increments to 76,000
pph. The five digit center window indicates
the fuel flow to the nearest 100 pph. The
indicator is not compensated for return flow
and indicates total fuel flow to engine, after-
burner and heat sink system. A positive in-
dication is normal during windmill operation
and the indicator will read high when the
mixer and temperature control valve is di-
verting cooling loop fuel back to tank 4.
During descent after high speed cruise both
high and low fuel flows and flow oscillations
may be indicated. Power for the indicators
is supplied by the No. 1 inverter.
Tachometers
A tachometer for each engine is mounted on
the right side of the instrument panel. The
tachometers indicate engine speed in revolu-
tions per minute. The main pointer is cali-
brated up to 10,000 rpm and the subpointer
makes one complete revolution for each
1000 rpm. The tachometers are self-
energized and operate independently of the
aircraft electrical system.
Engine Oil Pressure Gages
An oil pressure gage is provided for each
engine on the right side of the instrument
panel. The gages indicate output pressure
of the respective engine oil pump in pounds
per square inch. The gages are calibrated
from 0 to 100 psi in increments of 5 psi.
Power for the gages is furnished by the
No. 1 inverter bus through the 2.6-volt auto-
transformer.
Compressor Inlet Temperature Gage
A dual indicating compressor inlet tem-
perature gage is mounted on the upper
right side of the instrument panel. It is
calibrated in 50 increments from 0 C to
300C and 100 increments from 300�C to
500�C. The needles indicate the air tem-
perature forward of the first compressor
stage of each nacelle. The system uses
platinum resistance sensors and power is
furnished by the No. 1 inverter.
Compressor Inlet Air Static Pressure Gage
A dual indicating compressor inlet air static
pressure gage located on the upper center
of the instrument panel, measures absolute
pressure at the engine compressor inlet.
The gage is calibrated in one psi increments
and has marked red ranges from 0 to 4 psi
and 27 to 30 psi and a green radial mark at
7 psi. A white striped third pointer on the
CIP gage indicates pressure to be expected
when the inlets are operating normally if
over Mach 1.8 and 250 KEAS. The L and R
labeled pointers indicate actual inlet static
pressures. Power is furnished from the
No. 1 inverter.
1-18
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SECTION I
A- 12
AIR INLET SYSTEM
The air inlets for each nacelle are canted
inboard and down to align with the local air-
flow pattern. The inlet system consists of
the cowl structure, a moving spike to help
provide optimum internal airflow charac-
teristics, a spike porous centerbody bleed
and an internal cowl shock trap bleed for
internal shock wave position and boundary
layer flow control, forward and aft bypass
doors for control of airflow in the inlet and
to the engine, cowl exhaust louvers, sys-
tem controls, sensors, actuators and in-
strumentation. Suck-in doors are also pro-
vided in the aft nacelle area for ground
cooling. Nacelle cooling air is provided in
flight by air from the cowl shock trap bleed
and aft bypass. Normally, the spike and
forward bypass are operated automatically
by the air inlet control system. Inlet air-
flow is controlled so that the proper amount
of air is supplied to the engine and, at super-
sonic airspeeds, the positions of shock waves
ahead of the inlet and in the inlet throat are
controlled so as to provide maximum prac-
tical ram pressure recovery at the engine
face. Controls are provided in the cockpit
for incremental control of the aft bypass for
those conditions where additional bypass
area is required or where a reduction in
forward bypass flow is desired. Manual
controls are provided to override the auto-
matic spike and forward bypass control sys-
tems.
INLET SPIKE
The spike is locked in the forward position
for ground operation and flight below 30,000
feet. It is unlocked above this altitude and
is programmed during automatic operation
to move 1/4 inch off the forward stop at
Mach 1.4. Above Mach 1.6, the spike re-
tracts so as to increase the nacelle inlet
area and decrease the area of the throat or
narrowest portion of the duct. Spike posi-
tion is scheduled primarily as a function of
Mach number as sensed by the Rosemount
boom pitot static ports with biasing for
angle of attack and yaw angle. The spike
moves aft approximately 26 inches during
transition between Mach 1.6 and 3.2. The
inlet control also includes a shock expulsion
sensor (SES) and restart feature which can
operate automatically when speeds for inlet
scheduling are reached. It is effective
above approximately Mach 2.0. If an inlet
becomes unstable and expels the internal
shock, the shock expulsion sensor for that
inlet overrides the automatic spike and for-
ward bypass schedule. It causes the for-
ward bypass to open fully and the spike to
move forward as much as 15 inches. Spike
retraction is started automatically 3.75
seconds after the expulsion is sensed and,
when schedule position is reached, the for-
ward bypass is returned to automatic op-
eration. The SES reference pressure is
CIP, and the system is triggered when a
momentary decrease of CIP is 23% or more.
This is a characteristic CIP indication of
inlet unstart occurrence. However, it may
also operate as a result of pressure fluc-
tuations if CIP decreases rapidly below the
previous normal condition during compres-
sor stalls. The SES feature does not over-
ride a manually operated spike or forward
bypass control. Manual operation of a re-
start switch overrides the SES operation
for that inlet. Spike and forward bypass
door position changes may be observed
during SES operation on the spike and for-
ward bypass position indicators. Local
pitch attitude and yaw angle are sensed by
a pressure probe mounted on the Rosemount
pitot boom. The spike porous centerbody
bleeds boundary layer air from the inlet
throat to prevent flow separation. This air
Is ducted overboard through the supporting
struts and nacelle louvers. The spikes can
be fully controlled by use of cockpit controls
when hydraulic pressure is available.
Emergency spike forward
1-19
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SECTION I
A-12
switches provide pneumatic pressure to move
and lock the spikes forward in the event of
hydraulic system failure.
INLET FORWARD BYPASS
The forward bypass provides an exhaust for
inlet air which is not required by the engine,
and controls the inlet diffuser pressure so
as to properly position the inlet shock. It
consists of a rotating basket which opens
duct exhaust ports located a short distance
aft of the inlet throat. When the landing
gear is down, the forward bypass doors are
held open by an electrical override signal
from a landing gear door switch. The switch
is positioned to allow manual or automatic
control of the bypass when the landing gear
retracts. In automatic operation, the for-
ward bypass remains closed until a low,
supersonic speed is reached, then it mod-
ulates in accordance with air inlet control
system Mach and pressure schedules. The
inlet usually "starts" at Mach 1.4, that is,
the inlet shock is positioned near the cowl
shock trap bleed in the inlet throat area. As
speed is increased, the forward bypass
schedules as required to maintain the inlet
shock at the throat position.
The forward bypass position is controlled
by the ratio of inlet duct static pressure to
a reference total pressure. The inlet duct
static pressure is sensed by taps located
aft of the shock trap bleed.
The reference total pressure is sensed by
two external probes one located on the
lower inboard side of the nacelle and the
other at the top of the nacelle. The forward
bypass control also senses an unstart as a
result of the sudden decrease in pressure
at the engine face and controls the inlet
through a timed sequence. The minimum
Mach number at which the automatic re-
start actuates varies with the intensity of
the unstart but is generally in the vicinity
of Mach 2.0. An overriding switch holds
the forward bypass closed at speeds lower
than Mach 1.4.
I NLET AFT BYPASS
The aft bypass consists of a ring of adjust-
able peripheral openings allowing a maxi-
mum mass flow of approximately 3/4 of
that available from the forward bypass. The
ring is located just forward of the engine
face. These openings allow excess inlet
air to be bypassed around the engine. The
bypassed air joins cowl shock trap bleed
air and passes between the outside of the
engine and afterburner and the inside of the
nacelle. This flow augments the exhaust
gas in the ejector area. Each aft bypass
ring is positioned by a hydraulic actuator
which is powered by the respective L or R
hydraulic system and is controlled by the
cockpit switch. The bypass is held closed
during takeoff and landing by an electrical
signal from the nose gear downlock. It is
also closed during subsonic operation.
Position in flight is set manually in accor
dance with determined Mach number and
engine operating requirements.
1-20
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A-12
SECTION I
INLET AIRFLOWS
FORWARD BYPASS
AFT BYPASS
13001,
3,I'
, \
\I
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,
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1 ...,
,03,-,1
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I 011 i
IllIlDnn
1] II 11
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Figure 1-10 (Sheet 1 of 2)
8-31-65
F200-71(1)
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1-21
SECTION I
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A-12
INLET AIRFLOWS
POROUS BLEED
DUCT SHOCK TRAP BLEED
0011�0�
���
.cy---Iseleassa
'NOTE
DUCT SHOCK TRAP BLEED AIR FLOWING THROUGH THESE TUBES
REACHES NACELLE SECONDARY AREA AND EXHAUSTS THROUGH
EJECTOR.
Figure 1-10 (Sheet 2 of 2)
10-4-65
F200-71(2)(a)
1-22
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SECTION I
A-12
AIR INLET CONTROL SYSTEM
The air inlet control system incorporates a
computer which utilizes electrically trans-
mitted pneumatic pressure signals to auto-
matically schedule and reposition the spikes
and forward bypass. The computer also
serves as a calibrated path for the manual
spike and manual forward bypass control.
Major components for each inlet control are
the computer, pressure transducer, angle
transducer and two pressure ratio trans-
ducers. The spike and forward bypass con-
trols consist of four rheostat type knobs
and two inlet restart switches and an emer-
gency spike switch. Aft bypass control is
by means of two rotary type switches lo-
cated above the throttle. Three annunciator
panel lights are pertinent to the inlet control
system.
Nine different pressures are sensed for in-
let control. The Rosemount airspeed boom
provides pitot total and static pressures to
the pitot pressure transducer. The pitch and
yaw attitude probe on the left side of the
boom provides angle of attack and yaw angle
pressures for conversion to electrical sig-
nals by the attitude transducer. At each
nacelle local pitot pressure and two inlet
duct static pressures are sensed to enable
two sensors within the pressure ratio trans-
ducer to convert pressure ratios to elec-
trical signals which (1) direct forward by-
pass control, and (2) cause an automatic re-
start following shock expulsion. Some con-
trol functions are also accomplished within
the pressure transducer. Most of the elec-
trical outputs of the pitot pressure trans-
ducer, attitude transducer, and both pres-
sure ratio transducers are transmitted to
the computer. The computer also receives
a signal from the main landing gear doors
to assure that the forward bypass will be
open whenever the main gear is down.
Spike Controls
The L and R spike controls are located on
the lower left side of the instrument panel.
The controls are labeled AUTO, FWD, and
have labeled marks for 1.4, 1.8, 2.2, 2.6,
3.0 and 3.2 Mach numbers. Intermediate
marks for 0.1 Mach increments allow the
knobs to be positioned manually at any set-
ting from 1.4 to 3.2 Mach number. In the
detented AUTO position, spike position is
scheduled automatically by the inlet control
system. In the detented FWD position, the
spike will move to the full forward position.
The Mach numbered positions are used in
manual operation. Use of settings corre-
sponding to aircraft flight Mach number
moves the spike aft to the correct position
for proper inlet performance. The spike
control also biases the forward bypass as a
function of control knob position when the
bypass is being manually controlled. The
forward bypass position indicator and by-
pass control knob will not be in agreement
by the amount of bias. Control power for
the left spike is from the No. 2 inverter
and for the right spike the No. 3 inverter.
Forward Bypass Controls
The L & R BYPASS controls are located
just inboard of the spike controls. When a
control is turned full counterclockwise to
the detented AUTO position, operation of
the respective forward bypass is automat-
ically controlled by the inlet computer. As
the control is turned clockwise the first de-
tented position will position the forward by-
pass to the full open. As the control is
turned further clockwise the forward bypass
will incrementally move towards the closed
position and will be fully closed in the full
clockwise position. Markings from 0 to 100
in increments of 10 percent allow the con-
trol to be, positioned at any percentage of
full open. - Power for the circuits is from
the essential dc bus and No. 2 and No. 3
inverters.
1-23
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SECTION I
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A - 1 2
AIR INLET CONTROLS AND INDICATORS
BELOW
30,000 FT
MANUAL INLET
Figure 1-11
F no-79(0
1 -24
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SECTION I
A-12
Manual operation of the forward
bypass is permissible with the
spike operating on its automatic
schedule; however, when the
spike is operated manually, the
forward bypass must be operated
manually or the bypass will open
fully and will not schedule.
Inlet Restart Switches
Two 3-position toggle switches are located
on the left side of the instrument panel. The
L & R switches are labeled RESTART (up),
FWD DOOR OPEN (center) and OFF (down).
In the RESTART position the spike and by-
pass control settings are overridden, the
forward bypass is opened and the spike is
moved forward. In the center FWD DOOR
OPEN position the forward door is moved
to/or held open but the spike position re-
sponds to its control knob. In the OFF posi-
tion both the spike and forward bypass are
controlled by their respective controls.
Power for the restart circuit is supplied by
the essential dc bus.
Emergency Spike Switch
A single 3-position guarded switch, labeled
EMER SPIKE, is provided below the instru-
ment panel. The switch is guarded in the
center OFF position. After the guard is
opened the switch may be positioned in
either L or R positions as necessary. In the
event of L or R hydraulic failure, the one
shot emergency pneumatic bottle in the re-
spective nacelle is activated to drive and
lock the spike in the full forward position.
Power for the emergency spike circuit is
from the essential dc bus.
Inlet Aft Bypass Switches and Indicator Lights
The inlet aft bypass switches and indicator
lights are located above the throttle quad-
rant. They are four-position rotary type
switches equipped with concentric lever
handles. The switch positions from top to
bottom are labeled CLOSED, A (15% open),
B (50% open), OPEN (100%). Left and right
amber lights, located near the switch levers
Illuminate to indicate when an aft bypass
position and the switch setting do not cor-
respond. A light should illuminate each
time its switch is moved, then extinguish
as the bypass reaches the required position.
Approximately 5 seconds is required for the
aft bypass to move from full closed to full
open. The aft bypass actuator control cir-
cuits are powered by the essential dc bus.
Spike Position Indicator
A dual spike position indicator is located
on the lower right side of the instrument
panel. The L & R labeled pointers indicate
the position of the respective spike in inches
aft of the forward position. It is calibrated
In inches from 0 to 26 with 5, 10, 15, 20,
and 25 inch labeling. Power is furnished
from the No. 2 inverter for the left spike
and the No. 3 inverter for the right spike.
Forward Bypass Position Indicator
A dual forward bypass position indicator is
located on the lower right side of the instru-
ment panel. The L & R labeled pointers
indicate the opening of the respective for-
ward bypass in 10% increments. Labeled
positions are 20, 40, 60, 80 and 100 OPEN.
Power is furnished from the No. 2 inverter
for the left bypass and the No. 3 inverter
for the right bypass.
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SECTION I
A-12
FUEL QUANTITY DATA
TANK 1
Manual Inlet Indicator Light
TANK 2
TANK 3
11,
I iigafr
TANK 4
TANK 5
TANK 6
FUEL TANK CAPACITIES
Tank
Fuel
1 1.146 gal.
7, 390 lb.
2 1, 610 gal.
10,380 lb.
3 1,585 gal.
10,220 lb.
4 2, 135 gal.
13,770 lb.
5 2, 136 gal.
13, 780 lb.
6 1, 978 gal.
12, 760 lb.
TOTAL 10, 590 gal.
68, 300 lb.
'At average fuel density of 6.45 lb. /gal.
F200 -61(c)
Figure 1-12
The annunciator panel MANUAL INLET light,
when illuminated, indicates that one or more
of the four rotary spike and forward bypass
controls is not in the AUTO position or that
an inlet restart switch is not in the OFF
position. Power for the light is furnished
by the essential dc bus.
FUEL SUPPLY SYSTEM
There are six individual fuel tanks, iden-
tified from forward to aft as tanks 1, 2, 3,
4, 5, and 6. Interconnecting plumbing and
electrically driven boost pumps are utilized
for fuel feed, transfer, and dumping. Other
components of the system include pump con-
trols, nitrogen inerting, scavenging, pres-
surization and venting, a single-point re-
fueling receptacle, and a fuel quantity indi-
cating system. In addition to furnishing
fuel to the engines, automatic fuel manage-
1-26
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SECTION I
A-12
ment provides center of gravity and ,trim
drag control. The fuel is also used to cool
cockpit air, engine oil, accessory drive sys-
tem oil, and hydraulic fluid by means of the
fuel heat sink system.
FUEL TANKS
The integral, internally sealed, fuel tanks
are contained in the fuselage and wing root.
The tanks are interconnected by right and
left fuel manifolds and a single vent line.
Submerged boost pumps supply fuel through
the manifolds and transfer fuel for c. g. con-
trol. Forward transfer is accomplished by
manual control of the right manifold. Aft
transfer is accomplished automatically
through the left manifold. A fuel dump valve
is installed in each fuel manifold. Normal
sequence of tank usage is controlled by float
switches to automatically maintain an op-
timum c. g. for cruise. The left engine is
normally sequenced from tanks 1, 2, 3, and
4, the right engine is sequenced from tanks
1, 6, 5, and 4. Normal automatic tank se-
quencing is as follows:
L Engine
Tank 1 and 2
Tank 2
Tank 3
Tank 3
Tank 4
Tank 4
R Engine
Tanks 1-and 6
Tank 6
Tank 6
Tank 5
Tank 5
Tank 4
The fuel manifolds can be connected by de-
pressing the crossfeed switch. This operates
a motor operated valve between the fuel
manifolds and is mainly used during single
engine operation.
REFUELING AND DEFUELING
A single point refueling receptacle installed
on top of the fuselage aft of the air condi-
tioning bay is used for both ground and in-
flight refueling. Ground refueling is ac-
complished by use of an in-flight refueling
probe specially modified to utilize a hand
operated locking device so that refueling
may be done without hydraulic power. Fuel
from the receptacle flows through the fuel-
ing manifold to each tank. The use of a
different size orifice for each tank allows
all tanks to be filled simultaneously in ap-
proximately 15 minutes with a nozzle pres-
sure of 50 psi. Dual shutoff valves in each
tank terminate refueling flow when the tank
is full. A defueling fitting is installed on
the right fuel feed manifold in the lower
right side of tank 3. Tanks 2 and 3, which
feed the left manifold, are defueled by open-
ing the crossfeed valve.
Any fuel in tanks 5 and 6 must
be balanced with a like amount
of fuel in the other tanks during
ground fueling or defueling to
prevent the aircraft from rock-
ing down on the tail.
FUEL TANK CAPACITIES
See figure 1-12.
FUEL BOOST PUMPS
Sixteen single stage centrifugal ac powered
boost pumps are used to supply the fuel
manifolds. Tanks 1 and 4, which normally
feed both engines, are equipped with four
pumps and tanks 2, 3, 5 and 6 have two
pumps each. Either pump of a pair is cap-
able of supplying fuel to its manifold at a
rate sufficient for normal engine operation
in the event of a failure of the other pump.
The pumps in each tank may be operated
out of the normal sequence by use of the in-
dividual tank boost pump control switches
located on the right side of the instrument
panel. These switches supplement auto-
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E - an2TJ
�
�
uo
�
FUEL
TANK NO. 1
FUEL
TANK NO. 2
IA
�
FUEL
TANK NO. 3
FUEL
TANK NO. 4
�
10
FUEL
TANK NO. 6
FUEL
TANK NO. 5
1 FORWARD TRANSFER VALVE
2 RIGHT FUEL MANIFOLD
3 FUEL BOOST PUMP (16 TOTAL)
4 GROUND DEFUELING
5 GYRO CANS
6 TO MAIN AND A/B FUEL PUMPS
7 FLOW METER
8 FUEL FILTER
9 CHECK AND RELIEF VALVE
10 FUEL SHUTOFF VALVE
11 CROSSFEED VALVE
12 FUEL DUMP
13 JET PUMP (4 TOTAL)
14 LEFT FUEL MANIFOLD
15 AFT TRANSFER VALVE
�
�
�
I mou,Das
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SECTION I
A-12
matic tank sequencing if a tank fails to feed
in the proper sequence. It is necessary to
actuate the pump release switch to termi-
nate the manually actuated pumps when the
tank is empty. Normally, each pump (ex-
cept pumps 1-1 and 1-2 which are protected
by a common float switch) is protected by a
float switch that deactivates the pump when
the tank is empty. One of the float switches
in each tank illuminates the yellow tank
empty light contained in the respective
boost pump tank switch. For example, the
float switch for the number 4 pump in tank
4 is used to indicate that tank 4 is empty
and pump 4-4 is off. (The tank 4 light in-
dicates green when pumps 4-3 and/or 4-4
are on. When pump 4-4 is on and in auto-
matic sequencing, the green light may not
indicate the status of other tank 4 pumps
whose operation is affected by automatic
features of the ullage and refueling systems.)
The boost pumps that feed the left hand
manifold are normally powered from the
left generator bus and the pumps that feed
the right hand manifold are normally
powered from the right generator bus. In-
dividual circuit breakers for each pump are
located in the compartment behind the cock-
pit and are not accessible in flight.
Emergency Fuel Shutoff Switches
A guarded fuel shutoff switch for each
engine is installed on the lower right side
of the instrument panel. Each switch is
guarded in the down (fuel on) position. Fuel
is shut off in the OFF (up) position. Move-
ment of a switch causes a motor operated
valve in the respective engine feed line to
operate. Motor power is supplied from the
corresponding ac generator bus.
Fuel Boost Pump Switches and Indicator Lights
Six pushbutton type fuel boost pump switches
with green and yellow indicator lights are
installed in a vertical row on the right
side of the instrument panel. These switches
are provided for manual control of the fuel
boost pumps.
NOTE
Manual operation supplements,
but does not terminate the normal
automatic fuel tank sequencing.
The switches have an electrical hold and
bail arrangement that allows manual se-
lection of only one tank of tank group 1, 2,
3 and one tank of tank group 4, 5, 6 at the
same time. This feature is intended to
prevent more than eight boost pumps from
operating simultaneously.
NOTE
It is possible to operate more
than eight boost pumps at once by
a combination of automatic se-
quencing and manual actuation;
however, this condition will not
overload the electrical system
except when operating on a single
generator.
When a set of boost pumps is actuated,
either automatically or manually, a green
light will illuminate in the pushbutton. When
a tank is empty, a yellow EMPTY light in
the pushbutton illuminates. When depressed,
the boost pump switch will hold down elec-
trically until released by the pump release
switch. Power for the boost pump switch
circuit and lights is furnished by the es-
sential dc bus.
Pump Release Switch
A momentary pump release switch is in-
stalled on the instrument panel below the
fuel boost pump switches. The switch has
two positions; PUMP REL (up) and NORM
(down). When placed in the momentary
PUMP REL position, any boost pump
switch that has been depressed during
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SECTION I
A- 12
manual boost pump selection will be released
and automatic sequencing of the fuel tanks is
continued. Power for the circuit is furnished
by the essential dc bus.
A manually selected boost pump
should be released when a tank
indicates empty so that the pumps
in that tank will be shutoff; other-
wise, damage to the pump may
occur.
Crossfeed Switch
A pushbutton type crossfeed switch is lo-
cated above the boost pump switches on the
instrument panel. When depressed, it il-
luminates a green light in the switch, opens
a motor operated valve between the left and
right fuel manifolds, allowing operating
boost pumps to pressurize both fuel mani-
folds. The switch must be depressed a sec-
ond time to terminate crossfeeding. Power
for the circuit is furnished by the essential
dc bus.
Fuel Transfer Switch
A guarded three-position fuel transfer
switch is located on the right side of the in-
strument panel. The switch is guarded in
the OFF position. When the guard is raised
and the switch is moved to the FWD TRANS
position, the pumps in tank 1 are inactivated,
a valve at the forward end of the right fuel
manifold opens into tank 1 if fuel manifold
pressure is above approximately 8 psi and
fuel will transfer forward through the right
side fuel manifold as long as automatic or
manual pump sequencing continues. Trans-
fer will be automatically terminated by
closing of the forward transfer valve when
the tank 1 fuel level reaches 4000 pounds.
Tank 1 boost pumps will remain inactivated
until either tank 4 has approximately
800 lbs remaining or the transfer switch is
moved to the OFF (down) position. Tank 1
pumps will also start when the tank 1 pump
switch is pressed. The forward transfer
valve is not closed by manual selection of
tank 1 but right side boost pump pressure
makes forward transfer ineffective. The
lift-lock forward transfer switch may also
be pulled out and placed in the NO. 4
TRANS position. In this position, tank 1
pumps are inactivated, the right side pumps
in tank 4 are turned on, and tank 5 is turned
off if operative. The transfer is only from
tank 4, which prevents the accumulation of
hot fuel in tank 4 and puts the warmer fuel
into tank 1 where it will be used immediately
after an air refueling.
NOTE
Forward transfer should be dis-
continued before refueling is
started to restore normal tank
sequencing.
Transfer is automatically terminated when
the tank 1 4000 pound float switch operates,
and the tank 1 pumps remain off until either
tank 4 has 800 pounds remaining or the
transfer switch is moved to the OFF posi-
tion. Power for the transfer control cir-
cuits is furnished by the essential dc bus.
Those aircraft incorporating SIB 1141 are
modified to replace the Tank 4 Forward
Transfer position with an EMER forward
transfer position on these aircraft. When
the lift-lc switch is pulled out and replaced
in the EMER position, tank 1 pumps are in-
activated and the dual 4000 lb stop transfer
float switches in tank I are replaced by dual
7400# float switches. This allows forward
transfer to continue until tank 1 is full.
WARNING
The EMER posi ion is to be used
only in case of an aft c. g. emergency.
Fuel Dump Switch
A guarded 3-position lift-lock fuel dump
switch is located on the right side of the in-
strument panel. The switch is guarded in
the OFF (down) position. In the DUMP
(center) position dual type solenoid dump
valves in each manifold are opened and the
pumps in tank 1 are inactivated unless se-
lected manually. If fuel pressure is above
10 psi, all other tanks dump in normal
usage sequence until tank 4 is down to a
8000 pound remaining level. Dumping nor-
1-30
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A -12
SECTION I
.Add the following text to Fuel Transfer Switch
Those aircraft incorporating S/B 1141 are modified to replace the Tank 4
Forward Transfer position with an ,EMER forward transfer position. When the
lift-lc .switch is pulled out and placed in the EMER Position, tank 1 pumps
are inactivated and the dual 4000 lb stop transfer float switches in tank I are
replaced by dual 7400 # float switches. This allows forward transfer to
continue until tank 1 is full.
WARNING
The EMER position is to be used only in
case of an aft c, g. emergency.
1-30A
TDC 2
Page 2
26 Jan. 1968
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SECTION I
A-12
Fuel Quantity Low Light
mally stops at this point and, if fuel is in
tank 1, the tank 1 pumps will start unless
the forward transfer switch is in either the
FWD TRANS or NO. 4 TRANS position. The
switch knob must be pulled out to put the
switch through the DUMP position either to
the EMER or OFF position. In the EMER
position, the 8000 pound stop dump switch in
tank 4 is bypassed and fuel dumping will
continue from all tanks except tank 1. If
tank 4 is to be completely dumped, tank 1
should be pressed on before tank 4 empties
in order to avoid fuel pressure fluctuation
as tank 4 empties. Power for the circuit is
furnished by the essential dc bus.
WARNING
Emergency fuel dumping must be
terminated by placing the dump
switch to DUMP or OFF. All fuel
can be dumped with EMER dump
on and tank 1 selected manually.
Fuel Quantity Selector Switch and Quantity
Indicator
A fuel quantity indicator and a rotary seven-
position fuel quantity selector switch is in-
stalled on the lower right side of the instru-
ment panel. Positions on the selector
switch are marked for TOTAL and each of
the six tanks positions. The switch is ro-
tated to the individual tank or TOTAL posi-
tion to select the desired reading on the fuel
quantity indicator. The dial is calibrated in
5000 pound increments from zero to 70,000
pounds. The indicator has a digital read-
out window indicating to the nearest 100
pounds. Power for the circuit is furnished
by the No. 1 inverter.
A FUEL QTY LOW light on the annunciator
panel will illuminate when total fuel re-
maining in tank 4 is 5000 pounds or less.
Power for the light is furnished by the es-
sential dc bus.
Fuel Pressure Low Warning Lights
Fuel pressure warning lights, labeled L
and R FUEL PRESS LOW are located on
the annunciator panel. Illumination indi-
cates that engine fuel inlet pressure has
fallen below approximately 7 + 0.5 psi. The
light is extinguished by restoring fuel pres-
sure above approximately 10 psi. Power is
furnished by the essential dc bus.
NOTE
It is possible for a fuel pressure
low warning light to illuminate
when only two fuel pumps are
feeding an engine during high fuel
flows, especially with forward
transfer and/or fuel dump selected.
Test N and Tank Lights Switch
A test N and tank lights switch is installed
below the boost pump switches on the in-
strument panel. The switch has two posi-
tions, up and down (spring loaded down) and
is used to test the operation of the liquid
nitrogen indicators, nitrogen system an-
nunciator light, derichment light and fuel
boost pump lights. When the switch is
moved to the up position, the liquid nitrogen
indications will move down-scale toward
zero and the N QTY LOW annunciator light,
fuel boost pump lights and derichment light
will illuminate. Power for the circuit is
furnished by the essential dc bus.
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f7 I - I a an2T3
FUEL
TANK NO. 5
14
cr-
11 FUEL
13 TANK NO. 1
12
FUEL
TANK NO. 2
FUEL.
TANK NO.
FUEL
TANK NO. 4
� � � �
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FUEL
TANK NO. 6
1 OPEN VENT LINE (TANK 1)
2 SUCTION RELIEF VALVE
3 VENT LINE
4 FLOAT CHECK VALVES (6 TOTAL)
5 FLOAT CHECK AND RELIEF VALVE (5 TOTAL)
6 LIQUID CHECK VALVE
7 CHECK VALVE
8 VENT DRAIN VALVE
9 SECONDARY VENT PRESSURE RELIEF VALVE
10 PRIMARY VENT PRESSURE RELIEF VALVE
11 FUEL LINE TO SPRAY BARS ON LN2 SYSTEM
12 SUCTION RELIEF LINE (NOSE WHEEL WELL)
13 1112 FLOW FROM DEWAR TANKS
14 TO NITROGEN TANK PRESSURE SENSORS
�
�
NOII OHS
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SECTION I
A-12
FUEL PRESSURIZATION AND VENT SYSTEM
The fuel pressurization system consists of
two Dewar flasks, located in the nosewheel
well, and associated valves and plumbing to
the fuel tanks and indicators. These flasks
are equipped with automatic ac powered
heaters and contain liquid nitrogen. The
forward flask contains 75 liters and the aft
flask contains 106 liters. They supply ni-
trogen gas to the fuel tanks at 1.5 + .25 psi
above ambient pressure. This inerts the
ullage space above the fuel and will produce
some fuel flow to the engine-driven pump in
case of boost pump failure. The liquid ni-
trogen from the bottom of the flasks is
routed through submerged heat exchangers
in tanks 1 and 4 to ensure that the nitrogen
has become gaseous. The nitrogen gas is
then ported to the common vent line and to
the top of all tanks.
The venting system consists of a common
vent line through all tanks with two vent
valves in each tank except the No. 1 tank
which has only one vent valve and the open
forward end of the vent line. The forward
vent valve in tanks 2 through 6 is equipped
with a relief valve to relieve tank pressure
at 1.5 psi, and a float valve that closes the
vent valve when the tank is full. The float
shutoff is provided to keep fuel from enter-
ing the vent line. The aft vent valve is
similar to the forward except it has no re-
lief valve. The common vent line tees into
two lines in tank 6 and both go through the
rear bulkhead. In the tail cone area there
is a relief valve in each line with the left
valve set to relieve pressure at 3.25 +.25
psi above ambient pressure. In the event
of failure of this valve, the right valve will
relieve pressure at 3.85 to 4.15 psi. A
suction relief line and valve connects to the
common vent line in tank 1 and terminates
in a bell mouth fitting in the aft end of the
nosewheel well.
Two valves are provided in the vent system
to prevent fuel from surging forward in the
vent line during aircraft deceleration. A
check valve prevents fuel that is coming
forward from tank 6 from going farther
than tank 5. A python valve located in tank
3 prevents fuel coming from tank 4 from
going any farther than tank 3. This float
actuated valve closes the vent when fuel is
moving forward in the vent line and diverts
it into tank 3. Acceleration presents no
problem of fuel shift between tanks.
Liquid Nitrogen Quantity indicator
A dual liquid nitrogen quantity indicator is
installed on the right side of the' instrument
panel. The indicator displays the quantity
of liquid nitrogen remaining in each of the
two dewar flasks. The indicator is marked
in 5 liter increments from 0 to 110 liters.
Power for the indicator is furnished by the
No. 1 inverter bus.
N2 Quantify Low Light
An indicator light labeled N QTY LOW is
provided on the annunciator panel. The
light will illuminate when either hand on the
liquid nitrogen quantity gage reaches 1 liter
remaining. Power for the light is fur-
nished by the essential dc bus.
Fuel Tank Pressure Indicator
A fuel tank pressure indicator is installed
on the right side of the instrument panel.
The gage indicates the pressure existing in
the No. 1 fuel tank, and is marked from -2
to +8 in increments of 1/2 pound per square
inch. Power for the indicator is furnished
by the 26-volt instrument transformer.
1-33
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SECTION I A-12
FUEL HEAT SINK SYSTEM
TO EXHAUST
NOZZLE
ACTUATORS
PRIMARY
AIR COND
HEAT
EXCHANGER
SECONDARY
AIR COND
HEAT
EXCHANGER
HYDRAULIC
HEAT
EXCHANGER --e.
TO MAIN
BURNER
I s ao�oHEFFNi)
OPEN W
TO A/B
411F�
LJ
A/B
CONTROL
WINDMILL
AND
BYPASS
VALVE
ENGINE
VARIABLE MC) HYD
ORIFICE PUMP
MAIN
FUEL
CONTROL
ENGINE OIL
HEAT
EXCHANGER
so--
MAIN
FUEL
PUMP
2ND
STAGE
A/B
PUMP
MAIN
FUEL
PUMP
1ST
STAGE
FUEL TO
LH ENGINE
REMOTE
GEARBOX
HEAT
EXCHANGER
SPIKE
HEAT
EXCHANGER
Foin
THIS VALVE ALWAYS
PERMITS FLOW INBD,
BUT WILL PERMIT
FLOW IN BOTH
DIRECTIONS WHEN
CROSSFEED VALVE IS
OPEN �\
MIXING
VALVE
TEMP, CONTROL
(SMART) VALVE
IFLOWMETER
CIRCULATING
FUEL PUMP
Figure 1-15
FILTER
PRESSURE
OPERATED
TANK NO.4
RETURN
VALVE
RETURN TO
s TANK NO.4
I TO RH
MIXER
SWITCH
�
kss
LH ENGINE
FEED LINE
FROM BOOST
PUMPS
EMERGENCY
SHUTOFF VALVE
NORMALLY OPEN
SENSE
LINES
RH SYSTEM
IDENTICAL
CROSSFEED VALVE
(OPEN FOR SINGLE
ENGINE OPERATION)
AIRPLANE
FUEL TO
RH ENGINE
RH ENGINE
FEED LINE
FROM BOOST
PUMPS
Fno-mo
1-34
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SECTION I
A-12
Tank Pressure Low Light
A TANK PRESSURE LOW warning light is
located on the annunciator panel and will il-
luminate when the tank pressure reduces to
+.25 to +.10 psi. Power for the light is fur-
nished by the essential dc bus.
FUEL HEAT SINK SYSTEM
Fuel from the fuel manifolds is used as a
cooling medium for the air conditioning sys-
tems, the aircraft hydraulic fluid, and the
engine and remote gear box oil. Circulated
fuel from the engine fuel hydraulic system
is also used to cool the TEB tank and the
control lines which actuate the afterburner
nozzle. Engine oil is cooled by main engine
fuel flow through an oil cooler, located be-
tween the main fuel control and the windmill
bypass valve. This fuel is then directed to
the main burner section. The other cooling
is accomplished by fuel circulation through
several cooling loops. Hot fuel returning
from the remote gear box heat exchanger,
the primary and secondary air conditioning
heat exchangers, the hydraulic fluid heat ex-
changer, the spike heat exchanger and the
exhaust nozzle actuators is circulated
through a mixing valve and temperature
limiting valve (smart valve) and returned to
the main engine and afterburner fuel mani-
fold. lithe mixed fuel temperature is be-
low 265�F, all of the hot fuel will be burned
by the operating engine and afterburner. If
the temperature of the mixed cooling pop
and incoming engine fuel exceeds 265 1,
the smart valve starts to close and a por-
tion of the cooling loop fuel is prevented
from mixing with the incoming engine fuel.
A pressure operated valve routes the hot
fuel to tank 4. The smart valve is com-
pletely closed at 295�F and all cooling loop
fuel is returned to tank 4. If tank 4 is full,
the hot fuel will be diverted to the next tank
that has space for it. During single engine
operation with the inoperative engine
throttle in OFF, actuation of the fuel cross-
feed valve also allows the hot recirculated
fuel from the windmiLling engine to cross-
over and mix with the cooling loop and in-
coming fuel for the operating engine.
AIR REFUELING SYSTEM
The aircraft is equipped with an air refuel-
ing system capable of receiving fuel at a
flow rate of approximately 5000 pounds per
minute from a KC-135 boom type tanker
aircraft. The system consists of a boom
receptacle, doors,hydraulic valves, hy-
draulic actuators, a signal amplifier and
control switches and indicator light. Hy-
draulic power for the system is normally.
supplied from the L hydraulic system. If
the L hydraulic system is inoperative the
refuel system can operate from R hydraulic
pressure by selecting alternate steering and
brakes. Electrical power is supplied by the
essential dc bus.
Air Refuel Switch
An air refuel switch is installed on the
right side of the instrument panel. The
switch has three positions; READY, OFF
and MANUAL. When the switch is placed
in the READY (up) position hydraulic ac-
tuators open the refueling doors, the boom
latches are armed, the receptacle lights il-
luminate and the green READY light illum-
inates. The receptacle doors are opened
by spring action if hydraulic pressure is not
available. In the MANUAL (down) position
the latching dogs in the receptacle are
closed. They may be opened by holding the
disconnect (trigger) switch on the control
stick until the boom is seated. When the
disconnect switch is released the latches
1-35
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91-1 a an2T3
Cr,
�
�
�
FUEL
TANK NO. 1
FUEL
TANK NO. 2
FUEL
TANK NO. 3
�
FUEL
TANK NO. 4
FUEL
TANK NO. 6
FUEL
TANK NO. 5
1 AIR REFUELING RECEPTACLE
2 REFUELING MANIFOLD
3 PILOT VALVE (6 TOTAL)
4 FLOAT VALVE SHUTOFF (6 TOTAL)
I\1 01J, Das
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SECTION I
A-12
will close and hold the boom. The latches
will open to release the boom when the dis-
connect switch is depressed. This position
is used in the event of a malfunctioning am-
plifier. A3 secondtime delay is incorporated
to prevent nozzle damage if the manual posi-
tion is selected during refueling contact.
Air Refuel Reset Switch and Indicator Lights
A square dual indicator light and reset but-
ton, labeled 1FR PUSH TO RESET, is lo-
cated at the top left side of the instrument
panel. The top hall is labeled READY and
will illuminate green when the air refuel
switch is in the READY or MANUAL posi-
tion, and the refueling receptacle is open
and ready to accept the refueling boom. The
light will extinguish after the boom is en-
gaged. If the boom disconnects from the
fueling receptacle the lower half of the
switch will illuminate amber and show DISC.
If the air refuel switch is in the READY
position the light button is then pressed to
reset the system amplifier for another en-
gagement. If the air refuel switch is in the
MANUAL position the READY light will be
illuminated and manual engagement and dis-
connect are controlled by the disconnect
switch on the control stick. Power for the
switch and light is supplied by the essential
dc bus.
Disconnect Switch
A momentary contact trigger type switch is
installed on the forward side of the control
stick. Depressing the trigger switch will
normally initiate a disconnect. The dis-
connect switch is also depressed to open the
receptacle latches when the air refuel switch
is in the MANUAL position. Releasing the
disconnect switch will close the latches.
Disconnect
A disconnect may be accomplished in four
ways:
1. Automatically, if boom envelope limits
are exceeded ,(except when using man-
ual boom latching).
2. Automatically, when manifold pressures
reach 85-90 psi.
3. Manually, by the boom operator.
4. Manually, by actuating the disconnect
switch on the control stick.
Pilot Director Lights (Tanker)
Pilot director lights are located on the bot-
tom of the tanker fuselage between the nose
gear and the main gear. They consist of
two rows of lights; the left row for elevation
and the right row for boom telescoping. The
elevation lights consist of five colored
panels with strip green, triangular green
and triangular red colors and two illumi-
nated letters, D and U, for down and up
respectively. Background lights are lo-
cated behind the panels. The colored panels
are illuminated by lights that are controlled
by boom elevation during contact. The
colored panels that indicate boom tele-
scoping are not illuminated by background
lights. An illuminated white panel between
each colored panel serves as a reference.
The letters A for aft and F for forward are
visible at the ends of the boom telescoping
panel. The Air Refueling Director Lights
Profile (Figure 2-5) shows the panel illum-
ination at various boom nozzle positions
within the boom envelope. There are no
lights to indicate azimuth; however, a
yellow line is visible on the tanker to in-
dicate the centered position. When contact
is made, the panels automatically reflect
the correction the pilot must make to main-
tain position.
1-37
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Approved for Release: 2017/07/25 C06535936
Li-I airt2I3
RESET
GENERATOR OUT
L GEN
SELECTOR
SW ITCH
TRIP
L G N
CONTROL
TO GYRO
GROUND .011�
WARMUP
AC
EXT PWR
RECEPT
R GEN
CONTROL
,
.L XFMR RECTOUT
L XFMR RECT
200 AMP
at
L GENERATOR'BUS I
L GEN BUS SEL RELAY
NO. 1 N HEATER
L ENG FUEL SHUTOFF VALVE
BOOST PUMPS (8) (ODD)
PITOT HEATER
LANDING AND TAXI LIGHTS
PANEL LIGHTS
INSTRUMENT LIGHTS
INS EQUIP
HE AND TACAN EQUIP
L EGT TRIM MOTOR
UHF BLOWER AND HEATER
RCDR (INS - Q - BAY)
Q-BAY14__I
EQUIP
r. DC BUS
�
R GENERATOR OUT V..
DC
EXT PWR
RECEPT
INS
INS
MODE
SWITCH
�
INS BUS
ESS DC
BUS RELAY
EMER BAT ON
R XFMR RECT OUT
R XFMR RECT
200 AMP
4
R GENERATOR BUS
mini*Eimm&
TRIM PWR
OFF
R GEN BUS SEL RELAY
NO. 2 N HEATER
R ENG FUEL SHUTOFF VALVE
BOOST PUMPS (8) (EVEN)
R EGT TRIM MOTOR
MAN PITCH
AUTO PITCH
YAW
ROLL
(t)
TRIM
PWR
BUS
ON
INS
BAIT
6AH
BATE
@OFF
EXT
PWR
PWR
SWITCH
(CKPT)
ESS DC BUS
LAND R GENERATOR CONTROL
LG C I RCU ITS (31
FUEL PUMP CONTROL
FUEL XFER CONTROL (21
FUEL DUMP (4)
FUEL XFEED CONTROL
NLG STEER CONTROL
ENGINE FUEL SHUTOFF (2)
A/R CONTROL
EMER SPIKE CONTROL (2)
SPIKE OVERRIDE (2)
DRAG CHUTE ( 2 I
COCKPIT LIGHTS
TURN AND SLIP INDICATOR
DESTRUCT
UHF AND ADF
INTERPHONE
SPIKE SOLENOID (2)
ENG INLET AND BYPASS (2)
WARNING LIGHTS
BRAKE AND ANTI - SKID CONT
IGNITER PURGE
FACE PLATE HEATER
AIR CORD TEMP INDICATOR
AIR COND (2)
L AND R RUDDER LIMITER
L AND R HYD SYS CONTROL
TRIM CONTROL (2)
TACAN
AUTO PILOT
SAS (4)
FRS
NO. 1 AND NO. 2 N QTY IND
INV CONTROL (4),
DICTET
IFF/SIF
,PILOT VALVE CONTROL
SEAT ADJUST
Q BAY EQUIPMENT
RES HYD OIL CONTROL
HF RADIO AND SELCALL
PITOT HEAT CONTROL
DEFROSTER CONTROL
UHF INV PWR CONTROL (2)
BEACON LIGHTS
RCDR (INS - Q BAY)
PERISCOPE PROJECTOR
BATTERY ADF
RELAY BDH I
LAND R FUEL DERICH (2)
RAIN SPRAY
MAP DESTRUCT
SYST B-BW-RD 131
X BAND BEACON
ILS
CANOPY CAMERAS
LAND ROIL PRESS IND
FUEL TANK PRESS IND
A AND B HYD PRESS IND
LAND R SPIKE HYD PRESS IND
PITCH, ROLL, YAW, NAV. IND.
LAND R OIL TEMP
ADF
A, B, L, R, HYD QUANT
H NO. 4
INVERTER
H NO. 2
INVERTER
NO. 1
INVERTER
SWITCH
70
NORM OFF
INV
TRIM
ACTUATOR
XFMR
�
�
EMER
BATTERY #1
25 AMP-HR
�
EMER
BATTERY #2
25 AMP-HR
�
�
NO. 3
INVERTER
NO. 2
INVERTER
SW ITCH
EMER
INV
fir
NORM OFF EMER
INV INV
NO. 3
INVERTER
SWITCH
NORM OFF
INV
EMER
INV
=ammo DC POWER FLOW
IN��� AC POWER FLOW
PANEL LIGHTS
�
SAS PITCH A
SAS YAW A .
SAS ROLL A
FRS (2)
N QTY IND (2)
L AND R FUEL FLOW
LAND R EGT IND
FUEL QTY IND
LAND R EXH. NOZZLE IND
AIR COND-CKPT AND Q-BAY
LAND R CIT IND
STALL WARNING
FIRE WARNING
OXYGEN IND (2)
HE RAD 10
FLIGHT RECORDER
NO. 2 INV. OUTHI1
NO.2 INV BUS
STANDBY ATTITUDE INDICATOR
L SPIKE AND DOOR
SAS YAW B
SAS ROLL B
Q BAY EQUIP (3)
SAS PITCH B
FRS
OFF 42>
AUTO PI LOT
SELECTOR INS
SWITCH
3�INV,9PT....-
NO. 3 INV BUS
SAS P AND YAW MON
R SPIKE AND DOOR (21
MACH IND
INS (31
BEACON LIGHTS (3)
RECORDER (3)
AUTO PITCH
RO L AND PITCH SYNC.
AIR DATA COMPUTER
AIR DATA IND 1011
ATE GYRO AND IND
�
Approved for Release: 2017/07/25 C06535936
SECTION I
A-12
ELECTRICAL POWER SUPPLY SYSTEM
Three phase 115/220 volt ac power is pro-
vided by two engine driven generators rated
at 26 to 32 KVA depending on the installation.
Each generator supplies a separate ac bus
and a 200 ampere transformer rectifier.
Output of the transformer rectifiers is
paralleled and furnishes 28-volt ac power
to an essential dc bus and a monitored dc
bus and to a system of four 60 OVA in-
verters. In the event of a single generator
failure, a bus transfer and protection sys-
tem connects the two generator buses. Two
25-amp hour batteries are furnished to sup-
ply emergency power to the essential dc bus
in the event of complete power failure and a
smaller battery provides emergency power
to the INS and the No. 3 inverter.
AC ELECTRICAL POWER SUPPLY
Each engine drives an ac generator through
its remote gear box to supply 115/200 volt
3-phase power. There are no constant
speed drive units, so the ac frequency
varies directly with engine rpm; however,
the frequency is essentially constant at
scheduled engine speed during climb and
cruise. When the output of either generator
drops below 200 + 5 cps, it is automatically
tripped and the other generator automati-
cally provides power through the bus trans-
fer system. Generator cutout occurs at an
engine speed of approximately 2800 rpm.
Conventional switches are provided for
manual control of the generators.
EXTERNAL POWER SUPPLY
� The aircraft is equipped with two recepta-
cles for connecting ac and dc external power
sources to the aircraft electrical system.
These receptacles are located in the nose-
wheel well. When external power is con-
nected to the aircraft and the power switch
is in the EXT PWR position, the ac genera-
tors are automatically disconnected from
� their respective buses and the buses re-
ceive power from the ground power unit.
External dc power is paralleled with the
� dc output of the two aircraft transformer
rectifiers. External dc power and inverter
cooling air must be connected in order for
the external ac power to be available.
DC ELECTRICAL POWER SUPPLY
Electrical power for the essential and
monitored dc buses is normally suppliedby
the paralleled output of two 200-amp trans-
former rectifiers which are powered in-
dividually by the ac buses. The two 25
ampere-hour emergency batteries are fur-
nished to supply the essential dc bus with
power for a limited time when both trans-
former rectifiers or both generators are
inoperative.
AC INVERTER POWER SYSTEM
Fixed frequency ac power is supplied by
four 600 VA solid state air cooled inverters.
These inverters, located in the cheeks of
the nosewheel well, are controlled by cock-
pit switches and powered by the essential
dc bus. The No. 3 inverter is also con-
nected to the INS battery whenever the INS
mode switch is on. Normally the No. 1,
No. 2 and No. 3 inverters furnish power to
their respective buses. The No. 4 inverter
is normally off. Inverter power distribution
is so arranged that the No. 1 inverter bus
and its 26-volt instrument transformer
powers most of the flight and engine instru-
ments. The No. 3 inverter bus furnishes
ac power for the INS. In the event of in-
verter failure or other electrical system
malfunction, any one of the three inverter
buses may be operated from the No. 4 in-
1-39
Approved for Release: 2017/07/25 C06535936
SECTION I
Approved for Release: 2017/07/25 C06535936
A-12
CIRCUIT BREAKER PANELS (Typical)
0 0
z � z w cc �1
.5)
cc ce
(v)
ce 0
w LO - cc�
o3 >- C