A-12 FLIGHT MANUAL WITH TECHNICAL DATA CHANGE

• Document Type: FOIA
• Collection: FOIA Collection
• Document Number (FOIA) /ESDN (CREST): 00821248
• Release Decision: RIFPUB
• Original Classification: U
• Document Page Count: 459
• Document Creation Date: December 28, 2022
• Document Release Date: August 10, 2017
• Case Number: F-2014-00925
• Publication Date: June 15, 1968

11 Po'
Approved for Release: 2017/07/25 C00821248
IN
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO R F, INS F.; R T ED IN FRONT OF A-1 U TIL:TY
FLIGT1�1- MA Nil AT nATF,D 16 March 1968
Page I of 3
TDC NO. 1 1
10 May 1968
oyc-ey0-
SEC:1-1()N
PAC; F;
C,11AN(;11;
Lu
11
2-21
2-31
This TDC changes normal operation procedure for Descent and
Enc:ine Shutdown.
Insert:pao
Insert page 2-30A
The Abbreviated Checklist will be changed and replacing pages
furnished.
NOTE : The technical data information furnished herein is intended to be
used as INTERIM data.only. It will be replaced and superseded
at the time of issue of the next revision to the flight manual.
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A-12
Page 2 of 3
TDC No. 11
10 May 1968
NORMAL DESCENTS - Change to read as follows:
At Mach 1.5:
13. RPM - Check 6000 or above.
Maintain �,-_tt least 55C0 during remainder of descent to subsonic
speed..
Page 2 of 3
TDC; No. 11
10 1,,lay 1968
2-20A
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A-11
Page 3 of 3
TDC No. 11
10 May 1968
ENGINE SHUTDOWN - Change as follows:
CAUTION (same)
1. Wheel chocks - 7..nsted (same)
2. Canopy seal pressure levcr - OFF
3. Canopy - Open
4. INS - As briefed.
CAUTION
The INS should not be operated more than
5 minutes after opening the canopy to avoid
the possibility of excessive ".:.NS component
temperatures.
Balance of step procedure same.
Page 3 of 3
TDC No. 11
10 May 1968
2-30A
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COPY NO.
A-12 FLIGHT MANUAL
TECHNICAL DATA CHANGE
Page 1 of 1
TDC No. 9
16 March 1968
OXC-0364-61
COPY,2 OF 5/
This TDC transmits revised pages which replace and supersede
previously furnished pages for the Flight Manual dated 15 October
1967. Incorporation of previously furnished TDCIs provides ex-
panded performance which includes:
Revised Military Climb performance at various temper-
atures (1956 ARDC Atmosphere).
Revised Normal Climb performance at various temper-
atures for both 1956 ARDC and "Mean Tropic" Atmospheres.
Revised Cruise performance at various temperatures.
Revised Cruise Profiles covering:
Long Range Cruise'
High Altitude Cruise
Maximum A/B Ceiling Cruise
Additional descriptive and operating information has been incorporated
including Emergency forward transfer, updated engine time, EGT limits,
additional tire limits and a new drag chute deploy limits. The Pilot's
Abbreviated Checklist will be revised and issued to conform.
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COPY NO. 15
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LIST OF EFFECTIVE PAK, I
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SECTION I
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SECTION II
2-01
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pliptikk
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SECTION III
3-01
3-02
3-01
3-04
3-05
3-06
3-07
3-08
3-09
3-10
3-11
3-12
3-13
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Aluma
06-15-01
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:Page No! Issue
; 3-53
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SECTION IV
4-01
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. 4-30
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: 4-32
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e 4-34
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e 4-37
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* 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
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Changed 15 June 1968
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LIST OF EFFECTIVE PAGES
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SPCTION V
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PART I
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* 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. -II-Indi-
cates deletion of text.
Changed 15 June 1968
Approved for Release: 2017/07/25 C00821248
Issue Code C-2
LIST OF EFFECTIVE ,Approved for Release: 2017/07/25 C00821248
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T NDFX
I NDFX-01
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* The asterisk indicates pages changed, added, or deleted by NOTE: The portion of text affected by the change is indicated
the current change. Insert latest changed and/or added pages; by a vertical lino in the outer margins of the page. -I-Indi-
destroy superseded pages. cates deletion of text.
Issue Code C-2
IIIMIApproved for Release: 2017/07/25 C00821248
Changed 15 June 1968
Approved for Release: 2017/07/25 C00821248
A-12
TECHNICAL DATA CHANGE SUMMARY
TDC Date
Status
Superseded
by TDC 3
Inc.
Inc.
No. 1
No. 2
No. 3
10-16-67
1-26-68
2-05-68
_P2u2912
Est. Tropical Atmosphere Climb Performance
Emergency Forward Transfer
Revised Climb and Cruise Performance
(1956 ARDC Atmosphere & "MEAN
TROPIC" Atmosphere)
No. 4A
3-04-68
Time Limits ,8E 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 D/E
pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
�
COPY NO.
\.7-27Cd - e2.7f3 3-45. 1
, CO7Y 3 ,---�;-
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.
RETURN TO ARCHIVES Et RECORDS CENTER
IMMEDIATELY PEED USE
JOB Wi g�645/ BOX 7
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Al2
if4.4oSZOF
,t:11
ABLE OF CONTENTS
SECTION PAGE
I Description 1-1
II Normal Procedures 2-i
III 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
a:,ESCRIPTION
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 � RUDDER
SPIKE BLEED AIR OUTLET
FWD BYPASS
�SPIKE
�PITOT AND HF ANTENNA
� ADF LOOP ANTENNA
�TRANSLATOR-ARC-50 UHF ( R. H. SIDE)
� RCVR-XMTR-ARC-50 UHF ( L. H. SIDE)
�RETRACTABLE UHF ANTENNA
�R TACAN ANTENNA
AR RECEPTACLE DOORS
EX.PWR.RECEPT.
�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 TRANS CIEVER
�ANTENNA TUNING UNIT HF
�FRS COMPASS TRANSMITTER
�HF ANTENNA COIL
C")
1,1
rrl
�INBOARD ELEVON
ROLL AND PITCH MIXER
� YAW SERVOS RUDDER TRIM >
� EJECTOR FLAPS 70
7:1
rn
rrl
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
MOIIDaS
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A-12
SECTION I
INSTRUMENT PANEL
75
74
73
72
8
7 9
34
68 65 63 �60
71 70 17 69 67 66 64 59
1 AIR CONDITIONING CONTROL PANEL
2 AIRSPEED-MACH METER
3 BEARING DISTANCE HEADING
INDICATOR (BDHI
4 AN/ARC-50 RANGE 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 DI SPLAY INDI CATOR 38
23 IGNITER PURGE SWITCH
24
25
26
COMPRESSOR INLET
34
TEMPERATURE GAGE
35
ELAPSED TIME CLOCK
36
FIRE WARNING LIGHTS
37
58
57
56
55
54
11 13 15
10 12 14 16 1718
19 20 21 2223
0 0 0
o o 0 0 0 0
TACHOMETERS 39
EXHAUST GAS 40
TEMPERATURE INDICATORS 41
EXHAUST NOZZLE 42
POSITION INDICATORS 43
FUEL TANK SWITCHES
FUEL FORWARD TRANSFER SWITCH 44
QUAD HYDRAULIC QUANTITY
AIR REFUEL SWITCH 45
LIQUID NITROGEN QTY INDICATOR
FUEL TANK PRESSURE INDICATOR 46
RIGHT FORWARD PANEL
FUEL DUMP SWITCH 47
PUMP RELEASE SWITCH 48
FUEL QUANTITY INDICATOR 49
ILS PANEL 50
TEST N AND TANK LIGHT SWITCH 51
52
24 25
26 27
47 37
48 .42 40 38 35
43 41 39 36
49
53
54
55
51 56
57
58
59
60
53 61
62
63
FUEL FLOW INDICATORS 64
FWD BYPASS POSITION INDICATOR
65
66
67
68
69
50
52
OIL PRESSURE INDICATORS
SPIKE POSITION INDICATOR
HYDRAULIC SYSTEM (A AND 10
PRESSURE GAGE
HYDRAULIC SYSTEM (LAND R)
PRESSURE GAGE
COCKPIT PRESSURE SCHEDULE
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
70
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 DERICtiMENT 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
PITCH-ROLL-YAW TRIM INDICATORS
F200-14(1)
Changed 15 March 1968
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1-3
SECTION I
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A-12
COCKPIT LEFT SIDE
20
13
21
20
22
STBY AID
FAST ERECT
16
15
14
18
17
DEFOG
INCREASE
@HOLD
OFF
BCN ITS
'IFS BRIGHT
INSTR US
Ifte
OFF BRIGHT
PANEL ITS
12
ROLL TRUSS
r ADRS -RGE
RH RUDDER
SYNCHRONIZER
1-0--SEL -0,
Rt.CP1
10
ID-BAR SPECIAL PACKAGES PANEL
� � Z5
VOL
FRE()
�
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 IFF(S 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
F200-17(1)
1 -4
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A- 1Z
SECTION I
COCKPIT RIGHT SIDE
L 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. I NEEDLE SELECTOR SWITCH
& FLIGHT RECORDER SWITCH
9. FLOOD LIGHT SWITCH
10. FACE PLATE HEAT SWITCH
IL 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
ROLL
0%
;.EACH
1101D A ON ON
P KEAS AUTO HEADING
HOLD NAV HOLD
AUTOPILOT
TREE ROLL
t 11100 01
/T-3, PRESENT .-71-�
qp-o- POS 1E10% ...11-472.,D 1
i
, LAT
[...
di-A..... DESTP.A110% /FIX
N-17- .... j, .-
POS 1E10%
T
LA 2. I3,�133 14 1,5; 1
'= 7-np. VARIABLE , AilE0
"HAA.
I:. P DT
N LOT LONG
---j \,...L., 0" rR;C,:f %CY -,-----'
�
'
"/ N-.:,----
EU. 8-T: SIP
D TEST TEST ON
0` ON(
� OFF HIP OFF
CODE A ACTIVITY � CODE B -
Figure 1-4
FLIGHT
RECORDER
EtooD
FACE
w,.
10
F200-18(1)
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1-5
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1
5'1 aan2I3
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)
I Non 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
starting 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.
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1-7
SECTION I
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G
ENGINE AND A/B FUEL SYSTEM
COMPRESSOR
INLET TEMP
INDICATOR
CLOSE
4 $
STATIC PRESS
INDICATOR
MRUI
COMPRESSOR
BLEED ACTUATOR
STARTING
BLEED ACTUATOR
CLOSE
COMPR
BLEED
PILOT
-VALVE
ENGINE
DRIVEN
MAIN
FUEL PUMP
STARTING
BLEED
PILOT VALVE
5
FROM
FUEL
SYSTEM
BOOST
PUMP
311-
MAI N
GEARBOX
_Al If
FUEL/OIL
COOLER
MAIN FUEL
CONTROL
FROM
SMART
VALVE
FLOWMETER
FUEL FLOW
INDICATOR
WINDMILL
BYPASS AND'
DUMP VALVE
0 FILTER
A/B DETENT
OFF THROTTLE
HYDRAULIC
PUMP
� CODE
EZiGIEZZil
OIL
PRESSURE
INDICATOR
TO
SMART
VALVE
SOLENOID
VALVE
FUEL PRESS LOW
FUEL FLOW BURNERS
FUEL HYDRAULIC PRESS
FUEL DERI CH SYSTEM
ELECTR ICAL
Figure 1-6
EGT
INDICATOR
FUEL
DER I CH
ARM
EXHAUST
NOZZLE
ACTUATOR
noXICLOSE
EXHAUST NOZZLE
CONTROL VALVE
EXHAUST
NOZZLE
POSITION
INDICATOR
DERICH TO SMART VALVE
WARNING WHEN A/B IS OFF
LIGHTS (2)
ENGINE
PRESSURE
REGULATOR
A/B FUEL
CONTROL
A.AIR IN.
1:FROM
ENGINE
MAIN FUEL CONTROL COMPONENTS
PRESSURE REGULATOR VALVE
FUEL DENSITY SELECTOR
THROTTLE VALVE
PILOT VALVE
BURNER CAN LIMIT VALVE
AFTERBURNER FUEL CONTROL COMPONENTS
THROTTLE VALVE
PUMP REGULATOR
REC I RCULATING, BYPASS VALVE
PRESSURE REGULATOR VALVE
PEAK THROTTLE VALVE
rzoo-to(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 windmiLling 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 dear 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
pre s sure.
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 860 C 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 860oc 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 I
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 overspeed 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
100
Start And Bypass
Bleeds Closed
Ground
Idle
Start Bleeds
Bypass Bleed
Military Speed Schedule
....*
/ ..."'
/ /
..."' /
/ ...."
...." /
/
. /
Compressor Inlet Temperature �C
����
�
Bypass And Start
Bleeds Open
Windmill Band
100 200 300 400
Start Bleeds Exhaust To Nacelle Secondary Air Flow
7*Compres5or Bleed Air Bypass
Compressor Section
Burner
Turbine Section � Afterburner
Section Section
Figure 1-7
F 0 0 - 9 6
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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 incase
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
CIS DUMP
SOLENOID
FROM
FUEL VALVE
HYD
PUMP
ON
IGNITOR
PURGE SWITCH OFF
CODE
MAIN BURNER IGNITION �1�11111111111
MAIN IGNITION SIGNAL 22512151M72
DUMP SIGNAL
DUMP SIGNAL DRAIN ezazwzazara
FUEL COOLING IN cmunffor
TO A/B FUEL PUMP
CHEMICAL IGNITION SYSTEM
10U L)
0 u00000
Figure 1-8
COMPRESSOR
DISCHARGE PRESSURE
FUEL COOLING OUT
A/B IGNITION LINE
TURBINE DISCH PRESS
IGNITION SIGNAL
ELECTRICAL
F2.00-11(b)
1-14
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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 +15.6o + 3oC or
greater than 282.3C + 1 1�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 0oF 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
flarneholders 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 igniters.
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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1-15
SECTION I
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A-12
THROTTLE QUADRANT
1 THROTTLES
2 TRANSMIT BUTTON
3 MILITARY DETENT
4 THROTTLE FRICTION LEVER
5 MAX AFTERBURNER STOP
6 TEB SHOT COUNTERS
Figure 1-9
REV -11 -1d.56
FMA12713-(a)
1-16
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^
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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. Throtttle quadrants are
marked to indicate 82 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 OIFF 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 temparature. The
trim motors have a fuel flow or EGT travel
raw of about 150oC 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. They are calibrated
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 26-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 5)0 increments from 00C to
300C and 10 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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A-12
SECTION I
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
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Approved for Release: 2017/07/25 C00821248
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 betw. een 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
\
46s-,
FORWARD BYPASS
AFT BYPASS
\
us 1."
..-
---.
�
,,,,,e....._, ....
,.%_.g. \
---'s
���7
, Ve,.),,,,�.7.\ \
�
\
...../... .V.,00� �
1
.3.-
w:AZ \
)
/
-
..,-
oirl,..- /
,
:�-�.-
'
,-
,-
-- / ,-
401. ' -- ,-
-- \ 11 --
..
z ? ,
' -
/ \ ,-
/ I
}...-
..-.4\ ---
111111011M
CI
8-3t -65
F200-71(1)
Figure 1-10 (Sheet 1 of 2)
1 -21
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SECTION I
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A-12
INLET AIRFLOWS
,/w A
--47\ �
POROUS BLEED
DUCT SHOCK TRAP BLEED
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
F200-7 1(2)(a)
1 -22,
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A-12
SECTION I
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 ileR 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.
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SECTION I
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A-12
AIR INLET CONTROLS AND INDICATORS
Emmmommemos
RESTART
FWDD O O R 61) ^
2 OPEN OFF
MANUAL INLET
INLET
BELOW
30,000 FT
Figure 1-11
FM0-79
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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 itst 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
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. 10,2201b.
4 2, 135 gal. 13,7701b.
5 2, 136 gal. 13, 780 lb.
6 1, 978 gal. 12,7601b.
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
cQmponents 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 DEFUELI NG
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-
1-27
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�.1 - I aJn.213
FUEL
TANK NO. 1
FUEL
TANK NO. 2
14
FUEL
TANK NO. 3
FUEL
TANK NO. 4
10
15
13
FUEL
TANK NO. 5
FUEL
TANK NO. 6
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 CROSSFED VALVE
12 FUEL DUMP
13 JET PUMP (4 TOTAL)
14 LEFT FUEL MANIFOLD
15 AFT TRANSFER VALVE
Noi,Loas
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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. Whe3
a tank is empty, a yellow EMPTY light in
the pushbutton illuminates. When depressec
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-
strurnent 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 S/B 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 EM 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-
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A-12
SECTION I
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 I
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.
Fuel Quantity Low Light
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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ti-I a Inssm
12
11 FUEL
TANK NO. 1
FUEL
TANK NO. 2
FUEL
TANK NO.
1
FUEL
TANK NO. 4
FUEL
TANK NO. 5
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 Lk SYSTEM
12 SUCTION RELIEF LINE (NOSE WHEEL WELL)
13 LN2 FLOW FROM DEWAR TANKS
14 TO NITROGEN TANK PRESSURE SENSORS
mon pas
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SECT ION 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 Quantity 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 EXHAUST4
NOZZLE
ACTUATORS
PRIMARY
AIR COND
HEAT
EXCHANGER
SECONDARY
AIR COND
HEAT
EXCHANGER
HYDRAULIC
HEAT
EXCHANGER
SPIKE
HEAT
,EXCHANGER
�
TO
TO MAIN
BURNER
OPEN WHEN
A/B IS OFF
.TO A/B
��^�
CONTROL
4^1)� ENGINE
VARIABLE MC) HYD
ORIFICE MI PUMP
WINDMILL
AND
BYPASS
VALVE
MAIN
FUEL
CONTROL
ENGINE OIL MAIN FUEL
HEAT PUMP 2ND
EXCHANGER STAGE
A/B
PUMP
MAIN
FUEL
PUMP
1ST
STAGE
THIS VALVE ALWAYS
PERMITS FLOW INBD,
BUT WILL PERMIT
FLOW IN BOTH
DIRECTIONS WHEN
CROSSFEED VALVE IS
MIXING
VALVE
PRESSURE
OPERATED
TANK NO.4
RETURN
VALVE
RETURN TO
TANK NO.4
I TO RH
MIXER
1
CROSS SWITCH
FEED
s.
FUEL TEMP, CONTROL
PRESSURE (SMART) VALVE
0 SWITCH I
11
FUEL TO am .
1:1 ��!!!��
LH ENGINE
REMOTE
GEARBOX
HEAT
EXCHANGER
FLOWMETER
CIRCULATING
.FUEL PUMP
Figure 1-15
FILTER
�
SENSE
LINES
RH SYSTEM
I DENTI CAL
CliOSSFEED VALVE
(OPEN FOR SINGLE
ENGINE OPERATION)
�I FUEL TO
�
RH ENGINE
LH ENGINE
FEED LINE
FROM BOOST
PUMPS
EMERGENCY
SHUTOFF VALVE
NORMALLY OPEN
AIRPLANE
� PIM
RH ENGINE
FEED LINE
FROM BOOST
PUMPS
F200-38(d)
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. If the 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 ?op
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 295o 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 sidd of the instrument panel. The
switch has three positions; READY, 017
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
�ImillinmilmmimilmomApproved for Release: 2017/07/25 C00821248
1-35
917Z [Z9000 SZ/LO/LI.OZ :aseaia JOI ponaiddV
91-1 a .It1TJ
FUEL
TANK NO. 1
FUEL
TANK NO. 2
4 FUEL
TANK NO. 3
FUEL
TANK NO. 4
FUEL
TANK NO. 6
FUEL
TANK NO. 5
1 MR REFUELING RECEPTACLE
2 REFUELING MANIFOLD
3 PILOT VALVE (6 TOTAL)
4 FLOAT VALVE SHUTOFF (6 TOTAL)
1\101,1,03S
Approved for Release: 2017/07/25 C00821248
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 second time 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 IFR PUSH TO RESET, is lo-
cated at the top left side of the instrument
panel. The top half 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 100 + 5 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.
Changed 15 June 1968
Approved for Release: 2017/07/25 C00821248
1-37
Approved for Release: 2017/07/25 C00821248
LI-I aIn2I-4
kGENERATOR OUT
'
RESET
L GEN
SELECTOR
SW ITCH
TB I P
L G N
CONTROL
TO GYRO
GROUND ���---
WARMUP
AC
EXT PWR
RECEPT
RESET
R G N
CONTROL
� 1110
�
XFMR RECT OUT
L XFMR RECT
200 AMP
� � 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
HF RADIO
L EGT TRIM MOTOR
UHF BLOWER AND HEATER
RCDR (INS � Q - BAY)
O�BAYH
EQUIP
MON DC BUS
DC
EXT PWR
RECEPT
INS
INS
MODE
SWITCH
INS BUS
ESS DC
BUS RELAY
R XFMR RECT OUT
ViV
R XFMR RECT
200 AMP
� N�110
R GENERATI'OUT
R GENERATOR BUS
01
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
TRIM
PWR
BUS
ON
INS
BAIT
6AH
BATE
EXT
PWR
BATTERY
RELAYS
OFF
PWR
SWITCH
(CKPT)
TRIM
ACTUATOR
XFMR
EMER
BATTERY #1
25 AMP�HR
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, B. HYD QUANT
INST
XFMR
26V
EMER BAT ON
ESS DC BUS
NO. 1 INV OUT
LAND R GENERATOR CONTROL
LG CIRCUITS (3)
FUEL PUMP CONTROL
FUEL XFER CONTROL 121
FUEL DUMP (4)
FUEL XFEED CONTROL
NLG STEER CONTROL
ENGINE FUEL SHUTOFF (2)
H
NO. 4
INVERTER
411+ NO. 1 INV BUS
SAS PITCH A
SAS YAW A
AIR CONTROL
SAS ROLL A
EMER SPIKE CONTROL (2)
FRS (2)
SPIKE OVERRIDE (2)
N (NY IND (21
DRAG CHUTE (2)
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
NO. 1
INVERTER
SW ITCH
0
L AND R FUEL FLOW
L AND R EGT IND
FUEL QTY IND
LAND R EXH. NOZZLE IND
AIR COND�CKPT AND 0�BAY
LAND R CIT IND
STALL WARNING
FIRE WARNING
OXYGEN IND (2)
HF RADIO
FACE PLATE HEATER
FLIGHT
RECORDER
NORM OFF
EMER
AIR COND TEMP INDICATOR
INV OUTM
AIR COND (2)
INV
INV
NO. 2
L AND R RUDDER LIMITER
L AND R HYD SYS CONTROL �
NO. 2 INV BUS CIP
TRIM CONTROL (2)
���1
TACAN
ATTITUDE INDICATOR
AUTO PILOT
STANDBY
SAS (4)
NO. 2
L SPIKE AND DOOR
FRS
INVERTER
SAS YAW B
NO. LAND NO. 2 N QTY IND
SW ITCH
SAS ROLL B
INV CONTROL (4),
NO. 2
Q BAY EQUIP 131
DICTET
IFF/SIF
H
INVERTER
SAS PITCH B FRS
CI P OFF
PILOT VALVE CONTROL
SEAT ADJUST
�IP
AUTO P I LOT
SELECTOR iNS
NORM OFF
EMER
Q BAY EQUIPMENT
INV
INV
SWITCH
RES HYD OIL CONTROL
NO. 3
INV OUT
HF RADIO AND SELCALL
PITOT HEAT CONTROL
DEFROSTER CONTROL
UHF INV PWR CONTROL 121
BEACON LIGHTS
NO. 3
RCDR (INS -0 BAY)
PERISCOPE PROJECTOR
ADF
INVERTER
SWITCH
NO. 3
NO. 3 INV BUS
BDHI
INVERTER
SAS P AND YAW MON
LAND R FUEL DERICH (21
R SPIKE AND DOOR (21
RAIN SPRAY
MACH IND
MAP DESTRUCT
SYST B-BW�RD (31
X BAND BEACON
NORM OFF EMER
INV INV
INS (3)
BEACON LIGHTS (31
RECORDER (3)
IIS
CANOPY CAMERAS
DC POWER FLOW
AUTO PITCH
RO L AND PITCH SYNC.
1����� AC POWER FLOW
AIR DATA COMPUTER
AIR DATA IND (TDI)
EMER
BATTERY #2
25 AMP�HR
PANEL LIGHTS
All GYRO AND IND
rn cn
tx1
rn
C")
7, 0
C)
r-
-0
rn
7c)
C7Z)
Approved for Release: 2017/07/25 C00821248
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 600VA 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 supplied by
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
milmimmissmoommipproved for Release: 2017/07/25 C00821248
SECTION I
Approved for Release: 2017/07/25 C00821248
A-12
CIRCUIT BREAKER PANELS (Typical)
Z
Ce
0
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CD Lu <
1�..Z
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LL1
4.1 ce
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LL,
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0
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0
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1:Z
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1
0 z
cL 0 LL. U- 0
Z
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cc
DETAIL C LEFT CONSOLE
aCt
ESSENTIAL DC
7..JCANOPY
CAMERAS I. L.S.
00
ESSENTIAL DC
R-D HEAT
CHINE
PITOT PITOT
HEAT HEAT
TAXI LT TACAN AND HF
00
FLOOD � 0
UHF
0TACAN LoIGHTS 0RCDR HEATERS LDG S
L AC GEN
0 I
INSTR
PANEL
H-F BEACON
MUTE X BAND
- ESSENTIAL DC
DETAIL B
UPPER LEFT
UNLOCK
OFF
TRIM IND NAV r-OIL PRESS-1
PITCH ROLL 0 0 0 0 0
YAW IND R ' ADF
GEAR
GEAR
RELEASE
RELEASE
HYD PRESS
A CONT B L SPIKE R
0 0 a
HYD 0 IL FUEL TK r- 0 IL TEMP
cWY PRESS L
DETAIL D
LOWER INSTRUMENT PANEL
Figure 1-18 (Sheet 1 of 2)
F200-35(1)(g)
1 -40
MNIMIlApproved for Release: 2017/07/25 C00821248
MENWITO
la � �rove� or - e ease Se: A
SAS
PITCH A
FRS
FUEL QTY
(0)
ANGLE _
ATTK
NO.! NQTY L FUEL FLOW
0 0
AIRCONDc NO.1 OXYs
0 K (0)Y
:R iKE' PE
NO.2 OXY
0
0 (0)
FRS AIR DATA IND
0 0
L [CT IND R EGT IND
NO.2 N QTY
(
AIR CONDO
A
A
R-F I
NO.1 INV
R FUEL FLOW
0
0� PF
\\E
SAS SAS SAS _
YAW A SIP _ I ROLL A All IND � PITCH B
0 0
FRS AUTO PLT
(0)
LCIT IND R CIT IND
P-F
FLT REC
L [NP
INST XFNIR
AIR DATA
C)
RENT P
0-BAY EQUIP
L SPIKE
0PF
SAS �
YAW B
0-BAY EQUIP
0
L SPIKE
R 0 D
0PF
SAS _
ROLL B
0
Q-BAY EQUIP
0
ATT GYRO
0PE
CIPI SYS
0
NO.2 INV
INS W INS INS PE
0 (1()) 10)
BCN LTS BCN ITS BCN ITS
0 0 0
RCDR A RCDR RCDR
R SPIKE
PF
P 0
H SAS PITCH AUTO PIT
A
Y
A 0
E W
R SPIKE
AIR DATA IND
NO.3 INV
AIR DATA
NP'
0PE
A
MEMMELE
� �rove� or - e ease ii; A
SECTION I
Approved for Release: 2017/07/25 C00821248
A-12
LEFT AND RIGHT FORWARD PANELS
LEFT AND RIGHT FORWARD PANELS
12
GEAR AND WARN
LT TEST
2
1 OXYGEN QUANTITY GAGE
2 LANDING GEAR LEVER
3 FUEL QUANTITY INDICATOR SELECTOR SWITCH
4 INVERTER SWITCHES
5 GENERATOR SWITCHES
6 MAP DESTROY SWITCH
7 BATTERY SWITCH
8 FUEL QUANTITY INDICATOR
9 CIP AND OXYGEN TEST SWITCH
10 CABIN ALTIMETER
11 CABIN ALTIMETER SELECTOR LEVER
12 GEAR AND WARNING LIGHTS TEST BUTTON
F200 -54(f)
Figure 1-19
1 -42
iiiIIMI=IMMNI1Approved for Release: 2017/07/25 000821248
Approved for Release: 2017/07/25 C00821248
SECTION I
A- 12
verter power supply. Certain related equip-
ment is transferred from the No. 1 and No.
3 inverter by operation of the autopilot se-
lector switch to maintain the proper power
phase relationships. The AN/ARC 50 UHF
radio has its own rotary inverter supply.
Refer to Electrical Power Distribution dia-
gram this section.
ELECTRICAL SYSTEM CONTROLS AND
INDICATORS
Circuit Breakers
The cockpit circuit breaker panels are lo-
cated on the right and left consoles and be-
low the annunciator panel. The circuit
breakers are push to reset, pullout type
breakers for certain ac and dc circuits as
listed on the electrical power distribution
chart, figure 1-17. Circuit breaker panels
which are not accessible during flight, but
which should be inspected before flight, are
located in the air conditioning bay (just for-
ward of the refueling receptacle) and elec-
trical load center (left hand side of nose-
wheel well).
Generator Switches
A switch for each generator is located on
the right side of the instrument panel and is
powered by the essential dc bus. Each
switch has three positions; GEN RESET
(up), TRIP (down) and center (neutral). The
switches are spring loaded to the center
neutral position. Holding the switch in the
GEN RESET (up) position will return the re-
spective generator to normal operation if it
has been removed from the bus for any rea-
son other than complete generator failure.
In the TRIP (down) position, the generator
output will be removed from the generator
bus and the auto bus transfer system will
supply that bus from the other generator if
It is operating.
NOTE
The generators must be reset and
connected to the bus after the en-
gines are started and before the
ac ground power is removed.
Power Switch
A three-position battery-external power
switch is located on the right side of the
instrument panel. When in flight or on the
ground with ground power disconnected,
placing the power switch in the BAT (up)
position causes the emergency batteries to
supply power to the essential dc bus. In
the EXT PWR (down) position, the external
power sources furnish power for the elec-
trical systems. In the center OFF position,
external ac power is disconnectei but power
from the dc external receptacle will con-
tinue to supply the essential and monitored
dc buses and dc power will not be inter-
rupted by moving the power switch from the
EXT PWR to OFF positions.
Inverter Switches
Switches for No. 1, No. 2 and No. 3 in-
verters are located on the right side of the
instrument panel below the generator
switches. In the NORM (up) position, the
respective inverter is energized and sup-
plies power to its individual bus. In the
OFF (center) position the inverter is dis-
connected from the essential dc bus. In
the EMERG (down) position the No. 4 in-
verter is activated and connected to that
inverter bus. In the event of multiple in-
verter failure, the lowest numbered in-
verter switch that is placed in the EMERG
position receives power from the No. 4 in-
verter. Under this condition, a higher
numbered inverter can not receive power
even if its inverter switch is in the EMERG
1-43
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Approved for Release: 2017/07/25 C00821248
SECTION I
A-12
position. No. 3 inverter also may receive
dc power from the small INS battery if the
INS mode switch is not in the OFF position.
Generator Out Indicator Lights
The L and R GENERATOR OUT indicator
lights, located on the annunciator panel, il-
luminate when a generator is not furnishing
power to its ac bus.
Transformer-Rectifier Out Indicator Lights
The L and R XFMR-RECT OUT indicator
lights, located on the annunciator panel, il-
luminate to indicate that the respective
transformer-rectifier is not furnishing
power to the dc buses.
Inverter Out Indicator Lights
Three INVERTER OUT indicator lights are
located on the annunciator panel. When il-
luminated, the numbered light indicates
that the respective inverter bus voltage is
too low. An inverter switch must be placed
in the OFF position to disconnect that in-
verter from the bus. When a disconnected
inverter is switched to the EMERG position,
the No. 4 inverter is activated and will fur-
nish power to the respective inverter bus
and the light will be extinguished unless a
lower numbered inverter switch has already
been turned to EMERG.
Emergency Battery On Indicator Light
The EMER BAT ON light located on the an-
nunciator panel illuminates when the emer-
gency batteries are furnishing power to the
essential dc bus.
HYDRAULIC POWER SUPPLY SYSTEMS
Four separate hydraulic systems are in-
stalled on the aircraft, each with its own
pressurized reservoir and engine-driven
pump. The pumps for the A and L system
are driven from the left engine remote gear
box and the B and R system pumps are dri-
ven from the right engine remote gear box.
Hydraulic fluid is cooled by fuel-oil ex-
changers, using the aircraft fuel supply as
the cooling agent. The A and B hydraulic
systems provide power for operating the
flight controls. The L and R systems pro-
vide power for all other hydraulic require-
ments of the aircraft. Under normal op-
erating conditions, the systems are inde-
pendent of one another. The L hydraulic
system provides hydraulic power to the
left engine air inlet control, the landing
gear (including uplocks and door cylinders),
normal brakes, in-flight refueling door,
UHF retractable antenna, and normal nose-
wheel steering. The R hydraulic system
provides hydraulic power V) the right air
inlet control and also to the alternate
brakes, nosewheel steering, refueling door
and landing gear (emergency retraction
only) when the L hydraulic system has
failed. When the R hydraulic system sup-
plies power to the brakes, the anti-skid
feature is inoperatire.
Hydraulic System Pressures Gages
Two dual indicating hydraulic gages are in-
stalled on the lower center portion of the
instrument panel. The right hand gage in-
dicates hydraulic pressure of the A and B
(flight controls) systems, and the left hand
gage indicates hydraulic pressure of the L
and R systems. The gages are calibrated in
100 psi increments from 0 to 4000 psi.
Pressure indication on the gages is accom-
plished by means of remote transmitters in
the individual systems. Twenty-six volt ac
power is furnished by the instrument trans-
former and the No. 1 inverter.
1-44
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Approved for Release: 2017/07/25 C00821248
A - 1 2
SECTION I
A AND B HYDRAULIC POWER SUPPLY SYSTEM
,...1
ONE GALLON
LOW LEVEL
SWITCH
RETURN
CONNECTION
SHUTOFF l�
VALVE_
RES
PRESSURE
A RES
RELIEF VALVE RELIEF VALVE
QUANTITY
INDICATOR
L.
RETURN
FILTER
LI TEMPERATURE
CONTROL
HYD QUANTITY
GAGE
U.
II
I ILI HEAT
HEAT
EXCHANGER gri a
EXCHANGER
RELIEF
VALVE
1 TEMPERATURE
CONTROL
TO B
RESERVOIR
VENT VALVE
RETURN
FILTER
r an ow �
SEAL
DRAIN
es sal lie as s. es as as la as as a. .. 1, 0 1, ma ss *a 4. s aa es la a .1 ea e. la
gOSHUTOFFL...
VALVE
PRESS
rCONN
PRESSURE
FILTER
N2 PRESS
N2 FILL
N2 CYL
FIll PORT
'SHUTOFF
VALVE
HY
LOW1-
......
RESTRICTOR
PRESS
SWITCH
ACCUMULATOR
TO SURFACE
CONTROLS
EJD
PRESS TRANS
N2 FILLER
N2 GAGE
Ina.*
RESERVE
HYD TANK
OFF
HYD RES OIL
11��,�1
FROM SURFACE
CONTROLS
OVERBOARD
RELIEF
RELIEF e
VALVE
�
RESTR I CTOR
PRESS TRANS
N2 GAGE
PRESS
SW ITCH
ACCUMULATOR
ONE GALLON
LOW LEVEL
SWITCH
RETURN
CONNECTION
SHUTOFF
VALVE
RES
PRESSURE
j SHUTOFF
VALVE
PRESS
CONN _1
PRESSURE
FILTER
TO SURFACE
CONTROLS
moo:a A SYSTEM PRESSURE B SYSTEM PRESSURE . ELECTRICAL
enunaraza A SYSTEM RETURN raezmrae23. B SYSTEM RETURN RESERVE OIL SUPPLY
_
rrrrirrrwmwr CASE DRAIN mourom N2 PRESSURE
Figure 1-20
REG
N2 PI
N2 F
I
L
L
-
N2 CYti_
F200-20.11-)
Approved for Release: 2017/07/25 C00821248
1 -4 5
SECTION I
Approved for Release: 2017/07/25 C00821248
A-12
L AND R HYDRAULIC POWER SUPPLINHYD gEm
RELIEF
VALVE
.1.wA
RESERVOIR
PRESS IND
SHUTOFF r
VALVE
REG-
GAGE
N2
FILLER
N2
CYLINDER
HYD
PUMP
-Ft
- - - - -
FECIM
HEAT
EXCHANGER
AFT BYPASS
ACTUATOR
RETURN
BYPASS
ACTUATOR
AND SERVO
SPIKE
CONNECTION ACTUATOR
AND SERVO
(!)
CROSS-
OVER
VALVE
(RETURN)
L R
=MEN
ALTERNATE BRAKE
RETURN SELECTOR
VALVE
SYSTEM
RELIEF
VALVE
PRESS
TRANS
PRESSURE
CONNECTIONS
LO
IC:IS .0
HYDRAULIC
PRESSURE
SPIKE
R H D LO
N2 FILLER
ANTI-SKID ON
OFF
413XECE 8
HEAT
EXCHANGER
ONE GALLON
LOW LEVEL
SW ITCH
ALTERNATE
STEER AND I
BRAKE
I- - BRAKE'
I S /0 I
VAL
I
� --- -- �
�
NORM
BRAKE
S/O VALVE
NORM
BRAKE
SYST
ALT
BRAKE
BRAKE
SYST
mu4
1-1 CROSSOVER
PRESSURE
I SWITCH
N
_
-.ALT STEER I
SIO VALVE,
1
ALT STEER
amjg S/0 VALVE I
REFUELING
DOOR AND
PROBE
LANDING
GEAR
GEAR UP
STEERING
UNIT
mn.11
AFT BYPASS
ACTUATOR
BYPASS
ACTUATOR
AND SERVO
SPIKE
ACTUATOR
AND SERVO
CROSSOVER
VALVE (PRESSURE)
R L
GEAR DOWN
Figure 1-21
wwilam L SYSTEM PRESSURE
Kim L SYSTEM RETURN
crIcl= CASE FLOW LINE
anon�no R SYSTEM PRESSURE
GnmEnswzai R SYSTEM RETURN
---- ELECTRICAL
F200-21(0
1-46
Approved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
SECTION I
A-12
Hydraulic Warning Lights
Six hydraulic warning lights are located on
the annunciator panel. The A and B HYD
PRESS LOW lights will illuminate when the
pressure in the respective system drops be-
low 2200 +0 -150 psi. The A and B HYD
LOW light will illuminate when the quantity
is less than 1-1/4+ 1/8 gallons. The L and
R HYD LOW light will illuminate when the
respective reservoir quantity is less than
1-1/4+ 1/8 gallons. Power for the lights
I. furnished by the essential dc bus.
Hydraulic System Quantity Gage
A quadruple hydraulic fluid quantity indi-
cator installed on the right side of the in-
strument panel. The L and R concentric
needles are on the left side of the gage and
the A & B concentric needles are on the
right side of the gage. The dials are
marked in gallons. Power is furnished
from the 26 V ac instrument transformer.
HYDRAULIC RESERVE OIL SYSTEM
A reserve oil supply for the A and B hy-
draulic systems is contained in an 8.5 gal-
lon reserve tank mounted in the No. 4 fuel
tank. The reserve hydraulic oil is trans-
ferred by gravity flow and nitrogen pres-
sure through solenoid operated shutoff valves
to either the A or B hydraulic system.
Hydraulic Reserve Oil Switch
The hydraulic reserve oil switch is mounted
on the left side of the annunciator panel. It
is a three position switch, guarded in the
center OFF position. In the A (up) position,
solenoid operated shutoff valves are opened
to the A hydraulic system suction and tank
vent lines. This allows the reserve hy-
draulic fluid to supply the A system as
needed up to approximately 0.3 gallon per
minute. In the B (down) position the sole-
noid valves to the B system are opened and
the reserve fluid will supply the B system.
Power for the valves is furnished by the
essential dc bus.
WARNING I
Reserve hydraulic fluid is to be
used only to supply the operative
A or B system in the event of
malfunction of the other system.
FLIGHT CONTROL SYSTEM
The cockpit flight controls consist of a con-
ventional control stick and rudder pedals.
The delta wing configuration utilizes elevons
instead of separate aileron and elevator
control surfaces. The elevons, moving to-
gether in the same direction, function as
elevators and when moving in opposite di-
rections, function as ailerons. Each ele-
von consists of an inboard and outboard
panel with the inboard panel located between
the fuselage and the nacelle and the out-
board panel outboard of the nacelle. Both
panels on one side function as a single unit
with the servo input to the outboard elevon
connectea directly to the inboard elevon
surface. The dual canted rudders are full
moving, one piece, pivoting surfaces with
a small fixed stub at the junction of the
vertical surface and the nacelle. Deflection
and control of the elevons and rudders is by
means of dual, full hydraulic, irreversible
actuating systems.
Control surface travel limits are as follows:
Elevons Rudders
Pitch
10o Down
24o Up
_
Pitch plus
Roll
20� Down
35o Up
-
Yaw
-
20o Left
20� Right
Roll
12o Down
120 Up
-
IIM=11111=Approved for Release: 2017/07/25 C00821248
1-47
917Z [Z9000 SZ/LO/LI.OZ :aseaia JOI panaidd
CABLE QUADRANT
RUDDER PEDALS
CABLE TENSION
REGULATOR
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SERVO SURFACE LIMITER1
SOLENOID VALVE
r�lri SERVO SURFACE LIMITER
CYLINDERS
R P
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HYD. SYSTEMS
A AND B
ELECTRO-HYDRAULIC ENGAGE & TRANSFER VALVE
MOD PISTON WITH CENTERING SPRING
LINEAR TRANSDUCER
ACTUATING CYLINDERS
SURFACE LIMITER
SURFACE LIMITER CONT.)
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is SWITCHES FOR SERVO FM.
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POSITION WARNING
DUAL HYD. CONTROL VALVE
& BIAS SPRING
PUSH RODS
IN NACELLE
TERMINAL QUADRANT
IN WING
/SINGLE CAM SYSTEM
TO EACH RUDDER
FOLWWUP ROD
SERVO INPUT LEVER�.\
PUSH ROD IN FIN
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& TRIM POSITION TRANSMITTER
AIRCRAFT
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Approved for Release: 2017/07/25 C00821248
SECTION I
A-12
Manually operated mechanical stops are in-
corporated in the cockpit mechanism to limit
the surface movement at hig(/)a speedo. Elevon
travel in roll is limited to 7 up, j down
and rudder travel is limited to 10 right,
100 left. An additional stop is installed in
each rudder servo package to limit the rud-
der travel. These stops are electrically
controlled and hydraulically operated by
separate electrical and hydraulic systems.
If no electrical power is available, the rud-
ders will be limited to approximately 10 L
and R travel. If electrical power is avail-
able to on% stop, that rudder only will have
the full 20 L and R travel available. The
rudder cable must be stretched to obtain
this travel, causing a noticeable increase
in rudder pedal force.
CABLE SYSTEM
Cable systems are utilized to transfer con-
trol movements from the control stick and
rudder pedals to the flight control mechan-
isms. The pitch and roll axis cable sys-
tems are duplicated from the cockpit to the
mixing mechanism in the aft fuselage. The
rudder system has two separate closed
loop single cable systems, one to each rud-
der. Cable tension regulators and slack
absorbers are incorporated in the cable
systems.
TRIM CONTROL SYSTEM
Flight control trim is accomplished by de-
flecting the control surfaces through the
use of electrical trim actuators. The roll
and pitch trim actuators are located down-
stream of the feel springs so that stick
position remains neutral, irrespective of
the amount of trim. The trim actuator and
feel spring location is combined in the rud-
der mechanism and yaw trim is reflected
by rudder pedal position.
Travel limits of the trim system are 3-1/2�
down to 6-1/2� up in pitch, 4.5� up and
down (each side) in roll, and 10 left to 100
right in yaw. Trim position indicators are
provided for each axis. Trim rates are as
follows:
Pitch
Roll
Yaw
Max.
1.5�/sec
.954D/sec
Total
Diff.
1.5�/sec
Min.
0.67�/sec
.47�/sec
Total
Diff. _
0.67�/sec
Automatic pitch trim uses a separate, slow
speed motor for autopilot synchronization.
The automatic pitch trim rate is 0.15 /sec
maximum and 0.067 /sec minimum. Trim
power is normally furnished by the R gen-
erator bus.
RUDDER PEDALS
Primary control for the rudders consists
of conventional rudder pedals mechanically
connected by cables, bell cranks and push-
rods to hydraulic control valves at the rud-
der hydraulic actuators. The rudder pedals
are released for adjustment by pulling the
T-handle labeled PEDAL ADJ located be-
low the annunciator panel. Wheel brakes
are controlled conventionally by toe action
on the rudder pedals; refer to Wheel Brake
System, this section. Rudder pedal move-
ment also controls nosewheel steering;
refer to Nosewheel Steering System, this
section. The pedals are hinged to fold in-
board and upward, providing foot space on
the cockpit floor.
1-49
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pproved for Release: 2017/07/25 C00821248
I laatiS) �2-1 aInT3
aCc)
�
CABLE
DISCONNECTS
CABLE TENSION
\ .._
REGULATOR AND
SLACK ABSORBER
(PITCH)
a
SWITCHES FOR
SURF-LIMITER
WARNING
DUAL HYDRAULIC CONTROL VALVE
AND BIAS SPRING (INBD)
ROD FROM MIXER
TO INBOARD SERVO (R. H.)
SURFACE LIMITER
CONTROL
CONTROL
STICK
{ELECTRO - MECHANICAL
ROLL TRIM ACTUATOR WITH
POSITION TRANSMITTER
1,z)43
CABLE TENSION REGULATOR
AND SLACK ABSORBER (ROLL)
ra
ELECTRO-HYDRAULIC ENGAGE
AND TRANSFER VALVE (ROLL)
ELECTRO - HYDRAULIC ENGAGE
AND TRANSFER VALVE (PITCH)
OUTBOARD
CONTROL SURFACE
ACTUATING
CYLINDERS (6)
ANTI - BIAS
SPRING
ima,7
7-1\
!Col
PITCH MIXER
STOPS
ROD FROM MIXER
TO INBOARD SERVO
(L H.)
INBOARD
CONTROL
SURFACE
ROLL FEEL SPRING
PITCH FEEL SPRING
PITCH QUADRANT
IN TAIL CONE
ROLL QUADRANT
IN TAIL CONE
TRIM ACTUATOR -2 SPEED]ELECTRO-MECHANICAL
PITCH
FMA.12 -2 3
W31SAS10211NO3 1H9
Nois,Das
1-4
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A-12
FLIGHT CONTROL SYSTEMS (OUTBOARD ELEVONS )
DUAL HYD. CONTROL VALVE
AND BIAS SPRING (OUTB'D
TORQUE TUBE
IN NACELLE
PUSHRODS IN
WING ( INB'D
PUSHRODS IN
WING (OUTIVD)
TO TUBES
IN NACELLE
Figure 1-23 ( Sheet 2 of 2)
Section I
ACTUATING CYLINDERS (14)
SPRING CARTRIDGE LIMITER
DETENTED SPRING
CARTRIDGE
INB'D CONT.
SURFACE
3-30-66
F200-3(a)
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1-51
SECTION I
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A-12
CONTROL STICK GRIP
TOP VIEW
FRONT VIEW
SIDE VIEW
1 TRANSMITTER-INTERPHONE CONTROL SWITCH
2 PITCH AND YAW TRIM SWITCH
3 CONTROL STICK COMMAND-NOSEWHEEL
4 JAM O'RIDE SWITCH
5 EMERGENCY AUTOPILOT DISENGAGE SWITCH AND AIR
REFUEL DISCONNECT
Figure 1-24
FZ00-24(c)
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SECTION I
A-12
ARTIFICIAL FEEL SYSTEM
The use of a full power irreversible control
system for actuation of the surfaces prevents
air loads and resulting "feel" from reaching
the cockpit controls. Therefore, feel
springs are installed in each of the pitch,
roll and yaw axis mechanisms to provide an
artificial sense of control feel. The springs
apply loads to the pilot controls in pro-
portion to the degree of control deflection.
CONTROL STICK
The control stick is mechanically connected
by a torque tube, push rods and bell cranks
to the dual cable system which operates the
roll and pitch quadrants in the aft fuselage
tail cone. Mechanical push rod linkages
mix the control movements and position dual
hydraulic control valves. These valves di-
rect both A and B system hydraulic pres-
sure to the inboard elevon actuating cy-
linders.
Push rods, bell cranks and torque tubes
transfer inboard elevon deflection to posi-
tion the outboard dual hydraulic control
valves. These valves direct both A and B
system hydraulic pressure to the outboard
elevon actuating cylinders. A push rod
blowup system closes off the flow of hy-
draulic fluid to the actuators when the de-
sired elevon deflection is obtained. Lo-
cated on the control stick grip is a com-
bination pitch and yaw trim switch, an auto-
pilot control stick command, a nos ewheel
steering button, a microphone switch for
both interphone and radio transmission, a
combination autopilot disconnect and in-
flight refueling disconnect switch and a
jam override pushbutton.
Control Stick Command Switch (CSC)
Refer to Autopilot System, Section IV.
Pitch and Yaw Trim Switch
Pitch and yaw trim control is provided by
a spring-loaded, four position thumb ac-
tuated switch installed on the control stick
grip with a center OFF position. The
switch positions are LEFT, RIGHT, NOSE
UP and NOSE DOWN. The switch controls
trim motors powered by the right generator
bus through the 28-volt ac trim actuator
transformer and trim power bus.
NOTE
The trim power switch must be in
the ON position before the pitch,
roll and yaw trim switches will
operate.
Lateral movement of the switch to the left
corrects for right yaw and lateral move-
ment to the right corrects for left yaw.
Forward movement of the switch produces
down elevon operation of the trim motors
and actuators (aircraft nose down). Aft
movement moves the elevons up (aircraft
nose up).
Trim Power Switch
A trim power ON-OFF switch is located on
the annunciator panel. It enables the pilot,
if necessary, to disconnect power to all
trim motors quickly as the main trim
power ac circuit breaker is not available
to the pilot. To prevent inadvertent move-
ment the switch must first be pulled out be-
fore it can be moved from the ON to the
OFF position. In the ON position 200 volt
3 phase ac power from the right generator
bus is applied to the primary side of the
trim actuator transformer. Individual 28
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SECTION I
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A-12
ac circuit breakers for A and C phases of
the Manual Pitch, Auto Pitch, Roll and Yaw
trim circuits are located on the right con-
sole.
Roll Trim Switch
A three-position roll trim switch is located
just forward of the throttle quadrant. The
switch positions are indicated by L (left)
and R (right) arrows. The switch is spring-
loaded to the center off position. When the
switch is held in the R position, the roll
trim motor actuates to move the right ele-
vons up and the left elevons down. Actuation
of the switch to the L position moves the
right elevons down and left elevons up. 28-
volt ac power is furnished from the trim
power bus.
Rudder-Synchronization Switch
A three-position rudder synchronization
switch is installed just forward of the
throttle quadrant. The switch positions are
indicated by L (left), R (right) arrows. It is
springloaded to the center off position. In
the L and R positions the switch provides
electrical power to the right rudder trim
motor which moves the right rudder to
agree with the position of the left. Rudder
synchronization is obtained by superim-
posing the L and R needles on the yaw trim
gage. 28-volt ac power is furnished by the
trim power bus.
Roll, Pitch and Yaw Trim Indicators
Separate roll, pitch and yaw trim indicators
are located on the left side of the instrument
panel. The roll trim indicator uses a double
ended needle and displays the amount of roll
trim from 0 to 90 differential. The pitch
trim indicator displays the amount of pitch
trim from 5o nose down to 10o nose up, al-
1-54
though only 3-1/20 nose down and 6-1/20
nose up trim is available. The yaw trim
indicator displays the amount of yaw trim
from 10o left to 10o right for both rudders.
Rudder synchronization is obtained by super-
imposing the L and R needles on the yaw
trim gage. 26-volt ac power for the indi-
cators is furnished by the instrument trans-
former and the No. 1 inverter.
Surface Limiter Control Handle
A T-handle, labeled SURF LIMIT RELEASE,
is located on the left side of the annunciator
panel. When the handle is turned 90o
counterclockwise and released, the me-
chanical stops in the roll and yaw axis of
the cockpit control system are activated.
This action also opens an electrical switch
which de-energizes a solenoid operated
valve in each rudder servo package and
activates the servo package rudder stops.
Wien the handle is pulled out and rotated
90 clockwise, the mechanical stops in the
cockpit are released and the solenoid is
energized, releasing the servo package
stops. Power for the rudder limiting cir-
cuit is furnished by the essential dc bus.
Surface Limiter Indicator Light
When speed exceeds Mach 0.5, an indicator
light on the annunciator panel will illuminate
until the surface limiter handle is released.
If the speed is below Mach 0.5 and the sur-
face limiters are on, the indicator light
will illuminate until the surface limiter
handle is pulled out. Power for the lights
is furnished by the essential dc bus.
AUTOMATIC FLIGHT CONTROL SYSTEM
The automatic flight control system includes
stability augmentation, autopilot, and air
data systems, plus additional subsystems
furnishing attitude and navigational course
inputs for the autopilot. The air data sys-
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SECTION I
A-1Z
tern furnishes signals to the autopilot and
inertial navigation systems. The stability
augmentation system supplies signals to the
hydraulic servos that operate the control
surfaces. The inertial navigation system
supplies attitude and navigational course
inputs for the autopilot. Heading and atti-
tude reference signals for the autopilot are
also supplied by the Flight Reference Sys-
tem. The autopilot moves the aircraft hy-
draulic servos through the SAS. For fur-
ther information on the autopilot and in-
ertial navigation systems, refer to Section
IV.
STABILITY AUGMENTATION SYSTEM
The three axis stability augmentation sys-
tem is a combination of electronic and hy-
draulic equipment which augments the natu-
ral stability of the aircraft. It is designed
for optimum performance at the basic mis-
sion cruise speed and altitude, but also
provides improved stability for in-flight
refueling, landing and takeoff. The SAS is
part of the aircraft's basic control system
and is normally used for all flight condi-
tions.
Dual electronic channels are provided for
all axes and a third monitor channel is pro-
vided for both the pitch and yaw axis. Logic
circuits compare the functioning of each
pitch and yaw channel and automatically
delete a failed channel. The pilot is also
provided with a visual warning of a failed
channel.
In the roll axis, each channel controls the
elevons on only one side of the aircraft.
The pilot may select a single channel if de-
sired. Reliability is provided through dual
hydraulic and inverter supplies. Each
active channel in each axis is powered by
separate supplies so that the two halves of
each system are operated independently. A
separate gyro system is provided for each
channel in each axis. The design is such
that no single failure except overheating of
a complete gyro package can cause loss of
all channels in one axis. Even if this oc-
curred, it is unlikely that all of the gyros
in the package would fail simultaneously.
The SAS system compares the 3 electronic
systems and disengages a malfunctioning
A, B or M channel. Automatic gain in-
crease is applied to the remaining channels
so that control response remains the same.
A malfunctioning electronics channel is in-
dicated by illumination of the A or B and/or
M light.
STABILITY AUGMENTATION PITCH AXIS
The pitch axis SAS consists of two inde-
pendent active channels A and B and a
third monitor M channel. The two
independent active channels A and B provide
the desired control through two pairs of
tandem servos. There is one pair of
servos on each side of the aircraft. The
servos are in series with the autopilot and
the pilot's control movements. Damping
signals to the elevons do not move the
control stick. Each A and B channel
drives one servo on the left side of the
aircraft and one on the right side. A
channel uses A hydraulic system and B
channel uses the B hydraulic system. This
avoids loss of both channels in case of
failure of either the A or B hydraulic
systems. The sensors for the pitch axis
are rate gyros located in tank No. 3. The
gyros provide signals in proportion to the
rate of pitch attitude change of the air-
craft. Above 50,000 feet a "lagged" pitch
rate gain is programmed into the pitch
SAS electronic circuits. This pitch rate
signal changeover may be felt as an abrupt
pitch transient during a turn while climbing
or descending through the 50,000 foot level.
Refer to Section VII, Pitch Axis Character-
istics due to Lagged Pitch Rate Switching.
Phasing of the gyro signals is such that a-n
angular pitch motion produces elevon
movement to oppose and restrict attitude
change. The system will take corrective
action rapidly in the event of a gust
disturbance. Pilot inputs are also opposed;
however, the elevon motion produced by the
SAS is designed to aid the pilot in avoidi-g
overcontrol and improve the handling
qualities of the aircraft.
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A-12
SAS AND AUTO PILOT CONTROL PANEL
2
TRIM ROLL
TURN ON
TYPE "A" PANEL
1
ROLL CHANNEL DISENGAGE LIGHT
8
A/P ROLL TRIM SYNCRONIZATION
INDICATOR
2
SAS CHANNEL SWITCHES
9
A/P TURN CONTROL WHEEL
3
SAS RECYCLE INDICATOR LIGHTS
10
A/P PITCH TRIM SYNCRONIZATION
4
SAS LIGHT TEST SWITCH
INDICATOR
5
A/P HEADING HOLD SWITCH
11
A/P PITCH ENGAGE SWITCH
6
A/P AUTO NAV SWITCH
12
A/P PITCH CONTROL WHEEL
7
A/P ROLL ENGAGE SWITCH
13
A/P MACH/KEAS HOLD SWITCH
Figure 1-25
F200-25(c)
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A-12
SECTION II
The logic circuit is able to isolate a SAS
failure in either the electronics or the ser-
vos. When a malfunction is isolated, the
failed active channel will disengage and the
system continues in operation on a single
channel. Malfunctioning and disengageing
of channels is indicated by indicator lights.
The pitch axis can command a maximum
elevon surface travel of 2.5 up to 6.50 down.
Dual or single channel operation produces
the same corrective action of the elevon
surface. Power for A channel is from the
A phase of No. 1 inverter bus. Power for
B channel is from the A phase of No. 2 in-
verter. Monitor channel power is from the
B phase of the No. 3 inverter. Each power
source is protected by individual circuit
breakers in the cockpit.
STABILITY AUGMENTATION YAW AXIS
The yaw axis of the SAS is very similar to
the pitch axis, using two independent A and
B channels and a monitor channel. There
is one pair of hydraulic servos for each
rudder, each pair mounted in a whiffletree
arrangement. Damping signals to the rud-
der do not move the rudder pedals. Each
A and B channel drives one servo on each
side of the aircraft. The A hydraulic sys-
tem is connected to A channel and the B hy-
draulic system to B channel. The rate
gyro sensors for the three channels are
identical to the pitch rate gyros, except for
the physical orientation to sense yawing
motions. A "Hi Pass" filter circuit is in-
stalled to allow passage of normal short
term damping signals, but will stop the
signals when a deliberate turn is made. A
lateral accelerometer sensor is also used
in each channel of the yaw axis. This sen-
sor provides an input for high gain lateral
acceleration function to provide a more
rapid rudder response during engine failure
conditions. However, this function will op-
pose the pilot when he is purposely trying
to sideslip.
The logic circuit is identical to the pitch
axis and functions in the same manner. The
yaw axis can product?, a maximum rudder
travel of 8 left to 8 right. Corrective
surface motion is the same regardless of
one or two channel operation due to auto-
matic gain doubling if only one channel is
operative. Power for A channel is from the
B phase of the No. 1 inverter, B channel
from the B phase of the No. 2 inverter and
the monitor channel from the B phase of the
No. 3 inverter. The circuitry from each
power source is protected by individual
circuit breakers.
STABILITY AUGMENTATION ROLL AXIS
Roll a.xis reliability requirements are not
as severe as pitch and yaw; therefore, less
complicated circuitry and components are
used. The roll axis has two independent
channels, each operating the elevons on one
side of the aircraft. A channel positions
the left elevon surfaces and operates from
the A hydraulic system. B channel positions
the right elevon surfaces and operates from
the B hydraulic system. There is no moni-
tor channel. Each channel can be operated
individually. Although the system gain is
the same as two channel operation, roll
control is not symmetrical. Coupling into
the yaw and pitch axes is possible, but the
systems operating in those axes minimize
undesirable aircraft motion. Maximum
elevon travel in the roll axis is 2o up to 2o
down (each side), for a total of e differen-
tial with both systems operating. Power
for A channel is from C phase of the No. 1
inverter and B channel from C phase of the
No. 2 inverter.
STABILITY AUGMENTATION SYSTEM (SAS)
CONTROL PANEL
The SAS control panel on the right console
contains six channel switches, for A and B
channels of the pitch, roll and yaw axis.
The panel also contains a press-to-test
switch and six indicator lights for the A, B
and MON channels in the pitch and yaw axis.
Three guarded switches for the backup
pitch damper, pitch logic override and yaw
logic override are located on the right side
of the annunciator panel. A roll channel
disengage light is located between the roll
channel switches. Individual circuit
breakers are located on both right and left
consoles.
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SECTION I
A-12
Channel Switches
There are six toggle switches located on
the SAS control panel. There is one pair
for each axis; pitch, roll and yaw. The
forward switch of each pair is A channel
and the rear switch is B channel. The
switches have two positions; ON (forward)
and OFF (aft). When electrical power is
on the aircraft and the channel switches are
OFF, the SAS electronics are powered, but
the channel servos are not engaged into the
control system. Moving the switches to the
ON position engages the SAS servos pro-
viding the recycle light is extinguished. If
the recycle light is not extinguished it must
be depressed for engagement.
Recycle indicator Lights
Six indicator lights are located on the SAS
control panel adjacent to the pitch and yaw
channel engage switches. One light is pro-
vided for each A, B and MON channel in
the pitch and yaw axes. When the channel
switch is on and the light is not illuminated,
the channel is functioning properly. If the
light is illuminated, it indicates that the
channel has disengaged and the light may
be pressed to recycle the channel. In the
event the failure was momentary, this will
reengage the channel. If the light reillum-
inates, the channel is malfunctioning, but
it is not necessary to turn the channel en-
gage switch off because the light indicates
that automatic disengagement has occurred.
NOTE
The lighted recycle indicator light
should be pressed down firmly and
released. A control surface tran-
sient will occur if a hardover servo
exists in that channel. Refer to
Section III.
The six recycle lights will be illuminated
when electrical power is applied to the air-
craft. The channel switches must be on
and the recycle lights must be pressed to
engage the channel electronics to the servos.
When engaged and operating, the channel
lights will be out.
Roll Channel Disengage Light
A single roll channel disengage light is lo-
cated between the two roll channel switches.
When illuminated it indicates that both roll
channels have disengaged. Disengagement
results when the roll servo channel outputs
differ by more than an amount equivalent
to 0.6 surface deflection. When operating
on a single roll channel the light will, not be
illuminated and disengagement in the event
of a failure is not provided. The switch
must be ON for the active channel and OFF
for the malfunctioning channel.
Light Test Switch
A pushbutton light test switch is located in
the center of the SAS control panel. Press-
ing the button illuminates all SAS lights for
test.
Backup Pitch Damper Switch
A guarded BUPD switch is located on the
right side of the annunciator panel. It is
guarded in the OFF position. It is used in
case the SAS pitch channels are unusable
due to electronic malfunctions or over-
heating of the pitch gyro package. In the
ON position the backup gyro, located in the
electronic compartment, supplies pitch
rate signals through an independent elec-
tronic channel to either the A or B servos.
The pitch logic override switch must be
used to select their A or B servo operation.
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NOTE
The primary purpose of the BUPD
Is to provide an emergency system
for pitch stability augmentation dur-
ing refueling and landing approach.
The system is optimized for use at
light weight, aft center of gravity
and subsonic speeds. It is not in-
tended as an emergency backup sys-
tem during cruise. Refer to Section
III, Emergency Procedures.
SAS Pitch Logic Override Switch
A guarded, three-position SAS pitch logic
switch is located on the right side of the
annunciator panel. It is OFF in the center
guarded position and the logic circuit is op-
erative. Placing the switch in the A (up)
position deletes the logic circuit and selects
A channel operation. In the B (down) posi-
tion, the logic circuit is deleted and B chan-
nel is selected. The switch must be placed
in either the A or B position when the BUPD
is used. This selects operation of either
the A or B servos.
NOTE
The override switch is only used
as an emergency procedure. Refer
to Section ILI.
SAS Yaw Logic Override Switch
A guarded, three-position SAS yaw logic
switch is located on the right side of the
annunciator panel below the pitch logic
override switch. It is guarded in the OFF
position. The A (up) position deletes the
logic circuit and selects A channel operation.
The B (down) position deletes the logic cir-
cuit and selects B channel operation.
NOTE
The override switch is only used
as an emergency procedure. Refer
to Section III.
P1101-STATIC SYSTEMS
The pitot-static system supplies the total
and static pressure necessary to operate
the basic flight instruments and air data
system components. The pressures are
sensed by an electrically heated probe
mounted on the nose of the aircraft. The
probe and forward nose also serves as an
antenna for the high frequency radio. The
pitot orifice of the probe is divided inside
the head to provide two separate pressure
sources. It also has two circumferential
sets of four static pressure ports each.
One pitot and the aft set of static ports sup-
ply pressure signals to the air data com-
puter and inlet air control systems. The
other set of pickups supply pitot and static
pressure directly to the speed sensors on
the ejection seats, the altimeter, the rate
of climb and airspeed indicators. An offset
head on the left side of the probe provides
yaw and pitch pressure signals to the inlet
spike controls and to the stall warning light
sensor.
The heating elements of the probe are con-
trolled by the pitot heat switch located on
the left side of the annunciator panel.
Power is furnished by the left ac generator
bus.
An alternate heated pitot static source is
available from the Flight Recorder System.
Refer to Flight Recorder, Section IV.
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A-12
P ITOT STATIC SYSTEM
RATE OF
CLIMB
SHIP SYSTEM
SELECTOR VALVE
ANT TUNING
HF ANTENNA
DUAL PITOT COIL
STATIC TUBE
AND HF
ANTENNA
TO HEATED AREAS
HEATED
AREAr.
FLIGHT
RECORDER
SYSTEM
INV
CHINE
STATIC
PORTS
CHINE
P ITOT
ON
OFF
CHINE
STATIC
PORT
HEATED
AREA
FLIGHT RECORDER
(IN CHINE)
SPEED SENSOR
CHINE
P ITOT
AIR SPEED
INDICATOR
ALTIMETER
ESS DCC
BUS
TI-ON
P ITOT HEAT
LGENii OFF
BUS II
41 PITCH PROBE
STATICS
2 YAW PROBE STATICS
INDICATED
AIRSPEED
ALPHA-BETA
STATIC PROBE
SHIP SYSTEM
SELECTOR VALVE
,I I
ALTIMETER RATE OF TRIPLE
CLIMB DISPLAY -7 SEAT
INDICATOR
Figure 1-26
AIR DATA
COMPARTMENT
Q TRANSDUCER
EJECTION SEAT
SPEED SENSOR
(ANGLE OF ATTACK )
TRANSMITTER
TO INLET
P2
S2
AUTO PILOT
INS
ANGLE
TRANSDUCER
INLET
CONTROL
COMPUTER
AIR
DATA
COMPUTER
SYSTEM
F200 -74(b)
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Pitot-Heat Switch and indicator Light
A two-position toggle switch is located on
the left side of the annunciator panel. In
the ON (up) position ac power is applied to
the heating elements of the pitot-static
probe. The probe is grounded to the air-
frame in a manner which permits the HF
radio to be operated while pitot heat is on.
In the OFF (down) position ac power is dis-
connected from the probe heating elements.
The circuitry also incorporates an altitude
switch and a PITOT HEAT light located on
the annunciator panel. The pitot heat light
will be on when the switch is in the ON posi-
tion and the altitude is above 65,000 feet,
and also when the switch is in the OFF posi-
tion and the altitude is below 50,000 feet.
The light will be OFF if when below 50,000
feet and pitot heat is ON, and when above
65,000 feet with the switch in the OFF posi-
tion.
AIR DATA COMPUTER
The air data computer performs two func-
tions, computation and display. The total
and static pressures from the pitot-static
probe are converted to electrical signals
required for the pilot's triple display indi-
cator, compressor inlet pressure indicator
system, the automatic flight control and in-
ertial navigation systems. The ports which
supply pressure to the air data computer
are separate from those that furnish pres-
sure to the basic flight instruments. There-
fore, failure of the air data computer pres-
sure source will not leave the pilot without
the altitude, vertical velocity or airspeed
information. The air data computer con-
verts pitot-static pressures into propor-
tional rotary shaft positions which are
equivalent to pressure altitude and dynamic
pressure. These shaft positions are com-
bined in a mechanical analog computer
made up of cams, gears and differentials
to drive the output functions. Outputs of
the air data computer and the using equip-
ment are listed below:
OUTPUT SIGNALS
USING
EQUIPMENT
Pressure
Altitude
Equivalent
Air speed
Mach
._
Triple Display
Indicator
KEAS + MACH
Compressor Inlet
Pressure Indicator
KEAS
Mach
Mach Rate
Altitude
Dynamic Pressure
Autopilot
Pres sure
Altitude
Inertial Navigator
Computer
Power for the air data computer is furnished
either'by the No. 1 or No. 3 inverter de-
pending on the position of the autopilot se-
lector switch.
Triple Display Indicators
A triple display indicator is located on the
instrument panel to provide digital displays
of airspeed, altitude, and Mach number as
computed by the Air Data Computer. The
altitude indication range of the TDI is from
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.LA.- 1 L.
FLIGHT INSTRUMENTS
Figure 1-27
F 200-70(d)
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SECTION I
A-12
0 to 99,950 feet. At 100,000 feet the first
digit is dropped, indicating 09,950 feet at
109,950 feet pressure altitude, the maxi-
mum limit of the ADC signal to the instru-
ment. The Mach number display capability
range of each instrument is 0 to 3.99; how-
ever, the minimum indication at static con-
ditions normally ranges from 0.1 to 0.2
Mach number and the maximum indication
would be Mach 3.5 for a normally function-
ing instrument. This range corresponds to
the range of signals which the ADC is cap-
able of providing. The TDI displays air-
speed in knots equivalent airspeed (KEAS)
within an instrument capability from 0 to
599 KEAS; however, the minimum indi-
cation is normally 75 to 110 KEAS to cor-
respond with the minimum ADC signal pro-
vided. The maximum signal provided by
the ADC results in an airspeed indication
which decreases from 599 KEAS at sea
level to 523 KEAS at 66,800 feet and Mach
3.5, and then decreases further at high
altitudes to show the KEAS corresponding
to Mach 3.5 and the existing pressure alti-
tude. An off flag appears on the face of the
instrument if the ADC loses power. Power
for the instrument is from the No. 1 or No.
3 inverter.
NOTE
Indications of the triple display
indicator and the basic pitot-
static flight instruments should
be periodically cross checked to
confirm proper system operation.
Refer to figure A1-2, Appendix I.
The triple display indicator is
primarily used for aircraft con-
trol above FL 180 and to main-
tain proper airspeed control dur-
ing climbs to FL 180. Basic
pitot-static operated flight instru-
ments shall be used in the landing
pattern, during takeoff until pro-
per climb schedule is established
on the TDI, and during all simu-
lated or actual instrument flight
below FL 180.
. If KEAS indications oscillate be-
tween two values on the high end
of the range, it is an indication
that the indicator limit is being
approached.
INSTRUMENTS
For information regarding instruments that
are an integral part of a particular system,
refer to applicable paragraphs in this
section and Section IV.
Airspeed-Mach Meter
A combination airspeed and Mach meter
operating directly from pitot-static pres-
sure is located in the flight instrument
group. This is a special instrument with
airspeed and Mach number ranges com-
patible with aircraft performance capa-
bilities. Mach number and indicated air-
speed are read simultaneously on the win-
dow and outer index respectively. A limit
airspeed needle (white barred) shows the
airspeed limit of the aircraft. The actual
airspeed limit is in equivalent airspeed;
however, the needle varies with altitude to
read the indicated airspeed that converts
to equivalent airspeed.
Altimeter
A sensitive pressure altimeter is located
on the instrument panel. In addition to the
1000 foot and 100 foot pointers, it also has
a 10,000 foot pointer. This pointer extends
to the edge of the dial with a triangular
marker at its extremity. The center disc
has a cutout through which yellow and black
warning stripes appear at altitudes below'
16,000 feet. The barometric pressure
scale is in a cutout at the right side and is
set by a knob located at the lower left side
of the instrument.
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A-12
Attitude Indicator (MM-3)
The attitude indicator is located in the basic
flight instrument group on the instrument
panel. It provides constant visual indication
of nose and wing position in relation to the
earth's surface. Attitude indications are
presented by a spherical graduated dial, a
W reference line, a bank pointer, and a
marked outer ring. A horizontal line is
formed on the spherical dial by the meeting
of a gray, upper climb section and a black
lower dive section. The instrument shows
� attitude in climb or dive up to 85 degrees.
NOTE
At approximately 85 degrees climb
or dive, the attitude indicator will
flip but will not tumble. The 180
degree flip in roll will be very rapid
and the instrument will accurately
indicate pitch and roll attitudes im-
mediately thereafter. Some small
inaccuracies may develop after a
series of maneuvers beyond the 85
degree climb or dive attitude.
These inaccuracies will automat-
ically be cancelled out at the erec-
tion rate of .80 to 1.80 per minute.
The W reference line remains fixed with
the marked outer ring and represents the
aircraft in miniature. The spherical dial
moves up or down, or the whole spherical
dial assembly rotates within the instrument
case behind the W reference line and outer
ring to indicate aircraft attitudes. As the
dial assembly rotates, the bank pointer
moves with it to indicate degrees of bank
on the outer ring. The outer ring indicates
0 o
- 90 0 bank. The spherical dial and
pointer are capable of rotating a full 360
degrees of roll with the aircraft. Pitch
attitude of the aircraft is indicated by the
position of the horizon line in relation to
the miniature aircraft. A pitch adjustment
knob on the lower right side is used to
change the position of the spherical dial as
desired. During initial gyro erection, and
when power is off or is insufficient to keep
the gyro stabilized, a warning OFF flag
appears at the bottom of the indicator. The
autopilot'and attitude reference selector
switch is used to select pitch and roll atti-
tude signals from either the INS or FRS
stable platforms.
CAUTION
To avoid gross pitch attitude
errors the pitch adjustment
knob of the attitude indicator
should be adjusted to align the
index marks before the auto-
pilot and attitude reference
selector switch is changed in
flight.
NOTE
To determine a possible malfunction
of the attitude indicator, an occa-
sional accuracy check should be made
by comparing it against the standby
attitude indicator and other basic
flight instruments.
The system is powered by the No. 1 and
No. 3 inverter depending on the position of
the autopilot selector switch.
Standby Attitude Indicator
The standby attitude indicator located on the
lower left side of the instrument panel pro-
vides the pilot with an independent attitude
reference. It contains a sphere inscribed
with an artificial horizon and calibrated in
degrees of aircraft angle of pitch. The
globe is detailed to represent the sky and
earth areas, and is capable of rotating to
indicate pitch angles of + 82 degrees and
roll angles of 360 degrees. The bank angle
scale is marked on the lower periphery. A
pitch reference adjustment knob is provided
on the lower right corner of the instrument
for positioning the reference bar as desired.
A fast erect pushbutton is provided on a
small panel above the throttles.
1-64
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SECTION I
A-12
ANNUNCIATOR PANELS
NO. 1 OXY LOW
NO. 2 OXY LOW
Q-BAY HEAT HIGH
FUEL QTY LOW
N QTY LOW
TANK PRESSURE LOW
ANTI-SKID OUT
SURFACE LIMITER
SAS CHANNEL OUT
A HYD LOW
B HYD LOW
Do not hold fast erect button for
more than 45 seconds to prevent
overheating of fast erect motor.
STALL WARNING
MANUAL INLET
L OIL TEMP
L FUEL PRESS LOW
A HYD PRESS LOW
L GENERATOR OUT
L XFMR-RECT OUT
NO. 1 INVERTER OUT
NO. 2 INVERTER OUT
NO. 3 INVERTER OUT
PITOT HEAT
Figure 1-28
This instrument has its own self-contained
gyro and is not dependent on another re-
ference source. The OFF flag will be
visible whenever power to the indicator is
interrupted. Power is provided by the C
phase of the No. 2 inverter.
Vertical Velocity Indicator
A vertical velocity indicator is located on
the instrument panel and shows the rate of
change of altitude in feet per minute.
Changes in pressure due to changes in alti-
INS FIX REJECT
R OIL TEMP
R FUEL PRESS LOW
B HYD PRESS LOW
R GENERATOR OUT
R XFMR-RECT OUT
EMER BAT ON
R HYD LOW
L HYD LOW
Q-BAY EQUIP OUT
rz00-69(.)
tude aresensed by the static system and
transmitted to the indicator. Depending on
the instrument installed the instrument is
capable of indicating vertical speed of 0 to
+ 12,000 feet per minute or 0 to 6,000 feet
per minute. An over-pressure diaphragm
and valve prevent excessive rates of climb
or descent from damaging the instrument.
Turn and Slip Indicator
A turn and slip indicator is installed on the
instrument panel. The indicator is cali-
brated for either a two or four minute turn.
The indicator is powered by the essential
dc bus. An additional larger slip indicator
is mounted on the upper center instrument
panel beneath the CIP indicator.
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A - 1 2
LANDING GEAR SYSTEM
MANUAL LANDING GEAR
RELEASE HANDLE
CROSSOVER VALVE
S'RESSURD
CROSSOVER VALVE
(RETURN)
NOSE LANDING GEAR =I D
ACTUATING CYLINDER
MAIN LANDING GEAR
ACTUATING CYLINDER
PRESSURE �
SWITCH
LANDING GEAR
LEVER
11XCI F
DOOR
SELECTOR
VALVE
0 C
01:11:1
MAIN
LANDING GEAR
ACTUATING
CYLINDER
CCM
Er
MIMI IL
DOOR ACTUAL
CYLINDER (4 PLACES)
0
DOOR LATCH
CYLINDER (4 PLACES)
CABLE
ELECTRICAL CONNECTION
CHECK VALVE
RESTRICTOR VALVE (SMALL ARROW INDICATES
DIRECTION OF RESTRICTED FLOW)
FLOW REGULATOR
RESTRICTOR VALVE ( RESTRICTED FLOW
IN BOTH DIRECTIONS)
Figure 1-29
111Mi111112
IXOXECII
117X1EX1111
121111i1=1
111221122EZNI
R SYSTEM PRESSURE
R SYSTEM RETURN
L SYSTEM PRESSURE
L SYSTEM RETURN
MLG DOORS CLOSED
MLG DOORS OPEN
LANDING GEAR DOWN
LANDING GEAR UP
P}UU2 -26
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A-12
SECTION I
Clocks
An elapsed time clock is located on the in-
strument panel. It contains an elapsed time
mechanism that is started and stopped by
pushing the winding knob. An A13A clock
is also installed in the panel. The second
hand is started and stopped by the small
button on the upper right corner. The third
hand serves as a 60 minute recorder.
EMERGENCY EQUIPMENT
MASTER WARNING SYSTEM
An annunciator panel is mounted on the
lower instrument panel. The panel contains
individual warning lights that indicate mal-
functions or failures of equipment and sys-
tems. Illumination of any individual light
also illuminates an amber master caution
light on the upper portion of the instrument
panel. Once illuminated, the master
caution light can be extinguished (reset) by
depressing the light. The individual an-
nunciator panel light will remain illuminated.
Another malfunction again illuminates the
master caution light. Warning lights are
automatically dimmed when the instrument
panel lights are on. The master warning
system does not include the fire warning
and landing gear unsafe lights. Power is
furnished by the essential dc bus.
NACELLE FIRE WARNING SYSTEM
A fire warning system detects and indicates
the presence of a fire in the engine nacelles.
A hot spot anywhere along the length of the
detection circuit will illuminate the light of
that particular nacelle. The lights are lo-
cated on the pilot's instrument panel above
the respective column of instruments per-
taining to each engine.
Nacelle Fire Warning Lights
Left and right nacelle FIRE warning lights
located on the top right side of the instru-
ment panel, illuminate when nacelle tem-
perature at the turbine or at the after-
burner exceeds 1050�F + 50�. Flip down
glare shields are provided for night flying.
Power for the circuit is furnished by the
No. 1 inverter.
STALL WARNING LIGHT
A STALL WARNING light is located on the
annunciator panel which is illuminated when
the aircraft angle of attack reaches + 14 de-
grees and the nose landing gear scissor
switch is open. Pressure differences be-
tween the a 1 and 2 inlets on the pitch and
yaw probe are sensed by a pitch trans-
mitter unit to actu.ate this light. A steady
tone warning signal is also produced in the
pilot's earphone. Power for the stall warn-
ing light is furnished by the essential dc bus.
LANDING GEAR SYSTEM
The tricycle landing gear and the main
wheel well inboard doors are electrically
controlled and hydraulically actuated. The
main gea.r outboard doors and the nose gear
doors are linked directly to the respective
gear struts. Each three wheeled main gear
retracts inboard into the fuselage and the
dual wheel nose gear retracts forward into
the fuselage. The main gear is locked up
by the inboard doors and the nose gear by
an uplock which engages the strut. There
is no hydraulic pressure on the gear when
it is up and locked. Down locks inside the
actuating cylinders hold the gear in place
in the extended position. Hydraulic pres-
sure is also on the gear in the extended
position when L system pressure is avail-
able. The landing gear cylinders and doors
are actuated in the proper order by two se-
quencing valves. Normal gear operation is
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SECTION I
A-12
powered by the L hydraulic pump on the left
engine. Should pressure drop to 2000-2200
psi during retraction, the power source
automatically becomes the R hydraulic pump.
R hydraulic pressure will not extend the
gear in the event of an L system failure and
the manual landing gear release must be
used. Normal gear extension time is 12-
16 seconds.
LANDING GEAR LEVER
A wheel shaped landing gear lever is in-
stalled on the lower left side of the instru-
ment panel just forward of the throttle quad-
rant. The lever has two positions; UP and
DOWN. A locking mechanism is provided
to prevent the gear lever from being inad-
vertently placed in the DOWN position. A
button which extends upward from the top of
the lever must be pressed forward in order
to release the lock mechanism. An over-
ride button is installed just above the gear
lever and may be used to override the
ground safety switch should it become nec-
essary to raise the gear when the weight of
the aircraft is on the landing gear. Once
energized, the gear lever must be recycled
to the DOWN position in order to bring the
ground safety switch back into the circuit.
A red light installed in the transparent
wheel illuminates during cycling, or when
the gear is in an unsafe condition. Power
for the circuit is furnished by the essential
dc bus.
Manual Landing Gear Release Handle
A manual landing gear release handle la-
beled GEAR RELEASE is installed on the
annunciator panel. If the L hydraulic sys-
tem has failed but R hydraulic pressure is
available, the landing gear lever must be
in the DOWN position or the landing gear
CONT circuit breaker must be pulled out
before pulling the GEAR RELEASE handle.
Otherwise, the R system will retract the
gear. The gear extends by gravity force.
Approximately 9 inches of pull on the handle
is required since the uplocks are released
at different positions along the cable length.
The nose gear uplock is released first
followed by the right gear then the left.
Gear retraction is possible after being
lowered by the manual gear release handle,
provided L or R hydraulic system pressure
is available.
Gear and Warning Light Test Button
A gear and warning light pushbutton switch
is located on the left forward panel. When
depressed it illuminates the landing gear
lever red light, all annunciator panel lights,
the right and left nacelle fire warning lights,
and actuates the gear warning tone in the
headset. It is also used to test the three
green landing gear position lights when air-
borne.
Landing Gear Position Lights
Three green lights, located on the left side
of the instrument panel indicate the down
and locked condition of the landing gear.
The location of eacklight corresponds to
the respective wheel it monitors. Power
is from the essential dc bus.
Landing Gear Warning Light and Audible Warning
The red landing gear warning light in the
landing gear lever handle when illuminated
indicates:
1. Gear is cycling.
2. Gear system is not locked in the UP or
DOWN position.
3. Gear is UP and throttle settings are be-
low MILITARY and altitude is below
10,000 feet.
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SECTION I
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A pulsed tone warning signal is also pro-
duced in the pilot's earphones when the
throttles are retarded below approximately
1/3 the distance between the IDLE and MIL
throttle settings, the landing gear is not in
the down and locked position and aircraft
altitude is below 10,000 feet + 500 feet.
Power for the light and pulsed tone warning
Is furnished by the essential dc bus.
Landing Gear Warning Cutout Button
The audio gear warning signal can be elim-
inated by pressing the OR SIG REL push-
button switch on the instrument panel. The
circuit is reactivated when the throttles are
advanced above the minimum cruise setting.
Power is supplied by the essential dc bus.
Land Gear Ground Safety Pins
Removable ground safety pins are installed
in the landing gear assemblies to prevent
inadvertent retraction of the gear while the
aircraft is on the ground. Warning stream-
ers direct attention to their removal before
flight. An additional set of ground safety
pins is provided in a container behind the
seat.
LANDING GEAR STRUT DAMPER
A landing gear strut damper system is in-
stalled to control gear "walking" during
brake operation. The system is sensitive
to less than one g change in fore and aft
acceleration. The damping is controlled
through a g monitoring valve which auto-
matically increases or decreases the brake
pressure as required. Hydraulic pressure
for the damper system is provided by the L
system.
NOSEVVHEEL STEERING SYSTEM
The nosewheel steering system provides
power steering for directional control when
the aircraft weight is on any one gear. The
nosewheel is steerable 30 degrees either
side of center. Steering is accomplished
by a hydraulic steer-damper unit controlled
through a cable system by the rudder pedal
L hydraulic system pressure from the nose
landing gear down line is routed to the steez
ing system through a shutoff valve, which
is controlled by the nosewheel steering
(NWS) button on the control stick grip.
Steering is engaged by depressing the NWS
button and matching pedal position with
nosewheel angle. A holding relay circuit
allows the NWS button to be released after
it is once depressed and steering will stay
engaged. It is disengaged when the NWS
button is again pressed and released.
Steering is engaged at any time the NWS
button is held depressed. Nosewheel steer-
ing radius is approximately 75 feet. A me-
chanically operated centering cam auto-
matically centers the nosewheel when it re-
tracts. Power for the system is furnished
by the essential dc bus.
NOTE
Nosewheel steering is operable
only if essential dc bus power
is available and weight of the
aircraft is on any one gear.
If the L system pressure
should drop below 2000-2200 psi
alternate nosewheel steering
may be obtained by placing
the brake switch to ALT STEER
& BRAKE position.
WARNING 1
The landing gear side load
strength is critical. Side loads
during takeoff, landing and
ground operation must be kept
to a minimum.
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SECTION I A-12
BRAKE SYSTEM
LP
N2
FILLER
(VALVE)
LR
OVERBOARD
DRAIN
BRAKE
RESERVOIRS
NORM
BRAKE
PEDAL
I
��
ALT
MASTER
CYLINDERS
NORM
N2 PRESSURE
NORMAL
BRAKE
RELAY
VALVE
����1==� -----
ANTI - SKID
SHUTOFF
VALVE
RELIEF VALVE
STRUT DAMPER
Lp
=1=1�1 �1�1�11MM
LR
ALT
BRAKE
RELAY
VALVE
BRAKE BRAKE
RESTR I CTOR RESTR I CTOR
1�1�11=ill
COMM
BRAKE
PEDAL
ALT
�=1=I���
ANTI - SKID III
SHUTOFF
VALVE
NITROGEN
CYLINDER
ALTERNATE
BRAKE SHUTOFF
VALVE
RELIEF VALVE
BRAKE SHUTTLE
VALVES
ANTI - SKID till
GENERATORS
- GASEOUS NITROGEN
Imium:3�0 L SYSTEM PRESSURE
wAwlia� L SYSTEM RETURN
min MASTER CYLINDER SUPPLY
Figure 1-30
ANTI-
SK I D
NORMAL �
ALT STEER
AND BRAKE
a
roman BRAKE RELAY VALVE PRESSURE
lalxxxxx R SYSTEM PRESSURE (VALVE ENERGIZED)
- R SYSTEM RETURN
ELECTRICAL CONNECTION
F200-27(b)
1-70
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SECTION I
A-12
WHEEL BRAKE SYSTEM
The aircraft is equipped with artificial feel
hydraulically operated power brakes. De-
pressing the rudder pedals actuates the
four rotor brakes on each of the six main
wheels. The L hydraulic system furnishes
brake pressure with optional antiskid op-
eration. The hydraulic pressure to the
brakes is approximately 1200 psi. Should
the L hydraulic system fail, alternate brakes
are available. The alternate brakes operate
from an independent system using R hydrau-
lic pressure with no antiskid provision.
A small accumulator is incorporated in the
normal brake system which should provide
up to five brake applications after L and R
hydraulic failure provided accumulator
pressure has not been dumped by selecting
alternate brakes. Certain types of hydrau.-
lic system failures such as a broken line
could deplete the system fluid. Normal or
antiskid brakes are usable if left hydraulic
pressure is steady and above 2200 psi. Al-
ternate brakes are used if left hydraulic
system pressure is below this pressure.
Brake Switch
A three-position brake switch is located on
the left side of the instrument panel. In the
NORM (center) position, brake pressure
from the L hydraulic system is available,
but the antiskid system is not operative. In
the ANTISKID (up) position, the antiskid
system is operative. In the ALT STEER &
BRAKE (down) position, the brakes, NWS
and air refueling system are powered by
the R hydraulic system if left system pres-
sure is below 1250 psi. Power for the cir-
cuit is furnished by the essential dc bus.
WARNING I
Do not switch to alternate brakes
unless normal left hydraulic pres-
sure is unavailable or normal
brakes are inoperative. Pressure
may be trapped in the brakes after
the pedals are released, causing
grabbing or locking.
Anti-skid Out Indicator Light
Illumination of the ANTI-SKID OUT indi-
cator light on the annunciator panel indi-
cates that the anti-skid system is inoper-
ative. When the aircraft is on the ground,
the light will be illuminated when the brake
switch is in the NORM or ALT STEER &
BRAKE position. The light will be off when
the switch is in the ANTI-SKID position, if
the anti-skid control box and wheel gen-
erators are operative. If the fail safe cir-
cuit within the anti-skid control box is
tripped and power from the essential dc
bus is on the system, the light will illum-
inate. The light is off at all times when the
weight of the aircraft is not on the gear.
DRAG CHUTE SYSTEM
The drag chute system is provided to re-
duce landing roll and aborted takeoff roll
out distance. The 45-foot ribbon type para-
chute is packed in a deployment bag and
stowed in the upper aft end of the fuselage.
It rides free in the compartment and is
locked onto the airplane at the initial stage
of its deployment action. The neck of the
drag chute link incorporates a breakaway
section to protect against aircraft structural
damage if the chute is deployed at too fast
a speed. The chute deployment is actuated
electrically and power is furnished by the
essential dc bus.
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1-71
SECTION I
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A-12
COCKPIT PRESSURIZATION SCHEDULE
�
[24
0
TDI PRESSURE ALTITUDE
1-72
90
80
70
60
5
4
3
1
- I
tMAL SWITCH SETTING
10,000
FT. SWITCH
SETTING
SAFETY
VALVE
SETTING
0
1
,
10, 000 FT.
_Ai AdO i0ellk
NORMAL
1
7
)
oloosti.
1
0
5
10
15
20
25
COCKPIT PRESSURE ALTITUDE - 1000 FT.
Figure 1-31
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30
F200-97
Approved for Release: 2017/07/25 C00821248
SECTION I
A-12
Drag Chute Handle
The drag chute deploy and jettison handle is
located on the left edge of the instrument
glare shield. When pulled the handle acti-
vates micro switches which deploys the
drag chute. When turned 90 degrees counter-
clockwise and pushed in, the drag chute is
jettisoned. Power for the circuit is furnished
by the essential dc bus.
AIR CONDITIONING AND PRESSURIZATION
SYSTEM
Similar left and right hand air conditioning
and pressurization systems utilize high
pressure ninth stage compressor air from
each engine to pressurize and cool the cock-
pit and equipment compartments. System
shutoff valves allow compressor air to flow
when the engines are running and the system
switches are ON. Cooling is accomplished
by ducting the bleed air through a ram air
heat exchanger, primary and secondary
fuel/air heat exchangers, and through an
air cycle refrigerator. Temperature of the
air supplied by each system is modulated by
temperature control bypass valves located
upstream from the air cycle refrigerators.
The bypass valves are positioned by control
switches located in the cockpit.
A water separator is installed in each air
conditioning system downstream of the air-
cycle refrigeration units. Below an altitude
of approximately 36,000 feet a pressure
switch in the automatic temperature control
circuit limits the minimum outlet tempera-
ture of the air from the air-cycle refriger-
ation to 355oF to prevent freezing of water
in the separator. Using the manual tem-
perature controls will allow lower temper-
ature air to come from the refrigerator but
icing of the water separator may occur if
humidity is high. Above 36,000 feet the
altitude pressure switch opens the water
separator bypass valve and air does not
flow through the separator.
The left engine normally furnishes air for
the cockpit, nose compartment, ventilated
flying suit, inverters and INS platform.
The right engine normally furnishes air to
the E-bay where it mixes with cockpit dis-
charge air for ventilation of the E-bay,
Q-bay, and the aft equipment compartments.
A fixed orifice restriction and a duct divid-
ing into two outlets provide for a portion of
the right system air to flow to the upper
part of the cockpit. A crossover system
is provided to supply right engine system
air to the cockpit and equipment normally
supplied by the left engine system. The
operation of the crossover system will not
depressurize the Q-bay since the cockpit
air exhausts into the Q-bay; however, a
rise in temperature will occur in the Q-bay.
High pressure canopy and hatch seal air
and windshield defog air is furnished from
both right and left engine systems by ducts
connected downstream from the primary
fuel/air heat exchangers.
COCKPIT COOLING AND PRESSURIZATION
When the aircraft is at high altitude, the
pressurization systems maintain a constant
altitude of approximately 26,000 feet in the
cockpit and nose and 28,000 feet in the
Q-bay.
Cabin Pressure Schedule Switch
The cockpit pressure schedule switch is a
two position toggle switch labeled CABIN
PRESS located on the lower center of the
instrument panel. In the NORMAL (down)
position, the cockpit and Q-bay pressuri-
zation systems provide the normal pressure
schedule and will maintain constant altitudes
of 26,000 and 28,000 feet when the aircraft
is above 32,000 feet. In the 10,000 feet (up)
position, the cockpit pressure is regulated
to a 5 psi maximum differential and will
maintain a 10,000 foot cockpit altitude up to
1-73
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SECTION A-12
AIR CONDITIONING AND PRESSURIZATION SYSTEM
4th INVERTER
(STANDBY)
RH FWD CHEEK r- CROSSOVER HI LIMIT
h I CHECK SHUTOFF
SENSOR
VALVE VALVE
155�
TO E-BAY HE� =cm
DUCT
TEMPERATURE
SENSOR
TO INS COOLING
PLATFORM
TO INS PLATFORM
COOLING SYSTEM
XOVER
CHECK
VALVE-
TO COCKPIT
AND NOSE ,
J I
I
INV
COOLING I
9th STAGE
COMPRESSOR AIR
BLEED PORTS
HI LIMIT
SENSOR
155�
DUCT
TEMPERATURE
SENSOR
ALTITUDE
PRESSURE
SWITCH
WATER
SEPARATOR
SAFETY ZONE
DRAIN DRAINS
RH SYSTEM
SAME TO THIS
-4- PO INT
BYPASS CONTROL
CANOPY AND
SEAL GROUND
TEST CONNECTION
SEAL PRESSURE -4-6
DEFOG FLOW
n.
CHECK
VALVES21-
,--[
TURBINE
REFRIG BYPASS
VALVE (DUAL)
COMPRESSOR
SYSTEM SHUTOFF
VALVE
SECONDARY FUEL/AIR
HEAT EXCHANGER
( INTERCOOLER )
AUXILIARY
SYSTEM FUEL
CHECK VALVE
OCCURS ON
RH INSTL ONLY
14�)
FUEL
OUT
SENSE
LINE-
SENSE LINE
FUEL OUT
CHECK
VALVE
�
TOW
WINDSHIELD BLEED
DEICING ONLY PRESSURE
REGULATOR
WINDSHIELD VALVE
RAIN REPEUENT
TANK PRESSURE
PRIMARY AIR/FUEL
HEAT EXCHANGER
BYPASS
THERMOSTAT
U SAFETY ZONE
DRAINS
4-07/
AIR FROM ENGINE -I
AIR INLET DUCT
Figure 1-32 (Sheet 1 of 3)
r BYPASS
I VALVE
SPIKE TRANSDUCER
SHROUD
200 MESH
FILTER
FZ00 -7( I)(d
1-74
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A-12
SECTION I
AIR CONDITIONING AND PRESSURIZATION SYSTEM
TO UPPER COCKPIT
OUTLETS
LOOKING
FORWARD
GROUND )
AIR
CONNECT
NOSE WHEELWELL
ARC
50
ARC
50
NOSE AIR
SHUTOFF
VALVE
SUIT
POOPER
WSJ
IQ)
ELECTRIC
DEFOG
VALVE
PULL TO
CLOSE
VALVE
'
Ii
DEFOG AIR
SUPPLY
SEAT I E-BAY
DISCONNECT
si
iTv
GROUND
FIXED CONNECT E
ORIFACE INV.
RESTRICTOR COOLING
( 1,
i I
ALTERNATE
LOCATIONS
OF AFT
VALVES
TYPE II
Q=BAY
L
I GS
PLATFORM
SHROUD
A/C t
BAY ENS] NS
3NE
3
L. H. AND R.H.
FORWARD
CHEEKS
Figure 1-32 ( Sheet 2 of 3)
$ FROM
R. H.
_Is FROM
L. H.
NOTE
INVERTERS
L. H . SIDE ONLY
L. H. AND R. H.
AFT CHEEKS
A NO. 1
I' `INVERTER'
A
"INVERTER_
,
4\ NO.3
�iNVERTERI
EXIT THROUGH
NOSE WHEELWELL
F200-7(2)(d)
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1-75
Approved for Release: 2017/07/25 C00821248
SECTION I A - 1
AIR CONDITIONING AND PRESSURIZATION SYSTEM
CANOPY
SEAL
PRESSURE
TEST
MANUAL NOSE
AIR SHUTOFF
VALVE
�9'
...� NOSE COOLING
iAIR DUCT
NOSE HATCH SEAL
ARC ARC
50 50
74)
NOSE HATCH
SEAL SELECTOR
DE-ICE
SHUT-OFF
VALVE
REG.
HOT DEICE
AIR 200� MIN
FROM LH SYSTEM
CANOPY
SEAL
SELECTOR
\L.
Lis*
'1WINDSHIELDiejk
DEFOG tri
MANIFOLD
=
EJECTION
CUTTER
CKPT AIR
OUTLET
FOUR
PLACES
COCKPIT S ILL
OUTLET CHECK
(LAND R) IT IVALVE
(TYPICAL)
SUIT VENT BOOST
(TROMBONE)
CANOPY SEAL
4,- SUIT VENT HOSE alh-
CANOPY
SEAL 16- HOT DEFOG
PRESSURE COCKPIT/NOSE AIR (200�MIN.)
AIR SUPPLY
ELECTRIC DEFOG
DEFOG
VALVE
Figure 1-32 (Sheet 3 of 3)
HOT DEICE
AIR 200� MIN
FROM RH SYSTEM
FZ00-7(3)(d)
1-76
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SECTION I
A-12
26,500 feet. The 10,000 foot position is in-
tended for use during subsonic low altitude
ferry flights but is not restricted for use as
desired during climbs, descents and high
altitude cruise. A rate control is incor-
porated which limits the pressure change to
2500 ft/min when changing schedules.
NOTE
During descents from high altitude,
only the normal cockpit pressure
schedule will provide optimum
cockpit cooling. The 10,000 foot
schedule will not cool the cockpit
in descents as well as the normal
schedule due to increased turbine
back pressure.
Cockpit Air Switch
The cockpit air switch is a three position
switch with labeled positions of NORM (left)
OFF (center) and EMER (right). In the
NORM position the left system shutoff valve
is deenergized to open and the left engine
system furnishes air to the cockpit. In the
OFF position the left shutoff valve is en-
ergized to closed, shuting off the normal
cockpit air. In the EMER position left sys-
tem air is shutoff, the crossover valve in
the right system is energized closed and
the right system shutoff valve is deener-
gized to open and right system air is fur-
nished to the cockpit. The circuit is
powered by the essential dc bus.
NOTE
In the EMER position the Q-bay
system switch OFF position is in-
effective and right system air
must be shut off by moving the
cockpit air switch to the NORM
position.
Q-Bay System Switch
The Q-bay system switch has two positions
and is located on the upper left side of the
instrument panel. In the ON (up) position
the right engine system's shutoff valve is
deenergized to open so that right engine air
can flow to the E-bay. If the cockpit air
switch is in the crossover or EMER
position this air will be ducted to the cock-
pit and will enter the E-bay through the
cockpit regulator valving. In the OFF posi-
tion the shutoff valve is energized to off and
Q-bay system air is shutoff if the cockpit
air switch is in NORM position. The cir-
cuit is powered by the essential dc bus.
Temperature Control Selector Switches
Two selector switches, one for the cockpit
and one for the Q-bay and/or emergency
cockpit air, are installed on the upper left
instrument panel. Each switch has four
positions; AUTO (up), COLD (down left),
WARM (down right) and HOLD (center).
The switches are spring loaded to HOLD
from the COLD and WARM positions. The
switches will normally be in the AUTO posi-
tion; however, the pilot can manually over-
ride the automatic feature by moving the
switch td either the momentary COLD or
WARM position. The manual COLD control
will provide colder air, if required, than
the automatic control. The No. 1 inverter
powers the cockpit temperature control
system. The No. 2 inverter powers the
Q-bay and/or emergency cockpit air tem-
perature control system.
Temperature Indicator Selector Switch
A temperature indicator selector switch
located on the upper left instrument panel
allows the pilot to monitor cockpit or Q-bay
temperature. Cockpit temperature is indi-
cated when the switch is placed in the CKPT
(left) position and Q-bay temperature when
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1-77
SECTION I
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A-Li
AIR CONDITIONING CONTROL PANEL
OFF
NORM EMER
CKPT AIR-1
COLD WARM C01/1) WARM
NORM CKPT Q-BAY
AIR
AUTO AUTO ON
HOLD HOLD
COLD WARM COLD WARM OFF
NORM CKPT Q-BAY OR
AIR EMER CKPT AIR
PRESS
DUMP
PRESS
NORM
10
3
1 COCKPIT TEMPERATURE MONITOR SELECTOR SWITCH
2 COCKPIT AIR SWITCH
4 3 DEPRESSURIZATION SWITCH (DUMP)
4 TEMPERATURE INDICATOR
5 TEMPERATURE CONTROL KNOBS
5 6 TEMPERATURE CONTROL SELECTOR SWITCHES
7 Q-BAY SYSTEM SWITCH
8 CABIN PRESSURE SCHEDULE SELECTOR SWITCH
9 CABIN ALTITUDE GAGE
10 ALTITUDE INDICATOR SELECTOR LEVER
6
7
F200-50M
Figure 1-33
1-78
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Approved for Release: 2017/07/25 C00821248
A-12
SECTION I
the switch is placed in the Q-BAY (right)
position. Power for the indicator is fur-
nished by the essential dc bus.
NOTE
Up to a point, the insulation and
ventilation of the pressure suit
will keep the pilot comfortable in
a cockpit environment that is too
warm. The temperature indi-
cator is provided so as to allow
anticipation of a temperature
condition that might eventually
become too hot for comfort. If
the cockpit temperature approaches
140o, the suit will not keep the
pilot comfortable.
Temperature Control Rheostats
Two temperature control rheostats, one for
the cockpit and one for the Q-bay and/or
emergency cockpit air are installed on the
upper left instrument panel. Arrows indi-
cate the direction of rotation necessary to
increase temperature. Generally, it is
necessary to periodically rotate the re-
spective temperature control rheostat to-
ward the COLD position to maintain a com-
fortable temperature in the ventilated flying
suit and keep the Q-bay temperature in
tolerance. Electrical power for the cockpit
temperature control circuits is from the
No. 1 inverter. Q-bay and/or cockpit
emergency air control is powered by the
No. 2 inverter.
Pressure Altitude Gage
A cockpit and Q-bay pressure altitude gage
is located on the left forward panel and in-
dicates either cockpit or Q-bay altitude as
selected by the cabin-Q-bay selector.
Altitude Selector Lever
This switch type lever is located on the
left forward panel. It is labeled CABIN
ALT in the up position and Q-BAY ALT in
the down position and selects the respective
pressure altitude to be indicated on the
gage.
Depressurization (Dump) Switch
A two position lift-lock depressurization
switch labeled PRESS DUMP and PRESS
NORM is located on the upper left instru-
ment panel. When the switch is pulled out
and moved to the PRESS DUMP position,
both the cockpit and Q-bay will be depres-
surized by the opening of the safety valves.
When moved to the PRESS NORM position
the safety valves will close and the cockpit
and Q-bay will repressurize.
WARNING I
Depressurization and repressur-
ization will occur at an extremely
rapid rate.
Nose Hatch Seal Shutoff Lever
A nose hatch seal shutoff lever, located on
the forward right side of the cockpit, op-
erates the nose hatch seal shutoff valve.
It is normally in the ON position to allow
canopy seal pressure to inflate the nose
hatch seal. In the OFF position the nose
hatch seal is isolated from the canopy seal
system. This prevents the deflation of the
cockpit canopy seal in the event of excessive
nose hatch seal leakage.
Nose Air Shutoff Handle
A nose air shutoff T-handle is located at the
bottom of the annunciator panel. It is nor-
mally in the locked ON position. The handle
is turned counterclockwise to unlock and
then pulled out to shut off airflow to the
pressurized nose compartment.
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1-79
SECTION I
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LIQUID OXYGEN SYSTEM
13
SYSTEM 1 LIQUID OXYGEN CONVERTER
1-80
10
14
3
YSEM 2 LIQUID OXYGEN CONVERTER.
13
AIME=
10
=2=22)
:OXY QTY
OXY TEST
01 I
IND
1 PRESSURE SWITCH
2 DRAIN VALVE
3 HEAT EXCHANGER
4 CHECK VALVE
(5 PSI DIFFERENTIAL)
5 CONTAINER LIQUID OXYGEN
6 QUANTITY PROBE (110)
7 RELIEF VALVE 100-120 PSI
8 WARMING COIL
9 PRESSURE OPENING VALVE
(OPENING PRESSURE 88-92 PSI)
10 PRESSURE CLOSING VALVE
(CLOSING PRESSURE 73-75 PSI)
11 BUILDUP AND VENT VALVE
12 FILLER VALVE
13 OVERBOARD VENT
14 LIQUID OXYGEN
CONVERTER. ASSEMBLY
NO
1 OXY LOW'
Eto
ON
Figure 1-34
NO 2 OXY LOW
NOTE
SYSTEMS SHOWN IN
BUILDUP POSITION.
SYSTEM 2 VALUES AND
NOMENCLATURE IDENTICAL
TO SYSTEM 1.
F200 -55(b)
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A-12
SECTION I
OXYGEN SYSTEM AND PERSONAL EQUIPMENT
The aircraft is equipped with dual liquid
oxygen systems. Two liquid oxygen con-
verters located in the right side of the nose-
wheel well have a capacity of ten liters (2.6
gallons) each. The liquid oxygen flows, by
gravity, into the pressure buildup coil and
vaporizes because of exposure to ambient
temperature surrounding the coils. The gas
flows through the pressure closing portion
of the pressure control valve and the build-
up and gas ports of the fill valve and then
back into the top of the container where it
collects and develops into a higher pressure.
This cycle continues until the system op-
erating pressure is reached (80 + 2 psi) at
which time the pressure closing valve closes
and stops the flow of liquid oxygen through
the pressure buildup coil. The liquid oxy-
gen will now flow through the check valve
and out the converter supply port to the air-
craft heat exchanger. During periods of
shut down system pressure will continue to
rise because of normal liquid boil off. The
increase in pressure is sensed at the pres-
sure opening valve. At 90 + 2 psi this valve
opens dumping the gas back into the con-
verter. The pressure will continue to
slowly rise, due to boil off, until it reaches
reflief valve opening pressure of 100 to 120
psi. The excess pressure is vented over-
board through the relief valve. Two ON-
OFF levers for the two systems, are located
on the oxygen control installed on the left
console. The needles on the pressure gage
will fluctuate, indicating oxygen flow when
the pilot inhales. Liquid oxygen is warmed
and converted to gas for breathing bypass-
ing through a heat exchanger which consists
of additional length of tubing in the supply
line. The low pressure gage on the oxygen
control panel indicates a normal pressure
of 50-100 psi.
Liquid Oxygen Quantity Gage
The liquid oxygen quantity gage is located
on the left side of the instrument panel. It
is calibrated in 1/2 liter increments from
0 to 10. The quantity gage is a double
needle type and indicates the quantity of
liquid oxygen remaining in the No. 1 or
No. 2 systems. When visible, a red OFF
indicator at the bottom of the gage indicates
the gage is not receiving power from the
No. 1 inverter.
Indicator Test Switch
A red test button labeled IND TEST is lo-
cated on the left side of the instrument
panel. When this button is pressed the oxy-
gen quantity gage needles will reduce indi-
cations. As the oxygen needles approach
the 1 liter mark the OXY LOW warning light
will illuminate. When the button is re-
leased the gage needles will resume their
original position. The CIT and spike and
forward bypass position indicators are also
tested by this button.
Oxygen Low Indicating Lights
Two oxygen low warning lights are located
on the pilot's annunciator panel.. The lights
are labeled NO. 1 OXY LOW and NO. 2
OXY LOW. Each light will illuminate when
oxygen pressure drops to 58 + 3 psi or
when 1 liter or less remains in the system.
EMERGENCY OXYGEN SYSTEM
Two independent emergency oxygen systems
are installed in the pilot's parachute pack.
Each system consists of a 45 cubic inch,
2100 psi cylinder. The systems will supply
oxygen simultaneously during bailout and
when the aircraft oxygen systems fail. An
oxygen line is routed around each side of
the pilot's waist and connects to the suit
controller valve. Emergency oxygen flow
pressure is slightly lower than aircraft
system pressure. Oxygen duration of each
emergency system is approximately 15
minutes.
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1-81
SECTION I
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OXYGEN DURATION CHART
10
9
8
7
6
5
4
3
2
1
LIQUID
LITERS/
DEWAR
1 -82
1
1
I
t 1
1
1
I
I
I
I
BOILOFF
OXYGEN
DURATION
AVAILABLE:
I
.. I
I �
I 5
I S
I
i
4 SYS
DOES NOT
FLOW
TWO 10 LITER
ALT
CONVERTERS
5.22 PS
IA (26M)
DIVERGENCE
BETWEEN
CABIN
SUIT ALT
5.68 PS IA
�
�
�
�
�
�
SYSTEMS
�
�
�
�
�
�
� 1
\
<
5
\
MIDPOINT
FAILURE
(8.50 LITERS
LOST
.
30
AND
MIN GROUND
CLIMB TIME
I
V
�
�
�
�
�
25 LPM
(TWO SY
TEM)
�
�
� .0
V
LPM (ONE
SYSTEM)
I--4--CAB
ALT
%
�
�
I
I
AND SURPUS,
IN (26M)
st < 11.
Il
l�
ot HR
MIN
MISS ION
COMPLETE�\:
III \
%
I te7---
LOW
LEVEL LIGHT
COMES ON
I
I
I
I
I
I
I
I
I
I
I
�
�
\
�
0
2
4
6
8
10
12
TIME - HOURS
Figure 1-35
14
16
18
20
22
24
4-11-66
F200-84
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SECTION I
A-12
The emergency oxygen system is actuated
either manually by pulling the conventional
green apple, or automatically by the upward
motion of the seat during ejection. The
emergency oxygen system should be ac-
tuated if the aircraft systems are not de-
livering the desired amount of oxygen or
hypoxia or noxious fumes are suspected.
FULL PRESSURE SUIT
A full pressure suit is provided which is
capable of furnishing the pilot with a safe
environment regardless of pressure condi-
tions in the cockpit. The suit consists of
four layers, ventilation manifold, bladder,
link net, and heat-reflective outer garment.
The ventilation manifold layer allows vent
air to circulate between the pilotls under-
wear and the bladder layer. The bladder
provides an air-tight seal to hold pressur-
ized air in the suit. The link net is a mesh
which holds suit configuration in confor-
mance with the pilot's body. The outer
layer of heat-reflecting cloth provides some
protection from a hot environment. Air
pressure to the suit is regulated by a suit
controller valve, located on the front of the
suit just above the waist.
Pressure Suit Ventilation Air
Air for suit ventilation is provded by the
cockpit air-conditioning system. Temper-
ature of the ventilation air cannot be varied
except by changing cockpit inlet air temper-
ature. Ventilation airflow rate may be re-
gulated by a suit flow control valve installed
at the hose connection point on the suit.
Ventilation air and exhaled breathing air
are exhausted from the suit.
Suit Ventilation Boost Valve Lever
The suit ventilation boost valve lever, la-
beled SUIT VENTIL BOOST, is located on
the left console. The lever is marked
NORMAL (aft) and EMERG (forward). Op-
erating the lever positions a butterfly valve
in the cockpit air-conditioning air supply
line in such a way as to vary the pressure
of the air available to the suit system. In-
creased pressure results in more air to the
suit. Moving the lever toward EMERG
position progressively results in more
pressure to the suit system by constricting
the air-conditioning airflow to the cockpit;
in the NORMAL position (used when engine
rpm is high) the cockpit air-conditioning
line requires no constriction to provide
sufficient airflow to the suit. At IDLE
engine rpm the ventilation boost valve lever
must be kept at 2/3 of the way from NOR-
MAL to EMERG in order to provide suffi-
cient air for conditioning the suit and cool-
ing the INS platform and inverters in the
A/C bay. During takeoff and normal flight
the valve lever is kept in the NORMAL posi-
tion. If the pilot suffers discomfort, such
as might happen with a gradual climb to an
extreme altitude or during low-rpm descents,
the valve lever is gradually moved toward
the EMERG position until a comfortable
pressure and ventilation condition is at-
tained. The valve lever should not be
moved toward EMERG more than necessary
to provide pilot comfort; excessive suit
system pressure will unduly reduce the
available refrigeration.
Suit Controller Valve
All four aircraft and emergency oxygen sys-
tem lines enter the controller valve at the
front waist of the pressure suit. The con-
troller valve contains a sensor that pro-
grams airflow and oxygen to keep internal
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suit pressure at 3.5 psi (equivalent to pres-
sure at 35,000 ft) in the event of cockpit de-
pressurization. A press-to-test button for
each oxygen system is installed on the con-
troller valve, which allows the pilot to
check suit inflation.
Face Plate Heat Switch
A face plate heat switch is installed on the
right console of the cockpit. The switch has
four positions; OFF, LOW, MED and HIGH.
Heat may be regulated to defog the face plate
as required. Defogging is accomplished by
the combination of face plate heat and oxy-
gen flow. The face plate heater circuit is
powered by the essential dc bus.
HELMET
The helmet head area is divided into two
separate sections by a rubberized cloth face
seal. The front area between the face and
the face seal receives oxygen from either
the aircraft or emergency oxygen system
through regulators built into the helmet.
Oxygen flows across the face plate from the
inhalation valves inside the helmet and ac-
complishes some face plate defogging be-
fore it is inhaled. The rear area receives
vent air for helmet interior temperature
regulation. The face seal is not positive;
however, the pressure of the oxygen in the
front area is slightly higher to prevent vent
air from leaking forward. An external
crank on the helmet is provided for head
band adjustment. Buttons on each side of
the helmet, when actuated, will lower the
face plate and visor. The face plate is
opened by moving the buttons and dumping
the pressure, allowing the face plate to be
rotated upward. If the aircraft or emer-
gency oxygen supply to the helmet is inter-
rupted or exhausted; the regulators in the
helmet sense the drop in pressure and the
face plate seal deflates, allowing ambient
air to enter the helmet so the pilot will not
suffocate.
GLOVES
Leather gloves fasten onto the suit at the
wrist rings. The inner liner of the glove
is similar to the suit inner liner and will
retain pressure.
BOOTS
The sock or boot liner fastens onto the suit
at the thigh by means of a zipper. The
boots are made of white leather for heat re-
flection and fit snuggly over the socks. A
spur that fastens to the seat is attached to
each boot.
OXYGEN MASK AND REGULATOR
When permitted by appropriate regulations
a substitute oxygen mask assembly may be
used in place of a pressure suit for flights
at low or intermediate altitudes. The as-
sembly consists of a specially designed
oxygen mask and F2700 oxygen regulator,
anti-suffocation valve and two oxygen per-
sonal leads with connectors for both air-
craft and emergency oxygen systems. In
the event that the regulator should malfunc-
tion or the oxygen supply is exhausted, an
anti-suffocation valve installed between the
regulator and the mask will sense the drop
in oxygen pressure and allow ambient air to
enter the mask to prevent suffocation.
SURVIVAL KIT
A reinforced fiberglas survival kit container
fits into the seat bucket and attaches to the
parachute by snap attachments on each side.
A door on the top provides access to the
survival items stored inside. The kit con-
tains standard survival items such as radio,
flares, mirror, whistle, knife, matches,
rations, water, compass and first aid kit.
Various additional items depending on the
terrain and season may be
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provided. The kit is packed in a water
proof bag attached to a 20 foot retention
lanyard. If an overwater flight is antici-
pated, a life raft may be stowed on top of
the plastic bag and attached to the lanyard.
During ejection the life raft inflating device
is armed. Following ejection, the survival
kit release handle should be pulled before
reaching the ground. This action separates
the survival gear from the pilot and inflates
the life raft. The survival gear and life
raft remain attached to the parachute harness
by the retention lanyard. During a rapid
abandonment of the aircraft on the ground,
the survival kit release handle may be used
to free the pilot of the survival kit (including
the lanyard) without inflating the life raft.
PARACHUTE
A special parachute with a 35 foot canopy is
used. The large canopy provides a normal
descent rate with the bulky personal equip-
ment required for high altitude flight. A
small diameter, ribbon type stabilizing
drogue chute is also provided. Above
17,000 feet altitude, the drogue chute is de-
ployed first in order to stabilize free fall
of the pilot. The drogue is automatically
jettisoned at 15,000 (+ 400) feet after an
aneroid controlled opener deploys the main
chute. Below 16,200 feet the main chute
only deploys immediately. A manual T-
handle is also available for opening the
main chute. The chute pack is equipped
with conventional quick release buckles.
The emergency oxygen bottles are located
between the chute canopy and the pilotis
back. A combination hand squeezed bulb
and manually operated pressure relief valve
located adjacent to the suit. controller is
used to adjust cushion pressure as desired.
A red knob located on the left harness strap
is connected to the parachute timer arming
cable and is used to actuate the timer when
bailout is made.
WINDSHIELD
The windshield is composed of two glass
assemblies secured and sealed in a V-
shaped titanium frame. The glass surfaces
are coated with low reflective magnesium
fluoride. A collapsible vision splitter is
also installed on the windshield center line
to minimize reflections.
DEFOG SYSTEM
The windshield defog system delivers hot
air from both right and left air systems
through check valves to defog the windshield
and canopy. A plastic V-shaped air duct
runs along the lower edge of the windshield.
Hot defog air is supplied through this duct
when selected by a switch that is located on
the upper left console. The air is directed
to the windshield through a series of holes
on the upper surface of the duct. Holes are
also provided at the aft ends of the duct to
direct air toward the canopy glass.
Defog Switch
A three position defog switch is located at
the forward end of the upper left console.
When held in the momentary DEFOG IN-
CREASE (forward) position the motor
driven defog valve will open. Time of
travel to full open is approximately 7 to 13
seconds. In the HOLD (center) position
the valve will stop at any desired partial
open position; in the OFF position the valve
will completely close. The circuit is
powered by the essential dc bus.
LEFT WINDSHIELD HOT AIR DEICING SYSTEM
Hot air is ducted from the L & R pressuri-
zation supply downstream of the fuel air
heat exchanger and upstream of the pres-
sure regulator and air cycle refrigerator,
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CANOPY AND CONTROLS
1 CANOPY LATCH ROLLER BRACKETS
2 CANOPY LIFTING HOLE
3 CANOPY PROP ASSEMBLY AND UPLOCK
4 CANOPY PROP (GROUND HANDLING)
5 CANOPY EXTERNAL LATCH CONTROL
2
3
5
6
DETAIL B
6 CANOPY EXTERNAL JETTISON HANDLE
7 CANOPY INTERNAL JETTISON HANDLE
8 CANOPY LATCH HOOKS
9 CANOPY LATCH HANDLE
Figure 1-36
7
Mt12-36
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to a series of orifices located on the left
side of the outside center windshield support.
The system includes left and right solenoid
shutoff valves controlled by a switch in the
cockpit. Power is furnished by the essen-
tail dc bus.
Windshield Deice Switch and Indicator Light
The 3-position windshield deice switch is
located on the upper left instrument panel.
In the OFF (right) position the shutoff valves
are closed and no deicing air is supplied.
In the R ON (center) position the hot air is
furnished by the right pressurization sys-
tem and 1/2 flow is available for deicing.
In the LR ON (left) position both L & R shut-
off valves are opened and full flow is avail-
able to the windshield orifices. Power for
the switch and lights is furnished by the dc
essential bus.
NOTE
. A considerable amount of air is
used when operating the deicing
system in the L/R ON position.
This may reduce the cockpit and
Q-bay air supply when operating
in the lower ranges of engine rpm.
. The deicer indicator light, located
above the switch, will be illuminated
at any time the deice switch is not
In the OFF position.
WINDSHIELD RAIN REMOVAL SYSTEM
A rain removal system is provided for
clearing the windshield when operating the
aircraft in rain. It has a tank that is pres-
surized by air from the windshield deicer
system and the tank is connected to a spray
tube located on the left side of the wind-
shield center divider. A pushbutton switch,
located on the upper instrument panel, is
used to spray the rain removal fluid onto
the left windshield. Power is furnished by
the essential dc bus.
Do not apply rain repellent on a
dry windshield as prolonged
obscuration may result.
CANOPY
The canopy consists of two high temper-
ature resistant glass windows secured within
a reinforced titanium frame which is hinged
at the aft end of two hinge pins. Operation
of the canopy is completely manual. Small
holes in each side of the canopy are pro-
vided as lifting points from the outside. Nc
handles are provided on the inside of the
canopy for moving it up or down. A prop
assembly locks the canopy in the full open
position. The canopy is secured in the
closed and locked position by a four hook
interconnected latching mechanism. A ni-
trogen boost counterbalancing system is
provided to aid in the manual opening and
closing of the canopy. This nitrogen is
also used to force water into the map case
when the destruct system is actuated.
NOTE
Actuation of the destruct system
tends to deplete the nitrogen boost
counterbalance system and in-
crease the manual force needed to
open the canopy. Canopy jetti-
soning may be necessary for rapid
egress.
An internal latching handle is installed be-
low the right canopy sill, allowing the canopy
to be latched from the inside. An external
fitting located on the left side of the aircrail:
can be used to operate the latches from the
outside.
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The canopy should be opened or
closed only when the aircraft is
completely stopped. Maximum taxi
speed with the canopy open is 40
knots. Gusts or severe wind con-
dition should be considered as a
portion of the 40 knot limit.
Canopy Latch Handle
A canopy latch handle is located under the
right sill in the cockpit and rotates forward
to lock. The sill trim is cutout to expose
the action of the locking lugs and pins as
the handle is rotated forward. A cam over
center action allows the handle to remain
only in the latched or unlatched position.
No canopy unsafe warning light is provided.
Canopy External Latch Control
A flush mounted external latch fitting is
located on the left side of the aircraft and
permits the canopy to be opened from the
outside. The fitting accepts a 1/2 inch
square bar extension. Once the canopy is
unlocked, it may be raised manually until
the prop locks it in the open position.
Canopy External Jettison Handle
The canopy external jettison handle, located
beneath an access panel on top of the left
chine, permits ground rescue personnel to
jettison the canopy. Sufficient cable length
is provided to allow the operator to stand
clear of the fuselage during the jettisoning
procedur e.
Canopy Internal Jettison Handle
A canopy jettison T-handle is located on
the left console wall adjacent to the pilot's
leg. The handle can be used to jettison the
canopy without initiating the seat ejection
system. The handle is held in the stowed
position by a lockwire and a ground safety
pin. Storage for the canopy jettison and
seat safety pins is provided at the forward
end of the upper right console. Cable
travel is approximately six inches.
CANOPY SEAL
An inflatable rubber seal is installed in the
edge of the canopy frame. The seal seats
against the mating surfaces of the canopy
sill and windshield to provide sealing for
cockpit pressurization. The canopy seal
pressurization lever above the forward
right console operates the seal inflation
valve. A nose hatch seal shutoff lever is
also provided to prevent deflation of the
canopy seal in the event of nose hatch seal
leakage.
CANOPY JETTISON SEQUENCE
The canopy jettison system is designed to
unlatch and jettison the canopy from the
aircraft by means of explosive initiators
and thrusters. The system consists of two
initiators which are independently actuated
by either the ejection seat D-ring or the
canopy jettison handle, a canopy unlatch
thruster, a canopy removal thruster, a
canopy seal hose cutter, cable linkage and
gas pressure lines. Either the D-ring
initiator or the canopy initiator or the
canopy initiator will fire the unlatch
thruster which unlocks the canopy. This
thruster then activates the canopy seal hose
cutter and fires the canopy removal thruster
which jettisons the canopy. Whenever the
canopy is jettisoned by use of the canopy
jettison handle, the canopy jettison initiator
gas pressure positions a seat jettison safety
valve to prevent initiating the seat ejection
sequence until the D-ring is pulled. Pull-
ing the D-ring jettisons the canopy as the
initial step in the ejection sequence.
REAR VIEW PERISCOPE
A manually extended rear view periscope
is mounted in the top of the canopy to enable
the pilot to see the engine nacelles and rear
fuselage and rudder area. The periscope,
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tion. It is moved by using the white nylon
pad, mounted on the aft side of the viewing
tube, as a handle. Pushing the handle to the
left unlocks the tube, allowing the periscope
to be extended. Then, pushing the tube up-
ward to a spring-detented position makes
the rear view available. Cockpit pressure
tends to assist extension, and resists re-
traction. The diameter of the instantaneous
cone of view is approximately 100; however,
head movement extends the viewing cone to
approximately 30 total angle. When ex-
tended, the periscope can be rotated hori-
zontally to move the center of the viewing
arc up to 10o from the aft centerline. The
de-magnification ratio of the lens system is
1 to 0.5.
EJECTION SEAT
The ejection seat system utilizes an upward
catapult and rocket thrust to provide mini-
mum risk ejection capability at ground level
when airspeed is at least 65 KIAS. The
seat incorporates an ejection ring, headrest,
knee guards, automatic foot retractors,
automatic foot retention separation, a pilot-
seat separation device, shoulder harness,
inertia reel lock assembly, and an auto-
matic opening seat belt. A speed sensor
mounted on the fuselage behind the seat
automatically selects one of two seat se-
paration delays, depending upon airspeed
at ejection. (Refer to Ejection Sequence
this section.) Quick disconnect fittings in-
stalled on the seat rails and the floor of the
aircraft permit disconnection of the oxygen,
ventilated suit and electrical lines.
Seat Vertical Adjustment Switch
The seat may be adjusted vertically by
means of an electric actuator mounted on
the lower end of the catapult. The three-
position switch is located on the right side
of the seat bucket. The seat moves in the
direction the switch is moved. Power for
seat adjustment is furnished by the essen-
tial dc bus.
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Shoulder Harness Inertia Reel Lock Lever
SECTION I
A shoulder harness inertia reel lock lever
Installed on the left side of the seat bucket
is provided for locking and unlocking the
shoulder harness. The lever has two posi-
tions. LOCK and UNLOCK. Each position
is spring loaded to hold the lever in the se-
lected position. An inertia reel located on
the back of the seat will maintain a constant
tension on the shoulder straps to keep them
from becoming slack during backward
movement. The reel also incorporates a
locking mechanism which will lock the
shoulder harness when a 2 to 3 g force has
been exerted in a forward direction. When
the reel is locked in this manner, it will '
remain locked until the lever is moved to
the LOCK position and then returned to the
UNLOCK position.
Ejection (D) Ring
An ejection ring, located on the front of the
seat bucket, is the primary control for
ejection. An ejection safety pin is installed
in the ejection ring housing bracket.
Ejection 1-Handle
The aircraft are equipped with a backup
secondary seat ejection system. The T-
handle for this seat ejection system is un-
locked and made accessible only by first
pulling the ejection D-ring.
WARNING
The ejection seat must not be fired
by pulling the T-handle while the
canopy is still in place. The pilot
can not eject through the metal
canopy.
When the secondary ejection T-handle is
pulled a separate initiator fires the seat
catapult and seat separation and belt open-
ing initiator.
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EJECTION SEAT
1-90
2-089
1 MANUAL CABLE CUTTER RING
2 HEADREST
3 SHOULDER HARNESS
4 AUTOMATIC SEAT BELT
5 SHOULDER HARNESS INERTIA REEL LOCK LEVER
6 KNEE GUARDS
7 SEAT ADJUSTMENT SWITCH
8 EJECTION RING
6 9 EJECTION SEAT T AANDLE
10 FOOT RETRACTOR FMINGS
Figure 1-37
10
3-30-66
F200-28(b)
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Foot Spurs
Foot spurs, attached to the pilot's shoes,
are attached to the ejection seat by cables.
Normal foot movement is in no way re-
stricted since the cables, under a slight
spring tension, reel in and out freely.
When the ejection ring is pulled, the knee
guards rotate from their stowed position,
the cables to the foot spurs are reeled in
and the pilot's feet are retracted into the
foot rests. The foot cables are automati-
cally severed by a set of cutters as part of
the ejection sequence.
Manual Cable Cutter Ring
The ejection seat incorporates an emer-
gency means for cutting the foot retractor
cables. A D-ring, located to the right of
the seat headrest, will actuate the cable
cutters initiator if the automatic cable cut-
ter system fails or rapid abandonment of
the aircraft is required on the ground.
PILOT-SEAT SEPARATION SYSTEM
The ejection seat is provided with a pilot-
seat separation system which operates in
conjunction with the automatic seat belt re-
lease system. A windup reel is mounted be-
hind the headrest, and a single nylon web is
routed from the reel to halfway down the
forward face of the seat back. From this
point two separate nylon straps continue
down, pass under the survival kit, and are
secured to the forward seat bucket lip.
After ejection, as the seat belt is released,
an initiator actuates the windup reel which
winds the webbing onto a cross-shaft, pulls
the webbing taut, and causes the pilot to be
separated from the seat with a sling shot
action.
AUTOMATIC SEAT BELT
The ejection seat is equipped with an auto-
matic opening seat belt which facilitates
pilot separation from the seat following
ejection. Belt opening is accomplished
automatically as part of the ejection se-
quence and requires no additional effort on
the part of the pilot.
SEAT BELT-PARACHUTE ATTACHMENT
If the pilot is wearing an automatic opening
aneroid type parachute, the parachute lan-
yard anchor from the parachute aneroid
must be attached to the swivel link. As the
pilot separates from the seat, the lanyard,
which is anchored to the belt, serves as a
static line to arm the parachute aneroid.
The parachute aneroid preset altitude is
approximately 15,000 feet.
EJECTION SEQUENCE
Pulling the D-ring is normally the only ac-
tion required to initiate pilot ejection and
results in firing both the canopy jettison and
ejection seat systems. All resultant actions
will occur automatically and in a specific
sequence as explained below.
The D-ring cable fires the ejection se-
quence initiator, actuating the canopy jetti-
son system and the leg guard thruster. The
leg guard thruster rotates the knee guards,
retracts the pilot's feet, activates the cable
cutter backup initiator and locks the shoulder
harness. Movement of the canopy jettison
thruster (final step in canopy jettison se-
quence) actuates an initiator which fires a
0.3 second delay catapult initiator and arms
the speed sensor. The 0.3 second delay
assures complete canopy separation prior
to seat ejection. Gas pressure from the
catapult initiator fires the rocket-catapult,
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the 4-second seat separation delay initiator,
and enters the speed sensor. If airspeed is
below 295 KIAS, the gas pressure passes
through the speed sensor and fires the 1.0
second delay seat separation initiator. If
airspeed is above 302 KIAS, the pressure
is blocked by the speed sensor.
Initial seat movement upward on the rails
disconnects normal oxygen, ventilated suit
and electrical lines, and activates the emer-
gency oxygen supply. Between 295 and 302
KIAS either the 1 or 4 second delay may be
experienced because of the speed sensor
tolerance.
Either the 1.0 second delay initiator (below
295 KIAS) or the 4-second delay initiator
(above 302 KIAS) actuates the cable cutters,
releases the pilot's feet, opens the seat
belt and fires the seat separation system.
A static line attached to the seat belt is
pulled as the pilot separates from the seat
and activates the automatic parachute se-
quence.
If the normal D-ring ejection sequence was
not accomplished; the canopy must be jetti-
soned either by use of the canopy jettison
system or manually. Pulling the T-handle
initiates the secondary seat ejection se-
quence.
The T-handle backup ejection
sequence does not rotate the knee
guards nor retract the foot cables.
Seat separation delay time will be
4 seconds regardless of airspeed.
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Section II
ORMAL PROCEDURES
TABLE OF CONTENTS
Page
Page
Preparation For Flight
2-1
Cruise
2-18
Preflight Check
2-4
Prior To Descent
2-19
Starting Engines
2-6
Descent
2-19
Before Taxiing
2-9
Air Refueling
2-21
Taxiing
2-10
Before Landing
2-25
Before Takeoff
2-10
Landing
2-27
Takeoff
2-11
Go-Around
2-29
After Takeoff
2-15
After Landing
2-31
Normal Climb
2-15
Engine Shutdown
2-31
Alternate Climb
2-18
Abbreviated Checklist
2-32
PREPARATION FOR FLIGHT
FLIGHT RESTRICTIONS
Refer to Section V for Operating Restric-
tions and Limitations.
FLIGHT PLANNING
Refer to Appendix I.
TAKEOFF AND LANDING DATA
Refer to Appendix I for Takeoff and
Landing information.
WEIGHT AND BALANCE
Refer :10 Section V for Weight and Balance
Limitations. For detailed loading infor-
mation, refer to Handbook of Weight and
Balance Data. Before each flight, check
takeoff and anticipated landing gross
weights and weight and balance clearance
(Form 365F).
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SECTION II
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PERSONAL EQUIPMENT HOOKUP
0 HOOK UP SPURS
COOT SPURS WILL BE ATTACHED AND REMOVED
BY PILOT FROM A STANDING POSITION UPON
ENTERING AND LEAVING COCKPIT
CAUTION
PERSONAL EQUIPMENT TECHNICIAN WILL
ASSIST IN ATTACHING SPURS AND BALL
FITTING BY HAND IF REQUESTED
CONNECTED
DISCONNECTED
OCOMMUNICATIONS (FACE HEAT AND RADIO)
CONNECT HELMET CHORD TO PARACHUTE
EXTENSION CHORD
OTURN FACE HEAT ON LOW (CONTROL ON
RIGHT HAND CONSOLE)
ON RIGHT
CONSOLE
PANEL
I
0 SECURE OXYGEN PERSONAL LEAD HOSES
IN QUICK DISCONNECT (INSIDE FRONT OF
SEAT BUCKET)
a INSTALL NO. 2 HOSE CONNECTION
AND TURN PRESSURE ON
b INSTALL NO. 1 HOSE CONNECTION
AND TURN PRESSURE ON
C CHECK PRESSURE 65 TO 100 PSI
CONNECT PARACHUTE HARNESS, THREE PLACES
a CHEST STRAP (UNDER HELMET HOLD
DOWN LANYARD)
b RIGHT LEG STRAP (OVER PERSONAL
OXYGEN LEAD HOSES)
c LEFT LEG STRAP
ON LEFT
CONSOLE
PANEL
0 ADJUST KIT SEAT STRAPS; RIGHT AND LEFT SIDE
CONNECT EMERGENCY OXYGEN HOSES,
SLIDE KNURLED FITTING INTO PLACE,
INSERT SAFETY CLIP, PULL ON HOSE SLIGHT?
TO ASSURE OF LOCKED POSITION
NOTE
LEFT HOSE OVER HELMET HOLD DOWN STRAP
F200-12(I)(c)
Figure 2-1 (Sheet 1 of 2)
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SECTION II
PERSONAL EQUIPMENT HOOKUP
PULL TO
ADJUST
0 LAP BELT
SECURE SHOULDER HARNESS STRAPS AND
PARACHUTE TIMER ARMING KEY. LOCK
BELT AND ADJUST
ON LEFT
CONSOLE
PANEL
PRESS DOWN
--*". TO LOCK
CHECK EMERGENCY OXYGEN CABLE AND
REMOVE SAFETY PIN
CHECK PARACHUTE ARMING (RED KNOB)
KNOB IS SECURED INTO DETENT
CHECK ACCESSIBILITY OF EMERGENCY OXYGEN
ACTUATOR (GREEN APPLE) 1800 PSI MINIMUM BOTH
SYSTEMS. INSURE GREEN APPLE IS SNAPPED
SECURE INTO DETENT
OCHECK PARACHUTE MANUAL .r HANDLE.
INSURE HANDLE IS SNAPPED SECURE
INTO HOUSING
Figure 2-1 (Sheet 2 of 2)
0 CHECK (TWO) PARACHUTE CANOPY ROCKET
JET RELEASES. INSURE ROLL BAR PIN
IS IN DOWN (LOCKED) POSITION. PULL ON
EACH RELEASE TO INSURE LOCK POSITION
0
0
CHECK FACE HEAT, PLACE BACK OF HAND
ON VISOR
CONNECT HEAT PROBE (IF APPLICABLE)
PRESS TO TEST BOTH SUIT EMERGENCY
PRESSURIZATION SYSTEMS, (SEE ILLUSTRA-
TION NO. 7) ONE AT A TIME. CHECK PRESSURE,
APPROXIMATELY 65 TO 100 PSI AND
FLUCTUAT ING
CHECK ACCESSIBILITY OF SUIT FLOATATION
KNOB PULL TAB
READJUST LAP BELT
CHECK OXYGEN QUANTITY, BOTH SYSTEMS
CHECK FOOT REST GUARDS
CONNECT VENT HOSE
NOTE
THIS WILL BE ACCOMPLISHED AFTER ENGINES
ARE RUNNING UNLESS EXTERNAL AIR CONDITION
VENTILATION UNIT IS HOOKED TO AIRCRAFT
VENT SYSTEM. PULL DOWN ON VENT HOSE
CONNECTION TO INSURE LOCK POSITION
F200-72alf.)
momm=m1IMMIIMIIMMMM=Approved for Release: 2017/07/25 000821248
2-3
SECTION II
AIRCRAFT STATUS
Approved for Release: 2017/07/25 C00821248
A-12
5. Battery switch - EXT PWR.
Refer to Form 781 for engineering, ser-
vicing, and equipment status.
EXTERIOR INSPECTION
It is not practical for the pilot to perform
an exterior inspection while wearing a pres-
sure suit. The exterior inspection should
be accomplished by other qualified per-
sonnel.
PREFLIGHT CHECK
ENTRANCE
A ladder platform stand which overhangs
the chine is used to gain entrance to the
cockpit. The canopy is unlatched exter-
nally by rotating the external canopy con-
trol clockwise with an L-shaped 1/2 inch
square bar. The canopy is manually raised
to the full open latched position.
BEFORE ENTERING COCKPIT
1. Manual cable cutter ring - Secure.
2. Ejection seat and canopy safety pins
installed - Check.
6. Accomplish and check personal equip-
ment hookup. (Hookup will be per-
formed by personal equipment per-
sonnel). Refer to figure 2-1.
7. Suit vent boost lever - Set at 2/3 lever
travel.
Left Console
1. IFF - ON. Set to proper mode and
code.
2. Panel and instrument lights switches -
As desired.
3. COMM selector switch - UHF.
4. External light selector switch - OFF.
5. Defog switch - OFF.
6. HF radio - OFF.
7. UHF radio - OFF.
8. Throttle friction lever - As desired.
9. TEB counter - Check 12.
10. Aft bypass switches - Both CLOSED.
Instrument Panel
INTERIOR CHECK
1.
Cabin Q-bay altitude selector lever -
CABIN.
1.
All circuit breakers - In.
2.
Landing and taxi light switch - OFF.
2.
Foot retractors - Attach.
3.
Brake switch - ANTI-SKID.
3.
Throttles - OFF.
4.
Cockpit temperature switch - AUTO.
4.
Landing gear lever - DOWN.
5.
Q-bay temperature switch - AUTO.
6.
Q-bay air switch - ON.
2-4
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Approved for Release: 2017/07/25 C00821248
A-12
SECTION II
7. Cockpit and Q-bay auto temperature 26. Spike and forward bypass position in-
rheostats - As desired. dicators - Check.
8. Cockpit and Q-bay temperature indi- 27. Fuel transfer switch - OFF (guard
cator switch - Q-BAY. down).
9. Cockpit air switch - ON.
10. Pressure dump switch - OFF.
11, Drag chute handle - Stowed.
12. Windshield deicer switch - OFF.
13. Clocks - Check.
14, Compressor inlet temperature gage -
Check needles together and indicating
ambient temperature.
28. Fuel dump switch - OFF (guard down).
29. ILS receiver - OFF.
30. Air refuel switch - OFF.
31. Destruct switch - OFF (guard down).
Right Console
1. Nose hatch seal pressure lever - ON.
2. Pitot pressure selector lever -
NORMAL.
15. Igniter purge switch - OFF (down). 3. Canopy seal pressure lever - OFF.
16. Compressor inlet static pressure gage - 4. Stability augmentation switches - OFF.
Check needles together and indicating
barometric pressure. 5. Autopilot switches - OFF.
17. TDI - Check for proper indication. 6. Inertial navigation system panel - As
required.
18, Altimeter - Set.
19. Periscope MIR SEL handle - Full for-
ward - (Projector).
20. Fuel derichment arming switch - OFF.
9. TACAN switches - T/R and tuned to
21. Restart switches - OFF, desired station.
22. Spike knobs - AUTO. 10. ADF receiver switch - ANT.
23. Inlet air forward bypass knobs - AUTO. 11. Floodlight switch - As desired.
24. Emergency fuel shutoff switches - 12. Face plate heat switch - As desired.
Fuel On (guards down).
7. Autopilot and attitude reference selector
switch - As desired.
8. BDHI needle selector switch - TACAN.
13. Flight reference system (FRS) compass
25. Cockpit pressure schedule switch -As select switch - MAG.
desired.
14. Birdwatcher and SIP power switches -
OFF.
Z-5
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Approved for Release: 2017/07/25 C00821248
SECTION II
A-12
Lower Instrument Panel
1. Surface limit release handle - Pulled
out.
2. Pitot heat switch - OFF.
3. Hydraulic reserve oil switch - OFF
(guard down).
4. Trim power switch - ON.
5. Nose air conditioning handle - Stowed.
6. Backup pitch damper switch - OFF
(guard down).
7. Pitch logic override switch - OFF
(guard down).
8. Yaw logic override switch - OFF
(guard down).
9. Gear release handle - Stowed.
EQUIPMENT FUNCTION CHECK
1. Inverter switches - NORM.
2. N2 and tank lights switch - Test.
a. NZ quantity indicators should
decrease to zero.
b. N QTY LOW warning light should
illluminate.
3. Crossfeed and boost pump switches -
Press lights on.
4. Pump release switch - PUMP REL,
then release.
5. Tank boost pumps - Check 1, Z and 6
TANK lights on (automatic sequencing).
6. Crossfeed switch - Press (check light
off).
7. Fuel quantity indicating system -
Check.
a. Individual (1, 2, 3, 4, 5 and 6)
tank quantities - Check.
b. Total fuel quantity - Check.
8. Gear and warning lights test switch -
Press.
a. All warning and fire lights should
illuminate.
b. Landing gear unsafe warning horn
should sound.
9. IND TEST button - Press.
a. Oxygen quantity needles will move
to below 0.
b. CIT indicator will decrease to-
ward zero.
c. Spike and forward bypass position
indicators increase to maximum
forward indication on spike and
maximum open on forward bypass.
10. Headset plug and oxygen mask - Connect
(if pressure suit is not used).
11.
No. 1 and No. 2 oxygen systems - ON
(if pressure suit is not used). Check
system pressures.
12. Tape and flight recorders - ON.
STARTING ENGINES
Before starting an engine, deter-
mine that the wheels are firmly
chocked since brakes are in-
operable until hydraulic pressure
is available and no parking brake
is installed.
2-6
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Approved for Release: 2017/07/25 C00821248
SECTION II
A-12
Determine that intake and exhaust
areas are clear of personnel and
ground equipment. The ground
personnel using interphone com-
munication equipment will be in
position to observe the exhaust
nozzle and nacelle inspection
panels during starting.
Do not move the control stick
until at least 1500 psi hydraulic
pressure can be maintained on
the A or B system gages or a
control system inspection will
be necessary.
1. Check with INS crew prior to starting
engines.
2. Fuel low pressure lights - Off.
3. Engine instruments - Check.
4. Ground starting unit - Instruct ground
crew to rotate engine for start.
5. Throttle - IDLE when rpm is indicated.
6. Fuel flow - Check 1500-2000 pph.
7. Engine light up will be indicated in ap-
proximately 15 seconds by a continuous
rpm increase and by a rise in EGT.
8. EGT - Check for 540oC max during
acceleration.
NOTE
If engine does not accelerate
smoothly to 3550-3650 rpm, re-
tard throttle to OFF and then
quickly advance to IDLE. This
"double clutching" momentarily
leans the fuel:air mixture and
properly positions the flame
front in the burner cans. Count
as another TEB shot.
9. Ground starting unit - Signal ground
crew for starter OFF at 3200-3300
rpm.
10. Idle rpm - Check 3550-3650 rpm.
NOTE
Idle rpm increases 50 rpm per �C
above 32�C (90 F).
11. Engine and hydraulic pressure instru-
ments - Check normal.
a. Fuel flow - Check (approximately
3300 pounds per hour).
b. EGT - Check (350�-540�C).
c. Oil pressure indicator - Check.
Discontinue start if oil pressure
rise is not observed within 60
seconds from start of rotation.
d. Hydraulic system pressures -
Check.
12. UHF switch - BOTH.
13. Start other engine using above proce-
dure.
14. TEB counter - Check.
If throttle is inadvertently retarded
to OFF do not advance in an attempt
to restart engine. In case of false
start use engine clearing procedures,
this section. Afterburner duct
must be -visually checked and un-
burned fuel removed prior to at-
tempting another start.
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2-7
Approved for Release: 2017/07/25 C00821248
SECTION II
A - 12
TURNING DIAGRAM
NOTE: 151.9 FT MINIMUM RUNWAY WIDTH REQUIRED FOR 180-DEGREE TURN
(MAIN GEAR WHEELS ON EDGE OF RUNWAY AT START OF TURN ).
REV 12-21-611
FR-61(*)
2-8
Figure 2-2
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CLEARING ENGINE
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A-12
7.
When a false start occurs, trapped fuel and
fuel vapor may be removed from engine by
using the following procedure:
1. Throttle - OFF.
2. Ground starting unit - ON for approxi-
mately 1 minute. Then signal ground
crew for ground starting unit - OFF.
Do not rotate the engine with fuel
shut off (Emergency Fuel Shutoff
switch - UP, Guard up) except in
case of emergency, because damage
to the engine may result.
BEFORE TAXIING
1. UHF and IFF/SIF - Check.
2. IFF - As required.
3. Generator switches - RESET (mo-
mentary) at idle rpm. Check with INS
crew prior to resetting.
4. Battery switch - BAT (within 3 sec-
onds).
5. Generator out lights - Check Off.
NOTE
If the generator out warning
lights fail to extinguish, return
the battery switch to the EXT
PWR position and repeat steps
3 and 4 above.
6. INS DEST/FIX switch - VARIABLE
DEST.
SECTION II
INS mode switch - NAV. Check with
INS crew prior to actuating switch.
Press the STORE button and check
BDHI No. 2 steering needle for 100
right indication and Distance To Go in-
dicator for 122 nautical mile readout.
8. INS indications - Report Destination
Coordinates, Distance To Go and
Groundspeed when slewing is completed.
9. INS DEST/FIX switch - Select VARI-
ABLE FIX and press STORE button.
Check INS FIX REJECT light on.
10. INS DEST/FIX switch - Select VARI-
ABLE DEST and press STORE button.
Check INS FIX REJECT light off.
11. INS umbilical cord - Check discon-
nected (confirmed by INS crew).
12. External power - Signal for disconnect.
13. Inlet air forward bypass - Check open.
Ground crew will confirm open.
14. HF radio - ON.
15. SAS channel switches - All ON.
16. SAS i�ecycle lights - Press (all lights
should go out).
17. SAS light test switch - Press (all lights
should illuminate).
18. Autopilot pitch and roll engage switches-
ON.
19. Autopilot disengage switch (control
stick)- Press. Check that autopilot
disengages.
20. SAS channel switches - OFF. Pitch and
yaw A and B and Roll disengage lights
illuminate. Both MON lights must stay
out.
2-9
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Approved for Release: 2017/07/25 C00821248
SECTION II
A-12
21. Surface trim - Check for proper oper-
ation with ground crew and set to zero.
22. Control system - Check for proper di-
rection of movement. Individually
check each axis in both directions and
have ground personnel verify proper
deflection of control surfaces.
23. Package switches - As required.
24. Canopy and seat safety pins - Remove
and stow.
25, Canopy - Close and lock.
26. Canopy seal pressure lever - ON.
The canopy should be opened or
closed only when the aircraft is
completely stopped. Maximum
taxi speed with the canopy open is
approximately 40 knots. Gust or
severe wind conditions should be
considered as a portion of the
40 knot limit taxi speed.
27. Rear view periscope - Check.
28. Taxi clearance - Obtain clearance
from control tower.
29. Chocks and downlock pins - Signal for
removal. Observe ground crew for
clearance to taxi.
30. Nosewheel steering - Engage and check
operation.
TAXIING
1. Brakes - Check.
WARNING I
Do not switch to alternate brakes
with both L & R hydraulic systems
operative.
2. Flight instruments - Check.
3. Navigation equipment - Check operation
of ADF, TACAN, and INS.
All taxiing and turns should be ac-
complished at slow speeds so as
to limit side loads on the landing
gear. Fast taxiing should also be
avoided to prevent excessive brake
and tire heating and wear.
BEFORE TAKEOFF
1. Engine trim - As required.
NOTE
If engine trim run is required, EGT
values appropriate for ambient
temperature will be supplied during
preparation for flight.
During trim run at Military rpm:
2. Cockpit and Q-4pay auto temp controls-
Adjust if necessary.
NOTE
Adjust both controls toward in-
creasing temperature positions if
necessary, to eliminate cockpit
fog if fog is encountered at lower
temperature settings. 12:00 to
1:00 o'clock settings are normaLly
sufficient. Lower temperature
settings are desirable when local
humidity and ambient temperature
conditions permit, in order to
assure personal and equipment
cooling.
2-10
11=MEMMEMIIIMMIMEM=MINMINApproved for Release: 2017/07/25 C00821248
3.
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A-12
SAS channel switches - All ON.
C00821248
SECTION II
TAKEOFF
4.
SAS recycle lights - Press, if necessary
1.
Brakes - Hold.
(lights should go out).
2.
Nosewheel steering - Engaged.
5.
Surface trim indicators - Check for
zero setting.
3.
Throttles - Advance.
6. Tanks 1, 2 and 6 - Check ON.
7. INS - Check and fix as required. At
designated runway position, select cor-
rect STORED FIX position and fix.
Check INS FIX REJECT light off.
Select STORED MAN. Reset DEST/FIX
briefed initial destination position, and
store. Check distance to go after slew-
ing completed, then reset DEST FIX to
STORED AUTO if desired.
8. Compasses - Check. Check and syn-
chronize FRS and check INS if appli-
cable. Return INS mode selector switch
to desired position. Check Standby
Compass against runway heading.
9. Pitot heat switch - ON.
10. Warning lights - All Off.
11. External lights switch - BCN (if re-
quired).
12. Shoulder harness - Lock.
13. Flight controls - Cycle and check hy-
draulic pressures.
14. Suit vent boost lever - NORM.
15. Birdwatcher power switch - ON and
checked.
16. Fuel derich arming switch - ARM.
17. Elapsed time clock - Start.
Engine turbine life can be ap-
preciably decreased by too rapid
throttle movement. The time
for throttle advancement from
IDLE to MILITARY should be no
less than one second.
4. Brakes - Release at 6000 rpm.
The tires may skid if the brakes
are held on at high thrust.
5. Engine instruments - Check at MILI-
TARY thrust.
a. Tachometer.
b. Nozzle Position.
c. Oil Pressure.
6. Throttles - Advance to afterburner mid-
range position after engines reach
MILITARY rpm.
WARNING II
To prevent overspeed, afterburner
ignition must not be accomplished
before the engines reach MILITARY
rpm.
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2-11
SECTION II
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A-1 c
TAKEOFF
-
NOTE
ENGINE INSTRUMENT CHECKS SHOULD
BE MADE DURING THE INITIAL PORTION
OF TAKEOFF ROLL
THE TIRES MAY SKID WITH THE BRAKES
ON AT HIGH ENGINE THRUST
CONTINUE ROTATION TO
ASSUME TAKEOFF ATTITUDE AT
TAKEOFF SPEED.
BEGIN ROTATION AT COMPUTED
SPEED.
ACCELERATION-CHECK
USE NOSEWHEEL STEERING AS NECESSARY
FOR DIRECTIONAL CONTROL
ENGINE INSTRUMENTS - RECHECK
THROTTLES - ADVANCE TO MAX.
AFTERBURNER AFTER IGNITION.
THROTTLES - ADVANCE TO MID
AFTERBURNER WHEN AT
MILITARY RPM.
ENGINE INSTRUMENT - CHECK
THROTTLES - ADVANCE TO MILITARY
BRAKES - RELEASE AT 6000 RPM
THROTTLES - ADVANCE
NOS EWHEEL STEERING - ENGAGE
BRAKES - HOLD
Figure 2-3
Fno-4(e)
2-12
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Approved for Release: 2017/07/25 C00821248
A- 12
SECTION II
NOTE
. Afterburner ignition should occur
within 3 seconds.
. Abort the takeoff if one or both
afterburners do not ignite.
Advancing the power lever to initiate
afterburning results in momentary
nozzle excursion, and engine transient
speed oscillation may approach 250
rpm.
7. Throttles - Advance to MAXIMUM
THRUST.
The time for throttle advancement
should be no less than one second.
8. Engine instruments - Recheck at MAX-
IMUM THRUST.
NOTE
Exact readouts on these instru-
ments is time consuming. The
readout should be anticipated and
needle position checked against a
clock position. If there is any in-
dication of improper engine per-
formance during power advance-
ment, the takeoff should be aborted.
Monitor ground run distance and
airspeed during the takeoff roll. If
possible, any abort decision should
be made before the aircraft has
reached high groundspeed. Direc-
tional control can be maintained
with nosewheel steering up to nose-
wheel lift off speed.
9. Acceleration - Check indicated air-
speed against computed acceleration
check speed at selected acceleration
check distance. Refer to performance
data, Appendix I, for takeoff infor-
mation.
10. Rotation - Begin at computed airspeed
approximately five seconds before
reaching takeoff speed. Apply smooth,
constant back pressure on the stick so
that required stick deflection and rota-
tion to takeoff attitude occurs at take-
off speed. Refer to Appendix I for ro-
tation and takeoff speeds.
NOTE
Use indicated airspeed during
takeoff and climb until proper
climb schedule speed is reached
on the triple display indicator.
CROSSWIND TAKEOFF
During crosswind takeoffs the aircraft tends
to weather vane into the wind. This will be
noted when the nosewheel lifts off and nose-
wheel steering is no longer available. Rud-
der preseure must be held to counteract the
crosswind effect. A definite correction
must be made as the aircraft breaks ground.
Apply lateral control as necessary for wings
level flight. Both the directional and lateral
control applications are normal and no pro-
blems should be encountered when taking off
during reasonable crosswind conditions.
ROTATION TECHNIQUE
During takeoff, the maximum load on the
main wheel tires occurs during rotation to
takeoff attitude.
ImmommmimmimmimEMMEIMmmApproved for Release: 2017/07/25 C00821248
2-13
SECTION II
Approved for Release: 2017/07/25 C00821248
A-12
CLIMB SPEED SCHEDULES
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Figure 5-2
During Cruise,
Temperature May. be
Trimmed Within
This Band
5-4
Approved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
A-12 SEC TION V
INSTRUMENT MARKINGS
FUEL TANK PRESSURE GAGE
CODE
t.TIT
YELLOW
RED
GREEN
NOTE
LIMIT VAWE DENOTED BY EDGE OF RED
LINE SO THAT INDICATION WITHIN MARKED
RED RANGE EXCEEDS LIMIT VAWE
COCKPIT TEMPERATURE INDICATOR
COMPRESSOR INLET STATIC PRESSURE GAGE
HYDRAULIC SYSTEM PRESSURE GAGES
(A AND B - LAND R)
LIQUID NITROGEN GAGE
Figure 5-1 (Sheet 2 of 2)
4-20-66
F200-43(2)(d)
Approved for Release: 2017/07/25 C00821248
5-3
SECTION V
Approved for Release: 2017/07/25 C00821248
J.
ENGINE OPERATING LIMITS SUMMARY
Fuel: PWA 523E
Additive: PS 1-67A
100 pp.m by weight
Oil: PWA 52I8
� ,,,...�.1t/MX,IMUM ALLOWABLE STEADY
STATE ROTOR. SPEED
Rol: Op. Inst.
Dtd. 5-20 66
4;i-4 Tri4
0 100 200 300
COMPRESSOR INLET TEMPERATURE-PC
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5-4
pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
A-11. SECTION V
TIME LIMITS
YJT11D-20A and yj 1 engines may be operated continuously at all ratings
when viithin the nor/nal exhaust gas temperature limits; however, no more
than one hour may be accumulated with F,�;T in excess of the normal limit
schedule, and E:7T inn st bo reduced inlirlediatelv if an emergency limit tem-
perature is exceeded. (See EC;T Liniit. and figure 5-2.)
CAUTION
Continuous or accumulated operating time in the
emer(,ency 1-.;l operating zone for more than 15
minutes mav require engine removal.
EXI AUST CAS TEMPERATURE LIMITS
The nominal operating Land, normal limits and emergency exhaust gas
temperature Operating schedules are prescribed as a function of compressor
inlet temperature as shown in figure 5-2, Limit EC:;T's for continuous op-
eration are 805oC when conressor inlet temperature is above 600C, and
845�C when CIT is below 60 C. The setting at which the red warning light
on the ET gage illun-iinates and the fuel derichment system operates, if
armed, is 860 C, a val-te �xhich is above the normal operating temperature
limit schedule.
Note
At compressor inlet temperatures below 5�C, the
possibility of engine stall exists at EGT's between
the maximum permissible value and the nominal
operating band.
In the event that emergency engine operation is required, EGT maybe in- .
,
creased to 825'C when above 60oC Gil:, or to 865oC when below 60 C CIT;
however, an accurate accounting of operating time in the emergency op-
erating zone must be maintained.
Note
. Any operation in or above the emergency operating
zone requires special maintenance action.
. The permissible emergency EGT level at low CIT's -
is above the derich system actuation point; therefore,
the derich system must be disarmed if this level is to
be attained.
Page 2 of 2
TDC No. 4A
4 March 1968
5-4A
pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
SECTION V
A-12
TIME LIMITS
YJT11D-20A and YJ-1 engines may be oper-
ated continuously at all ratings when within
the normal exhaust gas temperature limits;
however, no more than one hour may be ac-
cumulated with EGT in excess of the normal
limit schedule, and EGT must be reduced
immediately if an emergency limit temper-
ature is exceeded. (See EGT Limits and
figure 5-2.)
Continuous or accumulated oper-
ating time in the emergency EGT
operating zone for more than 15
minutes may require engine re-
moval.
EXHAUST GAS TEMPERATURE
The nominal operating band, normal limits
and emergency exhaust gas temperature
operating schedules are prescribed as a
function of compressor inlet temperature as
shown in figure 5-2. Limit EGT's for con-
tinuous operation are 805oC whe%compres-
sorinlet temperature is above 60 C, and
845�C when CIT is below 60oC. The setting
at which the red warning light on the EGT
gage illuminates and the fuel derichment
system operates, if armed, is 860oC, a
value which is above the normal operating
temperature limit schedule.
NOTE
At comuessor inlet temperatures
below 5 C, the possibility of en-
gine stall exists at EGT's between
the maximum permissible value
and the nominal operating band.
In the event that emergency engine operation
is required, EGT may be increased to 825oC
when above 60�C CIT, or to 865�C when be-
low 60oC CIT: however, an accurate account-
ing of operating time in the emergency oper-
ating zone must be maintained.
NOTE
. Any operation in or above the
emergency operating zone re-
quires special maintenance
action.
The permissible emergency
EGT level at low CIT s is
above the derich system ac-
tuation point; therefore, the
derich system must be dis-
armed if this level is to be
attained.
COMPRESSOR INLET TEMPERATURE
The maximum allowable compressor inlet
temperature is 427 C. In addition, decel-
eration must be monitored so that engine
cooling rates will not be excessive. While
above an airspeed of Mach 1.8, the aircraft
maximum rate of descent should be such
that rate of deceleration does not exceed 1.0
Mach in three minutes. There is no limit-
ation on rate of deceleration while below
Mach 1.8.
COMPRESSOR INLET PRESSURE
The minimum pressure recommended for
airstarts from stabilized windmilling speeds
is 7 psi. This pressure is marked by a
green radial line.
ENGINE SPEED
Military and afterburning engine speeds are
the same and are automatically scheduled by
the fuel control as a function of Compressor
Inlet Temperature. The normal schedule is
shown by figure 5-2. Engine overspeed
above 7450 rpm requires a visual inspection
of the turbine. Notify the engine manufac-
turer if 7550 rpm is ever exceeded. Each
instance of overspeeding should be reported
as an engine discrepancy and should include
the maximum rpm attained.
Changed 15 March 1968
Approved for Release: 2017/07/25 C00821248
5-5
SECTION V
Approved for Release: 2017/07/25 C00821248
LIMIT FLIGHT SPEED AND ALTITUDE ENVELOPE
100
ALTITUDE - 1000 FT
MEIMMEMEM 11,0050,,PC1.401
MEEME
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NORMAL OPERATING CRUISE SPEED �
70-
50i
40
30
20
10
:.;
NOTE: ABOVE 50,000 Fl, MINIMUM AIRSPEED IS 300 KEAS.
MAXIMUM ALTITUDE RESTRICTION:
WITH DERICHMENT - 85,000 FT
idatiedrojatcloogi
.0".tmuerdimaro."1101
...,ww.sarloporponot,,�iitypipagini
'411
��
WITHOUT AUTOMATIC AUTOMATIC INLET OPERATION � 80,000 FT-
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pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
A-12
APPENDIX I
PART III
PENETRATION DISTANCE-NAUTICAL MILES
70
0 10 20 30 40 50 60 70
PENETRATION DISTANCE-NAUTICAL MILES
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SINGLE ENGINE TURNING DESCENT
MAXIMUM AB-350 KEAS-80,000 FT. TO 50,000 FT.
35 DEG. BANK-180 DEG. TURN
90,000 LB. INITIAL GROSS WEIGHT
NOTE: No service allowance included.
�END TURN,
40 50 60 0 1 2
FUEL USED
- 1000 LB.
Figure A3-24
INLET CONFIGURATION
ENGINE
SPIKE
FORWARD BYPASS
AFT BYPASS
OPERATING
AUTO
OPEN INITIALLY
CLOSED AT MACH 1.0
CLOSED
SHUT
DOWN
MANUAL
FORWARD
OPEN
OPEN
Changed 15 March 1968
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A3-37/A3-38
Approved for Release: 2017/07/25 C00821248
A-12
PART IV
SUBSONIC CRUISE PERFORMANCE
List of Illustrations
APPENDIX I
PART IV
Title Figure No.
TWO ENGINE OPERATION
Subsonic Long Range Cruise A4-1
Subsonic Maximum Range Cruise Climb - Mach 0.88 A4-2
Maximum Subsonic Specific Range Summary A4-3
Subsonic Range Factor Summary A4-4
Buddy Mission Cruise - Mach 0.77 and 28,000 ft A4-5
Subsonic Specific Range - Mach 0.77 A4-6
Loiter Performance A4-.7
Specific Range - 10,000 ft A4-8
Specific Range - 15,000 ft A4-9
Specific Range - 20,000 ft A4-10
Specific Range - 22,000 ft A4-11
Specific Range - 24,000 ft A4-12
Specific Range - 26,000 ft A4-13
Specific Range - 28,000 ft A4-14
Specific Range - 30,000 ft A4-15
Specific Range - 32,000 ft A4-16
Specific Range - 34,000 ft A4-17
Specific Range - 36,000 ft A4-18
Specific Range - 38,000 ft A4-19
Specific Range - 40,000 ft A4-20
SINGLE ENGINE OPERATION
Long Range Cruise - Afterburner Operation A4-21
Long Range Cruise - Military Thrust A4-22
Specific Range - Military Thrust A4-23
Single Engine Cruise Tabulation - Afterburner aL Military A4-24
INTRODUCTION
This part of the appendix supplies two engine
cruise and loiter performance date. and
single engine cruise performance data. The
material for two engine operation includes
a long range cruise chart, maximum spe-
cific range summaries for long range
cruise-climb and KC-135 buddy missions,
loiter performance, and specific range
charts for altitudes from 10,000 feet to
40,000 feet. The single engine data show
cruise climb range capability with and with-
out afterburner, and a specific range chart
for operation at Military thrust.
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A4-1
APPENDIX I Approved for Release: 2017/07/25 C00821248
PART IV
A-12
TWO ENGINE OPERATION
The two engine performance data applies to
operation with YJ or YJ-1 engines when
aircraft c. g. is at 25% MAC. Operation at
more forward c. g. conditions reduces spe-
cific range 1% for each one percent shift in
c. g., as noted on the specific range charts.
LONG RANGE CRUISE SUMMARY
Figure A4-1 presents the constant altitude,
maximum range cruise climb, and Military
thrust cruise climb capability of the air-
craft in terms of distance to go to 65,000 lbs
gross weight (approximately 10,000 lbs fuel
remaining). The additional distance avail-
able to lower gross weights is also provided.
Cruise speeds for constant altitude cruise
are tabulated on the chart. The chart can
be used on an incremental basis for any de-
sired start and end cruise condition.
Example:
Determine the range available at 25,000
feet, 30,000 feet, and by cruise climbing
with an initial gross weight of 120,000 lb if
cruise is to be terminated at 10,000 lbs fuel
remaining (approximately 65,000 lbs gross
weight). Figure A4-1 shows that by cruising
at 25,000 feet the range will be 1700 nmi.
This range increases to 1810 nmi by cruis-
ing at 30,000 feet. Maximum range is avail-
able by cruise climbing at 0.88 Mach number.
Under this condition cruise would be initiated
at 29,400 feet and ended at 41,900 feet at
10,000 lbs fuel remaining. Distance tra-
veled would be 1900 nmi.
MAXIMUM RANGE CRUISE CLIMB
Figure A4-2 presents the distance available
to 65,000 lbs gross weight (approximately
10,000 lb fuel remaining) for maximum
range cruise climb at 0.88 Mach number and
382,000 lb W/S . The chart can be used on
an incremental basis for any desired start
and end cruise condition.
A4-2
MAXIMUM SUBSONIC SPECIFIC RANGE
SUMMARY
Figure A4-3 presents the maximum specific
range summary for cruise climb at various
Mach numbers. Note that the optimum
cruise climb occurs at Mach 0.88. This
summary is obtained from the subsonic
range factor chart, figure A4-4, by the
equations Range Factor (instantaneous) =
Specific Range (instantaneous) x W (instan-
taneous) and & (and its corresponding
pressure altitude) = W/W/S . (Refer to
section on equations).
RANGE FACTOR
Figure A4-4 presents the subsonic range
factor for long range cruise climb at any
Mach number. The chart shows there is a
range factor and corresponding cruise
climb schedule (W/S ) for a given cruise
Mach number. This provides a quick means
for calculating best range available for any
given cruise Mach. The chart also shows
that the optimum range factor (3100,1b-nmi/1b)
occurs at Mach 0.88 and the corresponding
cruise climb schedule (w/6 ) is 382,000 lb.
Definition of Terms
W/S = Weight/pressure ratio, lb
W = Aircraft gross weight, lb
= Pressure ratio, P/Po, for the flight
pressure altitude (figure A1-8)
WF = Total fuel flow, lb per hour
KTAS = True airspeed, knots
Ln = Natural logarithm
Equations
Distance flown nmi
Specific Range (avg) -
Fuel Used lb
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A-12
APPENDIX I
PART IV
KTAS nmi
Specific Range (instantaneous) -
WF lb
Range Available = Specific Range (avg) x
Fuel Used, mill
Range Factor (avg) = Specific Range (avg) x
nmi
W (avg), lbx lb
Range Factor (instantaneous) = Specific
Range (instantaneous) x W (instantaneous),
nmi
x lb, or
lb
KTAS
x W)
Range Factor (avg) x
Range Available =
W (avg)
Fuel Used, nrni
Range Available = Range Factor (avg) x In
(initial W
)
final W , urn!
Distance flown
or Range Factor (avg) -
in initial WI
final W
whs
Example (1):
Determine the range available and the cruise
climb schedule for cruise at 0.80 Mach.
(Note that this is not the optimum cruise
speed.) The initial cruising weight is
100,000 lb, and 20,000 lb of fuel are to be
used. Assume a standard day with zero
wind.
a. Average gross weight is 90,000 lb.
b. From figure A4-4, at Mach 0.80, the
cruise climb schedule (W/S ) is
275,000 lb and the range factor is 2915
lb -
c. The range available = (2915 x 20,000/
90,000) = 648 nmi.
d. The initial pressure ratio, S , =
(100,000/275,000) = 0.3636.
The final pressure ratio, S , =
(80,000/275,000) = 0.2929.
e. Enter the standard atmosphere table,
figure A1-8, with the initial and final
pressure ratios, and determine the ap-
proximate initial and final cruise alti-
tudes as 25,500 ft and 30,500 feet, re-
spectively.
Example (2):
Determine the cruise fuel required and
cruise climb schedule for cruise at 0.75
Mach. The planned cruise distance is 650
nrni. Assume a standard day with zero
wind. Planned initial cruise gross weight
is 100,000 lb.
a. From figure A4-4, at Mach 0.75, the
cruise climb schedule (W/S ) is
227,000 lb and the range factor is
2730 lb - nmi/lb.
W (initial;
b. From section on equations, ln
W
Distance
in 100'000 650
Range Factor W (final) = 2730
100,000
or 0.2380; = 1.269; W (final)
W (final)
= (100,000/1.269) = 78,800 lb.
Therefore, cruise fuel required =
(100,000 - 78,800) = 21,200 lb.
c. Using the same method as in the pre-
vious example, the approximate initial
and final cruise altitudes are 21,000
feet and 26,500 feet, respectively.
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APPENDIX I
PART IV
A-12
BUDDY MISSION CRUISE
Figure A4-5 presents the distance available
to 65,000 lbs gross weight (approximately
10,000 lb fuel remaining) for Buddy Mission
cruise at Mach 0.77 and 28,000 feet. The
speed and altitude schedule is compatible
with KC-135 tanker performance charac-
teristics. The chart can be used on an in-
cremental basis for any desired start and
end cruise condition.
SPECIFIC RANGE - MACH 0.77
Figure A4-6 presents specific range data at
Mach 0.77. The Buddy Mission altitude is
listed on the chart. If desired, greater
range is obtained by cruise climbing.
LOITER PERFORMANCE
Figure A4-7 presents loiter performance as
minutes per 1000 lb of fuel used. The re-
commended speed schedule is listed in the
chart.
Example:
Determine the loiter time available at
20,000 feet for an initial gross weight of
70,000 lb. A planned 10,000 lb of fuel is to
be consumed. Enter figure A4-7 at 70,000
lbs gross weight and 20,000 feet and read
5.09 minutes per 1000 lb of fuel. Reenter
at 60,000 lbs and 20,000 feet and read 5.62
minutes per 1000 lb of fuel. The average
value is 5.35 minutes per 1000 lb of fuel.
This provides 53.5 minutes for the planned
10,000 lbs of fuel consumption.
SPECIFIC RANGE - CONSTANT ALTITUDE
The specific range charts (figures A4-8 thru
A4-15) present cruise data for various con-
stant altitudes (from 10,000 ft to 40,000 ft)
throughout the speed range from maximum
endurance to Military thrust. Each chart
presents nautical miles per 1000 lb of fuel
(nmi/K1b) as a function of Mach number and
gross weight with subs cales of KEAS and
KTAS for standard day. Also included are
an overlay grid of fuel flow per engine, the
maximum range speed schedule, and the
recommended loiter speed schedule.
SINGLE ENGINE OPERATION
The single engine performance data applies
to operation with YJ engines. A five per-
cent service allowance is included. Refer
to text for other items affecting the per-
formance results. The long range cruise
data for both Military and Afterburner op-
eration can be used in conjunction with the
single engine descent information in Part
III. Transition from end of descent (as in-
dicated in the single engine descent curves)
to start of single engine cruise is accom-
plished by drift down. Duration of drift
down is indeterminate and is largely de-
pendent on piloting technique. Drift down
consists of a slow sink period during which
fuel economy is above the corresponding
cruise values for the same weight as long
as the actual altitude is above the scheduled
cruise altitude. The difference in miles
per pound can be neglected in planning and
provides an operational contingency pad.
Refer to Section III for fuel management
during single engine cruise.
LONG RANGE CRUISE - AFTERBURNER
OPERATION
Figure A4-21 presents single engine long
range cruise performance for afterburner
operation in terms of distance to go to
60,000 lbs gross weight (approximately
5000 lbs fuel remaining). The chart is
based on zero wind distance without turns
at test day conditions. Test Eq. was
trimmed between 780�C and 810 C for CIT
range of -200C to +2.00C. The long range
A4-4
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A-12
APPENDIX I
PART V
PART V
SUPERSONIC CRUISE PERFORMANCE
List of Illustrations
Title Figure No.
Turning Performance A5-1
Mach 3.20 1
Specific Range, Ambient Temp. -66.5oC A5-2
II II III II -56.5�C A5-3
if if' It -53.0�C A5-4
if, II -43.0oC A5-5
Long Range Cruise 1956 ARDC Atmosphere
Fuel and Time Priofile qSheet 1 of 3 . A5-6
Climb - Cruise Intercept Points 'Sheet 2 of 3 .
I
Cruise Performance Sheet 3 of 3 .
High Altitude Cruise,- 1956 ARDC Atmosphere
Fuel and Time Profile .Sheet 1 of 3 . A5-7
I � i
Climb - Cruise Intercept Points I Sheet 2 of 3 .
Cruise Performance i Sheet 3 of 3 .
Maximum A/B Ceilirig Cruise Profile Sheet 1 of 2 . A5-8
(With STD DAY climb) .1Sheet 2 of 2 .
Long Range Cruise - MEAN TROPIC Atmosphere
Fuel and Time Profile ..Sheet 1 of 3 . A5-9
I
Climb - Cruise Intercept Points Sheet 2 of 3 .
Cruise Performance ,Sheet 3 of 3 .
High Altitude Cruise - MEAN TROPIC Atmosphere
Fuel and Time Profile .,Sheet 1 of 3 . A5-10
Climb - Cruise Intercept Points 1Sheet 2 of 3 .
Cruise Performanice iSheet 3 of 3 .
_
Maximum A/B Ceiling Cruise Profile Sheet 1 of 2 . A5-11
(With MEAN TROPIC climb) Sheet 2 of 2.
Mach 3.10
Specific Range, Ambient Temp.
1 -64.7�C A5-12
II II If It -56.5�C A5-13
1
If II II II -53.0oC A5-14
II II It II -43.5�C A5-15
Long Range Cruise - 1956 ARDC Atmosphere
Fuel and Time Profile Sheet 1 of 3 . A5-16
Climb - Cruise Intercept Points Sheet 2 of 3 .
Cruise Performance Sheet 3 of 3 .
High Altitude Cruise - 1956 ARDC Atmosphere
Fuel and Time Profile Sheet 1 of 3 . A5-17
Climb - Cruise Intercept Points Sheet 2 of 3 .
Cruise Performance Sheet 3 of 3 .
Maximum A/B Ceiling Cruise Profile Sheet 1 of 2 . A5-18
(With STD DAY climb) Sheet 2 of 2
Long Range Cruise - MEAN TROPIC Atmosphere
Fuel and Time Profile Sheet 1 of 3. A5-19
Climb - Cruise Intercept Points Sheet 2 of 3.
Cruise Performance Sheet 3 of 3 .
Changed 15 June 1968
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A5-1
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APPENDIX I
PART V A-12
List of' Illustrations (Con't)
Title
High Altitude Cruise - MEAN TROPIC Atmosphere
Fuel and Time Profile
Climb - Cruise Intercept Points
Cruise Performance
Maximum A/B Ceiling Cruise 131ofile
(With MEAN TROPIC climb) I
Mach 2.90
Specific Range, Ambient Temp.
II It It II
It
Figure No.
Sheet 1 of 3 . A5-20
Sheet 2 of 3 .
Sheet 3 of 3 .
Sheet 1 of 2 . A5-21
Sheet 2 of 2 .
-66.0�C A5-22
-56.5�C A5-23
-53.0�C A5-24
-42.5oC A5-25
Long Range Cruise - 1956 ARDC Atmosphere
Fuel and Time Profile
Climb - Cruise Intercept Points
Cruise Performance
Long Range Cruise - MEAN TROPIC Atmosphere
Fuel and Time Profile Sheet 1 of 3
Climb - Cruise Intercept Points Sheet 2 of 2
Cruise Performance Sheet 3 of 3
Performance Mission Planning FactIlors for Supersonic
Rapid Deployment to ARCP - 1956 ARDC Atmosphere
Sheet 1 of 3 . A5-26
Sheet 2 of 3 .
Sheet 3 of 3 .
A5-27
Cruise A5-28
A5-29
Profile of Rapid Deployment to ARC Sheet
(1956 ARDC Atmosphere) Sheet
Rapid Deployment to ARCP - MEAN TROPIC Atmosphere
Profile of Rapid Deployment to ARCP Sheet
(MEAN TROPIC Atmosphere) Sheet 2 of 2
1 of 2 . A5-30
2 of 2
TURNING PERFORMANCE
Figure A5-1 presents generalized turning
performance at constant Mach numbers for
various ambient temperatures and bank
angles. Turn radius, distance, and time
are plotted for a selected range of ivfach
numbers, ambient temperatures, bank
angles, and degrees of turn.
Example:
A5-31
1 of 2 A5-32
For a Mach 3.00 turnoat a forecast ambient
temperature of -56.5 C, 30o bank angle,
and a planned 180o of turn, find the turn
radius, distance, and time. As shown in
the chart,0 enter figure A5-1 at Mach 3.00
and -56.5 C ambient temperature and note
that true airspeedois 1720 knots. Proceed
horizontally to 30 bank angle and read turn
radius as 74.5 nautical miles. Proceed
downward to 1800 of turn and read turn dis-
tance as 235 nautical miles flown. Proceed
horizontally to 1720 KTAS and read the turn
time as 8.1 minutes.
A5-2
Changed 15 June 1968
�NImMINIM=INIIMINEMApproved for Release: 2017/07/25 C00821248
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APPENDIX I
A-12 PART V
SPECIFIC RANGE
Specific range charts are presented for
speeds of Mach 3.20, 3.10,1 and 2.90 and for
four ambient temperature Conditions at each
speed as shown by the list of illustrations.
The data is computed from Flight Test and
I Operational Testing results with YS-1
engines. Corrections for al range of bank
angles are included on each chart to show
the effect bank angle has oi specific range
and altitude capability whi1e turning. Sup-
plemental scales provide KEAS-altitude in-
formation and fuel flow conversions.
Example:
Refer to figure A5-13, Specific Range data
o
for Mach 3.10 cruise at -516.5 C ambient
temperature. Locate the Max Range cruise
schedule line. At long range cruise power
and 80,000 pounds gross weight the cruise
climb altitude is 78,150 feet and the zero
bank angle specific range is 61.0 nmi/1000
lb of fuel. For a turn at the same power
setting, using a 30 degree bank angle, the
specific range is 53.0 nmi/1000 lb of fuel
and the altitude is 75,100 fleet. The fuel
flow per engine is 14,600 lb/hr at zero bank
and 16,800 lb/hr at 30 degi�ee bank for a
-56.5 C ambient temperatiire day. At this
temperature, Mach 3.1 corresponds to 1777
KTAS as listed in the chart.
LONG RANGE AND HIGH ALTITUDE CRUISE
SUMMARIES
Long range cruise surrunar'les are presented
for Mach 3.20, 3.10, and 290. High altitude
cruise summaries are presented for Mach
3.20 and 3.10. The high altitude profiles
are based on the "90%" lines shown on the
Specific Range charts, except that the per-
formance shown conforms with the present
85,000 ft altitude restricticip. These data
are presented for both the :1956 ARDC At-
mosphere and the "MEAN TROPIC" Atmo-
sphere as shown in the list of illustrations.
The climb and cruise data are computed
from Flight Test and Operational Testing
results with YJ-1 engines. 1 Descent data is
based on Flight Test and OPerational testing
Changed 15 June 1968
at near standard temperatures. There are
three sheets for each figure. The first
sheet provides cruise summaries showing
distance and time from end AR at 30,000
feet through the climb, cruise, and descent
to 20,000 feet with either 5000 lbs or 7500
lbs of fuel reserve. The second sheet pre-
sents climb-cruise intercepts which are to
be used in conjunction with sheet 3. The
third sheet presents performance and flight
planning data. The initial conditions shown
are end AR at 30,000 feet, and brake re-
lease with either 64,000 lbs or 50,000 lbs
fuel remaining using the normal climb
schedule. The effect of various temper-
atures is shown for climb and cruise per-
formance. The descent performance shown
is based on operational testing and does not
include the effect of temperature. Descent
through a "Tropic" atmosphere may be ap-
proximated by increasing the presented de-
scent data by the following increments:
Distance - 30 miles
Time - 1 minute
Fuel used - 100 pounds
Use of the chart is illustrated by the follow-
ing example:
Example:
Refer to figure A5-7, sheet 2 of 3 and
sheet 3 of 3.
Find the total distance capability and time
required for a Mach 3.2 high altitude cruise
with a forecast ambient temperature condi-
tion of -56.5 C at cruise. A profile is
planned consisting of a heavyweight takeoff
at sea level with standard day climb, cruise
without turn, normal descent, and 7500 lb
fuel reserve at 20,000 feet. Planned fuel
load at brake release is 64,000 lb.
Enter figure A5-7, sheet 2 of 3, at 119,150
lb gross weight, sea level altitude6 standarc
day climb temperature, and -56.5 C cruise
temperature and read the cruise-climb in-
tercept as 80,100 feet. Read climb distance
as 345 miles, climb time as 20.1 minutes
A5-3
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APPENDIX I
PART V
A-12
and fuel remaining as 39,250 lb. Referring
to figure A5-7, sheet 3 of 3, the intercept
of the standard day climb line and the
-56.5�C cruise line is shown. The lower
portion of sheet 3 of 3 shows cruise distance
and cruise time to zero fuel remaining as a
function of fuel remaining and cruise re-
ference temperature. Entering the portion
of the curve at the fuel remaining value of
39,250 lb andoa cruise reference tempera-
ture of -56.5 C, read the cruise distance as
2655 miles and cruise time as 86.8 minutes.
Then read on the cruise line (from begin-
ning of the 7500 lb descent line) the fuel re-
maining as 8900 lb. Reading the distance
and time to zero fuel remaining, the dis-
tance is 740 miles and the time is 24 min-
utes. This gives the incremental cruise
distance as (2655 - 740) = 1915 miles and
the cruise time as (86.8 - 24) = 62.8 min-
U.
o 80-
LU
6�
:41 70-
U..
tn
CIL
-40-
-50-
-60-
-70-
-40--
-50--
-60--
-70-
A5-4
..6 6 .5 � C
- 5 6.5 C
_46.5*c
I I I
45 40 35
FUEL REMAINING -1000 LB
CRUISE DISTANCE-NAUTICAL MILES
CRUISE TIME-MINUTES
utes. The descent to 20,000 ft is 237 miles
and 13.8 minutes as shown by the vertical
scales at the right side of the profile portion
of the chart.
Distance and time from brake release at
sea level with 64,000 lb fuel to 20,000 feet
with 7500 lb fuel remaining is:
Distance = (345 + 1915 + 237) = 2497
miles
Time = (20:1 + 62.8 + 13.8) = 96.7
minutes
PRESSURE ALTITUDE - 1000 FT
85, 000 -
20,000 -
-40-
-50-
-60-
-70--
- 40 -
-50-
-60 -
-70-
10
7.5 5
FUEL REMAINING -1000 LB
CRUISE DISTANCE-NAUTICAL MILES
CRUISE TIME-MINUTES
Changed 15 June 1968
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A-12
APPENDIX
PART V
Figures A5-30 and A5-32 (sheet 1 of 2)
show standard and tropic day mission pro-
files for five representative Mach numbers,
and portray the climb, cruise and decel-
eration segments of the missions. Figures
A5-30 and A5-32 (sheet 2 of 2) show the
corresponding time and fuel remaining for
the presented profiles.
Figures A5-29 and A5-31 give the neces-
sary detail information for planning a flight
of specific length. These curves present
the overall mission time from brake release
to ARCP, cruise Mach number, altitude to
initiate constant Mach climb, cruise altitude
and the DTG to start deceleration to arrive
at 29,000 feet at a point 20 miles from the
ARCP. Mach 1.25 is the minimum super-
sonic cruise Mach recommended, as this
speed is the "break point" for minimum
time between subsonic and supersonic flight
plans. For a mission distance of less than
130 miles, the flight should be made at 0.91
Mach. Missions longer than 130 miles would
be flown at the Mach number given by fig-
ures A5-29 and A5-31.
Example:
To select flight plan for minimum time to
ARCP, with Mean Tropic day temperatures,
and ARCP 300 miles from takeoff point.
Refer to figure A5-31, "Rapid Deployment
to ARCP".
Mission time from brake release to ARCP
is 23.5 minutes.
Cruise Mach = 2.31.
Start constant Mach climb = 55,300 feet.
Cruise altitude = 67,000 feet.
DTG at start decel = 117 miles.
Changed 15 June 1968
A5-4C/A5-4D
IIIMMIIIM=11111�1111mApproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
A-12
APPENDIX I
PART V
MAXIMUM A/B CEILING CRUISE-SUMMARIES
Maximum A/B Ceiling Cruise summaries
are presented for Mach 3.20 and 3.10 as
shown in the list of illustrations. The data
were calculated from Flight Test and Op-
erational Testing results with YJ-1 engines.
There are two sheets for each figure. The
first sheet presents cruise summaries
showing distance and time from end AR at
32,000 feet through the climb, cruise, and
descent to 20, 000 feet with either 5000 lbs
or 7500 lbs fuel reserve. The second sheet
presents cruise summaries which are in-
dexed at 10,000 lb fuel remaining at altitude
(zero distance and time). The initial con-
ditions shown are end AR at 30,000 feet and
brake release with 64,000 lbs fuel remain-
ing using the normal climb schedule. Dis-
tance and time allowances for reserves of
5000, 7500, and 10,000 lbs at 20,000 feet
are shown in the charts. To obtain the
total distance and time, add the two dis-
tances and times for the desired profile.
Example:
Refer to figure A5-18, sheet 2 of 2, and
the example figure on the following page.
Find the total distance and time for a 3.10
Mach maximum A/B ceiling cruise at
forecast ambient temperature of -56.5�C at
cruise. A profile is planned consisting of
a heavyweight takeoff at sea level with
standard day climb, cruise without turns,
and 7500 lb reserve at 20,000 feet. Planned
fuel load at brake release is 64,000 lb.
Enter figure A5-18, sheet 2 of 2, at the
climb line for the sea level 64,000 lb fuel
remaining case and read distance and time
as 1809 nrni, and 1 hr, 09.5 min. Reenter
at the 7500 lb reserve descent line at 20,000
feet and read distance and time as 310 nmi
and 16.7 min. Add the distances and times
and obtain 2114 nmi and 1 hr, 26.2 min.
If forecast temperatures indicate standard
day climb and cold day cruise, -64.5 C,
the distance will be increased by two small
increments. The cruise distance will be
longer due to the colder temperature, and
the climb distance will be longer due to the
climb to higher altitude. Referring to the
text illustration below, which is for 119,150
lb gross weight and 64,000 lb fuel remaining
at brake release, the shaded triangles show
where the standard day climb intercepts the
four cruise lines. The cold day intercept
shows a distance of 1635 nmi. Extend the
climb curve to the altitude where the cold
day cruise begins and read a distance of
1475 nmi. The difference between these
distances (1635 - 1475 = 160) is the increase
in range due to cold day cruise conditions.
The corresponding time increment is 4.3
min. for the additional 160 nmi of cruise.
This results in a total range and time of
2279 nmi and 1 hr, 30.5 min.
MISSION PLANNING FACTORS TABLE
A Mission Planning Factors Table is pro-
vided on figure A5-28 for quick reference
in mission planning.
RAPID DEPLOYMENT TO ARCP
Figures A5-29 thru A5-32 present the data
for a minimum time profile from brake re-
lease to ARCP.
The profile is defined as:
1.
50,000 pounds fuel remaining at brake
release.
2. Normal climb schedule to cruise Mach
number.
3. Climb to cruise altitude at constant
Mach number.
4. Cruise for two minutes at 82� PLA.
5. Normal deceleration to 300 KEAS.
6. Normal 300 KEAS descent to reach
29,000 ft at a point 20 miles from
ARCP.
The data are presented for both the 1956
ARDC and Mean Tropic atmospheres.
Changed 15 June 1968
Approved for Release: 2017/07/25 C00821248
A5-4A
Approved for Release: 2017/07/25 C00821248
A- 12
EXAMPLE FIGURE
REFER TO FIGURE A 5-18 H EET '2 OF 2
AND PAGE tA5 -4 A
TIME - HR : MIN
-40-
-50 �
-60 �
-70
85-
80 �
75
55
4.3 MINUTES
40,000 LB
FUEL REMAINING
:45
-64.5�C
.53.0�C
III
1800 1700 1600 1500 1400 1300 1200 1100
A5-4B
Changed 15 June 1968
pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
SECTION V
A-12
TIME LIMITS
YJT11D-20A and YJ-1 engines may be oper-
ated continuously at all ratings when within
the normal exhaust gas temperature limits;
however, no more than one hour may be ac-
cumulated with EGT in excess of the normal
limit schedule, and EGT must be reduced
immediately if an emergency limit temper-
ature is exceeded. (See EGT Limits and
figure 5-2.)
Continuous or accumulated oper-
ating time in the emergency EGT
operating zone for more than 15
minutes may require engine re-
moval.
EXHAUST GAS TEMPERATURE
The nominal operating band, normal limits
and emergency exhaust gas temperature
operating schedules are prescribed as a
function of compressor inlet temperature as
shown in figure 5-2. Limit EGT's for con-
tinuous operation are 805oC when compres-
sor inlet temperature is above 60 C, and
845�C when CIT is below 60�C. The setting
at which the red warning light on the EGT
gage illuminates and the fuel derichment
system operates, if armed, is 860 C, a
value which is above the normal operating
temperature limit schedule.
NOTE
At comuessor inlet temperatures
below 5 C, the possibility of en-
gine stall exists at EGT's between
the maximum permissible value
and the nominal operating band.
In the event that emergency engine operation
is required, EGT may be increased to 825oC
when above 60�C CIT, or to 865�C when be-
low 60�C CIT: however, an accurate account-
ing of operating time in the emergency oper-
ating zone must be maintained.
NOTE
. Any operation in or above the
emergency operating zone re-
quires special maintenance
action.
The permissible emergency
EGT level at low CIT s is
above the derich system ac-
tuation point; therefore, the
derich system must be dis-
armed if this level is to be
attained.
COMPRESSOR INLET TEMPERATURE
The maximum allowable compressor inlet
temperature is 427 C. In addition, decel-
eration must be monitored so that engine
cooling rates will not be excessive. While
above an airspeed of Mach 1.8, the aircraft
maximum rate of descent should be such
that rate of deceleration does not exceed 1.0
Mach in three minutes. There is no limit-
ation on rate of deceleration while below
Mach 1.8.
COMPRESSOR INLET PRESSURE
The minimum pressure recommended for
airstarts from stabilized windmilling speeds
is 7 psi. This pressure is marked by a
green radial line.
ENGINE SPEED
Military and afterburning engine speeds are
the same and are automatically scheduled by
the fuel control as a function of Compressor
Inlet Temperature. The normal schedule is
shown by figure 5-2. Engine over speed
above 7450 rpm requires a visual inspection
of the turbine. Notify the engine manufac-
turer if 7550 rpm is ever exceeded. Each
instance of overspeeding should be reported
as an engine discrepancy and should include
the maximum rpm attained.
Changed 15 March 1968
Approved for Release: 2017/07/25 C00821248
5-5
SECTION V
Approved for Release: 2017/07/25 C00821248
LIMIT FLIGHT SPEED AND ALTITUDE ENVELOPE
ALTITUDE - 1000 FT
100
90
4
1.
:71Tr
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NORMAL OPERATING. CRUISE SPEED
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NOTE: ABOVE 50,000 FT , MINIMUM AIRSPEED IS 300 KEAS.
MAXIMUM ALTITUDE RESTRICTION:
WITH DERICHMENT - 85,000 FT
1-,
Ay.
MACH NUMBER
Figure 5-3
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3.5
5 - 6
Changed 15 March 1968
Approved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
SECTION V
A-12
FUEL
The approved fuel is PWA 523E. The P&W
approved source of lubricity additive, PSJ-
67A, must be mixed with the fuel in the ratio
of 0.29 gallons per 4000 gallons of fuel.
Fuels such as JP-4, JP-5, and JP-6 may be
used only for emergency requirements such
as air refueling when standard fuel is not
available and air refueling must be accom-
plished or risk loss of the aircraft. Oper-
ation with emergency fuels should be re-
stricted to speeds below Mach 1.5.
OIL
The approved oil is PWA 524B. If neces-
sary because of low ambient temperatures,
it may be diluted with Trichloroethylene,
Federal Specification O-T-634, Type 1, in
accordance with Maintenance Manual pro-
cedures.
Oil Pressure
Oil pressures below 35 psi are unsafe and
require that a landing be made as soon as
possible, using minimum thrust required
to sustain flight until a landing can be ac-
complished. Normal oil pressure is from
40 to 55 psi. Except at IDLE throttle set-
tings, oil pressures between 35 psi and 40
psi are undesirable and should be reported
after flight. A gradually increasing oil
pressure up to 60 psi is acceptable at high
Mach numbers provided the indication re-
turns to normal values after aircraft decel-
erates to subsonic speed.
Oil Temperature
Oil temperature must be at least 60�F
(15 C) prior to starting unless previously
diluted with Trichloroethylene (PWA 9003).
Engine oil temperatures above 290 C are
unsafe and a landing should be made as soon
as possible if the temperature cannot be
maintained below this value. An engine
should not be restarted after windmilling5 at
subsonic speed when CIT is less than 15 C
(60 F) for more than 5 minutes. If re-
started, operation above IDLE with OIL
TEMP warning light illuminated shall be
as brief as possible.
MAXIMUM WEIGHT LIMITS
Maximum gross weight is not limited ex-
cept by takeoff performance capabilities.
Base maximum takeoff weights on infor-
mation provided in Part II of the Appendix.
MAXIMUM ALTITUDE
Maximum altitude with derichment installed
and operational is 85,000 feet; maximum
altitude without derichment is 75,000 feet.
LIMIT AIRSPEEDS
(Refer to figure 5-3 for the limit flight
speed and altitude envelope.)
MINIMUM AIRSPEED RESTRICTION
The stall warning light on the annunciator
panel and the master caution lige illuminate
when angle of attack reaches 14 in flight.
A tone is also produced in the pilot's head-
set. When above 135 KIAS, the speed at
which stall warning occurs is the minimum
airspeed restriction for the existing vehicle
weight, c. g., and load factor unless oper-
ation is governed by a higher value of mini-
mum KEAS as displayed by the Triple Dis-
play Indicator. Minimum airspeed is 300
KEAS above 50,000 feet.
INDICATED AIRSPEED
The Mac,h-airspeed indicator limit hand is
set to indicate airspeed (KIAS) correspond-
ing to 500 KEAS. However, the 500 KEAS
limit applies only at altitudes above :)400
feet, and at airspeed below Mach 2.6. Be-
low 9400 feet, limit airspeed decreases
linearly with altitude from 500 KEAS at
9400 feet to 450 KEAS at sea level. Above
Mach 2.6, limit airspeed decreases linearly
from 500 KEAS at Mach 2.6 to 450 KEAS at
Mach 3.2. See figure 5-4 for variation of
KIAS with altitude for KEAS.
Note
Maximum recommended operating
speeds are at least 50 KEAS less
than limit airspeeds. 450 KEAS
(Mach 0.9) is not recommended
below 14,800 feet.
Changed 15 March 1968
NIMMMIIMEMNI==MEMMEllApproved for Release: 2017/07/25 C00821248
5-7
SECTION V
Approved for Release: 2017/07/25 C00821248
LIMIT AIRSPEED VS ALTITUDE
NOTE: FOR, MODIFIED ADC'S DESIGNATED DHG 72A5, DHG 72J5 AND SUBSEQUENT MODELS
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5-8
pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
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Noiiml Approved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
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Figure A3-9
(Sheet 3 M 3)
Approved for Release: 2017/07/25 C00821248
A5 -25/A5
Approved for Release: 2017/07/25 C00821248
A-12
SECTION V
MINIMUM AIRSPEED LIMITS FOR 14� ANGLE OF ATTACK
NOTE: MASTER CAUTION AND ANNUNCIATOR PANEL STALL WARNING LIGHTS
ILLUMINATE AND WARNING HORN SOUNDS WHEN 14� 01 REACHED IN FLIGHT
surarma:6-ArnIntrnEratriAtIMMU
111111=M111:
Figure 5-5
Approved for Release: 2017/07/25 C00821248
5-9
SECTION V
Approved for Release: 2017/07/25 C00821248
A- 12
LIMIT LOAD FACTOR DIAGRAM
5ymmetricnI Maneuve
120,500 LB OR LESS
........
k9U1V-�AiONTAIR4siib�KNOTSW
kft:?.wzge
135 KIAS M NIMUM AIRSPEED RESTRICTION':
SIVZOM:Ar;
14 MAXIMUM !!...N.OLE OF ATTACK LIMIT -..KEAS VARIES
" WaltariAfe%,
KEAS DESIGN HIGH SPEED AT SEA LEVEL, MACH 0.6, AND 74,000 FEET,
Ann,',AA,V0/1 ,5.'"UniaMenkelalintUffe*
4:LC
456 KEAS DESIGN LIMIT SPEED AT SEA LEVEL, MACH 0.68, AND 69,500 FEET, MACH
ALSO: DESIGN HIGH SPEED (VH) FROM MACH 0.9 TO MACH 2.6
ALSO: DESIGN LIMIT DESCENT SPE!DTGRDECEERTIONFRMCROIS
Vt: 500 KEAS DESIGN LIMIT SPEED FROM MACH 0.9 TO MACH 2.6
Rolling Mon�u.ers
120,500 LB
OR LESS
�%*
Figure 5-6
5-10
pproved for Release: 2017/07/25 C00821248
SECTION V
Approved for Release: 2017/07/25 C00821248
A- 12
RATED TIRE SPEED
PRESSURE ALTITUDE -1000 FT
INDICATED AIRSPEED - KNOTS
5-1Z
GOODRICH 27.5 x 7.5 x 16 SILVER TIRES
239 KNOTS (275 MPH) MAXIMUM GROUND SPEED RATING
ROSEMOUNT PITOT STATIC
Figure 5-7
pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
SECTION V
A-12
A red radial line at 135 KIAS represents the
minimum subsonic speed restriction below
30,000 feet when the stall warning light is
off.
EQUIVALENT AIRSPEED
The triple display indicator is not marked
however, the limit equivalent speeds are
as follows unless:
a. The Mach-airspeed instrument indicated
airspeed equals either the limit airspeed
hand indication or the minimum (135
KIAS) restriction.
b. The stall warning light illuminates or
the stall warning tone is heard.
Maximum TDI Airspeed
The limit airspeed is 450 KEAS at sea level,
increasing linearly with altitude to 500 KEAS
at 9400 feet pressure altitude; then 500
KEAS between 9400 feet and the altitude for
Mach 2.6. Limit airspeed then decreases
linearly with Mach number to 450 KEAS at
Mach 3.2. Normal operation cruise speed
is 3.1 Mach.
Minimum TDI Airspeed
The minimum airspeed restriction varies
linearly with Mach number from 135 KEAS
(Mach 0.38) at 30,000 feet to 300 KEAS
(Mach 1.34) at 50,000 feet, and is then a
constant 300 KEAS to 85,000 feet (Mach 3.1).
LOAD FACTOR LIMITS
The maximum allowable positive load
factor is 2.5 grs in symmetrical maneuvers
and 2.0 g's in roll maneuvers as des-
cribed by figure 5-6. The maximum nega-
tive load factor is -1.0 when below 400 KEAS
varying from -1.0 to 0 g's at higher air-
speed as shown by figure 5-6.
To avoid exceeding a safe angle of attack
positive g's are limited to 1.5 es when op-
erating above 2.5 Mac(1)-1. (This is equivalent
to approximately a 45 bank level turn.)
PROHIBITED MANEUVERS
The aircraft shall be operated in a manner
to avoid full stalls, spins, and inverted
flight. Normal bank angle when operating
above Mach 2.5 is 30 degrees.
RATE OF DESCENT LIMITATION
Rates of descent must be limited so as to
maintain positive fuel tank pressure when
sustained cruise speeds have exceeded
Mach 2.8.
CENTER OF GRAVITY
The aircraft shall be operated within a c. g.
range from 19% to 25% MAC while subsonic.
The c. g. must be forward of 25% MAC for
takeoff and should be as near to 19% MAC
as possible with existing fuel for landing.
The aft c. g. limit is 28% MAC while super-
sonic. This limit results from stability
considerations at high Mach number. Ade-
quate stability exists at farther aft centers
of gravity between Mach 1.2 and Mach 2.6
but for simplicity the aft limit is not
changed. The purpose of elevon trim limits
imposed in this Mach region is to alert the
pilot of a major malfunction in the fuel sys-
tem.
On those aircraft incorporating S/B 1141, if
an aft c. g. emergency exists and EMER for-
ward transfer is operated to place more than
4000 lbs in tank 1 and total fuel is less than
30,000 lbs, the aircraft should be limited to
maneuvers causing not more than 1.5 g.
As elevon trim can be used as an indication
of abnormal c. g. condition, the following
pitch trim limits apply:
While subsonic - no more than 10 nose down.
Changed 15 March 1968
IIIIIIMMENIApproved for Release: 2017/07/25 C00821248
5-l1
Approved for Release: 2017/07/25 C00821248
A-12
LONG RANGE CRUISE PROFILE
APPENDIX I
PART V
PRESSURE ALTITUDE -1000 FT.
FUEL REMAINING -1000
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600
I 0111k
200 400
I 1 2tir
800 1000 noo 1100 1600 1800 2000
DISTANCE NAUTICAL MILES
2200 2100 2600 2800
3000
3200
3400
MACH 3.20
STANDARD DAY CLIME
CRUISE REF. TEMP.
-56.5�C
'Approved for Release: 2017/07/25 C00821248
Figure A5-6
pproved for Release: 2017/07/25 C00821248
APPENDIX I
PART V
A-12
PROFILE CHART: CLIMB - CRUISE INTERCEPT POINTS
1956 ARDC ATMOSPHERE
LONG RANGE CRUISE - MACH 3.20
INITIAL
GR. WT .
LB.
INITIAL
ALTITUDE
FT.
CUMB
TEMP
�C
CRUISE
TEMP.
�C
CUMB - CRUISE INTERCEPT
ALTITUDE
FT.
DISTANCE
N. MI.
TIME
MIN.
FUEL REM.
LB.
122,450
30,000
STD -10
STD
STD +10
-66.5
-56.5
-46.5
-66.5
-56.5
-46.5
-66.5
-56.5
-46.5
77,050
75,296
75,296
77,500
75,500
75,296
78,350
76,450
75,296
258
237
237
326
302
259
438
416
402
14.6
13.9
13.9
17.4
16.6
16.5
22.0
21.3
20.8
48,410
48,875
48,875
45,620
46,150
46,205
41,070
41,570
41,875
119,150
S.L.
STD -10
STD
STD +10
-66.5
-56.5
-46.5
-66.5
-56.5
-46.5
-66.5
-56.5
-46.5
78,200
76,200
75,296
78,600
76,750
75,296
79,350
77,400
75,800
267
244
233
326
305
287
421
398
379
16.9
16.1
15.8
19.5
16.8
18.2
23.6
22.8
22.1
42,130
42,665
42,905
39,650
40,145
40,530
36,075
36,595
37,020
105,150
S.L.
STD -10
STD
STD +10
-66.5
-56.5
-46.5
-66.5
-56.5
-46.5
-66.5
-56.5
-46.5
80,400
78,550
76,900
80,800
79,050
76,400
81,450
79,750
78,000
254
233
213
301
281
249
375
355
335
15.3
14.5
13.9
17.3
16.6
15.5
20.4
19.7
19.0
30,980
31,475
31,915
28,915
29,380
30,085
26,280
26,735
27,200
Approved for Release: 2017/07/25 C00821248
Figure A5-6
(Sheet 2 of 3)
Approved for Release: 2017/07/25 C00821248
SECTION V
A-12
While climbing - 2-1/2o nose down from
Mach 1.4 to Mach 2.6,
3-1/20 nose down above
Mach 2.6.
At initial cruise the trim limit is 3-1/2�
nose down at 28% c. g. As altitude increases
and KEAS decreases, the 28% c. g. trim
limit becomes approximately 2 more nose
up per 50 KEAS decrease from 450 KEAS.
(In addition, expect approximately 10 more
nose up trim for each percent that c. g. is
forward of 28% MAC).
FUEL LOADING LIMITATIONS
These limits to be supplied at the operating
site.
AIRCRAFT SYSTEM LIMITATIONS
STABILITY AUGMENTATION SYSTEM
The SAS shall be on for all takeoffs and
landings.
INLET SPIKE AND BYPASS CONTROLS
The spike and forward bypass controls
must be operated in the AUTO mode at all
times when above 80,000 feet. When inlet
controls must be operated manually, maxi-
mum allowable speed is Mach 3.0.
CANOPY
The canopy shall be opened or closed only
when the aircraft is completely stopped.
Maximum taxi speed with the canopy open
is 40 knots. Gusts or strong winds should
be considered as a portion of the 40 knot
speed limit.
LANDING GEAR SYSTEM
Landing Gear
Do not exceed 300 KEAS or Mach 0.9 with
a maximum of 5o sideslip with gear ex-
tended. When sideslip angle exceeds 5�,
operation with gear extended is limited to
Mach 0.7 or 300 KEAS. Operation at super-
sonic speed with gear extended is prohibited.
The landing gear is designed for landing
sink speeds at touchdown which decrease
from 9 FPS at 57,000 pounds to 5 FPS at
123,600 pounds. Side loads during takeoff,
landing, and taxiing must be kept to a mini-
mum, as landing gear side load strength is
critical during ground maneuvering.
Tires
The maximum taxi speed recommended is
40 knots for Goodrich 27.5 x 7.5 x 16 "sil-
ver" tires. The rated ground speed limit
is 239 knots. At 4500 feet elevation, 239
knots corresponds to 210 KIAS with 108oF
ambient temperature on a calm day, and
226 KIAS at 32�F ambient temperature.
Limit indicated airspeed on the ground de-
creases by the amount of tailwind compon-
ent along the runway and increases by the
headwind component. Refer to figure 5-7
for rated speeds at other altitudes and
temperatures.
Taxiing Restrictions
A heat check is required for tires, wheels,
and brakes:
a. Prior to takeoff when taxiing has
exceeded one statute mile.
b. When continuous taxi distance has
exceeded 5 statute miles.
When clear of the runway after an
aborted takeoff or a heavy weight
landing.
Changed 15 March 1968
Approved for Release: 2017/07/25 C00821248
5-13
SECTION V
Approved for Release: 2017/07/25 C00821248
INITIAL BRAKING SPEED FOR STOP USING RATED BRAKE CAPACITY
PRESSURE ALTITUDE - 1000 FT
ONE STOP CAPABILITY
118,800,000 FOOT - POUND CAPACITY DRY AND HARD RUNWAY
6 x 4 ROTOR BRAKES ZERO WIND, ZERO SLOPE
ROSEMOUNT PITOT STATIC NOSE DOWN
5
4
3
2
o'
�11'
:-+41
biPleY
-T- +- -4
. 190 K1AS DRAG CHUTE
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Data as of 1J ugly 1967
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230
220
210
200
190
180
170
160
150
140
130
120
110
100
MAXIMUM INITIAL BRAKING SPEED
WITH DRAG CHUTE -KIAS
MAXIMUM INITIAL BRAKING SPEED WITHOUT DRAG CHUTE - KIAS
Figure 5-8
= - 1 117 .rAir Milr7aVilWar --, 1 uW
5-14
Changed 15 March 1968
pproved for Release: 2017/07/25 C00821248
Approved for Release: 2017/07/25 C00821248
1
2.
0Ep(oyAnr-u-r To AeC.P rkogir crir
Normal Climb Performance From Brake Release 44 A &IVA
(1956 ARDC Atmosphere)
Phase I, Subsonic Climb A3-1
Phase II, Transonic Maneuver A3-2
Phase III, Std Day, -10 C AT, Supersonic Climb A3-3
Std Day, Supersonic Climb A3-4
Std Day, +10 C AT, Supersonic Climb A3-5
Normal Climb Performance After Air Refueling
(1956 ARDC Atmosphere)
A3-6
Normal Climb Performance From Brake Release
("MEAN TROPIC" Atmosphere)
Phase I, Subsonic Climb
A3-7
Phase II, Transonic Maneuver
A3-8
Phase III,Mean Tropic Day -10�C AT, Supersonic Climb
A3-9
Mean Tropic Day, Supersonic Climb
A3-10
Mean Tropic Day +10 C AT, Supersonic Climb
A3-11
Normal Climb Performance After Air Refueling
("MEAN TROPIC" Atmosphere)
A3-12
Military Thrust Climb Performance, Std Day -10�C AT
A3-13
Std Day
A3-14
Std Day +100 C AT
A3-15
Std Day +24.5�C AT
A3-16
Normal Descent Performance
A3-17
350 KEAS Descent Performance
A3-18
Alternate 350 KEAS Descent Performance
A3-19
Single Engine Descent Summary - Max A/B
A3-20
Single Engine Descent - 300 KEAS
A3-21
- 350 KEAS
A3-22
- 400 KEAS
A3-23
Single Engine Turning Descent
A3-24
NORMAL CLIMB PERFORMANCE
FROM BRAKE RELEASE
Figures A3-1 through A3-5 and A3-7 through
A3-11 present normal climb performance
from brake release to cruise altitudes for
supersonic operation with 1956 ARDC At-
mosphere and "MEAN TROPIC" Atmospheric
conditions, respectively. The data is c-Drn-
puted from results of Flight Test and Op-
erational Testing with YJ-1 engines. The
climb is segmented in three phases and in-
cludes the effects of varying gross weigts
and air temperatures on fuel used, time.
and distance. Phase I is the subsonic
portion of the climb from brake release at
sea level to 30,000 feet and 0.90 Mach. Cor-
Changed 15 March 1968
pproved for Release: 2017/07/25 C00821248
A3-1
Approved for Release: 2017/07/25 C00821248
APPENDIX I
PART III
A-12
rections for time, fuel, and distance are
listed in the chart for takeoffs from other
field elevations. Phase LI is the transonic
acceleration portion of the climb from
30,000 feet and 0.90 Mach to 30,000 feet and
1.25 Mach utilizing the climb and dive tech-
nique. Phase III is the supersonic portion
of the climb from 30,000 feet and 1.25 Mach
to the altitude at which cruise Mach number
is first attained. Phase ILLA. is the constant
Mach portion of the climb from the end of
Phase III to the altitude for start of cruise.
The following is a tabulation of the average
results of flight tests for Phase MLA.
Cruise
Profile
Scheduled
Long Range
High Altitude
Example:
Avg.
RIG
FPM Power
2500 Cruise
4000 Max AB
Avg. Total
Fuel Flow
Lb/Min.
666
(20,000/
PPH/eng)
900
(27,000/
PPH/eng)
Obtain the time, distance, and fuel required
from brake release for takeoff at a field
elevation of 4500 feet to 3.10 Mach and
73,000 feet for a standard day. Fuel load
at brake release is 64,000 lb after sub-
tracting ground fuel allowances for normal
ground operation. (See Appendix, Part II,
for ground allowance computation procedure.
Find the initial gross weight at brake re-
lease by adding the zero fuel weight and the
fuel load remaining. If the zero fuel weight
is 55,150 pounds, the initial gross weight is
119,150 pounds. Enter figure A3-1 at the
initial gross weight at brake release and
read fuel used, time and distance for Phase
I as 6150 pounds, 4.8 minutes, and 34 nau-
tical miles, respectively.
From the table in figure A3-1, for the 4500
foot field elevation, reduce time, fuel, and
distance by 0.30 minutes, 500 lb, and 1.6
nmi, respectively. Therefore, fuel used,
time, and distance for Phase I is 5650 lb
(6150-500), 4.5 min. (4.8 - 0.3), and 32.4
nn-ii (34 - L6), respectively. Recompute
the gross weight at end of climb Phase I as
113,500 pounds (119,150 - 5650). Enter
figure A3-2 as the recomputed gross weight
and read fuel used, time, and distance for
Phase LI as 3100 pounds, 2.9 minutes, and
27 nautical miles, respectively. Recompute
the gross weight at end of Phase LI as
110,400 pounds (113,500 - 3100). Enter
figure A3-4 with the recomputed gross
weight and at 73,000 feet and Mach 3.10,
read fuel used, time, and distance for
Phase III as 13,500 pounds, 9.9 minutes,
and 205.6 nautical miles, respectively.
Add all three phases and obtain fuel used,
time, and distance as 22,250 pounds, 17.3
minutes, and 265 nautical miles, respec-
tively. Fuel remaining at 73,000 feet is
41,750 pounds (64,000 - 22,250).
Service allowances and/or allowances for
deviations from the normal climb schedule
can be applied to an affected phase when re-
quired. (For example, a subsonic cruise
operation prior to reaccelerating might be
scheduled in the flight plan.) The effect of
such an allowance must be accounted for
when computing the initial weight to be used
for the next phase of the climb.
AFTER AIR REFUELING
Figures A3-6 and A3-I2 present normal
climb performance from the end of 30,000
foot refuel (refueling with one AB on) to the
altitudes at which cruise Mach number is
reached for 1956 ARDC Atmosphere and
"MEAN TROPIC" Atmospheric conditions,
respectively. Adjustments which should be
used for other end A/R altitudes are listed
in the charts. The data is computed from
Flight Test and Operational Testing results
with YJ-1 engines. The assumed fuel load
at the end of A/R is 67,300 lb. Phase MA
results are identical to the tabulation in the
previous discussion.
A3-2
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A-12
APPENDIX I
PART III
PHASE III CLIMB WITH TURNS
Turns during climb are not recommended,
however, if mission requirements include
a turn, compensation for range lost due to
the turn must be included in the flight plan.
For examople, consider a 45o heading change
with a 30 bank at an initial altitude of
45,000 feet.
To minimize the rate of climb loss due to
turning, the recommended procedure is to
advance power to Maximum A/B during the
turn and maintain the speed schedule of
450 KEAS. Resume the normal climb pro-
cedure on completion of the turn.
Comparison of straightaway climb and turn-
ing climb on time, fuel and distance results
in an overall range loss of 32 miles for a
Mach 3.20 profile. On completion of the
45o turn, at 49,000 feet; time, fuel and dis-
tance to that altitude will be 0.35 min, 730
lb, and 6.4 ml greater than for a normal
climb with no turn.
MILITARY THRUST CLIMB PERFORMANCE
Figures A3-13 thru A3-16 present Military
climb performance for a schedule of 300
knots equivalent airspeed (KEAS) while be-
low 33,300 feet and 0.90 Mach number when
higher altitudes are attained. This power
and speed schedule provides the most climb
distance for the fuel consumed when sub-
sonic cruising flight plans, such as for ferry
or buddy missions, are used.
Example (1):
Find the time, distance and fuel required to
climb to 30,000 feet from S. L. on a std
-10oC day with an initial gross weight of
105,000 lb. Enter figure A3-13 at 30,000
ft, and at 105,000 lb initial gross weight
read 8.5 min, 56.6 miles and 4200 lbs.
Adding takeoff allowances results in time,
distance and fuel values of 9.4 min. (8.5 +
0.9), 59.2 miles (56.6 + 2.6) and 6000 lb
(4200 + 1800) for climb from sea level to
30,000 feet.
Example (2):
Find the time, distance and fuel required
to climb to 30,000 feet from 4500 foot take-
off on a std day with an initial gross weight
of 105,000 pounds. Enter figure A3-14 at
4500 feet and at 105,000 pound initial gross
weight; read 0.4 min, 3.8 miles and 550
pounds. Reenter figure A3-14 at 30,000
feet and an adjusted initial gross weight of
105,550 pounds; read 8.6 min, 58.0 miles
and 4300 pounds. Adding takeoff allowances
and subtracting values for climb from sea
level to 4500 feet results in time, distance
and fuel values of 9.1 min (8.6 + .9 - .4),
56.8 miles (58.0 + 2.6 - 3.8) and 5550
pounds (4300 + 1800 - 550) for climb from
takeoff at 4500 to 30,000 feet.
TWO ENGINE DESCENT PERFORMANCE
On course descent performance is shown on
figures A3-17, A3-18, and A3-19. Figure
A3-17 presents descent performance for the
normal 300 KEAS schedule. Figures A3-18
and A3719 present 350 KEAS descent per-
formance with forward bypass doors in the
automatic and open positions respectively.
SINGLE ENGINE DESCENT PERFORMANCE
Figures A3-20 through A3-24 present single
engine descent performance from 80,000
feet and Mach 3.10 (337 KEAS). The data
is based on flight test with the inlet con-
figuration as listed in the charts. Time,
distance, and fuel required are plotted
versus altitude. A pushover at constant
Mach is required to increase airspeed from
337 KEAS to the 350 KEAS or 400 KEAS
schedule. Better range is obtained when
the 300 KEAS schedule is used, reducing
airspeed to 300 KEAS while maintaining
constant altitude. These effects are in-
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APPENDIX I
PART
A-12
cluded in the performance data. Specific
range begins to decrease rapidly near
50,000 feet; therefore, the charts are in-
dexed to an altitude of 50,000 feet so that a
power reduction technique can be used and
the resultant change in performance can be
determined. The effect of changing KEAS
at the indexed 50,000 feet has not been de-
fined by flight testing and is not included in
the data.
Figure A3-20 summarizes the effect of air-
speed on Maximum AB descent performance
for constant values of 300, 350, and 400
KEAS. Figures A3-21 through A3-23 pre-
sent the effects of decreasing power at the
index altitude of 50,000 feet for constant
airspeeds of 300, 350, and 400 KEAS, re-
spectively. Fivre A3-24 presents the
effects of a 180 turn at 35 bank angle on
a 350 KEAS descent. Approximately 23,000
feet of altitude is required to complete the
180o turn. For convenience in mission
planning, a ground track profile is also pro-
vided.
Example (1):
Find the time, distance, and fuel required
to descend on course from 80,000 feet and
Mach 3.10 using the 300 KEAS descent
schedule and Minimum AB below 50,000
feet. Enter figure A3-21 at 80,000 feet and
read the time, distance, and fuel required
to 50,000 feet as 8.6 minutes, 176 nautical
miles, and 2800 pounds of fuel. Reenter at
the final altitude of 31,500 feet on the Min-
imum AB line and read time, distance, and
fuel required as 5 minutes, 52 nautical
miles and 1600 lb of fuel. Add the results
and obtain 13.6 minutes, 228 nautical miles,
and 4400 pounds of fuel.
Example (2):
Find the track time, distance, and fuel re-
quired to descend from 80,000 feet and Mach
3.10 using the 350 KEAS descent schedule.
A 90o turn is to be completed above 50,000
feet, and Minimum AB is to be used below
50,000 feet. Enter figure A3-24 and noteo
on the penetration distance curve that 90
of turn is completed at 65,000 feet altitude.
Read time at that altitude as 3.0 minutes
and fuel used as 1200 pounds. On the
ground track Profile note that the distance
traveled is 80 nautical miles. Enter figure
A3-22 at 65,000 feet (end of turn altitude)
and read time, distance, and fuel required
to 50,000 feet as 4.1 minutes, 73 nautical
miles, and 1950 pounds. Reenter at the
final altitude of 28,000 feet on the Minimum
AB line and read time, distance, and fuel
required as 3.4 minutes, 41 nautical miles
and 1200 lb of fuel. Add the incremental
readings and obtain 10.5 minutes, 194 nau-
tical miles, and 4350 pounds of fuel.
A3-4
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A-12
APPENDIX I
PART LEI
as 22,200 pounds, 17.3 minutes, and 265
nautical miles, respectively. Fuel remain-
ing at 73,000 feet is 41,800 pounds (64,000 -
22,200).
Service allowances and/or allowances for
deviations from the normal climb schedule
can be applied to an affected phase when re-
quired. (For example, a subsonic cruise
operation prior to reaccelerating might be
schedule in the flight plan.) The effect of
such an allowance must be accounted for
when computing the initial weight to be used
for the next phase of the climb.
Example (2):
Obtain the MEAN TROPIC day time, dis-
tance, and fuel required from refuel at
29,000 feet to 0.90 Mach at 38,000 feet
(start of Phase II) . Enter fig.. A3-8 at 29,000
feet and read fuel used, time, and distance
for Phase IA as 3740 pounds, 3.60 minutes,
and 31.6 nautical miles, respectively. The
recomputed gross weight for entering Phase
LI will be 118,710 pounds (122,450 - 3740).
PHASE III CLIMB WITH TURNS
Turns during climb are not recommended,
however, if mission requirements include
a turn, compensation for range lost due to
the turn must be included in the flight plan.
For examople, consider a 45 heading change
with a 30 bank at an initial altitude of
45,000 feet.
To minimize the rate of climb loss due to
turning, the recommended procedure is to
advance power to Maximum A/B during the
turn and maintain the speed schedule of
450 KEAS. Resume the normal climb pro-
cedure on completion of the turn.
Comparison of straightaway climb and turn-
ing climb on time, fuel and distance results
in an overall range loss of 32 miles for a
Mach 3.20 profile. On completion of the
45o turn, at 49,000 feet; time, fuel and dis-
tance to that altitude will be 0.35 min, 730
lb, and 6.4 mi greater than for a normal
climb with no turn.
MILITARY THRUST CLIMB PERFORMANCE
Figures A3-13 thru A3-16 present Military
climb performance for a schedule of 300
knots equivalent airspeed (KEAS) while be-
low 33,300 feet and 0.90 Mach number when
higher altitudes are attained. This power
and speed schedule provides the most climb
distance for the fuel consumed when sub-
sonic cruising flight plans, such as for ferry
or buddy missions, are used.
Example (1):
Find the time, distance and fuel required to
climb to 30,000 feet from S. L. on a std
-10�C day with an initial gross weight of
105,000 lb. Enter figure A3-13 at 30,000
It, and at 105,000 lb initial gross weight
read 8.5 min, 56.6 miles and 4200 lbs.
Adding takeoff allowances results in time,
distance and fuel values of 9.4 min. (8.5 +
0.9), 59.2 miles (56.6 + 2.6) and 6000 lb
(4200 + 1800) for climb from sea level to
30,000 feet.
Example (2):
Find the time, distance and fuel required
to climb to 30,000 feet from 4500 foot take -
off on a std day with an initial gross weight
of 105,000 pounds. Enter figure A3-14 at
4500 feet and at 105,000 pound initial gross
weight; read 0.4 min, 3.8 miles and 550
pounds. Reenter figure A3-14 at 30,000
feet and an adjusted initial gross weight of
105,550 pounds; read 8.6 min, 58.0 miles
and 4300 pounds. Adding takeoff allowances
and subtracting values for climb from sea
level to 4500 feet results in time, distance
and fuel values of 9.1 min (8.6 + .9 - .4),
56.8 miles (58.0 + 2.6 - 3.8) and 5550
pounds (4300 + 1800 - 550) for climb from
takeoff at 4500 to 30,000 feet.
TWO ENGINE DESCENT PERFORMANCE
On course descent performance is shown on
figures A3-17, A3-1`8, and A3-19. Figure
A3-17 presents descent performance for the
normal 300 KEAS schedule. Figures A3-18
and A3-19 present 350 KEAS descent per-
formance with forward bypass doors in the
automatic and open positions respectively.
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APPENDIX I
PART III
A-12
SINGLE ENGINE DESCENT PERFORMANCE
Single Engine Descent data is presented for
Military, Minimum afterburning and Maxi-
mum afterburning power at 300, 350 and
400 KEAS with 1956 ARDC and Mean Tropic
Atmosphere conditions. Refer to figures
A3-20 through A3-23B.
Allowances For Deceleration To Descent Speed:
When cruising at a higher KEAS than the
desired descent schedule, the constant alti-
tude deceleration is made at the same power
setting as the constant KEAS descent. The
constant Mach lines show the beginning
point of the deceleration for each Mach
number. In the situation where the cruise
KEAS is less than the desired descent
KEAS, the constant Mach descent is made
with Maximum afterburning power. The
constant Mach lines show the descent for
different Mach numbers.
Comparison Of Descent Power and Speed Schedules:
The Maximum afterburning descent, as
compared to the Minimum afterburning and
Military power descents, results in a longer
distance, a longer elapsed time and more
fuel used. The 400 KEAS descent as com-
pared to the 350 and 300 KEAS descents re-
sults in a slightly longer distance, less
elapsed time and more fuel used. Maxi-
mum overall range results if a descent
speed of 300 KEAS is used and if Military
power is used in the descent and for cruise.
There will be little overall range loss if
either Minimum afterburning or Maximum
afterburning descent power is used as long
as the cruise is accomplished in Military
power. The charts are indexed to an alti-
tude of 50,000 feet so that a technique of
power or airspeed change can be used and
the resultant effect after power change in
performance can be determined. The effect
of changing KEAS at the indexed 50,000 feet
has not been defined by flight testing and is
not included in the data.
CAUTION
When making a single engine de-
scent with the operating engine in
Military power, the Mach rate
limit of 1.0 Mach in three minutes
will be exceeded.
Single Engine Turning Descent
Figure A3j24 presents the effects of a 180�
turn at 35 bank angle on a 350 KEAS de-
scent. Approximately 23,000 feet of altitude
is required to complete the 180o turn. For
convenience in mission planning, a ground
track profile is also provided.
Sample Use Of Charts
Example (1):'
Find distance, time and fuel to descend
from 80,000 feet to 29,000 feet, using Mini-
mum afterburning power and 300 KEAS.
Initial speed is Mach 3.1 (337 KEAS). Nor-
mal (ARDC Standard) atmosphere conditions
are expected. Refer to Figure A3-20.
Enter the chart at 80,000 feet and located
the Minimum afterburning line for the 3.1
Mach, (337 KEAS) condition, and read dis-
tance, time and fuel to 50,000 feet.
Distance = 137 miles
Time = 6.8 minutes
Fuel = 1200 pounds
A3-4
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A-12,
APPENDIX I
PART III
Enter the same chart at 29,000 feet and
read distance, time and fuel from 50,000
feet to 29,000 feet.
Distance = 75 miles
Time = 8.5 minutes
Fuel = 2400 pounds
Add the above values to obtain distance
time and fuel from 80,000 feet and 3.1 Mach
to 29,000 feet in Minimum afterburning at
300 KEAS.
Distance = 212 miles
Time = 15.3 minutes
Fuel = 3600 pounds
Example (2):
Find the track time, distance, and fuel re-
quired to descend from 80,000 feet and Mach
3.10 using the 350 KEAS descent schedule.
A 90o turn is to be completed above 50,000
feet, and Minimum AB is to be used below
50,000 feet. Enter figure A3-24 and noteo
on the penetration distance curve that 90
of turn is completed at 65,000 feet altitude.
Read time at that altitude as 3.0 minutes
and fuel used as 1200 pounds. On the ground
track profile note that the distance traveled
is 80 nautical miles. Enter figure A3-21 at
65,000 feet (end of turn altitude) and read
time, distance, and fuel required to 50,000
feet as 4.1 minutes, 73 nautical miles, and
1950 pounds. Reenter at the final altitude
of 28,000 feet on the Minimum AB line and
read time, distance, and fuel required as
3.4 minutes, 41 nautical miles and 1200 lb
of fuel. Add the incremental readings and
obtain 10.5 minutes, 194 nautical miles,
and 4350 pounds of fuel.
Changed 15 June 1968 A3-4A/A3-4B
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A-12
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