TA-12 TRAINER FLIGHT MANUAL

• Document Type: FOIA
• Collection: FOIA Collection
• Document Number (FOIA) /ESDN (CREST): 06230172
• Release Decision: RIFPUB
• Original Classification: U
• Document Page Count: 285
• Document Creation Date: December 28, 2022
• Document Release Date: August 10, 2017
• Case Number: F-2014-00925
• Publication Date: October 16, 1967

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5
OX /3v.7-6
PPY 3 OF
Page I c
TDC No. 20
16 October 1967
TECHNICAL DATA CHANGE
TA-I2 Trainer Flight Manual
This TDC, dated 16 October 1967, accomplished the following:
I. Revises the Abort Procedure
2. Revises the Brake Description and adds a maximum initial
Braking Speed Chart.
3. Revises the Landing Field Length Requirements description and
adds several Landing Distance performance charts.
The pilot's abbreviated checklist is supplied separately.
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15
COPY NO
TRAINER
RIGHT MANUAL
PUBLISHED UNDER AUTHORITY OF THE SECRETARY OF THE AIR FORCE
31 MARCH 1967
Changed 16 October 1967
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SECTION VI
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APPENDIX I
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PART I
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NOTE: The portion of text affected by the change is indicated
by a vertical line in the outer margins of the page.:
cotes deletion of text.
Issue Code A-1
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October 1967
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APPENDIX I
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APPENDIX I
PART III
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APPENDIX I
PART IV
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WTheashortsh Indkates pages changed, added, ordelimmlby; NON:Thepettlemetexteffectodbythechangehtindiceted
.06ecunentchempeAnsertkomackengedandloreddedpagek byltvertkleffitteintheoWNwrimerensofflopage.Allndi-
dufrov suPerai"a totes deletion ef text.
_ _
Chanted 16 October 1967
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Is sue Code A-1
�-�
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TA-12
SECURITY INFORMATION
SPECIFIC INSTRUCTIONS FOR SAFEGUARDING THIS INFORMATION
1.
This document contains information affecting the national defense of
the United States within the meaning of the Espionage Laws Title 18,
USC Section 793 and 794. The transmission or the revelation of its
contents in any manner to unauthorized persons is prohibited by law.
The nature of this document is such that dissemination and handling
will be carried out with strict adherence to the following policies:
a. Distribution will be controlled on a strict, officially established
"need-to-know" basis.
b. Strict accountability of each document will be maintained.
c. This document will be controlled in such a fashion to prevent its
loss, destruction, or falling into the hands of unauthorized
persons.
2. In the event this document is lost or is subject to unauthorized dis-
closure or other possible subjection to compromise of classified
information, such fact will be promptly reported to the authority
responsible for the custody of the material for appropriate action.
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�
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TA-12
TABLE OF CONTENTS
SECTION I
SECTION II
SECTION Ill
SECTION IV
SECTION V
SECTION VI
SECTION IX
Page
DESCRIPTION
' NORMAL PROCEDURES � 2-I
EMERGENCY PROCEDURES 3-1
AUXILIARY EQUIPMENT 4-I
OPERATING LIMITATIONS 5-I
FLIGHT CHARACTERISTICS 6-I
'ALL WEATHER OPERATION 9-I
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COPY NO.
TECHNICAL DATA CHANGE
TA-12 Trainer Flight Manual
ICrAC- 0605-C7
COPY 3 OF 3
Page 1 of 1
TDC No. 19
31 March 1967
This TDC, dated 31 March 1967, is a reissue of the TA-12
Trainer Flight Manual. The flight manual has been updated to
the latest configuration and replaces the TA-12 Trainer Flight
Manual dated 1 May 1964.
The pilot's abbreviated checklist is supplied separately.
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46424:4:04-SEA
pproved for Release: 2017/07/25 C06230172
.T6 I-3E INSERTED IN FRON"i' OF
,:-..0p.y. f....,.0-..
TA- IZ Trainer Flight
FLIGHT 1\4(\ NUAL.1)A�ED 1 May 1964, ;changed 1965
S.1 CJI()N.
Stf
GE:
.-._,.
_
'CFI A NG.F.7.', .
;
.
;�
'rill TDC, compri.sing the changed andier-.added pages listed
.
On pages A_ and F3,.i-ni:ikes the :following changes:
f
1 . � Reflects installation of alternate nosewbeel: Steerinp, �
. ,
Changes coc1;:pit illuStrations to sho\'v late.st instrufnent.
...
,
a r ran:�:-,:enient s .
3. Adds an Ii-qS Steering Characteristics and Destination.
Yt
IV.
Reject Pattern illustration to. Section777
,_
.
,.
,
Adds a 'Mach Hold Engag. ement procec,iure.
5. Updates Inertial Navigation Sy Stern description .and�
ocedure.
.1b
Clzanges landino:, and penetration speeds.
:-�
,
,
Adds X-13and radar beacon- operation to normalopeat
pr 0 C e Cl U 0 S .
-1:,�
7.
.1'
-,,
-,
,
,.
Fey sec checi-z_1i!3i: page.s are supplied separately.
Note .
._.
Cancel ".ED.0 No. 17:
s --
.:
..
_.
..
...
,
,
� , '
;:
�
NOTE:

The technical data information furnished herein is intended to be
used as INTERIM data only. It wtll be replaced and superseded
at the time of issue of the next revision to the flight ms:nual.
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,FFM.
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�
10 B.P.;
lelANUAL IDA�E',D 1 N.O.r_erri1De1:1963 (Changed 15 Dec i 1963)
f:
�T.06NO. 7
Lc/
CHANGE'
The new pages and, title page eit4�ted 1E2, DeceMber 19.63 ,.supplied by
this inc .replace and -supercede.!'affetted pages previoUsryissuca
� The new A page can be used a:9.a referenee to rnake�certain...that
the Flight Manual contain's all effective pages.
No checi(list.c.i�anges are necessary because of these new
. changed pages
�NOTE The technical data inforrnation furnished herein is intended to be
used as INTERIM data only. ,lt will be replced and superseded
:at the time Of i.satie of the next, revision to the flight rnanUal..-
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''W*576 -gt;WitrAVIMgegaUrr',AgilitfAUZ4,
proved for Release: 2017/07/25 C06230172F-
,Ntiv.:,,
3 7577''
copr3
IlF,:IN5.F...-.RTF,D IN, FRONT OF Train,er
FLICJI 1.1 AN-11 A 1, 1-.)ATED 1 November 1962 (Changed 1 Noverriber:.q:i
CIT A Nic,F,
The neW pages and title page. dated 1 November 19.63
by this Tpc replace and supeIsi.:7.de affected pages previousl.y.
iSsued. The new "A" page can Lic used as.a �reference'to:-
make certain that the Flight Manual contatris all effective
pa a es.
Changed procedures are the result of operational e,xperienc,e
and reCommendations by the using agency.
Changed check list pages dated 11-1-63 are consisterit with
this TDC.
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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:Approved for Release: 2017/07/25 C06230172,,
==.�
T BE INSERTED IN FRONT OF A42 Trainer COPY
FLIGHT MANUAL DATFD 1 November 1962 (Changed 15 August 1963
O
P r riON
,
PAUF
CHANGE
1. The new pages and title page dated 15 August 1963 supplied by
this �TDC replace and supersede affected pages previously
issued in the basic Flight Manual, and those pages changed per
TDC's No. 1 and 2.
2.. Changes to Sections II, III, and V dated 15 August 1963 have
already been issued under TDC's No. 3 and 4.
3. An t'A" page is provided with this change and may be used as
a reference ttx) make certain that the Flight Manual contains
all of the latest changed pages.
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 .inannal. �
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-..Approved for Release: 2017/07/25 C06230172
.,� �
-
-.TEGH.N:l.p4.. DATA -C.HA GE
FLIGHT -MANUAL.
TO BE INSERTED IN FRONT OF A-12. Trainer
FLIGHT MANUAL 1)A TED 1 November 1962
1 September 196,
S.-46y
COPY OF
SECTION
PAGE
CHANGE.
New pages supplied by this TDC replace affected pages in Section II
dated 15 August 1963.
This new material reflects procedures developed as a result of
Trainer aircraft operation which include:
a. Positioning of aft cockpit battery switch during preflight
check
b. Redefinition of minimum airspeeds for drag chute jettison
with crosswind components of 5 to 12 knots. ,
Pages supplied by this TDC are marked YChanged 1 September 19631�
at the lower inside corner of the page.
List of pages supplied:
Section II
'
2.-5
2-31
,r-1h
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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----
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C,0 PY Na 15
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A.-12 Trainer
FLIGHT MANUAL DATED 1 November 196?
Page 3. of 2
TDC NO. 3
15 August 196!
SECTION
PAGE
., CHANGE
New pages supplied by this TDC replace all original
pages in Sections II and III and Section V Limitations.
pages 5-11 and 5-12 of A-12 Flight Manual dated ,
l November 1962. Changes to affected pages authorized
by TM's No. 1 and 2 are superceded. The Pilots
Abbreviated Check List dated 1 November 1962 is also
supereeded by this material and revised cheek lists
will be issued as soon as possible
New material supplied by this TDC reflects procedures
developed as a result of trainer aircraft operation$
.and Oquipment changes which Include:
a. Inverter switching rearrangement (SB 351)
b. Operational anti-skid braking
e. RMI needle switching rearrangement
d. 4-rotor brake installation and "silver" tires
Pages supplied by this TDC are marked "Changed
15 August 1963" at the lower inside corner of the
,
page.
NOTE The technical data information furnished herein is intended to he
used as INTERIM data only. It will he replaced and superseded
at the time of issue of the next revision to the flight manual.
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Page .2 of:
TDC NO. 3
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BF: IN5ERT.ED IN FRONT OF A-12
.. �
FLIGHT M ANU A PA T ED I NOVeiriber 1962
15 August 196"
SFCTION
PAC;F;
CHANGE;
List of changed pages supplied:
Section V
Section 11
Section III
2-1
2...1
through
,
3-1
3-1
through
�
5-11
5-12
.NOTE : The technical data information furnished herein is intended to be
used as INTERIM data only. It will he replaced and superseded
at the time of issue of the next revision to the flight manual.
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AL DATA CHANGE
FLIGHT MANUAL
TO RE INSERTED IN FRONT OF A-12 Trainer
ELIC;IIT MANUAL, DATE]) 1 November 1962
,
ION
PAGE
CH A NC; E
-.38
Under Fuel Shutoff Switches, after the 3rd sentence add a
,
WARNING as follows:
'
WARNING
-
-
When a fuel shutoff switch is actuated to the EMER
-
,
position, a minimum of 5 seconds delay must be observed
prior to moving the switch back to the ON position
to allow for full travel of the shutoff valve.
.
,
Attempting to recycle the valve within 5 seconds
may cause the circuit breaker in Air Conditioning
Bay to open.
_
,
-
I ,
1-41
Under Crossfeed Switches, before the last sentence, add a
CAUTION as follows:
...,
A
-
c.
CAUTION
-
When the crossfeed switch is depressed, a mininum
;
,
of 5 seconds delay must be observed prior to
depressing the switch a second time to allow for
full travel of the crossfeed valve. Attempting to
,�
t,
_-...
i
1

� 1-1,2
recycle the valve within 5 seconds may cause the
circuit breaker in a Air Conditioning Pay to open.
Under Fuel Dunn Switches, after the 3rd sentence add a WARNING
as follows:
,-,
,
,:.
,
WARNING
When the fuel dump switch is actuated to the MIT
, position, a minimum of 5 seconds delay must be
observed prior to moving the switch back to the
OFF position to allow the full travel of the dump
valves. Attempting to recycle the valves within
5 seconds may cause the circuit breakers in the
.
Air Conditioning Bay to open.
_
.
-
,
,
.
1-85
Under Landina. Gear Warning Light and Audible Warning, revise
as follows:
,
L
,
,
,
,
.
,
Change the 1st sentence after step 3 to read: An audible
warning is produced in the pilot's earphones when the
throttles are retarded below minimum cruise setting, the
landing gear is not in the dcwn and locked position and
altitude is below 10,000 (4- 500) feet. ,
,
The technical data information furnished herein is intended to be
9 A. sN R
.TT--,,- data only. -
IM '. Itt revision7,i4b replaced and superseded
t the time of issue of the 'Approved for Release: 2017/07/25 C06230 72
Approved for Release: 2017/07/25 C06230172,
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED 1 November 1962
, .
SEC 1 JON
,
PAGE
CHANGE
I
107
Under EJECTION SEQUENCE, in the 2nd and 6th sentences change
the numeral 2 to 4.
;
107
At end of Section 1 after EJECTION SEQUENCE add the following:
V
EGRESS (Bail Out) SYSTEM
,V
,
,
An egress light system installed in the aircraft permits
bailout coordination between pilots in addition to normal
interphone communication or in the event that interphone
communication is interrupted. With this system the
aircraft commander always has the capability to issue and
check compliance with a bailout signal, regardless of
which cockpit he may be occupying. Power for the system
is furnished by the essential dc bus. See EFERGENCY
ESCAPE IN FLIGHT, Section III for further information.
I '
107
Egress Lights and Switches
,
,
,
- ,
The forward cockpit lower right instrument panel contains a
guarded toggle switch labeled BAIL OUT (up) and two lights
which read BAIL OUT (red) and AFT COCKPIT EJECTED (amber)
when illuminated. The aft cockpit lower instrument panel
contains a guarded switch labeled BAIL CUT (up) and a light
which reads BAIL OUT (red) when illuminated. Both forward
and aft cockpit switches are safety wired tc the off (guard
down) position. Actuation of a BAIL OUT switch illuminates
the BAIL OUT light in the opposite cockpit. The AFT
CCCKPIT EJECTED light is wired directly to the aft cockpit
ejection seat tracks and will illuminate when the aft seat
is ejected.
.
.
,
:
-
,
ie technical data information fUtniehed herein ie. iiitepcie0 to be.
�
as INTERIM data only. It will
be replaced and
P.4.13edeciVVtrncSofi8uVe of the next revl8ion to the flight manjl.V
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ar
�T DC NO:
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT M A NU A L DA TED 1 NoVember 1962'
SEC; 110N
PAGE
CHANGE
l'-II
2-5
Under Instrument Panel, revise as follows:
Change step 12 to read: Forward cockpit system switch - ON.
I/ II
2-9
Under Instrument Panel, revise as follows:
1. Delete step 13.
2. Renumber steps 14 thru 39 to 13 thru 38.
II
2-20
Under PRE-TAKEOFF AIRCRAFT CHICK, revise as follows:
/.----
1. Delete steps 5 and 6.
2. Renumber steps 7 thru 12 to 5 thru 10.
_ II
2-26
Under AFTER TAITOFF CLIMB, revise as follows:
�

1. Delete steps 5 and 6.
2. Renumber step 7 to step 5.
,
;
The technical data information furnished herein is :intended to be
used as INTERIM data only. iI6v4I.elsre'plaeedper'superse4d
me of Issue of the next revision to the flight manual
'Approved for Release: 2017/07/25 C062301724:
Approved for Release: 2017/07/25 C06230172
TDC NO: ,
TECHNICAL DATA CHANGE
FLIGHT MANUAL
,TO INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUA T., DATED 1 November 1962
SEC }MN
PAGE
. CHANGE
III
..) J -24
Under EJECTION, in 2nd sentence under item d., change to read:

y//
,
There is a 0.6 second delay on seat separation below 265-300
KIAS, and a 4 second delay above 265-300 KIAS.
4,, III
3-24
Under EMERGENCY ESCAPE IN FLIGHT, temporarily delete existing
'
steps 1 and 2 and replace with the following:
Aircraft Commander flying in forward cockpit.
1. If possible, notify aft cockpit of decision to eject.
2. Actuate bailout switch.
3. Observe AFT SEAT EJECTED light illuminated.
After aft cockpit seat ejects,
4. Pull ejection "D" ring with both hands.
5. After parachute is open and before touching down, pull
�
survival kit release handle to reduce touchdown weight
and avoid leg injury on landing.
�
Aircraft Commander flying in aft cockpit.
1. If possible, notify forward cockpit pilot of decision to eject.
2. Actuate bailout switch.
3. Observe forward cockpit seat ejection.
After forward cockpit seat ejects,
4. Same as step 4 above.
5. Same as step 5 above.
III
3-39
Under FUEL DUMPING PROCEDURE, revise as follows:
3-39a
1. After step I add the following WARNING:
WARNING
'Allow a minimum of 5 seconds before moving the dump
switch back to the OFF position.
, .
The technical data information furnished herein is intended to be -
used as iST-lklivt,data only. It NvEff..:be..replaced and superseded
the tiMe. of the: nect, revision to the 'flight manual
-
Approved for Release: 2017/07/25 C062301721-
Approved for Release: 2017/07/25 C06230172
TDC NO.
TECHNICAL DATA CHANGE
FLIGHT MANUAL
� TO .FIE INSERTED IN FRONT OF A-12 Trainer
FUG! TT MA NU AI., DATED 1 November 1962
SEC I ION
PA(;E
CHANGE
III
3-39
WARNING (Contld)
3-39a
(Cont'd)
*Do not attempt to dump fuel when the fuel level in
tank 3 is below 4000 lbs. indicated. If dumping
has been initiated, terminate dumping when fuel
level reaches 4000 lbs.
'
. Revise the sentence under step 6 to read:
When fuel level in tank 3 reads 4000 lbs,
. Delete existing NOTE under step 7 and replace with the
following:
i
NOTE
If a power failure should occur in the dump circuit,
it is possible for the normal motor driven valves to
fail in the open position; however, solenoid operated
back up dump valves installed in the dump lines will
close to stop fuel dumping.
III
3-39a
Under FORWARD FUEL TRANSFER AND FUEL DUMPING PROCEDURE, after
step 1, add the following WARNING:
WARNING
'Allow a minimum of 5 seconds before moving the
dump switch back to the OFF position.
.D0 not attempt to dump fuel when the fuel level
in tank 3 is below 4000 lbs. indicated. If dumping
has been initiated, terminate dumping when fuel
level reaches 4000 lbs.
V
4-8
Under NORMAL OPERATION delete the WARNING.
5-10
Delete theAIR CONDITIONING SYSTEM limitation.
,.
>
Thp technical data information fdini...heci herein is iritencieci to be
used ag INTF;RIM data Only. will be replaced and ,super8eded
me of q3�B�u.'e�:'gt�-:,t:fiernext reyl ion to the fligIt manual
� " " -
_,AApproved for Release: 2017/07/25 006230172.4,
Approved for Release: 2017/07/25 C06230172
UVV.1. 'Mt 1.
..TECHNICAL DATA CHANGE -
FLIGHT -MANUAL,
TO .TIE.lNSEIZTED IN FRONT OF: A-12 Tre.iner
FLIGHT NIA NU AI, DA TED 1 November .1962
SECTION
PAGE
CHANGE
I
1-1
Under AIRCRAFT GROSS WEIGHT, delete both sentences and add a 11011
sentence to read: The approximate ramp gross weight of the aircraft
1Tith fuel load for present operating restrictions, water, two pilots
and equipment is 87,000 pounds.
vr
1
1-20
Under AFTERBURNER IGNITION SYSIEM, in 1st sentence change 91%to 93%.
1
1-25
Add new paragraph after Constant Speed Drive Oil Reservoir as follows:
Constant Speed Drive Oil Low Level Lights. Constant Speed Drive t-v
oil low level lights on the annunciator panel (labeled L OIL QTY
LOW and R. OIL QTY LOW) will illuminate when respective CSD oil level
has depleted below approximately I quart.
1
1-31
Under Inlet Air Bypass Door ST4tches, change 5th sentence to read
as follows: When the aft cockpit switches are placed in the OPEN
or CLCSED position they will override the forward cockpit switches.
i.-----
1
1-32
Under Emergency Snike Switches, in 2nd sentence delete the words
1
1-48
"and closes the bypass doors.
Under DC EIECTRICAL POWER SUPPLY in 3rd sentence delete the words,
1-49
"reverse current relay and."
1
1-79
Under MACE TRIM SYSTEM, revise 5th sentence as follows: The trim
system operates between 0.2 and 1.5 Mach number on a schedule of
approximately 80 per Mach number. It operates Only within the 871�
�
nose up and 50 nose down trim limits of the elevons. s..""
1
1-80
Under PINT-STATIC SYSTEM, revise 6th sentence as follows: The
other Set of pickups supplles the speed sensors on the ejection
seats, the altiraeters� TAB indicators, and irertical velocity
indicators. /.../.
1
1-82
Under LAMING GEAR SYSTEM, delete entire discussion down to the
paragraph on LANDING GEAR LEVERS and substitute with the following:
The tricycle type landing gear and the .main wheel well inboard doors
are electrically controlled and hydraulically actuated. The main
gear outboard doors and the nose gear doors are linked directly to
the respective gear struts. Each three wheeled min 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 unlock which engages the strut.
-,
NOTE : The technical data information furnished herein is intended to be
used as I.NTRIM data only. 'It w_ill,:he.iepladed and superseded
at the tirneotisane of the next revis ion to the flight mandal.
Approved for Release: 2017/07/25 CO6230172� '
Approved for Release: 2017/07/25 C06230172
Page 2 of 13
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A.,12 Trainer
FLIGHT MANUAL DATED 1 November 1962
SECTION
PAGE
CHANGE
I
1-82
There is no hydraulic pressure on the gear when it is up and locked.
(Contld)
Down locks inside the actuating cylinders hold the gear in place in
the extended position. Hydraulic pressure is on the gear in the
extended position when L system pressure is available. The
landing gear cylinders and doors are actuated in the proper order
by two sequencing valves. Normal gear operation is powered
hydraulically by the L hydraulic pump on the left engine. Should
pressure drop to 1200 psi during retraction, the power source
automatically becomes the R hydraulic pump. R hydraulic pressure
will not extend the gear in the event of the L system failure and
the marmal landing gear release must be used.
I
1-84
Under Manual Landing Gear Release Handles, delete the 2nd sentence
and substitute with the following: If the L hydraulic system 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.
1
146
Under NOSE WHEEL STEERING SYSTEM, delete the words, "nose wheel"
at the end of the let sentence and in the NOTE and substitute
the words, "main gear." e,,,--
1
1-87
Under WHEEL BRAKE SYSTEM, revise as follows:
1. 2nd sentence should be changed to read: ,
Depressing the rudder pedals actuates 4-rotor brakes on
each of the six main gear wheels. 4/-
2. Delete the 3rd and 4th sentences. k----
1
1.-87
Under Brake Switches, revise as follows:
1. In the 2nd sentence, change the word SKID OFF tb.NORMAL.
2. In the 4th Sentence change the words EMER BRAKES to
ALTERNATE. i,...-
3. Change the 5th sentence to read:
When the aft cockpit switch is placed in the ANTI-SKID or
ALTERNATE position it is capable of overriding the forward cockpit
switch. �/""
NOTE : The technical data information furnished herein is intended to be
used as INTERIM data only. It will be repla.ced and superseded
at the time of issue of the next revision to the flight manual.
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
Page 3 �f.!_
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED 1 November 1962
SECTION
PAGE
CHANGE
1
1-89
Under Anti-Skid Out Indicator Lights in 2nd sentence change the
words SKID OFF to NORMAL and DER BRLES to ALTERNATE. /..../
1
1-90
Under DRAG CHUTE SYSTEM, revise 1st sentence on page to read:
The chute mechaniam incorporates a sbear section in the yoke which
ruptures if the chute is deployed above limit airspeed. Refer
to Section V for further information. 2.,---
1
,
1-90
Under Drag Chute Switches, revise as follows:
1. Change 3rd sentence to read:
Deployment is accomplithed by placing either switch to the
DEPLOY position.
2. Add a new 6th sentence to read:
The aft cockpit switch is capable of overriding the
forward cockpit switch. i�..7'
3. Delete the WARNING. tV
1
1-106
Under EJECTION SEQUENCE, after the 1st sentence add a note to read:
The ejection seat cannot be fired until the canopy jettison system
has fired. This design feature is necessary to prevent pilot
ejection thru the metal canopy.
1,,,-
1
Replace the following illustrations with new attached illustrations:
GENERAL ARRANGEMENT Figure 1-1 ''Page 1-2
LOWER INSTRUMENT PANEL Figure 1-6 ''Page 1-13
1./
'Page
AFT COCKPIT - LEFT SIDE Figure 1-9 1-16
AFT COCKPIT - RIGHT SIDE Figure 1-10 v/Page 1-171.
FUEL SUPPLY SYSTEM Figure 1-13 Page 1-34.
----
ELECTRICAL POWER DISTRI .BUTION Figure 1-15 Page 1-47
,..----/
CIRCUIT BREAKER PANELS Figure 1-16 'I'age 1-50
"
BRAKE SYSTEM , Figure 1-25 Page 1-88 ...
_
./.
CANOPYS AND CONTROLS Figure 1-26 Page 1-1001--'
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.
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
� Page 4 of 13
TDC NO. 1
TECHNIQA . DATA' CHANGE
FLIGHT.MANUAL
TO TIE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED 1 November 1962
SECTION'
PAGE
CHANGE A -
II
2.4
Under Instrument Panel, revise as follows'.2..5
1. Change Change step 0 read: Brake switak... NORMAL 4�7-
-
to: ,
2. Change step 7 to read: Forward 0O0 it temperature
rheostat -r 12 Oclock position.- t�,/' T-
1
3. Change step 9 to read: Aft -cockpit temperature rheostat -
,
12 o'clock position. c/ .. -\ -
4., Change step 10 to read: Aft cockpit air system switch - ON
5* Change step 12 to read: Cockpit system crossover switch -
i./
6 Ada a Step 29 to read: Battery switch - OFF
II
2-6
Under Lower Instrument Pane4 revise as follows;
1. Delete step
2. Renumher step 8 to step '7., V---
II
2-8
Dnder Instrument Panel, revise As follows:
2.9
2711
1. Delete the words "at 4900C" from step 2.
2. Change step 6 to' read: Brake switch'.. NORMAL P'---
3. Change step 10 to read: a. Forward cockpit .....0N
b. Aft cockpit-- ON 1.---
;.4. Change step 13 to tei4: Cockpit system Crossover
switch... ON. t.,..
_
-5. Delete step 40i
11
2-43
-Under :STARTING ENGINES, revise as follows:
2..44
1 . Change step 2 to read.; Boost pumps -jCbeCk tanks 1 2 .
and 6-indiCator lights .(green) illnriineted'. .1.-"-.. '
2. Delete note- tn4ftt. step 2.
NOTE The : 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.
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
Pagea_of 13
TDCNO 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED 1 November 1962
SECTION
PACTE
CHANGE
II
2-16
Under ENGINE CHECKS, delete 1st two sentences of CAUTION. fr-----
II
2-16
Under EMERGENCY ring. $YSTEM CHECK, revise as follows:
1. Change step 3 to read: Tachometer - Check that tachometer
stabilizes at a new value. /...----
2. Delete the 2nd Sentence of NOTE. 17-
II
2-17
Under BEFORE TAXIING, revise as follows
2-18
1. Change step 11 to read: Inlet air bypass doors AUTO
(Ground crew will check doors open). p...."'`
2. Change step 16 to read: Brake switch - NORMAL
II
2-20
Under PRE-TAKEOFF AIRCRAFT CHECK, delete the WARNING. -1,-
II
2-21
Under PRE-TAKEOFF ENGINE CHECK, change step 5 to reads Tanks
2 and 6 - Check lights ON.
11
2-27
Under CLIMB, add a new sentence after the 3rd sentence to read:
Begin the rotation sufficiently in advance of reaching climb
Speed to avoid exceeding the recommended airspeed schedule.
11
2-30
Under DESCENT, add a NOTE after step 2 to read: The landing gear
may be lowered in order to increase the rate of descent, provided
that the airspeed is first reduced below 300 KEAS (gear limit speed).
-
II
II
2..30
Under BEFORE LANDING, revise as follows:
2-31
1. Add a new step 3 to read: Cross feed switch - Depress
(Check light . ON) L-
2. Renumber existing steps 3 and 4 to 4 and 5.v/
3. Add a 2nd sentence to NOTE under new step 4 to read:
Forward fuel flow during transfer may he increased by holding the
aircraft in a moderate nose down attitude. V�
114 Add a new step 6 to read: Cross feed switch -0 Depress
(Check light - OW
nil 1,
NOTE i 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.
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
Page 6 of 13:
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED 1 1\14:Walther 1962
SECTION
PAGE
CHANGE
11
2-30
5. Change new step 8 to read: Brake switch � NORMAL. 6/'
2-31
(Contld)
6. Renumber existing steps 5 thru 18 to 7 thru 20. 1/
7. Under hew step 17 add a sentence to read: Normal gear
extension time is approximately 16 seconds.
/I
2-34
Under NORMAL LANDING, revise as follows:
1. Delete 2nd sentence of step 4. 1../
�
2. Change 3rd sentence of step 4 to read: Steering will not
engage until rudder pedals align with nose wheel position
(straight ahead) and weight of aircraft is on the main gear. ir-
3. Add a step 7 to read: Drag chute . JETTISON V
�
4. Below new step 7 add a CAUTION to read:
�
CAUTION
WheneVer possible, the drag chute should be jettisoned above
20 mph. This provides sufficient pull for the socket to open
and the ball to clear the aircraft. Below this speed the ball
may damage the upper fuselage. If the chute is not jettisoned,
the elevons should not be moved during taxiing since the shroud
�
lines could jam between the inboard elevons and the fuselage and
cause structural damage. t.'
II
2.34
Add a new paragraph before CROSSWIND LANDING to read as follows:
,
AFT COCKPIT LANDING TECHNIQUE
�
Approach apeint 1/4 to 3/a of a nile from the end of the runway
at approximately El. 30 degree angle. As this point is approached,
start a shallow turn to line up with the runway. Pick up each
side of runway as reference points. Point of vision should be
approximately 15 to 20 degrees to either side of a point dead
ahead. Flare and touchdown are normal. 1.....
I es"
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.
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
Page 7
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT -MANUAL
TO BE. INSERTED IN FRONT OF A-12Tr�ir
FLIGHT .M.A.N11.AL DATED. 1 Novet*er 1962
SF,C;TION
PAGE
CHANGE
II
2,-34
Add a new paragraph after CROSSWIND LANDING to read as follows:
GCA APPROACH AND LANDING
The following procedure is reconnaended when msktng a GCA approach
and landing.
1. Approach the radar pickup point in a clean configuration.
2. Adjust throttles to establish 250 ICUS on base leg.
3. Lower landing gear on base leg, maititaining 250 MS.
4. Decrease airspeed to 230 KIAS during turn on final approach.
5. Decrease airspeed to 180 KIAS minimum on final approach.
6. Adjust throttles to maintain 1200 1400 feet per minute
on glide slope. i-
II
2-38
Under AFTER LANDING, revise as follows:
1. Delete step 1. t./'
2. Renumber steps 2 thra 6 to 1 thru 5, L---'
II
2-38
Under ENGINE SHUTDOWN, revise as follows:
1. Delete existing steps 4 and 5. '--
2. Add -the words "provided that the starting csrts are
connected" to 1st sentence of NOTE. v---- ,
3.. Change existing step 7 to step 4. t-v
4. Change existing step 9 to step
5. Change existing step 8 to step 7.
II
ABBREVIATED CHECKLIST
,
Al]. numbered check list items changed in this section must also-
be changed it the pilot's abbreviated checklist.
NOTE : The technical dat�nformation furnished herein DA 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.
intIL Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
Page 8 of 13
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A,-12. Trainer
FLIGHT MANUAL DATED 1. November 196
SE("rJON
PAGE
CHANGE
III
3-.6
Under ENGINE FAILURE DURING FLIGHT, revise as follows:
1. Add a step 6 to read: Inlet air bypass doors - OPEN
(To reduce buffeting) is
2. After,new step 6 add: If the left engine fails, V--
3. After above statement add a new step 7 to read: Cockpit
system crossover switch - CROSSOVER t,,--
III
3-7
Under MEDIATE AIRSTART ATTEMPT, in step 2, delete all -words in
the parenthesis after the word IDLE.
III
3-10
Under NORMAL AIRSTART, revise as follows:
1. Renumber existing step 4 to step 5. V/
1/-
2. Renumber existing step 5 to step 4.
111
3-13
Under LANDING WITH ROTH ENGINES INOPERATIVE, delete the words I-7
"is not recommended" and replace with "should not be attempted."
III
3-24
'Under EJECTION, item d, change * to 2. 1'
III
3-25
Under EMERGENCY ESCAPE IN FLIGHT, revise as follows:
1. Change step 3 to step 4. v"-
\
2. Change step 4 to ster, 3 1"--
3. After step 7 insert a new paragraph to read: ''
DELAYED EJECTION
If an emergency arises (smoke or fire in cockpit, etc.) which
requires jettisoning the canopy with a subsequent need for the
pilot to eject, proceed as follows: .
,
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.
Approved for Release: 2017/07/25 C06230172
inok
Approved for Release: 2017/07/25 C06230172
Page 9 of 13
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED. 1 November 1962
SECT ION
PAGE
CHANGE
III
3-25
1., Canopy - Jettison
(Contld)
If the canopy does not jettison, _
2. Canopy latch haulle Pull
3. Pull ejection seat "D" ring with both hands when or
if necessary. %,--
III
3-35
Under SPIKE INLET CONTROL SYSTEM MALFUNCTIONS, change the WARNING
to read: If the emergency helium system is used, the spike will
move forward and its position cannot be changed for the remainder
of the flight. t7
III
3-36
Under IMPROPER FUEL SEQUENCING, revise as follows:
3-37
1. Delete Delete 3rd sentence from WARNING.
2. Delete NOTE. yr"'
III
3-42
Under UTILITY HYDRAULIC SYSTEM FAILURE revise as follows:
1. In 2nd sentence change 1900 psi to 1200 psi. P'.-
2. In the 4th sentence add: "the L water injection system"
after "aerial refueling system."
3. In the 5th sentence add: "R water injection system"
after "right inlet spike.' ii,---
III
3-51
Under LANDING GEAR EMERGENCY EXTENSION, revise as follows:
1. Add a new step 1 to read: Landing gear CONT circuit
breaker - Pull. p''''
2. Renumber existing steps 1 and 2 to 2 and 3.
III
3-52
Under BRAKE SYSTEM EMERGENCY OPERATION change step 1 to read:
irake switch - ALTERNATE. br
III
-
ABBREVIATED CHECKLIST -
7
,
All numbered Checklist items changed in this section must also
be changed in the pilot's abbreviated checklist.
nel
NOTE : The technical data information furnished herein is intended to be
used as INTEIPM.data only, It will be replaced and superseded
at the time of issue of the next revision to the flight manual.
Approved for Release: 2017/07/25 C06230172',
Approved for Release: 2017/07/25 C06230172
Page 10 of 13
TDC NO. 1
TEQHNICA.L. DATA 'CHANGE
:FLIGHT 'MANUAL:
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED 1 November 1962
SECTION
PAGE
CHANGE
IV
4-9
'Under DEFOG SYSTEM, add the following after the NOTE:
,
Defog St4tches
Defog switches are located on the left side of each
instrument panel. The switches have three positions;
ON (up), neutral (center), OFF (down), and are spring
loaded from the neutral to the ON position. Each switch
controls a motor driven shut off valve in the respective
eockpit defog duct. When the switch is held in the ON
position, the valve moves toward full open. Releasing the
switch to the neutral position stops the valve travel.
When the switch is placed in the OFF position the shut off
valve travels to the full closed position. Power for the
switches is furnished by the essential dc bus. 1,...--""
IV
4-20
Figure 4-6 (IFF and SIP CONTROL PANELS)
Replace existing illustration with new attached illustration.
IV
4-22
Under Coder Group (SIP) Control Panel, change the 2nd sentence
to read: The panel is installed on the forward cockpit right
console.
IV
4-23
Under AN/AIC-10 INTERPHONE CONTROL PANYii, change 1st sentence to
read: Both AN/AIC-10 interphone control panels are located on a
shelf inside of lower batch below the aft cockpit. 4.---
IV
4-49
Under posgtion of the MA-1 CompaSs System, revise as follows:
I. Delete step 1. (..--
IV
4-50 4-50
2. Change steps 2 tbru 5 to l thru 4.
Under INERTIAL NAVIGATION SYSTEM, change last sentence of INS
description to read: Power for the system is furnished by
the No. 3 inverter, the LH generator and the monitored de bus.
IV
4-63
Under Inadvertent Selection of Present Destination, revise
4-64
as follows:
1. Delete lest sentence on page 4-63. 17
2. Delete items 1, 2 and 3 on page 4-64.
NOTE : The technical data information furnished herein is intended to be
used as INTERIM data only. It will be replaced and superseded
at time of issue of the nea.ci revision to the flight manual. �
. Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
Page 11 of 2.1_
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED I November 1962
SECTION
PAGE
S CHANGE
V
5-11
Under Landing, delete existing text and replace with the following:
Normal Landing
There is no limitation on airspeed for brake application at
normal landing weights when the drag chute is used. The maximum
speed for brake application without the drag chute is 125 KIAS
in calm air, or 125 KIAS plus the runway wind component when
headwinds exist.
V
5-11
Under Aborted Takeoff, delete existing text and replace with
the following:
At 85,000 pounds the maximum airspeed for brake application is
approximately 100 KIAS in calm air without the drag chute. The
airspeed is 135 KIAS when the drag chute is used. These airspeeds
can be increased by an amount equal to the runway wind component
when headwinds exist) but they will burn out before the aircraft
is stopped if applied at higher speeds.
,
1nnl.
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.
� Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
Page 12 of
TDC NO. 1
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer
FLIGHT MANUAL DATED 1 November 1962
SECTION
PAGE
CHANGE
VI
6-3
Under CLIMB, revise as follows:
1. Add the words �particularly when piloting from the
aft cockpit� at the end of the 3rd sentence./.�/
2. Add a new sentence after the 4th sentence to read:
An increase in the level of wind noise may be noticed in the
forward cockpit at 380 KEAS./
VI
6-6
Under SUPERSONIC ACCELERATION revise as follows:
1. Change last sentence to read: The throttles should
be retarded very slowly at the end of acceleration to avoid
afterburner flameout and engine stall. L.----
2. Add the following sentence at the end of above sentence:
Yaw maneuvers should not be made at maximum speed unless the yaw
damper is on. Return to neutral after rudder kick is slow with
the yaw damper off. 1,'
VI
6-7
Under APPROACH AND LANDING revise as follows:
6-8
,
1. Add the following sentence after the 2nd sentence:
The transfer rate is subptantially higher during descent than
during level flight. V"
2. Change first sentence on page 6-8 to read: Simulated
and actual single engine landings can be made using the
procedures outlined in Section III. pe"..'
3. Change 4th sentence on page 6-8 to read: Normally,
adequate control is available, but caution should be observed
and higher than normal approach speeds used when landing in
extremely turb ent air where maximum control rates may be
required.
VI
6-12
Under STABILITY AND CONTROL CHARACTERISTICS, change 7th sentence
on page 612 to read: The breakout forces are considered to be
exceptionally good, altho h they are somewhat greater than for
single place aircraft.
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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Page 13 of
TDC NO.2
TECHNICAL DATA CHANGE
FLIGHT MANUAL
TO BE INSERTED IN FRONT OF A-12 Trainer.
FLIGHT MANUAL DATED 3- November 1962
SECTION
PAGE
CHANGE
VT
./Appendix
6_15
Replace the-following illustration with new attached illustration:.
oops WEIGHT-CENTER OF.GRAWTY VARIATION, Figure 6-Page 6-13.
Replace exiSting'4pendix with ew attached Appendix- (Parts I
thru IV)
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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TA-12
SECTION I
DESCRIPTION
TABLE OF CONTENTS
Page
Page
The Aircraft
1-2
Instruments
1-52
Engine (J-75)
1-2
Emergency Equipment
1-56
Afterburner System
1-8
Landing Gear System
1-57
Engine Air Inlet System
1-19
Nosewheel Steering System
1-59
Fuel Supply System
1-21
Wheel Brake System
1-60
Air Refueling System
1-28
Drag Chute System
1-60
Electrical Supply System
1-29
Air-conditioning and Pressurization
Hydraulic Power Supply Systems
1-34
System
1-62
Flight Control System
1-37
Oxygen System and Personal
Equipment
1-65
Automatic Flight Control System
1-45
Windshields
1-69
Stability Augmentation System
1-45
Canopies
1-70
Mach Trim System
1-49
Ejection Seats
1-72
Pitot-Static System
1-50
Ejection Sequence
1-75
Air Data Computer
1-50
1-1
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SECTION I TA-12
THE AIRCRAFT
The A-12 trainer is a delta-wing, two-place
aircraft powered by two axial-flow turbojet
engines with afterburners. The aircraft is
designed to operate at high altitudes and
supersonic speeds. Some notable features
of the aircraft are very thin delta wings,
twin canted rudders mounted on top of the
engine nacelles, and a pronounced fuselage
chine extending from the nose to the leading
edge of the wing. The surface controls,
comprising the elevons, and twin rudders,
are operated by irreversible hydraulic ac-
tuators with artificial pilot control feel. A
single-point pressure refueling system is
installed for both ground and in-flight re-
fueling. A drag chute is provided to reduce
landing roll.
The following switches and instruments are
installed in this aircraft to simulate their
arrangement and location in the A-12 air-
craft. They are not wired for use. Refer
to the A-12 Utility Flight Manual for infor-
mation regarding their operation and pur-
pose.
Spike switches - located on the instrument
panel in forward cockpit.
Dual spike indicator - located on the instru-
ment panel in forward cockpit.
Dual nozzle position indicator - located on
the instrument panel in forward cockpit.
Inlet aft bypass switches - located above the
throttle quadrant in each cockpit.
Restart and forward by-pass switch - located
on the right throttle in each cockpit.
Dual compressor inlet pressure gage - lo-
cated on the instrument panel in forward
cockpit.
Quad-hydraulic quantity gage - located on
the instrument panel in forward cockpit.
AIRCRAFT DIMENSIONS
The overall aircraft dimensions are as
follows:
Wing span
Length
Height
Tread
55.62 feet
98.75 feet
18.45 feet
16.67 feet
AIRCRAFT GROSS WEIGHT
The approximate ramp gross weight of the
aircraft, with fuel load for present operating
restrictions, INS, two pilots and equipment
is 85,000 pounds.
ENGINE (J-75)
The aircraft is powered by two Pratt &
Whitney J-75-19W(S)A turbojet engines
equipped with afterburners. This engine
has a sea level standard day installed static
thrust rating of approximately 19,500 pounds
at Maximum thrust. The engine is a con-
tinuous-flow gas turbine, incorporating an
eight-stage low pressure compressor, a
seven-stage high pressure compressor,
eight radially positioned combustion cham-
bers, a split three-stage turbine, and an
afterburner incorporating a two-position
exhaust nozzle. The rotor systems are
mechanically independent of each other.
The high pressure compressor is driven
by a hollow shaft from the first stage tur-
bine wheel. The low pressure compressor
is driven from the second and third-stage
turbine wheels by a shaft rotating within
the hollow, high-pressure compressor shaft.
The throttle controls the rpm of the high
pressure rotor only. The main engine ac-
cessory section is gear-driven from the
high pressure rotor and provides reduction
gearing and mounting pads for the engine.
fuel control, the fuel pump, the oil boost
pump, the tachometer generator, two hy-
draulic pumps, and the external starter
drive.
1-2
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SPIKE BLEED AIR OUTLET �
INLET BYPASS AIR OUTLET �
SPIKE �
ROLL RATE GYRO AND
YAW ACCELEROMETER �
IFR RECEPTACLE DOORS
AIR CONDITIONING BAY
a SR-3 FLIGHT REFERENCE SYSTEM
b INERTIAL NAVIGATION COMPONENTS
TACAN ANTENNA
UHF ANTENNA�
ADF LOOP ANTENNA�
BOOM
EJECTION SEATS
RUDDER�
DRAG CHUTE
RECEPTACLE�
OUTBOARD ELEVON-
-PITCH AND YAW RATE
ACCELEROMETER
EXTERNAL POWER RECEPTACLE
BATTERY
TACAN ANTENNA
�LANDING LIGHT
� E BAY
a AUTOPILOT COMPONENTS
b BACK-UP PITCH RATE GYRO
c AIR DATA COMPUTER
d STABILITY AUGMENTATION SYSTEM COMPONENTS
IFF
ANTENNA
ADF SENSE
ANTENNA
rn
�INBOARD ELEVON rT1
ROLL AND PITCH MIXER
1:6
YAW SERVOS
RUDDER TRIM
XI�
0
EJECTOR FLAPS m
TERTIARY DOORS-
ELEVON ACTUATORS
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SECTION I TA-12
ENGINE FUEL CONTROL SYSTEM
Desired engine rpm is established by throttle
movement. The fuel flow to the engine is
regulated by the engine fuel control system.
Note
The engine fuel control systems
are identical and completely
separate for each engine.
The system includes the engine-driven fuel
pump unit, the engine fuel control unit, the
fuel pressurization and dump valve unit,
and the afterburner fuel system.
Engine-Driven Fuel Pump Unit
The engine-driven fuel pump unit contains
four pumps. It supplies fuel, at the pres-
sure required, to the engine and after-
burner systems. A centrifugal pump re-
ceives fuel from the airplane fuel system
and forces it into three gear-type pumps.
One gear-type pump is the engine stage fuel
pump; the other two are the afterburner
stage fuel pumps. The engine stage fuel
pump furnishes fuel to the engine fuel con-
trol unit which regulates fuel flow to the
combustion chambers. The afterburner
stage fuel pumps furnish fuel to the after-
burner fuel control which regulates fuel
flow to the afterburner when afterburner
operation is selected. When the afterburner
is not operating, the output from the after-
burner stage fuel pumps is returned to the
discharge stream of the centrifugal pump.
If the engine pump output pressure drops
below approximately 50 psi, the emergency
transfer valve in the engine-driven fuel
pump unit automatically opens to allow fuel
from the output side of one of the after-
burner fuel pumps to flow to the hydrome-
chanical fuel control unit.
Note
There is no direct indication of
engine fuel pump failure.
During this condition, fuel can be supplied
to both the engine and the afterburner sys-
tems, but may be insufficient for full after-
burning thrust at low altitude. No thrust
loss will occur in the MILITARY thrust
range.
Engine Fuel Control Unit
The engine fuel control unit regulates fuel
to the combustion chambers and incorpo-
rates normal and emergency fuel control
systems. The normal fuel control system
contains a mechanical computer, a gover-
nor, and temperature and pressure sensing
elements which control the main throttle
valve. The computer, in addition to sensing
throttle position, senses changes.in flight
conditions and regulates fuel flow to insure
optimum engine operation for the selected
thrust setting. During rapid engine accel-
erations, the normal fuel control system
regulates fuel flow to prevent overspeed,
overtemperature, compressor stalls, and
flameouts. The normal fuel control system
also maintains a minimum fuel flow to pre-
vent engine flameout at high altitudes and
during rapid decelerations. When the
throttles are retarded to OFF, a mechan-
ically controlled cutoff valve in the fuel
control unit cuts off all fuel to the combus-
tion chambers. The emergency fuel control
system provides an alternate system of re-
gulating fuel flow to the combustion cham-
bers in event of failure of components with-
in the normal system. The emergency fuel
system is capable of supplying sufficient
fuel to obtain at least 95% Military thrust
on a 100oF day at low altitudes and at least
80% Military thrust at altitudes up to 30,000
feet at standard, plus 40 F ambient con-
ditions.
1-4
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TA-12
SECTION I
ENGINE FUEL AND IGNITION SYSTEM
FROM
FUEL SYSTEM
FUEL PRESS LOW
MR 1--
w
au,
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GRD c..,
ON
AFTERBURNER
SW ITCH
OFF
CIT
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FUEL
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HYDRO
DRIVEN
PUMP
FILTER
ENGINE DRIVEN
FUEL PUMP UN IT
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MAIN ENGINE FUEL
CONTROL UN IT
BOOST PUMP
N. a,* ��� are ittr a. a. IN* �Aw ��� ����� %Iv 4, NI. �� ��� AB RANGE
POWER LEVER LINKAGE % 4.�
01 EXCHANGERFUEL-0 IL HEAT
EMERGENCY
TRANSFER
SOLENOID
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ID LE
Cur-OFF
AB SW ITCH
EMERGENCY
EMERGENCY FUEL
CONTROL SW ITCH
NORMAL
PRESSURIZING
AND
DUMP VALVE
BURNER
PRESS.
PRIMARY MANIFOLD
SEDCONDARY MANIFOLD
DUMP
OVERBOARD
AB FUEL
CONTROL
2
AB
IGNITER
AB I GN. TUBE
#3 BURNER CAN
AB SPRAY
BARS
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N GEARBOX
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OIL TEMP
COMPOSITOR
47407�74!������,�:%!���������7�74vii
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IGNITION PLUG: (2)
I IN 4 & 5 CANS
TACH'
�
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ION IN. =NENE.
AB
ACTUATOR
MOTOR
EXHAUST
NOZZLE
CONTROL
PC
CLOSE
I I
OPEN
ACTUATOR
F201-16 (6)
Figure 1-2
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1-5
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SECTION I TA-12
Note
When operating on the emergency
fuel system, full obtainable thrust
may result in lower rpm and ex-
haust gas temperature than that of
normal fuel system operation.
The engine may be started on the emergency
system, either in flight or on the ground.
The afterburner may be operated normally
on the emergency fuel system.
CAUTION
. The emergency fuel control system
should be used only during an
actual in-flight emergency or
flight check.
When operating on the emergency
fuel control system, rapid throttle
movements must be avoided to
prevent overspeed, overtemperature,
compressor stalls, and flameouts.
Fuel Pressurizing and Dump Valve
The fuel pressurizing and dump valve is lo-
cated in the fuel control system between the
fuel cutoff valve and the combustion cham-
bers. The unit controls fuel flow to the pri-
mary and secondary injector nozzles in the
engine combustion chambers. To facilitate
starting, fuel at relatively low pressure is
directed through the primary manifold, and
spring tension on the pressurizing valve
keeps the port to the secondary manifold
closed until increasing engine speed builds
up fuel pressure high enough to overcome
the spring tension and open the valve. When
this happens, fuel flows through both pri-
mary and secondary manifolds to the com-
bustion chambers. When the engine is to be
shut down, the cutoff valve in the fuel con-
trol unit is closed by throttle movement and
the dump valve automatically opens to per-
mit residual fuel in the manifolds of the
main combustion system to drain overboard.
Emergency Fuel Control Switches
The two-position emergency fuel control
switches are located on the left console in
each cockpit. In the NORM (aft) position,
the normal fuel control is in operation.
The EMER (forward) position is used in the
event of failure of the normal system, and
fuel flow to the engine is controlled by a
separate emergency throttle valve connected
directly to the throttle. Power for the cir-
cuits is furnished by the essential dc bus.
THROTTLES
Two throttle levers, one for each engine,
are located in a quadrant on the left forward
console in each cockpit. The forward cock-
pit and aft cockpit throttles for each engine
are interconnected and mechanically linked
to the engine fuel control units, which di-
rectly govern engine thrust. The quadrant
positions are labeled OFF, IDLE, and
AFTERBURNER. Moving the throttles for-
ward from OFF to IDLE mechanically opens
the fuel cutoff valve. At the IDLE position,
each throttle drops over a hidden ledge,
which prevents inadvertent engine cutoff
when the throttles are retarded to IDLE.
Slow throttle movement ensures that the
throttles will follow the curve of the quad-
rant. When returning the throttles from
IDLE to OFF, the throttles must be pulled
upward in order to clear the ledge. Above
IDLE, forward throttle movement past a
raised detent indicates that the afterburner
micro-switches inside the throttle quadrant
are about to close.
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TA-12
SECTION I
THROTTLE QUADRANT (Both Cockpits)
RESTART AND FORWARD
DOOR SWITCH
OFF STOP
FORWARD COCKPIT - AFT COCKPIT
LEFT THROTTLE
RIGHT THROTTLE
TRANSMIT SWITCH
FRICTION LOCK LEVER
MAX A/I3 STOP
F201-4(a)
Figure 1-3
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SECTION I TA-12
Throttle Friction Levers
The throttles are prevented from creeping
by interconnected friction levers, located
on the inboard sides of the throttle quadrants.
When the levers are fully aft, the throttles
are free; moving the levers forward pro-
gressively increases the amount of friction
to hold the throttles in position.
AFTERBURNER SYSTEM ,
The afterburner fuel control meters fuel
flow to the afterburner spray bars on de-
mand, as a function of engine burner pres-
sure. The control incorporates a metering
valve, shutoff valve, pressure regulator
bypass valve, and a burner pressure me-
chanical metering linkage. Thrust with
afterburner can be varied approximately 50
percent through throttle modulation between
minimum and maximum afterburner position.
AFTERBURNER IGNITION SYSTEM
The afterburners are ignited electrically at
engine rpm above approximately 93 percent
rpm by moving the throttles through the de-
tent to the AFTERBURNER position, and
then placing the afterburner switches to the
ON position. A "hot streak" igniter valve
supplies a streak of burning fuel which
passes through the turbines and ignites the
afterburner fuel. The igniter valve also
recirculates fuel when the afterburner is
shut off.
Afterburner Switches
Each afterburner is actuated by individual
afterburner switches located on the instru-
ment panel in each cockpit. Either the for-
ward or aft afterburner switches will initiate
afterburner operation. In the ON (up) posi-
tion, the afterburners will light when either
throttle is forward of the afterburner micro-
switch setting on the throttle quadrant.
Afterburner operation is terminated by mov-
ing the forward and aft cockpit afterburner
switches to the OFF (down) position.
Note
Both forward and aft cockpit
afterburner switches must be
moved to the OFF position to
terminate afterburner operation.
Power for the circuit is furnished by the
essential dc bus.
Afterburner Emergency Shutoff
Should either of the electrically operated
afterburner switches (afterburner switch or
throttle microswitch) fail to terminate after-
burner operation, an afterburner shutdown
may be accomplished by retarding the re-
spective throttle to a point slightly aft of
the afterburner detent. The afterburner
fuel control shutoff valve is manually
closed and shuts off all fuel to the after-
burner.
Note
If afterburner operation is ter-
minated by manually closing the
afterburner fuel shutoff valve in
the manner described above and
electrical power has been re-
stored, the afterburner switch
must be recycled after 5 seconds
in order to regain afterburner
operation.
EXHAUST NOZZLE SYSTEM
The engines are equipped with an iris type,
two-position afterburner primary nozzle
comprised of segments which are operated
by a cam and roller mechanism and pneu-
matic actuators. The primary nozzle is
enclosed by a fixed contour convergent-
divergent ejector nozzle followed by free
floating trailing edge flaps. Secondary air
is provided by the inlet and bypasses around
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.TA-12 SECTION I
the engine to cool the engine and ejector. A
series of free floating doors ahead of the
nozzle provide tertiary air into the nozzle.
The trailing edge flaps and tertiary doors
open and close with varying internal nozzle
pressure which is a function of Mach number
and engine power. The tertiary doors will
open to provide additional air which is re-
quired during low speed flight.
ENGINE OIL SUPPLY SYSTEM
The engine oil system is a dry sump recir-
culating pressure type system from a 5.5
US gallon oil tank mounted on the left side
of the engine compressor section. Usable
oil capacity is 4.5 US gallons. Oil flows
from the tank to a gear-type boost pump,
then through a fuel oil cooler to the main
pump which supplies pressure through the
main oil strainer to the engine gears and
bearings. The strainer is equipped with a
full flow bypass valve. Engine main oil
pressure is governed by a pressure regulat-
ing valve located downstream of the filter.
An oil scavenging system with four scavenge
pumps returns oil to the tank. An engine
oil breather pressurizing valve (aneroid
type) regulates pressure in the bearing com-
partments, breather system and oil tank.
The valve is open at sea level and regulates
to hold a constant breather system altitude
of approximately 35,000 + 4000 feet when
aircraft operation is above 35,000 feet. The
breather system vents overboard.
Engine Oil Temperature Lights
Two oil temperature lights, labeled L OIL
TEMP and R OIL TEMP, are located on
each annunciator panel. These lights will
illuminate when engine oil temperature is
less than 4�C + 3�C or greater than 282oC
+11�C.
CONSTANT SPEED DRIVE UNIT (CSD)
A constant speed drive unit (CSD) mounted
on the front of each engine is driven by the
low pressure rotor. The unit converts the
variable speed of the rotor to maintain
constant speed rotation of the A. C. gen-
erator. The CSD consists of a hydraulic
pump, separate reservoir, constant dis-
placement hydraulic motor which turns
faster or slower as the pump forces more
or less oil into it, and a governing system
which controls pump flow, thereby control-
ling the speed of the motor.
Constant Speed Drive Oil Reservoir
The CSD unit has a separate pressurized
oil reservoir which supplies oil to the hy-
draulic pump. Reservoir capacity is 9
quarts, with a normal operating fluid level
of 7.2 quarts.
Constant Speed Drive Oil Pressure Low Lights
Two CSD oil pressure low lights are in-
stalled on the annunciator panel in each
cockpit. The lights are labeled L CSD OIL
PRESS LOW and R CSD OIL PRESS LOW
and illuminate whenever the respective
CSD oil pressure is less than approximately
125 psi.
IGNITION SYSTEM
Individual ignition systems are installed on
each engine. Each system has two separate
exciter units and each exciter feeds an in-
dividual iv-liter plug. The igniter plugs are
located in No. 4 and No. 5 burner cans. A
single exciter is sufficient for making an air
or ground start.
1-9
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SECTION I
TA-12
Engine Start Switches
The engine start switches for both engines
are located on the lower left side of each
instrument panel. Each toggle switch con-
trols the ignition for one engine. The
switches are momentary contact, three-
position switches with a center OFF position.
In the GRD (down) position, ignition power
for both exciters iS furnished through a
single 15-ampere circuit breaker, from the
essential dc bus. In the AIR (up) position,
ignition power is furnished to each exciter
from the battery bus through separate 10-
ampere circuit breakers. The aft cockpit
switches are capable of overriding the for-
ward cockpit switches.
STARTER SYSTEM
An air turbine starter is provided for
ground starts. An external air supply fur-
nishes the necessary power. There are no
aircraft controls over this system, being
turned on and off by the ground-crew ac-
cording to pilot signals. Air starts do not
require a starter but are made by a wind-
milling engine.
ENGINE INSTRUMENTS AND INDICATOR LIGHTS
Exhaust Gas Temperature Gages
Two exhaust gas temperature (EGT) gages,
one for each engine, are mounted on the
right side of each instrument panel. They
are calibrated from 0 to 1200 C and indi-
cate the temperature sensed by the turbine
discharge thermocouples. The four digital
windows at the top of the gages indicate the
exhaust gas temperature to the nearest de-
gree. An ON-OFF window at the bottom of
each dial indicates instrument operational
status. Power is furnished by the No. 1
inverter.
Fuel Flow Indicators
Fuel flow indicators, one for each engine,
are mounted on both instrument panels.
The indicator dial is calibrated in incre-
ments of 2000 pounds per hour to 76,000
pph. A digital indication is also provided
by each indicator in a center window which
shows fuel flow to the nearest 100 pph.
Power for the indicators is supplied from
the No. 1 inverter.
Tachometers
Two tachometers, one for each engine, are
mounted on the right side of each instru-
ment panel. The tachometers indicate per-
centage of high pressure rotor rpm based
on 8732 rpm as 100 percent. The main
pointer is calibrated to 100 percent rpm
and the subpointer makes one complete re-
volution for each 10 percent change in rotor
rpm. By using the subpointer, up to 110
percent rpm can be read. The tachometers
are self-energized and operate independently
of the aircraft electrical system.
Engine Oil Pressure Gages
Two oil pressure gages are provided, one
for each engine, on the right side of each
instrument panel. The gages indicate out-
put pressure of the respective engine oil
pump. The gages are calibrated from 0 to
100 psi in 5-psi increments. Power for the
gages is furnished by the inverter No.3 bus
through the 26-volt auto-transformer.
Compressor Inlet Temperature Gages
A dual-indicating compressor inlet tem-
perature gage is located on the right side
of each instrument panel. The forward and
aft cockpit indicators are independent of
each other. The gages are calibrated in
100 increments from -500 to +500 and from.
1-10
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TA-12
SECTION I
INSTRUMENT PANEL (Forward Cockpit)
8 9 10 II 12 13 14 15 16
82
80
79
78
77
76
75
74
73
72
71
70
69
68
67 65 63
66 64
61 59 57
62 60 58 56 54
52
1 INS DESTINATION AND SELECT PANEL
2 AFT COCKPIT TEMPERATURE SWITCH
3 COCKPIT TEMPERATURE RHEOSTATS
4 STANDBY ATTITUDE GYRO
5 AIR SPEED - MACH METER
6 DISTANCE TO GO - GROUND SPEED
IND ICATOR
7 ATTITUDE INDICATOR
8 AIR REFUEL READY -DISCONNECT LIGHT
AND SWITCH
9 DRAG CHUTE HANDLE
10 RAIN REMOVAL SPRAY BUTTON
11 MASTER CAUTION LIGHT
12 HORIZONTAL SITUATION INDICATOR
13 PERISCOPE VIEWING SCREEN
14 RATE OF CLIMB INDICATOR
15 TRIPLE DISPLAY INDICATOR
16 FIRE WARNING LIGHTS
17 COMPRESSOR INLET TEMPERATURE GAGE
18 COMPRESSOR INLET PRESSURE GAGE
19 TACHOMETERS
20 EXHAUST NOZZLE POSITION INDICATOR
21 EXHAUST GAS TEMPERATURE GAGES
22 AIR REFUEL READY SWITCH
73 FUEL DUMP SWITCH
24 FUEL FLOW METERS
25 FUEL TRANSFER SWITCH
26 EMERGENCY FUEL SHUTOFF SWITCHES
27 LIQUID NITROGEN QUANTITY GAGE
28 PUMP RELEASE SWITCH
29 FUEL QUANTITY SELECTOR SWITCH
30 BAILOUT SWITCH
51
50
49
6 6 6
48 47 46
41 40 39 38 37 36
42
43
44
45
31 AFT SEAT EJECTED LIGHT
32 BAILOUT LIGHT
33 BATTERY SWITCH
34 INVERTER SWITCHES
35 GENERATOR SWITCHES
36 FUEL BOOST PUMP SWITCHES
37 FUEL TANK PRESSURE GAGE
38 HYDRAULIC QUANTITY GAGE
39 ENGINE OIL PRESSURE GAGES
40 HYDRAULIC PRESSURE GAGES
41 FUEL QUANTITY INDICATOR
42 BACKUP PITCH-DAMPER SWITCH
43 FLIGHT INSTRUMENT CONTROL PANEL
44 PITCH LOGIC O'RIDE SWITCH
45 YAW LOGIC O'RIDE SWITCH
46 LANDING GEAR RELEASE HANDLE
47 ANNUNCIATOR PANEL
48 SURFACE LIMITER HANDLE
49, HYDRAULIC RESERVE OIL SELECTOR
SWITCH
50 TRIM POWER SWITCH
51 PITOT HEAT SWITCH
52 FORWARD BYPASS POSITION INDICATOR
53 FORWARD BYPASS CONTROL SWITCHES
17
18
19
20
21
22
23
24
25
26
27
28
29
30
35 33 31
34 32
54 SPIKE POSITION INDICATOR
55 YAW TRIM INDICATOR
56 ROLL TRIM INDICATOR
57 SPIKE CONTROL SWITCHES
58 ELAPSED TIME CLOCKS
59 PITCH TRIM INDICATOR
60 ENGINE START SWITCHES
61 BRAKE SELECTOR SWITCH
62 LANDING GEAR DOWN LIGHTS
63 COCKPIT PRESSURE DUMP SWITCH
64 LANDING GEAR SELECTOR HANDLE
65 COCKPIT PRESSURE ALTITUDE GAGE
66 OXYGEN CYLINDER PRESSURE GAGE
67 CABIN ALTIMETER SELECTOR SWITCH
68 AFTERBURNER SWITCHES
69 INDICATOR AND LIGHTS TEST SWITCH
70 LANDING GEAR WARNING CUTOUT BUTTON
71 LANDING AND TAXI LIGHT SWITCH
72 PERISCOPE MIRROR SELECT HANDLE
73 ALTIMETER
74 COCKPIT TEMPERATURE INDICATOR
75 PERISCOPE PROJECTOR FILM LIGHT
76 FORWARD COCKPIT AIR SYSTEM SWITCH
77 PERISCOPE PROJECTOR LIGHT RHEOSTAT
78 COCKPIT TEMPERATURE INDICATOR
SWITCH
79 SUN COMPASS SWITCH
80 FWD COCKPIT TEMPERATURE SWITCH
81 AFT COCKPIT AIR SYSTEM SWITCH
82 PERISCOPE MAGNIFICATION CONTROL
HANDLE
F201-15(d)
Figure 1-4
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1-B
SECTION I
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TA-12
INSTRUMENT PANEL (Aft Cockpit)
68
67
65
64
63
62
61
60
59
58
57
5
4
3
2
55 54 53 52 51 50
48 I 46 44
49 47 45
43
42 �
41
40
1 STANDBY ALTITUDE GYRO
2 AIR SPEED-MACH METER
3 DISTANCE TO GO-GROUND SPEED INDICATOR
4 ALTITUDE INDICATOR
5 DRAG CHUTE HANDLE
6 AIR REFUEL READY-DISCONNECT LIGHT AND S
7 MASTER CAUTION LIGHT
8 HORIZONTAL SITUATION INDICATOR
9 STANDBY COMPASS 27
10 RATE OF CLIMB INDICATOR 28
11 TRIPLE DISPLAY INDICATOR 29
12 FIRE WARNING LIGHTS 30
13 COMPRESSOR INLET TEMPERATURE GAGE 31
14 TACHOMETERS 32
15 EXHAUST GAS TEMPERATURE GAGES 33
16 AIR REFUEL READY SWITCH 34
17 FUEL DUMP SWITCH 35
18 FUEL FLOW METERS 36
19 FUEL TRANSFER SWITCH 37
20 EMERGENCY FUEL SHUTOFF SWITCHES 38
21 PUMP RELEASE SWITCH 39
22 BAILOUT SWITCH ao
23 FUEL QUANTITY SELECTOR SWITCH 41
24 BAILOUT LIGHT 42
3 BATTERY SWITCH 43
26 INVERTER SWITCHES 44
7 8.9 10 11 12
SELIP BLM110 ALIAP
rokttCN
0
6,..p�-; I))
39 38 37 36
WITCH
32 31 30 29 28
GENERATOR SWITCHES
FUEL BOOST PUMP SWITCHES
FUEL TANK PRESSURE GAGE
ENGINE OIL PRESSURE GAGES
HYDRAULIC PRESSURE GAGE
FUEL QUANTITY INDICATOR
BACKUP PITCH DAMPER SWITCH
FLIGHT INSTRUMENT CONTROL PANEL
PITCH LOGIC O'RIDE SWITCH
YAW LOGIC O'RIDE SWITCH
LANDING GEAR RELEASE HANDLE
ANNUNCIATOR PANEL
SURFACE LIMITER HANDLE
HYDRAULIC RESERVE OIL SEL SWITCH
TRIM POWER SWITCH
PITOT HEAT SWITCH
FORWARD BYPASS POS INDICATOR
YAW TRIM INDICATOR
33
34
35
13
14
15
16
17
18
19
20
21
22
27 26 25 24
23
45 FORWARD BYPASS CONTROL SWITCHES
46 ROLL TRIM INDICATOR
47 ELAPSED TIME CLOCKS
48 PITCH TRIM INDICATOR
49 ENGINE START SWITCHES
50 LANDING GEAR BYPASS SWITCH
51 COCKPIT PRESSURE DUMP SWITCH
52 INDICATOR AND LIGHT TEST SWITCH
53 COCKPIT PRESSURE ALTITUDE GAGE
54 CABIN ALTIMETER SELECTOR SWITCH
55 OXYGEN CYLINDER PRESSURE GAGE
56 BRAKE SELECTOR SWITCH
57 LANDING GEAR LOCK WARNING LIGHT
58 LANDING GEAR DOWN LIGHTS
59 LANDING AND TAXI LIGHT SWITCH
60 AFTERBURNER SWITCHES
61 LANDING GEAR WARNING CUTOUT BUTTON
62 ALTIMETER
63 COCKPIT TEMPERATURE INDICATOR
64 FORWARD COCKPIT AIR SYSTEM SWITCH
65 COCKPIT TEMPERATURE INDICATOR SWITCH
66 AFT COCKPIT AIR SYSTEM SWITCH �
67 FWD AND AFT COCKPIT TEMPERATURE
SWITCHES
68 COCKPIT TEMPERATURE RHEOSTATS F201-14(d)
1-12
Figure 1-5
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Approved for Release: 2017/07/25 C06230172
TA-12
SECTION I
A CONT B L SPIKE R
ANNUNCIATOR PANEL (Typical)
FORWARD COCKPIT
SURF LIMIT RELEASE
PUL
P I TOT
HEAT ON
OFF
NO. IWO( LOW
NO. 2 OXA LOW
Q-BAY HEAT HIGH
FUEL OTT LOW
N WA LOW
TANK PRESS LOW
ANTI-SKID OUT
SURFACE LIMITER
SAS CHANNEL OUT
A HAD LOW
8 HAD LOW
STALL WARNING
L OIL TEMP
L RAP DR NOT OPEN
L RAP DR MAO OPEN
RAP DR MAN CLOSED
LEND BLEED OPEN
L FUEL PRESS LOW
I. COD OIL PRESS LOW
L GENERATOR OUT
L XFMR-RECT OUT
NO.-1 INVERTER OUT
NO. 2 INVERTER OUT
NO. 3 INVERTER OUT
PIVOT HEAT
ROIL TEMP
R RAP DR NOT OPEN
R RAP DR MAN OPEN
R BAP DR MAN CLOSED
R ENG BLEED OPEN
R FUEL PRESS LOW
R CUD OIL PRESS LOW
R GENERATOR OUT
R XFMR-RECT OUT
ENTER BAT ON �
R HOD LOW
L HOD LOW
Q-BAY EQUIP OUT
AFT COCKPIT
RELEASE
TRIM IND NAV r- OIL PRESS
0PITCH 0ROLL 0YAW ND 0L 0R 0ADF
HYD PRESS FUEL TANK r--- IND
PRESS BYPASS ATI
0 0 0
OFF
YAW
F201-10(c)
Figure 1-6
Approved for Release: 2017/07/25 C06230172
1-13
SECTION I
Approved for Release: 2017/07/25 C06230172
TA-12
FORWARD COCKPIT (Left Side)
STE All
FAq`ERICT
�
BEACON
0-BAND
ON
IFR PHONE
ON
OH RUDDER
SYNCHpoNIZER
ROLL -
TRIM
CO1
oN
OX`i �s
SYS I
PSI
(
. Cit) Cro
MANUAL
r- COOT TRANS
TA.,b11 NSTR qi) UHF
COOT TRANS ---i
FUEL CONTROL
L EMER R
(5)
L NORM R
FUEL COOT FUEL OTY
0 CO
�
TEA iiR C.
OFF ATE
AIR
COND AFT
CONT
TRANS
FIND CT/PT
NORA1
Figure 1-7
FWD COCKPIT
1 ROLL TRIM SWITCH
2 THROTTLE QUADRANT
3 UHF CONTROL PANEL
4 CONTROL TRANSFER PANEL
5 EMERGENCY FUEL CONTROL SWITCHES
6 IFF CONTROL PANEL
7 IFR INTERPHONE SWITCH
8 RADAR BEACON SWITCH
9 OXYGEN CONTROL PANEL
10 STANDBY ATTITUDE FAST ERECT SWITCH
11 RIGHT HAND. RUDDER SYNCHRONIZER SWITCH
F201-12 (d)
1-14
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Approved for Release: 2017/07/25 C06230172
TA-12
SECTION I
FORWARD COCKPIT (Right Side)
FWD COCKPIT
1 SAS CONTROL PANEL
2 AUTOPILOT CONTROL PANEL
3 INERTIAL NAVIGATION CONTROL PANEL
4 TACAN CONTROL PANEL
5 FRS CONTROL PANEL
6 DEFOG AND FACE HEAT CONTROL PANEL
7 LIGHTING CONTROL PANEL
8 TRIM POWER CIRCUIT BREAKER PANEL
Figure 1-8
Approved for Release: 2017/07/25 C06230172
1-15
SECTION I
Approved for Release: 2017/07/25 C06230172
TA-12
AFT COCKPIT Left Side)
5
VIEW A
6
7
Of.
ANY
SYS I
on
COOT TRANS
TACANONSTR ART UHF
COOT TRANS
FUEL COST FUEL oil
ROLL
TRIM
AIR
GOOD
COST
TRANS
FUEL CONTROL
L
CO- CB)
L ITOR"I R
5
9
AFT COCKPIT
1 FAST ERECT BUTTON FOR
STANDBY ATTITUDE � INDICATOR
2 RUDDER SYNCHRONIZATION
SW ITCH
3 ROLL TRIM SW ITCH
4 THROTTLE FRICTION LOCK
5 THROTTLES
6 OXYGEN CONTROL PANEL
7 UHF CONTROL PANEL
8 CONTROL TRANSFER PANEL
9 EMERGENCY FUEL CONTROL
SW ITCHES
F201-7 (b)
1-16
Figure 1-9
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
TA-12
SECTION I
AFT COCKPIT (Right Side)
AFT COCKPIT
1 SAS CONTROL PANEL
2 LOGIC OVERRIDE PANEL
3 TACAN CONTROL PANEL
4 ADF CONTROL PANEL
5 FRS CONTROL PANEL
6 LIGHTING CONTROL PANEL
a�
PITCH O
A,
NORM (13)
OFF : �
-----}?-\
B ON
\
0 NORM 0
OFF
MON
STAB AUG
ON
0
ROLL
0 '
oN. , '
(e) ..
LITE TEST
YAW �I
ON ;A/
id), NORM
OFF. 0
/
(TN E'
i
0 NORM
OFF
MON
0 A A OVERRIDE � Nij.
CONTROL
FWD
PITCH YAW im .
0 B B 0 keci DAMPER
LOGIC OVERRIDE
T
A
� C
A
CHAR AA
III REC --
/
OFF
VOL
A
D OFF
r
. ,00n Bic)
0114).
FREQUENCY
OFF
ADF ANT.
\ ' 1 LOOP
'
�.�� � ��.
:4t) .4
!9 90 09
s
TORT
o NJ
e ..�.,
_ .0_,�
,E�
,
0 1 .,, ,
. �
2
�
t
-
7"r,-
- cr.
E - -
0
e = v:�
5
6
F201-6 (d)
Figure 1-10
Approved for Release: 2017/07/25 C06230172
1-17
SECTION I
Approved for Release: 2017/07/25 C06230172
TA-12
LEFT AND RIGHT FORWARD PANELS
2 3 4
11
10
FORWARD COCKPIT
AFT COCKPIT
7
1
CABIN ALTIMETER SELECTOR SWITCH
8
BAILOUT LIGHT
2
INDICATOR AND WARNING LIGHTS TEST BUTTON
9
BATTERY-EXTERNAL POWER SWITCH
3
OXYGEN CYLINDER PRESSURE GAUGE
10
GENERATOR SWITCHES
4
CABIN ALTIMETER
11
INVERTER SWITCHES
5
LANDING GEAR LEVER
12
GEAR NOT LOCKED LIGHT
6
FUEL QUANTITY INDICATOR SELECTOR SWITCH
13
LANDING GEAR SWITCH
7
BAILOUT SWITCH
F201-5(c)
1-18
Figure 1-11
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
TA-12 SECTION I
400� to 460�C. Each gage also has two di-
gital readout windows for the compressor
inlet temperature. Power is furnished by
the No. 1 inverter.
ENGINE AIR INLET SYSTEM
The engine air inlets are canted inward and
downward to align with the local airflow
pattern. Air is bled from the spike and
cowl to prevent boundary layer separation.
The porous bleed on the spike centerbody
exhausts overboard through the supporting
struts and louvres. The cowl bleed supplies
ejector secondary air for cooling the engine
and ejector. Ground cooling suck-in doors
are also provided in the aft nacelle area.
Inlet airflow is controlled by the inlet by-
pass, a rotating basket which opens ports in
the duct a short distance downstream of the
inlet throat. On the ground, the bypass is
open and the spike is full forward.
Note
The spikes are locked forward in
this aircraft for all operations.
BYPASS CONTROLS AND INDICATORS
Inlet Bypass Switches
Two three-position toggle switches (aft
cockpit) and two rotary switches (forward
cockpit) are located on the lower left side
of the instrument panel. The switches pro-
vide manual control of the inlet air bypasses.
The aft cockpit switches are labeled OPEN
(up), FWD CKPT (center) and CLOSED
(down). The FWD CKPT position allows the
forward cockpit control of the inlet air by-
passes. When the switches are in the OPEN
or CLOSED position the forward cockpit by-
pass controls are overridden. The forward
cockpit rotary switches are labeled OPN,
HOLD, and CL. The rotary switches are
turned either clockwise or counterclockwise
for the desired positioning of the bypasses.
Inlet Bypass Position Indicators
A dual inlet bypass position indicator is lo-
cated on the lower left side of each instru-
ment panel. The pointers indicate the
amount of inlet bypass opening that has been
selected with the inlet bypass switches, and
do not indicate actual door position. The
indication is in 10 percent increments and
the labeled positions are 20, 40, 60, 80,
and 100 percent.
Inlet Bypass Not Open Indicator Lights
Two indicator lights, one labeled L BYP
DR NOT OPEN and the other R BYP DR
NOT OPEN, are located on each annun-
ciator panel. The light, when illuminated,
indicates that the bypass is not open when
the landing gear is down. Power for the
lights is furnished by the essential dc bus.
Inlet Bypass Manually Open Indicator Lights
Two indicator lights, one labeled L BYP DR
MAN OPEN and the other R BYP DR MAN
OPEN, are located on each annunciator
1-19
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SECTION I
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IM-IL
AIR FLOW PATTERNS
� MACH NO. � 0.0
CENTERBODY BLEED SUCK-IN DOORS OPEN
if I
ra 'III! ttrfr,rir 4j_t_71
4 re
;(k "
SPIKE FORWARD .�../4.
COWL BLEED SUPPLIES
ENGINE COOLING AIR
BYPASS DOORS
OPEN
_.,svar/rft"
M MO
SPIKE FORWARD
COWL BLEED SUPPLIES,/
ENGINE COOLING AIR
SPIKE FORWARD
COWL BLEED SUPPLIES
ENGINE COOLING AIR
TERTIARY DOORS OPEN
� MACH NO. 0.9 �
CENTERBODY BLEED
OVERBOARD
ler
BYPASS DOORS
CLOSED
CENTERBODY BLEED
OVERBOARD
/./A//
BYPASS DOORS
MANUALLY OPEN
SUCK-IN DOORS CLOSED
//-7_//777./
-7-7a
MACH NO. 1.35 �
� MACH NO. 1.7
CENTERBODY BLEED
OVERBOARD
/w A A
\\
SPIKE RETRACTING
z.2>.
\\k
BY PASS DOORS
MANUALLY OPEN
TERTIARY DOOR OPEN
SUCK-IN DOORS CLOSED
TERTIARY DOORS CLOSED
SUCK-IN DOORS CLOSED
r'L-4-1 � `4-1-X
174.=,, 7 / 11,1,1.41
/ /
, _?L. , � 4
ft,/,
I
I
EJECTOR
FLAPS
CLOSED
EJECTOR
FLAPS
CLOSED
EJECTOR
FLAPS
CLOSED
TERTIARY DOORS CLOSED
EJECTOR FLAPS CLOSED
FZ01-17
1-20
Figure 1-12
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
TA-12 SECTION I
panel. The lights, when illuminated, indi-
cate that the bypass has reached the full-
open position. Power for the lights is fur-
nished by the essential dc bus.
Inlet Bypass Manually Closed Indicator Lights
Two indicator lights, one labeled L BYP DR
MAN CLOSED and the other R BYP DR MAN
CLOSED, are located on each annunciator
panel. The lights, when illuminated, indicate
that the bypass has reached the fully closed
position. Power for the lights is furnished
by the essential dc bus.
FUEL SUPPLY SYSTEM
The aircraft fuel supply system consists of
six integral fuel tanks with interconnecting
plumbing and electrically actuated boost
pumps for fuel feed, transfer, and dumping.
Other components of the system include
nitrogen inerting, pressurization and vent-
ing, single-point refueling, and fuel quantity
indication. In addition to furnishing fuel for
the engine, automatic fuel management pro-
vides center-of-gravity and trim drag con-
trol at cruise speed. The fuel is also used
as a heat sink to cool cockpit air, engine
oil, CSD oil, and hydraulic fluid.
FUEL TANKS
The six integral, internally sealed fuel tanks
are contained in the fuselage and wing stub
extensions. The tanks are numbered 1
through 6, fore to aft, and are interconnected
by right and left fuel manifolds and a single
vent line. Electrically actuated, submerged
boost pumps are contained in all tanks, two
each in tanks 2, 4, 5, and 6 and four each in
tanks 1 and 3. Fuel manifolds, fed by the
fuel boost pumps, route fuel to the engines,
transfer fuel to tank 1 for cg control, or to
the fuel dump valves where it can be dumped
overboard in an emergency. Normal se-
quence of tank usage is controlled by a
float switch for each pump to automatically
maintain an optimum center of gravity for
cruise. The left engine normally uses fuel
in a sequence of tanks 1, 2, 4, and 3; the
right engine uses fuel in a sequence of
tanks 1, 6, 5, and 3. Normal automatic
tank sequencing Is as follows:
L ENGINE
R ENGINE
Tanks 1 & 2
Tank 2
Tank 4
Tank 4
Tank 3
Tanks 1 & 6
Tank 6
Tank 6
Tank 5
Tank 5
Tank 3
Use of an electrically operated crossfeed
valve and the boost pump switches makes
it possible for any tank to feed any engine.
REFUELING AND DEFUELING
A single-point refueling receptacle, installed
on top of the fuselage just aft of the rear
cockpit, is used for both ground and air re-
fueling. Ground refueling is accomplished
by use of a probe especially modified to
utilize a hand-operated locking device so
that refueling may be done without hydraulic
power. Fuel from the receptacle flows
through the fueling manifold to each tank.
The use of a different size orifice for each
tank allows all tanks to be filled simulta-
neously in approximately 12 minutes, with
a refueling pressure of 50 psi. Dual shutoff
valves in each tank shut off fuel flow when
the tank is full. A defueling fitting is in-
stalled on the right fuel-feed manifold in the
lower right side of tank 4. Tanks 2 and 4,
�which feed the left manifold, are defueled
by opening the crossfeed valve.
CAUTION
Any fuel in tanks 5 and 6 must be
balanced with a like amount of
fuel in the other tanks when fueling
or defueling to prevent the aircraft
from rocking down on the tail.
1-21
Approved for Release: 2017/07/25 C06230172
ZLI.O�2900 SILO/L10Z :aseaia JOI penaidd\of
FUEL QUANTITY
FUEL QTY
IND SELECTOR
�
FUEL
SHUTOFF
�
CROSS- H
FEED
I #1 TANK I 4
EMPTY
FWD FUEL
I#2 TANK TRANS TRANS
EMPTY
I #3 TANK
I EMPTY
EMPTY
I #4 TANK _I
OFF
I_ 5 TANK
EMPTY
I_ #6 TANK I
EMPTY
FORWARD
� im � 10311N � sa(DLI��. 1m
6-1 5-1
6-22
L
EMER
NORM
OFF
FUEL DUMP
4-1
� W � II
4-2 �
3-3
I..
3-4111
IIpIEI
LUIUUIIIIU
FUEL
FLOW
RIGHT
ENGINE
I-N2
������� VENT
FUEL
IN mil REFUELING
0 FLOAT OPERATED SHUTOFF VALVE
ED FLOAT OPERATED SHUTOFF VALVE WITH PRESSURE RELIEF
FUEL TANK
PRESS
PRESSURE RELIEF VALVE E5 FUEL TANK
PRESSURE SENSOR
FUEL QUANTITY SENSOR
LN2
al
WESAS Aldan 13nA
Approved for Release: 2017/07/25 C06230172
TA-12
SECTION I
FUEL CONTROL PANEL (Both Cockpits)
1 TANK
EMPTY
2 TANK
EMPTY..
3 TANK
EMPTY'
4 TANK-
EMPTV1-;
5j A'�
-� EMPTY
&TANK:
EMPTY-.
EMER
FUEL
SqUTOFE
.6-;.; TOTAL '
AFT COCKPIT
FWD COCKPIT
,
AIR
REFUEL
READY.
FUEL:,
D.41?
OFF EM,E1R
NORM
.1 2 TANK-
EMPTY;:'.:
-3.TANK
Er.,,IPTY
'EATIPATi
6 TANK
EMPTY
FUEL QTY SELECT
5 6 TOTAL
F201-5(a)
Figure 1-14
Approved for Release: 2017/07/25 C06230172
1-23
SECTION I
Approved for Release: 2017/07/25 C06230172
TA-12
Fuel Tank Capacities
Tank
Capacity
Fuel Loading Limit*
1
1,125 gal.
3,000 lbs
1,580
6,000
3
1,571
10,200 (full)
4
2,125
3,300
5
2,150
3,700
6
1,945
8,600
Total
10,496 gal
34,800 lbs*.4,
* Automatic shutoff float switches set to
restrict maximum weight.
** At 6.45 lb/gal fuel density.
FUEL BOOST PUMPS
Sixteen single-stage, centrifugal ac-powered
fuel boost pumps are used to feed the fuel
manifolds. Tanks 1 and 3, which normally
feed both right and left engines, are equipped
with four boost pumps, and tanks 2, 4, 5,
and 6 have two pumps each. A single pump
in each tank is capable of supplying fuel to
the engine in the event of failure of the other
pump. The pumps in each tank may be op-
erated out of the normal sequence by ac-
tuating the individual tank boost pumps
switches, located on the right side of each
instrument panel. These switches supple-
ment automatic tank sequencing if a tank
fails to feed in the proper sequence. It is
necessary to actuate the pump release
switch to terminate any manually actuated
pump when the tank is empty. Normally,
each pump (except 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
emptied in sequence. One of the float
switches in each tank illuminates the yellow
tank-empty light contained in the respective
boost pump tank switch. The boost pumps
that feed the left-hand manifold are nor-
mally powered froin'ifie left generator bus
and the pumps that feed the right-hand mani-
fold are normally powered from the right
generator bus. Individual circuit breakers
for each pump are located in a compartment
behind the aft cockpit and are not accessible
in flight.
Emergency Fuel Shutoff Switches
A guarded fuel shutoff switch for each en-
gine is installed on the right side of each
instrument panel. The switches are
guarded in the open, or ON, position. When
the switches (forward or aft cockpit) are
moved to the EMER (up) position, power
from the ac generator bus closes motor-
driven valves in the engine feed lines. Each
switch in the forward cockpit is safety-wired
to the ON position for dual flight.
Fuel Boost Pump Switches and Indicator Lights
Six fuel boost pump switches are installed
in a vertical line on the right side of each
instrument panel. These switches are
plastic, self-illuminated pushbutton-type,
and control manual operation of the fuel
boost pumps in each tank. The switches
read out 1 TANK through 6 TANK when the
respective tank boost pumps are operating.
1-24
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Approved for Release: 2017/07/25 C06230172
TA-12 SECTION I
. 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 selec-
tion of only one tank of tank group 1, 2, and
4 and one tank of tank group 3, 5, and 6 at
the same time. This feature is intended to
prevent more than eight boost pumps from
operating simultaneously if one engine gen-
erator is inoperative.
Note
It is possible to operate more than
eight boost pumps at once by a
combination of automatic sequencing
and manual actuation; 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 the pushbutton and the
number of the tank involved. When a tank
is empty, a yellow light in the pushbutton
illuminates EMPTY. When depressed, a
boost pump switch will hold down electri-
cally until released by the pump release
switch. Power for the boost pump switch
circuits and lights is furnished by the es-
sential dc bus. (Refer to description of
forward and aft cockpit control transfer
panels in Section IV for further information.)
Pump Release Switches
A momentary pump release toggle switch is
installed on each instrument panel below
the fuel boost pump switches. The switch
has two positions, PUMP REL (up) and
NORM (down). When placed in the PUMP
REL position, any boost pump switch that
has been depressed manually will be re-
leased and automatic tank sequencing will
resume. Power for the circuit is furnished
by the essential dc bus;
CAUTION
A manually selected boost pump
should be released when a tank
indicates empty so that the pumps
in that tank will be shutoff.
Crossfeed Switches
A pushbutton-type crossfeed switch is in-
stalled at the top of the column of boost
pump switches on each instrument panel.
When depressed, it illuminates a green
light in the switch, opens a motor-operated
valve between the left and right fuel mani-
folds, allowing the right manifold to feed
the left engine and the left manifold to feed
the right engine. The switch must be de-
pressed a second time to stop crossfeeding.
Power for the circuit is furnished by the
essential dc bus.
Fuel Transfer Switches
A guarded fuel transfer switch is installed
to the right of the boost pump switches on
each instrument panel. When the switch is
in the FWD TRANS (up) position, a valve
is opened in the forward end of the right-
hand fuel supply manifold, the boost pumps
in tank 1 are inactivated, and fuel will
transfer forward from tank 3, 5, or 6.
Transfer is automatically terminated by a
float switch when the quantity in tank 1
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SECTION I
TA-12
reaches approximately 3000 pounds. This
setting precludes the possibility of encoun-
tering gusts with more than the limit fuel
quantity in tank 1. Tank 1 boost pumps will
remain inactivated until 800 pounds of fuel
remaining in tank 3, or the transfer switch
is moved to the OFF (down) position. Power
for the circuit is furnished by the essential
dc bus.
Fuel Dump Switches
A guarded three-position lift-lock fuel
dump switch is installed on the right side of
each instrument panel. The three positions
are EMER (up), NORM (center) and OFF
(down). When the switch is moved to the
NORM position, the pumps in tank 1 are in-
activated to maintain a forward cg and all
other tanks will dump in normal usage se-
quence. Fuel dumping will stop when the
fuel level in tank 3 reaches 5000 lbs re-
maining. If there is any fuel remaining in
tank 1, the boost pumps in tank 1 will start
when tank 3 is down to 5000 lbs or dumping
is terminated. When the switch is moved
to the EMER position, the stop dump switch
in tank 3 is bypassed and fuel dumping con-
tinues until all tanks, except tank 1, are
empty. Power for the circuit is furnished
by the essential dc bus.
WARNING
Emergency fuel dumping must be
terminated by moving the fuel dump
switch to either the NORM or OFF
position; otherwise, all fuel, except
fuel in tank 1, will be dumped.
Fuel Quantity Selector Switch and Quantity
Indicator
A quantity indicator and a rotary fuel quan-
tity selector switch is located on the right
side of the instrument panel in each cockpit.
Positions on the selector switch are marked
for each of the six tanks, and TOTAL posi-
tion. The switch is rotated to the individual
tank or TOTAL position for the desired
reading on the fuel quantity indicator. The
indicator is calibrated in 1000 pound in-
crements from zero to 75,000 pounds. It
also has a digital readout window indicating
to the nearest 100 pounds the amount of fuel
remaining. Power for the circuit is fur-
nished by the No. 1 inverter.
Fuel-Quantity-Low Lights
Fuel-quantity-low lights, labeled FUEL
QTY LOW, are located on each annunciator
panel. The lights are illuminated by the
closing of a low level (5000 pound) float
switch in tank 3. Power for the lights is
furnished by the essential dc bus.
Fuel Pressure Low Warning Lights
Fuel pressure low warning lights labeled L
FUEL PRESS LOW and R FUEL PRESS
LOW, are located on the annunciator panel
in each cockpit. Illumination of a light in-
dicates 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.
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Approved for Release: 2017/07/25 C06230172
TA-12 SECTION I
FUEL PRESSURIZATION AND VENT SYSTEM
The fuel pressurization system consists of
two liquid-nitrogen-filled Dewar flasks, lo-
cated in the nosewheel well, and associated
valves and plumbing to the fuel tanks and
indicators. The Dewar flasks supply nitro-
gen gas to the fuel tanks at 1.5 (+ 0.25) psi
above ambient pressure, which inerts the
ullage space above the fuel and provides
pressure to ensure fuel flow to the engine-
driven pump in case of boost pump failure.
When Dewar flasks are full, the nitrogen
supply is sufficient for approximately 9
hours of flight, including two refueling op-
erations. The liquid nitrogen from the
bottom of the flasks is routed through sub-
merged heat exchangers in tanks 1 and 3 to
ensure that the nitrogen has become gaseous.
The nitrogen gas is then ported to the com-
mon 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 tank 1. Tank 1
has only one vent valve and the open for-
ward end of the vent line. The forward vent
valves in tanks 2, 3, 4, 5 and 6 are 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 that it has no
relief valve. The common vent line tees
into two lines in tank 6 and both go through
the rear bulkhead. In the tail-cone area is
a relief valve in each vent line with the left
valve set to relieve pressure at 3 (+ 0.25)
psi above ambient pressure. ,In the event
of failure of this valve, the right valve will
relieve pressure at 3.5 (+ 0.25) 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 formth.rd in the vent line when the
aircraft is decelerated. A check valve
prevents fuel, that is coming forward from
tank 6, from going beyond tank 5. A valve,
located in tank 3, prevents fuel coming
from tank 4 from going beyond tank 3. This
float-actuated valve closes the vent when
fuel is moving forward in the vent line and
diverts it into tank 3. Tank 2 fuel can go
forward into tank 1. Acceleration presents
no problem of fuel shift between tanks.
Liquid Nitrogen Quantity Indicators
A dual liquid nitrogen quantity indicator is
installed on the right side of the forward
cockpit instrument panel. The indicator
displays the quantity of liquid nitrogen re-
maining in each of the two Dewar flasks.
The indicator is marked in 5-liter incre-
ments from 0 to 110 liters. Power for the
indicator is furnished by the essential dc
bus and the No. 1 inverter bus.
N2 Quantity Low Indicating Light
An indicator light labeled N QTY LOW is
located on the annunciator -panel in each
cockpit. The light will illuminate when either
liquid nitrogen quantity gage reaches 1 liter
remaining. Power for the light is furnished
by the essential dc bus.
Fuel Tank Pressure Indicators
A fuel tank pressure indicator is installed
on the right side of each instrument panel.
The indicators read the amount of gaseous
nitrogen pressure existing in fuel tank 1,
and are marked from -2 to +8 in increments
of 1/2 psi. Power for the indicators is
normally furnished by the No. 2 26-volt
instrument transformer.
Tank Pressure Low Indicating Light
This light labeled TANK PRESSURE LOW
is located on the annunciator panel in each
cockpit and will illuminate when the tank
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SECTION I
Approved for Release: 2017/07/25 C06230172
TA-12
pressure reduces to +. 25 to +.10 psi.
Power for the light is furnished by the es-
sential dc bus.
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, receptacle doors, hydraulic
valves, hydraulic actuators, a signal am-
plifier, control switches and indicator lights.
Hydraulic power for the system is normally
supplied from the L hydraulic system. If
the L system is inoperative the refueling
system can be operated by R hydraulic pres-
sure by selecting ALT STEER & BRAKE.
Electrical power is supplied by the essential
dc bus.
Air Refuel Switches
An AIR REFUEL switch is located at the top
of the right instrument panel in each cock-
pit. The switch in the aft cockpit has three
positions labeled READY, FWD, and OFF.
When the switch is in the READY position
the refueling doors are hydraulically actu-
ated open, the boom latches are armed, the
fueling receptacle lights are illuminated,
the green READY portion of the air refuel
reset light in each cockpit is illuminated,
and the forward cockpit AIR REFUEL switch
is made inoperative. When the AIR REFUEL
switch in the aft cockpit is placed in the
FWD position, the forward cockpit AIR RE-
FUEL switch is operative, and when placed
in the OFF position the forward cockpit
switch is inoperative.
The AIR REFUEL switch in the forward
cockpit has three positions labeled READY,
OFF, and MANUAL. In the READY position
the system is readied for automatic latching.
In the OFF position the doors are closed
and electrical power is removed from the
system. In the MANUAL position the doors
are open, the green READY portion of the
reset switch is illuminated, and the fueling
receptacle latches are closed. The latches
may be opened to accept the probe by hold-
ing the A/R DISC trigger switch on the con-
trol stick grip. When the A/R DISC dis-
connect trigger is released the latches will
close and hold the boom. The latches will
open to release the boom when the A/R
DISC trigger is depressed. MANUAL posi-
tion is used in the event of a malfunctioning
amplifier.
Air Refuel Reset Switches and Indicator Lights
A square dual indicator light and reset
button is located on the top of each instru-
ment panel on the left side. The top half
is labeled READY and will illuminate green
when an air refuel switch is in the READY
or MANUAL position, and the refueling re-
ceptacle is open and ready to accept the re-
fueling boom. The lights will extinguish
after the boom is engaged. If the boom dis-
connects from the fueling receptacle for any
reason when automatically latched, the
lower half of the switches will illuminate
amber and show DISC. The light may then
be pressed to reset the system amplifier for
another engagement. The DISC lights do not
illuminate if a disconnect occurs while man-
ually latched. Power for the system is sup-
plied by the dc essential bus.
Disconnect Trigger Switches
A momentary contact, trigger-type switch,
marked A/R DISC, is installed on the for-
ward side of each control stick. Depress-
ing either trigger switch will initiate a boom
disconnect. The trigger is also depressed
to open the receptacle latches when the air
refuel switch is in the MANUAL position;
releasing the trigger will close the latches.
1-28
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Approved for Release: 2017/07/25 C06230172
TA-12 SECTION I
Disconnect
A refueling disconnect may be accomplished
in one of the following ways:
1. Automatically.
a. If boom envelope limits are ex-
ceeded.
b. When manifold pressure reaches
85-90 psi.
2. Manually.
a. By the boom operator.
b. By depressing the A/R DISC trig-
ger on the control stick grip.
Pilot Director Lights (On 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 ele-
vation, and the right row for boom tele-
scoping. The elevation lights consist of
five colored panels with green strips, green
triangles, and red triangles to indicate re-
lative position. Two illuminated letters, D
and U for down and up, respectively, indi-
cate elevation correction. Background lights
are located behind the panels. The colored
panels are illuminated by lights controlled
by boom elevation during contact. The
colored panels which 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. Figure 4-16 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 indicate the
centerline. When the contact is made, the
panels automatically reflect the correction
required by the pilot to maintain position.
ELECTRICAL SUPPLY SYSTEM
The basic ac electrical system consists of
a 30-KVA, constant-speed ac generator on
each engine, furnishing 3-phase power to
two ac buses through an automatic bus
transfer and protection system. DC power
is obtained by two 200-amp transformer-
rectifiers, one from each ac bus. The
parallel output from these transformer-
rectifiers supplies the essential dc bus and
a monitored dc bus. Three 600-VA inver-
ters, powered by the essential dc bus, fur-
nish fixed-frequency ac power to three se-
parate ac buses. Three instrument trans-
formers furnish 26-volt ac power; one is
powered from inverter 1, the other two
from inverter 3. A battery bus is furnished
to provide power for air starting.
AC GENERATOR POWER SUPPLY
Each engine drives a 30-KVA generator
through a constant-speed drive. This is
the primary source of ac electrical power
for the aircraft. The generators supply
115/200 V, 3-phase, constant-frequency
power to the aircraft electrical system.
Either generator will provide through the
automatic bus transfer system in the event
one generator fails. Conventional switches
are provided for manual control of the
generators.
INVERTER POWER SUPPLY
Three 600-VA, solid-state inverters fur-
nish constant-frequency ac power to in-
dividual buses. Inverter 1 and 3 buses
supply power for the entire inverter load
except for the B SAS channel and one stand-
by attitude gyro which is supplied by the in-
verter 2 bus. In the event that either in-
verter 1 or 3 fails, its load can be trans-
ferred to bus 2 and full operation continued.
If inverter 2 fails, the B SAS channels and
standby attitude gyro will be inoperative and
inverter 1 or 3 load cannot be transferred.
Should inverters 1 and 3 fail simultaneously,
1-29
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Approved for Release: 2017/07/25 CO6230172
[0ENEkATOR'OUV'
. .
RESET RESET
LEFT 6>
GENERATOR TRIP
SW ITCHES (FWD)
LEFT
GENERATOR
30 KVA
TRIP
(AFT)
TO GYRO
GROUND �als
WARM-UP
LEFT
GENERATOR
CONTROL
RIGHT
GENERATOR
30 KVA
EXTERNAL
POWER
RECEPTACLE
RIGHT
GENERATOR
CONTROL
RIGHT
GENERATOR
SWITCH
TRIP
RESET
RESET
TRIP
(FWD) (AFT)
GENERATOR OUT
TRIM
POWER
ON SWITCHES
OF4F?(FWD)
TRIM ACTUATOR
TRANSFORMER
03
0
LEFT GENERATOR BUS
LEFT GENERATOR BUS S EL
NO, 1 N HEATERS
FUEL CROSSFEED VALVE
LEFT ENGINE FUEL S/0 VALVE
BOOST PUMPS (8) ODD NO.'S
1. FUEL DUMP VALVE
PITOT HEATERS
LANDING AND TAXI LIGHTS
PANEL LIGHTS
FLOOD LIGHTS
INSTRUMENT LIGHTS
INS EQUIPMENT
IFF AND TACAN EQUIP.
L. XFMR RECT OUT-: .
R. XFMR_ kid OUT
RIGHT GENERATOR BUS
RIGHT GENERATOR BUS SEL
ON NO. 2 N HEATERS
RIGHT ENG. FUEL S/0 VALVE
BOOST PUMPS (8) EVEN NO. 'S
OFF R FUEL DUMP VALVE
(AFT)
TRIM POWER
BUS
LEFT ENGINE AIR START
RIGHT ENGINE AIR START
LEFT XFMR
RECTIFIER
200 AMP
EMER BAT 0.1\1,
NO. 1
INSTR
XFMR 26V
NAV IND
ADF
ESSENTIAL DC BUS
ESSENTIAL DC
BUS RELAY
NO .1
.
INV. OUT�
L AND R GENERATOR CONT.
NO.
1
ENGINE FUEL CONTROL
ENGINE FUEL SHUTOFF
INVERTER
FUEL TRANSFER
FUEL DUMP
FUEL CROSSFEED
INFLIGHT REFUEL ( IFR)
OFF
NORM
(600 VA)
NO. 1
NVERTER
BATTERY
ER SWITCH
EMERGENCY SPIKE
RELAY
SPIKE OVERRIDE
DRAG CHUTE
NO 1 NV
COCKPIT LIGHTS
XFER REL
NORM
INS
WARNING LIGHTS
TURN RATE
UHF RADIO
NO. 2
INTERPHONE
INV. OUT
BRAKE AND SKID CONT.
INS - GOE
FACE HEAT
SWITCH TEMP INDICATOR
NORM
AIR CONDITIONING
NO. 2
RUDDER LIMITS
INVERTER
MON
DC BUS
L AND R HYDRAULIC SYST
TRIM CONTROL
OFF
(600 VA)
AUTOPILOT
NO. 2
INS
SAS
INVERTER
IN
ADF
SWITCH
1No.3
REL
XFER
NO. 1 AND NO. 2 N QTY
NO.
INVERTER
3
INLET BYPASS DOORS
INVERTER CONTROL
NORM
RIGHT XFMR
MACH TRIM
(600
VA)
RECTIFIER
200 AMP
IFF
PILOT VALVE CONTROL
OFF
NO. 3
L AND R GROUND START EMER.
RAIN REMOVAL NO.
INVERTER
FIRE WARNING LIGHTS
SWITCH
TACAN
3
INV. OUT
RESERVE HYD OIL LAND R
BATTERY
25 AMP - HR
BAT BAT
02E._
EXT. PW R. EXT. PVV R.
(FIND) (AFT)
BATTERY
EXTERNAL
POWER
SWITCHES
X-BAND BEACON
INLET BYPASS INDICATOR
ENGINE WATER PURGE
AFT FUEL TRANSFER
STEER AND IFR RELAY
LANDING GEAR WARNING LT.
L, R-A/B POWER
A/B CONTROL
SEAT ADJUST
TRANSFER PANELS
L. G. INDICATOR
L. G. CONTROL
BEACON LIGHTS
PERISCOPE PROJECTOR
NO. 3
IN SIR
XFMR 26V
An IND
4,..1NO . 1
INVERTER BUS
SAS A CHANNEL
(PITCH, YAW, ROLL)
FRS
NO. 1 AND 2 N QTY
HSI
L AND R FUEL FLOW
FLIGHT RECORDER
L AND R EGT IND
FUEL QTY
ANGLE OF ATTACK
AIR COND.
OXYGEN IND
LAND R FIRE WARN
L AND R CIT IND.
L AND R ENP
NO 2
INV BUS
SAS B CHANNEL
(PITCH, YAW, ROLL)
STANDBY An GYRO
NO. 2
IN SIR
XFMR 26V
INLET BYPASS IND
FUEL TANK PRESS
LAND R OIL PRESS
A AND B HYD PRESS
TRIM INDICATORS
NO 3
INV BUS
MACH TRIM
SAS PITCH AND YAW
MON
AIR DATA IND
AUTO PILOT
AIR DATA COMP
INLET BYPASS IND
INS
ATTITUDE IND
BEACON LIGHTS
a 213MOd 1VD
Approved for Release: 2017/07/25 C06230172
TA-12 SECTION I
the load of only one of the inverters can be
transferred to the No. 2 inverter. (See
figure 1-15, Electrical Power Distribution
diagram.)
EXTERNAL POWER SUPPLY
The aircraft is equipped with a receptacle
for connecting an external ac power source
to the aircraft electrical system. This re-
ceptacle is located in the nosewheel well.
When the external power source is connected
to the aircra_ft and the BAT-EXT PWR switch
is in the EXT PWR position, the generators
are automatically disconnected from their
respective buses and both buses receive
power from the ground power unit.
DC ELECTRICAL POWER SUPPLY
DC electrical power for the dc essential bus
and the dc monitored bus is supplied from
a 200-amp transformer-rectifier from each
ac bus. The two transformer rectifiers are
in parallel to supply the essential dc bus
and the monitored dc bus. The 25-ampere-
hour battery for emergency use will only
supply current to the essential dc bus when
both transformer-rectifiers are inoperative
and the battery switch is ON.
CIRCUIT BREAKERS
The circuit breaker panels in the cockpit
are located on the right and left consoles
and below the annunciator panel, and con-
tain push-to-reset, pullout-type breakers
for certain ac and dc circuits. Circuit
breaker panels which are not accessible
during flight, but which should be inspected
before flight, are located in the air condi-
tioning bay and in the electrical load center
(left-hand side of nosewheel well).
Generator Switches
A generator switch for each generator sys-
tem is located on the right side of each in-
strument panel and is powered from the
essential dc bus. Each switch has three
positions, GEN RESET, TRIP, and center
(neutral). The switches are spring loaded
to the center neutral position. Placing
either switch up to the GEN RESET position
will return the respective generator to nor-
mal operation if it has been removed from
the bus for any reason other than complete
generator failure. In the down, TRIP, posi-
tion the automatic bus transfer system will
supply that bus from the other generator if
it is operating.
The generators must be reset and
connected to the bus after the
engines are started and before
the ground power is removed.
The BAT-EXT PWR switches
must be moved to the OFF posi-
tion within 5 seconds after the
generator is reset or the gen-
erator will trip.
Battery-External Power Switches
A three-position, center-off, battery-
external power switch is located on the
right side of each instrument panel. In the
BAT (up) position the 25-ampere-hour bat-
tery is connected to the essential dc bus if
neither 200-amp transformer-rectifier is
furnishing power to the essential dc bus.
The BAT-EXT PWR switch should be in
the BAT position during flight so that the
battery will be automatically connected to
the essential dc bus if both transformer-
rectifiers fail. In the EXT PWR position,
the external power source, if connected and
operating, furnishes power for the entire
electrical system. In the center OFF posi-
tion, neither external nor battery power is
supplied.
1-31
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I' � �rove� or 'e ease: � � �
II I :1 1aa/111M
0 SAS
PITCH A
NO.1 N QTY
:FUEL QTY
L CIT IND
0r�AS
FRS YAW A
L FUEL FLOW
0
AIR COND
R CIT IND
L ENP IND R ENP IND
0
L FIRE WARN
�I
PF
0 SAS
FRS ROLL A
NO.2 N QTY R FUEL FLOW
- L EGT IND
� OXY IND
Y 1
R EGT IND B c HS I
0
FLT REC
0
wPM
OXY IND AP HSI
0 SVIOV2
E
AIR COND PF PF
A
ANGLE ATTK
0 Qr�Q
NAV T---' P & YAW 1
N T
S R
T I
R WI
RAIN REM ROLL c
O.
INLET DR.
XFER �
PANEL
DRAG
PROJ � CHUTE
BEACON
X-BAND IFF BCN ITS
0
R FIRE WARN INSTR XEN1R
NO. 1 INV
UHF�
�
INTPH
ADF
TACAN
0 0 0
0
NO.
C SAS 0
FRS --I PITCH B-1
0 C A
A A
j.
SAS YAW B SAS ROLL Bi
B
INS
BCN ITS
PF
PF- p p INSTR XFMR
'0; 0
A
CANOPY
MON c CAMERAS
LOGIC
SEAT
ADJ
AUTO
PILOT
TURN
RATE
0
AL9A
NO.
2
S STDBY ATT S INLET DOOR
G E P
Y 0
R S
NO. 2 INV &-11
IJI
WARN
CKPT
0
G
H
T
S
WARN
FIRE WARN NO. 1
0
INS � INS �
0
BCN LTS BCN LTS
0
A SAS PITCH AUTO PLT
P A
y 0 M 0
0
HW N
AIR DATA A AIR DATA
I C
0
D M
ATT IND E'
,.--.......
0
INSTR XEN1R
..,----.
B
ENG I y
N P
DRAG ET
CHUTE S
Q
ENG WATER NO. 2 T
G
FACE HT c FACE HT c
F A
K K
W F
D P T P
T T
ESSENTIAL 0 DC
1
-J
AFT CKPT-1
0 Al
FWD CKPT C
0 ON
TEMP
IND
NO.3 INV
L
la 0
Of
AFT FUEL �
XFER
0
0
CONT
FWD XFER
DUMP
X-FEED
IFR
� ' 0 0
0 C 0 0
RES OIL1 WARN 1
0 yH
0 ys
EMER
FUEL
ENG
R OFF
(C)
BRAKE 8
SKID
0
0 k 0 0
CONT
IND
A
A
STEER 8
IFR RLY]
1.42ISIUME
Figure 1-16 (Sheet 1 of 2)
LIMN= ATM 04 MriaKei V/LIMAIMOWIli VA
Approved for Release: 2017/07/25 C06230172
TA-12
SECTION I
CIRCUIT BREAKER PANELS
FWD COCKPIT
ROLL A AUTO P
TRIM POWER
DETAIL A
CIRCUIT BREAKER PANEL TRIM POWER RH CONSOLE
LDG LTS
CHINE
TACAN PITOT HT P 'TOT HT
..�----,.....� . ,�....,��. ......",....
0
I NSTR , PANEL
LTS �LTS TAXI JS
L. AC GEN
DETAIL
CIRCUIT BREAKER LH CONSOLE F/S 318
GEAR RELEASE
TRIM IND NAV -OIL PRESS -1
PITCH ROLL YAW IND L R ADF
0 0 0 0 0 0 0
HYD PRESS
A CONT B L SPIKE R
FUEL TANK r---- IND
PRESS BYPASS ATE
0 0 0 0 00
DETAIL B
CIRCUIT BREAKER PANEL CENTER INSTR PANEL
0
1201-9(2)(d)
Figure 1-16 (Sheet 2 of 2)
Approved for Release: 2017/07/25 C06230172
1-33
SECTION I
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TA-12
Inverter Switches
A switch for each of the three inverters is
located on the right side of each instrument
panel. The No. 1 and No. 3 inverter
switches have three positions: NORM, OFF,
and EMER. For normal operation the in-
verter switches are placed in the NORM
position. The No. 2 inverter switch is
placed in the ON position for normal oper-
ation. When either the No. 1 or No. 3 in-
verter switch is placed in the EMER posi-
tion, the respective inverter load is trans-
ferred to the No. 2 inverter. If both the
No. 1 and No. 3 inverter switches are
placed in the EMER position the No. 1 in-
verter load only will transfer to the No. 2
inverter. When the aft cockpit switches
are placed in the EMER position they will
override the forward cockpit switches.
Indicator and Light Test Pushbutton Switch
A pushbutton switch, labeled IND & LT TEST
is located on the left forward panel in each
cockpit. The pushbutton switch, when de-
pressed, illuminates the landing gear lever
red light, all annunciator panel lights, the
right and left nacelle fire warning lights,
fuel boost pump lights, and actuates the
gear warning tone in the headsets. This
switch is also used to test the operation of
the dual liquid nitrogen indicator which is
located in the forward cockpit. When the
aircraft is airborne, depressing the push-
button switch illuminates the three green
landing gear position lights for test.
Generator Out Indicator Lights
The L GENERATOR OUT and R GENER-
ATOR OUT indicator lights are located on
the annunciator panels and illuminate when
the respective generator is not furnishing
power to the respective generator bus.
Transformer-Rectifier-Out Indicator Lights
The L XFMR-RECT OUT and R XFMR -
RECT OUT indicator lights are located on
the annunciator panels and illuminate to
indicate that the respective transformer-
rectifier is not furnishing dc power to the
dc buses.
Inverter Out Indicator Lights
Three inverter out lights, labeled NO. 1
INVERTER OUT, NO. 2 INVERTER OUT
and NO. 3 INVERTER OUT, are installed
on the annunciator panel in each cockpit.
When illuminated, the respective light in-
dicates that the inverter bus voltage is be-
low minimum. When NO. 1 INVERTER
OUT or NO. 3 INVERTER OUT light illum-
inates, the inverter load may be transferred
to the No. 2 inverter if operative, by placing
the respective failed inverter switch to the
EMER position. The light will extinguish
after load transfer is accomplished. If
both NO. 1 INVERTER OUT and NO. 3 IN-
VERTER OUT lights illuminate, one in-
verter load only can be transferred to the
No. 2 inverter.
Emergency-Battery-On Indicator Lights
The emergency-battery on lights, labeled
EMER BAT ON, are located on the annun-
ciator panels. The lights illuminate when
the battery is furnishing power to the es-
sential 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. Hydraulic fluid is cooled by fuel-
oil heat exchangers, using the aircraft fuel
supply as the cooling agent. The A and B
hydraulic systems provide power for op-
erating the flight controls. The L and R
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TA-12
SECTION I
A & B HYDRAULIC POWER SUPPLY SYSTEM
ONE GALLON
LOW LEVEL
SWITCH
QUANTITY
IND I CATOR
RETURN
CONNECTION
SHUTOFF L������
VALVE SEAL
SEAL
Wiej P. 4. SO S. II
DRAIN I
e e e .2. . PO 2. 12 SO .� ON .2 .9 .. .� .2 St �.
4.L DRAIN
� a FILL PORT
11111 OVERBOARD
RELIEF
mixml
RELIEF VALVE
A RES=
RELIEF VALVE
',HEAT
EXCHANGER HEAT rl
L.EXCHANGER
Fire
PRESSURE
A
HYD
PUMP
80 SHUTOFF L__)___
VALVE
������
PRESS
CONN
0 REG
PRESSURE
FILTER
1\pcimie
N2 PRESS
02 FILL
N2 CYL
RETURN
FILTER
a
a
"IN RELIEF
\ VALVE
�
i �
SHUTOFF
VALVE
RESTR I CTOR
PRESS
SW ITCH
ACCUMULATOR
TO SURFACE
CONTROLS
TEMPERATURE
CONTROL CONTROL
TEMPERATURE
HYD L
PRESS TRANS
N2 FILLER
N2 GAGE
TO B
RELIEF
RESERVOIR
VALVE
VENT VALVE
�������
RESERVE
HYD TANK
OFF
HYD RES OIL
RETURN
FILTER
RELIEF s'
VALVE j
SI
� SHUTOFF
VALVE
B HYD L
PI In MI OM =1
A.
FROM SURFACE
CONTROLS
RESTRICTOR
PRESS TRANS
N2 GAGE
PRESS
SWITCH
ACCUMULATOR
ONE GALLON
LOW LEVEL
SW ITCH
RETURN
CONNECTION
SHUTOFF
VALVE
RES
PRESSURE
j SHUTOFF
VALVE
PRESS
CONN
4in
REG
PRESSURE
FILTER
TO SURFACE
CONTROLS
maim A SYSTEM PRESSURE B SYSTEM PRESSURE . � ELECTRICAL
mom A SYSTEM RETURN ammo B SYSTEM RETURN RESERVE OIL SUPPLY
um3Lamix CASE DRAIN marnamaisans N2 PRESSURE
I Figure 1-17
N2 PRESS
N2 FILL
N2 CYL
F201-29(b)
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SECTION I TA-12
L & R HYDRAULIC POWER SUPPLY SYSTEM
RELIEF
VALVE
H.wv\
RESERVOIR
PRESS IND
DUMP
VALVE
REG
GUAGE
0 N2
FILLER
N2
CYLINDER
�1333:EL
HEAT
EXCHANGER
(AFT CKPT)
OPEN
FUEL COOLING
CIRCULATING
PUMP
FWD
CKPT �1;1 INLET AIR
CLOSED BYPASS
AND CONTROL
RETURN
CONNECTION
NMI
SPIKE E.
MAIN
CONTROL
SPIKE
ACTUATOR
AND CONTROL
CROSS-
OVER
VALVE
(RETURN)
L R
ALTERNATE BRAKE
RETURN SELECTOR
VALVE
REFUELING
DOOR AND
PROBE
CHECK
VALVE
SYSTEM
RELIEF
VALVE
GROUND
PRESSURE
CONNECTIONS
PRESS
TRANS
L HOD Co
HYDRAULIC
PRESSURE
SPIKE
R HOD CO
N2 FILLER
"IIG
F201.47
Figure 1-19 (Sheet 2 of 3)
1-39
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(E JO E 4"9S) 611 e4^61:J
CABLE TENSION REGULATOR
AND SLACK ABSORBER (PITCH)
SURFACE LIMITER CONTROLS-
0
FWD
COCKPIT
DUAL HYDRAULIC CONTROL VALVE
AND BIAS SPRING (INBD)
ROD FROM MIXER
TO INBOARD SERVO (R.H.)
CABLE TENSION REGULATOR
AND SLACK ABSORBER (R011)-1
AFT
COCKPIT
1�AMEIREIREIREari
0
SWITCHES FOR
SURF LIMITER
WARNING
SIM
FWD
CONTROL
STICK
ELECTRO-MECHANICAL)_ROLL �
TRIM ACTUATOR
IMN
1 P
AFT
CONTROL
STICK
0
ELECTRO-HYDRAULIC ENGAGE
AND TRANSFER VALVE (ROLU
ELECTRO-HYDRAULIC ENGAGE
AND TRANSFER VALVE (PITCH)
SEE Fl G.1- 21
FOR OUTBOARD
CONTROL SURFACE
9
ungi-
ANTI -B I AS
SPRING
ACTUATING
CYLINDERS (6)
IFmsairr
ww,
PITCH MIXER
STOPS
ROD FROM MIXER
TO INBOARD SERVO
(L H.)
INBOARD
CONTROL
SURFACE
ROLL FEEL SPRING
PITCH FEEL SPRING
PITCH QUADRANT
IN TAI L CONE
ROLL QUADRANT
IN TAIL CONE
FLIGHT CONTROL SYSTEM (Outboard Elevons)
ELECTRO-MECHANICAL }PITCH
TRIM ACTUATOR-2 SPEED
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TA-12
SECTION I
the servo input to the outboard elevon con-
nected 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:
Pitch
Roll
Pitch plus Roll
Yaw
Elevons
11 deg down
24 deg up
12 deg down
12 deg up
20 deg down
35 deg up
Rudders
20 deg left
20 deg right
Manually operated mechanical stops are in-
corporated in the cockpit mechanism to
limit the surface movement at high speed.
Elevon travel in roll is limited to 7 degrees
up, 7 degrees down, and rudder travel is
limited to 10 degrees right, 10 degrees left.
An additional stop is installed in each rud-
der servo package to limit rudder travel.
These stops are electrically controlled and
hydraulically operated by separate electrical
and hydraulic systems. If no electrical
power is available, the rudders will be
limited to approximately 10 degrees L and
R travel. If electrical power is available
to one stop, that rudder only will have the
full 20 degrees L and R travel available.
The rudder cable must be stretched to ob-
tain this travel, causing a noticeable in-
crease in rudder pedal force.
CABLE SYSTEM
Cable systems are utilized to transfer con-
trol movements from the control sticks and
rudder pedals to the flight control mechan-
isms. The pitch and roll axis cable sys-
tems are duplicated from the aft cockpit
only to the mixing mechanism in the aft
fuselage. The rudder system has two sep-
arate closed loop single cable systems, one
to each rudder. Cable tension regulators
and slack absorbers are incorporated in the
cable systems.
ARTIFICIAL FEEL SYSTEM
The use of a fully powered, irreversible
control system for actuation of the surfaces
prevents air loads and resulting "feel"
from reaching the cockpit controls. There-
fore, feel springs are installed in each of
the pitch, roll, and yaw axis mechanical
control mechanisms to provide an artificial
sense of control feel. The springs apply
loads to the pilot controls in proportion to
the degree of control deflection.
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 stick position
remains neutral, irrespective of the amount
of trim. The trim actuator and feel spring
location is combined in the rudder mechan-
ism and yaw trim is reflected by rudder
pedal position.
Travel limits of the trim system are 3-1/2
degrees down to 6-1/2 degrees up in pitch;
4.5 degrees up and down (each side) in roll;
and 10 degrees left to 10 degrees right in
yaw. Trim position indicators are provided
for each axis. Trim rates are as follows:
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SECTION I
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TA-12
Pitch
Roll Yaw
1.120 /sec. .40 /sec. 1.10 /sec.
Automatic pitch trim uses a separate, slow
speed motor for auto trim when the autopilot
is engaged and mach trim when the autopilot
is not engaged. This trim motor operates
at one-tenth the rate of the manual trim
motor, or .112 /sec.
CONTROL STICKS
The control sticks are mechanically con-
nected by torque tubes, pushrods, bell-
cranks, and cables to the dual cable system
which operates the roll and pitch quadrants
in the aft fuselage tailcone. Mechanical
pushrod linkages mix the control movements
and position dual hydraulic control valves.
These valves direct both A and B system
hydraulic pressures to the inboard elevon
actuating cylinders.
Pushrods, bellcranks, 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 pushrod
followup system closes off the flow of hy-
draulic fluid to the actuators when the de-
sired elevon deflection is obtained. Located
on each control stick grip are pitch and yaw
trim switches, a combination nosewheel
steering and autopilot control stick command
button, a microphone switch for both inter-
phone and radio transmission, an autopilot
disconnect switch, and an in-flight refueling
disconnect switch.
RUDDER PEDALS
Primary control for the rudders consists
of conventional rudder pedals mechanically
connected by cables, bellcranks, 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 ADS, located at
the bottom of the respective cockpit lower
Instrument panel. Wheel brakes are con-
trolled conventionally by toe action on the
rudder pedals; refer to Wheel Brake System,
this section. Rudder pedal movement also
controls nosewheel steering; refer to Nose-
wheel Steering System, this section.
The pedals in the forward cockpit are hinged
to fold inboard and upward, to provide ad-
ditional foot space on the cockpit floor.
Pitch and Yaw Trim Switches
Pitch and yaw trim control is provided by
a spring-loaded, four-position, thumb-
actuated switch installed on each control
stick grip. The switch positions are center
OFF, LEFT, RIGHT, NOSE UP, and NOSE
DOWN. The switches control trim motors
powered by the right generator bus through
the 28 volt trim actuator transformer and
trim power bus.
Note
The trim power switches must be
in the ON position before the pitch,
roll, and yaw trim switches will
operate.
The aft cockpit trim switch is capable of
overriding the forward cockpit switch.
Lateral movement of either switch to the
left corrects for right yaw and lateral move-
ment to the right corrects for left yaw.
Forward movement of either switch produces
down elevon operation of the trim motors
and actuators (aircraft nose down). Aft
movement moves the elevons up (aircraft
nose up).
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TA-12
SECTION I
CONTROL STICK GRIP (Both Cockpits)
TOP VIEW
FRONT VIEW
SIDE VIEW
1 TRANSMIT1ER - INTERPHONE CONTROL SWITCH
2 CONTROL STICK COMMAND - NOSEWHEEL STEERING BUTFON
3 PITCH AND YAW TRIM SW ITCH
4 EMERGENCY AUTOPILOT DISENGAGE SWITCH
AND AIR REFUEL DISCONNECT
� F201-27(b)
Figure 1-20
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SECTION I
TA-12
Trim Power Switches
The trim power switch, installed on the
annunciator panel in each cockpit, has two
positions, ON and OFF. Both the forward
and aft cockpit switches must be in the ON
position for the system to be operative. To
prevent inadvertent movement, the switches
must first be pulled out before they can be
moved from ON to OFF. When in the ON
position, the right generator bus power is
provided to the roll, pitch, and yaw trim
actuators. The trim power circuit breaker
is located in the electrical load center and
is not available to the pilots.
Roll Trim Switches
A three-position roll trim switch is installed
just forward of each throttle quadrant. The
switch positions left and right are indicated
by arrows. The switch is spring-loaded to
center. When either switch is held in the
right position, the roll trim motor actuates
to move the right elevons up and the left
elevons down. Actuation of the switch to
the left position moves the right elevons
down and left elevons up. The aft cockpit,
switch is capable of overriding the forward
cockpit switch. 28-volt ac power is fur-
nished by the trim power bus.
Rudder-Synchronization Switches
A three-position rudder synchronization
switch is installed just forward of each
throttle quadrant. The switch positions
(left and right) are indicated by arrows.
The switches are spring-loaded to center.
When in the left and right positions the
switches provide 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 superimposing the L and R pointer on the
yaw trim gage. The aft cockpit switch is
capable of overriding the forward cockpit
switch. 28-volt ac power is furnished by
the trim power bus.
Roll, Pitch, and Yaw Trim Indicators
Separate roll, pitch and yaw indicators are
installed on the left side of each instrument
panel. The ROLL trim indicators use a
double-ended pointer to display the amount
of differential roll trim from 0 to 9 degrees.
The PITCH trim indicators display the
amount of pitch trim from 5 degrees nose-
down to 10 degrees nose-up, although only
8-1/2 degrees nose-up trim is available.
The YAW trim indicators use two separate
pointers, one for each rudder and marked
R and L, to display the amount of yaw trim
from 10 degrees left to 10 degrees right.
Rudder synchronization is obtained by
superimposing the L and R pointers on the
indicators. 26-volt ac power for the indi-
cators is normally furnished by the No. 2
instrument transformer and the No. 3 in-
verter.
Surface Limiting Control Handles
Interconnected T-handles are located on the
annunciator panel. When either handle is
turned 90 degrees counterclockwise and re-
leased, the mechanical stops in the roll and
yaw axis of the cockpit control system are
activated. This action also opens an elec-
trical switch which de-energizes a solenoid-
operated valve in each rudder servo pack-
age and activates the servo package rudder
stops. When either handle is pulled out and
rotated 90 degrees clockwise, the mechan-
ical stops in the cockpit are released and
the solenoid is energized, releasing the
servo package stops.
Surface Limiter Indicator Lights
When speed exceeds Mach 0.5, the SUR-
FACE LIMITER indicator lights illuminate
on the annunciator panels until either sur-
face limiter handle is released. If the
speed is less than Mach 0.5 and the surface
limiters are on, the SURFACE LIMITER
indicator lights illuminate until either, sur-
face limiter handle is pulled out. Power
for the lights is furnished by the essential
dc bus.
1-44
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TA-12 SECTION I
AUTOMATIC FLIGHT CONTROL SYSTEM
The automatic flight control system includes
stability augmentation, autopilot, Mach
trim, and air data systems, plus additional
subsystems furnishing attitude and navi-
gational course inputs for the autopilot. The
air data system furnishes signals to the
autopilot, Mach trim, and inertial naviga-
tional systems. The stability augmentation
system supplies signals to the hydraulic
servos that operate the control surfaces.
The Mach trim system furnishes signals to
the slow-speed motor on the pitch trim ac-
tuator. The inertial navigation system sup-
plies attitude and navigational course inputs
for the autopilot. Heading and attitude re-
ference signals for the autopilot are also
supplied by the FRS. The autopilot moves
the aircraft hydraulic servos through the
stability augmentation system. For further
information on the autopilot and inertial
navigation systems, refer to Section IV.
STABILITY AUGMENTATION SYSTEM
The three-axes stability augmentation sys-
tem (SAS) is a combination of electronic
and hydraulic equipment which augments the
inherent stability of the aircraft. It is de-
signed for optimum performance at the
basic mission cruise speed and altitude, but
it also provides improved stability for in-
flight refueling, landing, and takeoff. The
SAS is part of the aircraft basic control sys-
tern and is normally used for all flight con-
ditions.
Dual electronic channels are provided for
all axes, and a monitor channel is provided
for both the pitch and yaw axes. Logic cir-
cuits compare the functioning of each pitch
and yaw channel and automatically eliminate
a failed channel. The pilots are provided
with a visual warning on the annunciator
panel of a failed channel. The monitor
channels for the pitch and yaw axes are
powered by inverter 3.
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 simulated logic circuit is provided for
the roll channel to warn of a malfunction
and to disconnect the two channels. A sep-
arate 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.
STABILITY AUGMENTATION PITCH AXIS
Two independent active channels termed 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. Damp-
ing 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. The A channel
uses the A hydraulic system and the B chan-
nel uses the B hydraulic system. This
avoids loss of both channels in case of
failure of either the A or B hydraulic sys-
tems. 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 aircraft.
Phasing of the gyro signals is such that an
angular pitch motion produces elevon move-
ment 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 de-
signed to aid the pilot in avoiding overcon-
trol and improve the handling qualities of
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SECTION I
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TA-12
SAS AND AUTOPILOT CONTROL PANEL
17
16
15
14
21
JI
MACH AUTO HEADING .
HOLD NAV ON HOLD-
AUTOPILOT )
. . :
FORWARD COCKPIT
13
12
11
971-1
4
5
1 SAS CHANNEL ENGAGE SWITCHES
2 ROLL CHANNEL DISENGAGE LIGHT
3 SAS RECYCLE INDICATOR LIGHTS
4 SAS LIGHT TEST SWITCH
5 A/P HEADING HOLD SWITCH
6 A/P ROLL ENGAGE SW ITCH
7 A/P ROLL TRIM SYNCHRONIZATION INDICATOR
8 BACKUP PITCH DAMPER CONTROL INDICATOR LIGHT
9 OVERRIDE POWER TRANSFER SWITCH
10 YAW LOGIC OVERRIDE CONTROL INDICATOR LIGHTS
11 PITCH LOGIC OVERRIDE CONTROL INDICATOR LIGHTS
12 A/P AUTO NAV SWITCH
13 A/P TURN CONTROL SWITCH
14 A/P PITCH TRIM SYNCHRONIZATION INDICATOR
15 A/P PITCH ENGAGE SWITCH
16 A/P PITCH CONTROL WHEEL
17 A/P MACH HOLD SWITCH
STAB AUG,
"B B'
NORM t-. Ncou4
., ....,,
OFF OFF
UTE i 1 TEST
MON MON
PITCH YAW
LOGIC OVERRIDE
AFT COCKPIT
10
B/U DAMPER
0
AFT
OVERRIDE CONTROL
972-1
4
F201-18(b)
1-46
Figure 1-21
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TA-12
SECTION I
the aircraft. The logic circuit is able to
isolate a SAS failure in either the electronics
or the servos. When a malfunction is iso-
lated, the failed channel will disengage and
the system continues in operation on a single
channel. Malfunctioning and disengaging of
channels is indicated to the pilots by indi-
cator lights. The pitch axis can command
a maximum elevon surface travel of 2.5 de-
grees up to 6.5 degrees down. Dual or sin-
gle 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 chan-
nel is from the A phase of No. 2 inverter
and MON channel power is from the B phase
of the No. 3 inverter. Each power source
is protected by individual circuit breakers
in the forward 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
hydraulic 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 sig-
nals when a deliberate turn is made. A
lateral accelerometer sensor is also used
in each channel of the yaw axis to minimize
steady-state sideslip caused by an engine
failure until the pilot can retrim the rud-
ders. The logic circuit is identical to the
pitch axis and functions in the same mariner.
The yaw axis can produce a maximum rud-
der travel of 8 degrees left to 8 degrees
right (each surface). Corrective surface
motion is the same regardless of one or
two-channel operation. Power for the A
channel is from the B phase of inverter 1,
for the B channel from the B phase of in-
verter 2, and for the monitor channel\ from
the B phase of inverter 3. The circuitry
from each power source is protected by in-
dividual circuit breakers.
STABILITY AUGMENTATION ROLL AXIS
The reliability requirements for the roll
axis are not as severe as for 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. The
A channel positions the left elevon surfaces
and operates from the A hydraulic system;
the B channel positions the right elevon
surfaces and operates from the B hydraulic
system. Each channel can be operated in-
dividually. There is no monitor channel as
such; there is, however, a simulated logic
circuit to disengage both channels and il-
luminate a light on the SAS panel if a roll
channel malfunctions. Although the system
gain is the same as for two-channel oper-
ation, roll control is not symmetrical.
Coupling into the yaw and pitch axes is pos-
sible, but the systems operating in those
axes minimize undesirable aircraft motion.
Maximum elevon travel in the roll axis is
2 degrees up to 2 degrees down (each side),
for a total of 4 degrees differential with
both systems operating. Power for A chan-
nel is from the C phase of inverter 1, and
power for the B channel from the C phase
of inverter 2.
STABILITY AUGMENTATION SYSTEM (SAS)
CONTROL PANELS
The SAS control panel on each right console
contains six channel engage switches, A and
B channels for the pitch, roll and yaw axes.
The panels also contain a press-to-test
switch and six indicator lights for the A, B,
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SECTION I
TA-12
and MON channels in the pitch and yaw axes.
Three guarded switches for the backup pitch
damper, pitch logic override, and yaw logic
override are located on the lower instru-
ment panel and right side of the center con-
sole. Individual circuit breakers are lo-
cated on both right and left consoles.
Aft Cockpit Stability Augmentation System
Override Panel
The aft cockpit SAS panel also contains an
additional panel with five lights and a toggle
switch. The switch is labeled OVERRIDE
CONTROL and has two positions, FWD (up)
and AFT (down). The five lights are iden-
tified as follows: one each for A and B
channels of the pitch logic override circuit,
one each for the yaw logic override circuit,
and one for the backup pitch damper. When
the switch is in the FWD position, it allows
the aft cockpit pilot to determine the posi-
tion of the pitch logic override, yaw logic
override, and backup pitch damper switches
in the forward cockpit by observing which
of the lights are illuminated. When the aft
cockpit pilot moves the switch to the AFT
position, control of the BUPD and logic
override circuits is transferred to the aft
cockpit and the five lights indicate aft cock-
pit switch positions.
Channel Engage Switches
There are six channel engage toggle switches
on each SAS control panel. One pair is pro-
vided for each axis, pitch, roll, and yaw.
The forward switch of each pair controls
the A channel and the rear switch controls
the B channel. The forward cockpit switches
have two positions, ON (forward) and OFF
(aft). The aft cockpit switches have three
positions, ON (forward), NORM (center),
and OFF (aft). The NORM position on the
aft cockpit switches allows the forward
cockpit pilot to assume control of the chan-
nel engage switches. When the aft cockpit
switches are in the ON or OFF position
they override the forward cockpit switches.
When electrical power is on the aircraft
and the channel engage switches are OFF,
the SAS electronics are powered but the
channel servos are not engaged with the
control system. Moving the switches to
the ON position engages the SAS servos,
provided that the recycle light is extin-
guished. If the light is on, the light must
be depressed before engagement is possible.
Recycle Indicator Lights
Six recycle indicator lights are located on
the SAS control panel on each right console
adjacent to the pitch and yaw channel en-
gage switches. One light is provided 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 chan-
nel is functioning properly. If the light is
illuminated, it indicates that the channel
has disengaged and the light may be de-
pressed to recycle the channel. If the
failure is momentary, the channel will re-
engage; if the light reillurninates, it indi-
cates that the channel is malfunctioning.
(It is not necessary to turn the channel en-
gage switch off in a malfunctioning channel
because the light indicates automatic dis-
engagement.)
Note
The recycle indicator light should
be pressed down firmly and re-
leased. If the recycle light is
held down, a control surface
transient will occur if a hardover
servo condition exists in that
channel. Refer to Section III for
additional information.
The six recycle lights will illuminate when
electrical power is first applied to the air-
craft. The channel switches must be on
and the recycle lights must be pressed to
engage the channel electronics with the
servos. When engaged and operating, the
channel lights will be out.
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TA-12 SECTION I
Roll Channel Disengage Light
A single roll ehannel disengage light is lo-
cated between the two roll channel switches
on the forward cockpit SAS panel, and the
forward of the two switches on the SAS
panel in the aft cockpit. When illuminated,
the light indicates that both roll channels
have disengaged. This disengagement re-
sults when the roll servo channel outputs
differ by more than an amount equivalent to
0.6 degree surface deflection. When op-
erating 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 each SAS control panel. De-
pressing the pushbutton illuminates the six
recycle lights and one roll disengage light
for test.
SAS Pitch Logic Override Switch
The SAS pitch logic override switch is a
guarded, three-position switch, located on
each annunciator panel. Placing the switch
in the A (up) position eliminates the logic
circuit and selects A-channel operation. In
the B (down) position, the logic circuit. is
eliminated, and B-channel operation is se-
lected. When the switch is in the center,
guarded OFF position, the logic circuit
functions normally. The override switch is
only used as an emergency control. The
switch must be placed in either the A or B
position when the BUPD is used.
SAS Yaw Logic Override Switch
The three-position SAS yaw logic override
switch is located on each annunciator panel.
The switch is guarded in the OFF position.
The A (up) position eliminates the logic cir-
cuit and selects A-channel operation. The
B (down) position eliminates the logic cir-
cuit and selects B-channel operation. The
override switch is only used as an emer-
gency procedure.
BACKUP PITCH DAMPER
The primary purpose Of the backup pitch
damper (BUPD) is to provide an emergency
system for pitch stability augmentation dur-
ing refueling and landing approach. It is
used in case the SAS pitch channels are un-
usable due to electronics malfunction or
overheating of the pitch gyro package. The
system is optimized for use at light weight,
aft center of gravity, and subsonic speeds
from 0.3 to 0.8 Mach number; it is not in-
tended as an emergency backup system dur-
ing cruise.
Backup Pitch Damper Switch
A guarded BUPD switch is located on each
annunciator panel. It is guarded in the OFF
position. When in the ON position, the
backup gyro located in the electronics com-
partment supplies pitch rate signals through
an independent electronic channel to either
the A or B servo, depending on which is
selected by the pitch logic override switch.
MACH TRIM SYSTEM
In the transonic region in this aircraft, the
variation of elevon angle with Mach number
is such that it would normally require the
pilot to use nose-up trim with increasing
Mach number. This characteristic is re-
ferred to as "speed instability". To com-
pensate for this, the Mach trim system is
incorporated in the aircraft control system
to slowly drive the trailing edge of the
elevons upward as Mach number increases,
thus providing artificial stability by re-
quiring the pilot to apply nose-down trim as
Mach number increases. The system op-
erates between Mach 0.2 and 1.5 on a sched-
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SECTION I TA-12
ule which varies with Mach number in the
8-1/2 degrees nose-up and 5 degrees nose-
down trim limits range of the elevons. The
trim change rate is 15 degrees per Mach be-
tween 0.95 and 1.30 Mach, and 5 degrees
per Mach between 0.2 and 0.95 Mach and be-
tween 1.3 and 1.5 Mach. Signal input to the
Mach trim system is obtained from the air
data computer and the electronic components
associated with the system are located in
the autopilot electronic component assembly.
The system is operative whether or not any
SAS channel is engaged; however, the Mach
trim system does not function when the pitch
autopilot is engaged. The only controls over
the system are the circuit breakers in the
forward cockpit which should be pulled in the
event of an air data computer malfunction to
prevent undesirable Mach trim effects.
Power for Mach trim is furnished by the
No. 1 or No. 3 inverters, and the essential
dc bus.
PITOT-STATIC SYSTEM
Three-pitot-static systems supply the total
and static pressures necessary to operate
the basic flight instruments and air data
system components. Normally, the pres-
sures are sensed by an electrically heated
probe mounted on the nose of the aircraft.
The pitot orifice of the probe is divided in-
side the head to provide two separate pres-
sure sources. It also has two circumfer-
ential sets of four static pressure ports
each. One pitot and the aft set of static
ports supply pressure signals to the air
data computer system; the other set of
pickups supply the normal ship system pitot
and static pressure directly to the speed
sensors on the ejection seats, the altimeters,
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 stall warning light sensor. An alternate
pitot-static source is available from the
flight recorder system for flight instruments
in the forward cockpit.
Pitot Heat Switches
The heating elements of the nose and flight
recorder probes are controlled by two OFF-
ON pitot heat switches, located on the an-
nunciator panel in each cockpit. Power is
furnished by the left ac generator bus.
Pitot Pressure Selector Lever
The pitot pressure selector lever is located
on the forward cockpit right trim panel. It
Is normally safety wired in the NORMAL
position. In the event of a malfunction of
the normal pitot-static position system, the
lever may be moved to ALT position. This
furnishes pitot-static pressure from the
flight recorder system. to the altimeter,
the rate of climb and the airspeed indicator
In the front cockpit only.
Pitot Heat Indicator Lights
A pitot heat indicator light labeled PITOT
HEAT, is located on each annunciator panel.
When illuminated, the light indicates that
the pitot heat switch is not in the correct
position for the aircraft altitude. Power
for the lights is furnished by the essential
dc bus.
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 into the electrical
signals required for the pilot triple display
indicators and for the automatic flight con-
trol and inertial navigation systems. The
ports which supply pressure to the air data
computer are separate from those that fur-
nish pressure to the basic flight instruments;
therefore, failure of the air data computer
pressure source will not leave the pilot with-
out the necessary altitude, vertical velocity,
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TA-12
SECTION I
PITOT STATIC
FLIGHT RECORDER
AFT SPEED SENSOR
ANGLE OF ATTACK TRANSMITTER
FWD SPEED SENSOR
PITOT STATIC SELECTOR
z
FLIGHT RECORDER SOURCE ' /
FWD
COCKPIT
FLIGHT RECORDER SOURCE
AFT INSTRUMENT PANEL
AIR DATA COMPUTER
SAS TRANSDUCER
FORWARD INSTRUMENT PANEL
PITOT MAST
RATE OF CLIMB ALTIMETER INDICATED
AFT AIRSPEED
COCKPIT
PITOT STATIC
SELECTOR
VALVE
ALT NORM
PITOT STATIC PRESS
J i
TRIPLE
DISPLAY INDICATOR
RATE OF CLIMB ALTIMETER INDICATED TRIPLE
AIRSPEED DISPLAY INDICATOR
FWD
SEAT
AFT
SEAT
EJECTION SEAT
SPEED SENSOR
Sil PI
ANGLE OF
ATTACK
TRANSDUCER
f PS
IS
SAS
TRANSDUCER
SCHEDULER
,
F201-71
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SECTION I TA-12
or airspeed information to fly the aircraft.
The air data computer converts pitot-static
pressures into proportional rotary shaft
positions which are equivalent to pressure
altitude and dynamic pressure. These shaft
positions are combined 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 equipment are as listed below:
OUTPUTS
USING EQUIPMENT
Pressure Altitude
Equivalent Airspeed
Mach
Triple-Display
Indicator
Mach
Mach Rate
Altitude
Dynamic Pressure
Autopilot
Mach
Mach Trim System
Pressure Altitude
Inertial Navigator
Computer
Power for the air data computer is furnished
either by inverter 1 or 3, as selected by the
autopilot selector switch.
Triple-Display-Indicators
Triple-display indicators (TDI) are installed
on each instrument panel. The indicators
present digital indications of altitude in 50-
foot increments, Mach number in 0.01-
Mach increments, and equivalent airspeed
in 1-knot increments. Altitude readout
range is -1000 to 110,000 feet; Mach range
is 0.2 to 3.5; and speed range is 100 to 560
KEAS at sea level (decreasing to 466 KEAS
at Mach 2.5 and 460 KEAS at Mach 3.2.) If
the ADC loses power, an OFF flag appears
on the face of each indicator.
Note
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.
WARNING
The digital speed and altitude in-
dications are primarily used for
aircraft control above FL 180 and
to maintain proper airspeed con-
trol during climbs to FL 180.
Pitot-static instruments shall be
used in the landing pattern, during
takeoff until proper climb schedule
is established on the TDI, and
during all simulated or actual in-
strument flight below FL 180.
During subsonic flight pitot-static
instruments should be consulted
frequently to confirm correct air
data system operation.
INSTRUMENTS
For information regarding instruments that
are an integral part of a particular system,
refer to applicable paragraphs in this sec-
tion and Section IV.
Airspeed-Mach Meter
A combination airspeed and Mach meter
operating directly from pitot-static pres-
sure is installed in the basic six flight in-
strument group on each instrument panel.
This is a special instrument with airspeed
and Mach number ranges compatible with
aircraft performance. Mach number and
airspeed are simultaneously read on the
window and outer index, respectively. A
limit airspeed needle (white-barred) shows
the airspeed limit of the aircraft. The
actual airspeed limit is an equivalent air-
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TA-12 SECTION I
speed; however, the limit is shown as indi-
cated airspeed, the needle varying with alti-
tude to read the indicated airspeed that
converts to limit equivalent airspeed.
Altimeter
A sensitive pressure altimeter is installed
on each instrument panel. In addition to the
1000-foot and 100-foot pointers, it also has
a 10,000-foot indicator. This pointer ex-
tends to the edge of the dial with a triangular
marker at its extremity. The center disc
has a cutout through which black and yellow
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.
Attitude Indicator
An attitude indicator, located on the instru-
ment panel in each cockpit, combines the
functions of an attitude indicator and a turn
and slip indicator. Pitch and roll signals
from the INS or FRS are connected to each
indicator through an ATT/AP select switch
that is located on each instrument panel.
Control is transferred from one cockpit to
the other by moving the TACAN/INSTR trans-
fer switch on the left console in either cock-
pit. The incoming signals are used to posi-
tion an attitude sphere that has unrestricted
motion, allowing pitch and roll presentation
through 360 degrees. The sphere moves
behind a miniature aircraft silhouette fixed
at the center of the instrument. A pitch
trim knob allows manual positioning of the
sphere in pitch with relation to the miniature
aircraft. Pitch angle is displayed by the
relationship of the miniature aircraft to
markings located on the sphere. The sphere
is marked with a horizon line, small dots
for 5 degree increments, short lines for 10
degree increments, numeral markers for
each 30 degree increments, and large dots
to indicate the poles. Bank angle is shown
at the bottom circumference of the instru-
ment. Ten degree graduations are provided
for angles to 30 degrees, and 30 degree
graduations for angles up to 90 degrees of
bank. The turn and slip indicator is mounted
at the bottom of the attitude indicator, and
is centered with the vertical axis. A de-
flection of one needle width indicates a four
minute 360 degree standard turn. The rate
of turn transmitter receives power from the
essential dc bus. Bank and pitch steering
bars and a glideslope needle which are
visible when the instrument is deenergized,
are not used and are out of view when the
instrument is energized.
Standby Attitude Indicators
A standby attitude indicator located on each
Instrument panel provides 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 de-
grees and roll angles of 360 degrees. The
bank angle scale is marked on the outer
periphery. A pitch reference adjustment
knob is provided on the lower right corner
of the instrument for positioning the re-
ference bar as desired. A fast erect push-
button is located adjacent to the throttles
In each cockpit.
CAUTION
Do not hold fast erect button for
more than 45 seconds to prevent
overheating of fast erect motor.
The OFF flag will be visible whenever
power to the indicator is interrupted. This
instrument has its own self-contained gyro
and is not dependent on another reference
source.
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SECTION I TA-12
NAVIGATION INSTRUMENTS
ATTITUDE INDICATOR
NOTE
THE ATT/AP SELECT SWITCH DETERMINES WHICH
SYSTEM, NAV OR FRS SUPPLIES PITCH AND ROLL
TO THE ATTITUDE INDICATOR AND PITCH, ROLL
HORIZONTAL SITUATION INDICATOR
MODE SELECT
MAG BEARING ATT/AP
NAV TACAN SELECT SELECT
TACAN . FRS
FLIGHT INSTRUMENT CONTROL PANEL
... ....._ ._ . .._._._ ._ .. _
STEERING .S.IGNAIS ARE ONLY AVAILABLE IN THE
INS POS ITION 1
DISPLAY MODE SELECTOR SWITCH
NAV.
MAG
TACAN
INDICATOR
INDICATOR FUNCTION
BEARING SELECT SWITCH
BEARING SELECT SWITCH
BEARING SELECT SWITCH
TACAN 1 ADF
TACAN 1 ADF
TACAN 1 ADF
HORIZONTAL
SITUATION
INDICATOR
NAV STEERING
NAV STEERING
MANUALLY SET
HEADING MARKER II ll
BEARING POINTER 1).
TACAN
1 ADF
TACAN
1 ADF
TACAN
1 ADF
COURSE ARROW f
SERVOED TO
LUBBER LINE
SERVOED TO
LUBBER LINE
MANUALLY SET TO
SELECT.TACAN COURSE
COURSE DEVIATION I
CENTERED
CENTERED
LEFT -RIGHT
TACAN COURSE
COMPASS CARD
TRUE
MAGNET IC
MAGNETIC
TO-FROM �
OUT OF VIEW
OUT OF VIEW
TACAN
RANGE INDICATOR IIII
DISTANCE
TO SELECTED TACAN STATION
K SHUTTER
USED --vP-
-
NOT
0.-
7P
D 1ST. SHUTTER
OUT OF VIEW IF TACAN DISTANCE VALID
DIGITAL
COURSE DISPLAY
SERVOED TO
LUBBER LINE
SERVOED TO
LUBBER LINE
MANUALLY SET TO
SELECT.TACAN COURSE
ATTITUDE
INDICATOR
BANK DIRECTOR NEEDLE
OUT OF VIEW - NOT USED
PITCH DIRECTOR NEEDLE
OUT OF VIEW - NOT USED
GLIDE SLOPE SLOPE INDICATOR
OUT OF VIEW -NOT USED
VP.-
LOCALIZER FLAG
OUT OF VIEW - NOT USED
GLIDE SLOPE SLOPE FLAG
OUT OF VIEW - NOT USED
POWER FLAG FLAG
OUT OF
VIEW IF ATTITUDE REFERENCE
VALID
F201 -70
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TA-12 SECTION I
HORIZONTAL SITUATION INDICATOR (HSI)
There is a horizontal situation indicator lo-
cated on the instrument panel in each cock-
pit. Each HSI visually presents information
from Tacan, ADF, INS and FRS. Display
functions for both indicators are selected by
the use of the display MODE SELECT switch
and BEARING SELECT switch in either
cockpit. Control with these switches is
transferred from one cockpit to the other by
using the TACAN/INSTR transfer switch on
either control transfer panel. Power Ior
the HSI is supplied by the No. 1 inverter.
The various components of the indicators
are described below.
Rotary Compass Card
The compass card is a rotating azimuth
ring read at a stationary lubber line at the
12-o'clock position. The card displays
true heading from the INS source when the
display MODE SELECT switch having con-
trol is in the NAV position. When the dis-
play MODE SELECT switch having control
is in the MAG or TACAN position the card
displays magnetic heading from the FRS
source.
Bearing Pointer
The bearing pointer is a small arrow on
the outer periphery of the rotary compass
card, and indicates the bearing to either
the TACAN or ADF station as selected with
the BEARING SELECT switch on the instru-
ment panel which has control.
Heading Marker
The heading marker is a rectangular
marker located just outside of the rotating
compass card. When the display MODE
SELECT switch having control is in NAY
or MAG position the heading marker dis-
plays navigational steering. When the dis-
play MODE SELECT switch having control
is in the TACAN position the heading marker
can be set manually with the HEADING SET
knob on the lower left corner of the HSI
instrument.
Course Arrow and Course Deviation Bar
The course arrow and course deviation bar
are located inside the rotating compass
card. The course arrow points to the lub-
ber line and the course deviation bar is
centered when the display MODE SELECT
switch having control is in the NAY or MAG
position. When the display MODE SELECT
switch having control is in the TACAN posi-
tion, the course arrow may be manually set
to the desired tacan course with the course
set knob on the lower right corner of the
HSI instrument, and the course deviation
bar will indicate deviation left and right of
the selected course.
Digital Course Display
A digital course display located in the upper
right corner of the HSI displays at all times,
the same course indicated by the course
arrow on the compass card.
To-From Arrows
The to-from arrows are located on a radial
near the center of the HSI instrument, in
line with the course arrow. One or the
other arrow will be exposed to indicate the
direction to the station when TACAN mode
is selected and reliable tacan signals are
being received. At any other time, both
arrows will be masked from view.
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SECTION I
TA-12
Range Readout Window
The range readout window located in the
upper left corner of the HSI instrument is
labeled MILES and displays the slant range
in nautical miles to a selected tacan station
regardless of the position of the display
MODE SELECT switch.
Vertical-Velocity Indicators
A vertical-velocity indicator is installed on
each instrument panel and shows the rate
of change of altitude in feet per minute.
Changes in pressure due to changes in alti-
tude are sensed by the static system and
transmitted to the indicator. The instru-
ment is capable of indicating vertical speeds
from 0 to + 12,000 feet per minute. An
over-pressure diaphragm and valve prevent
excessive rates of climb or descent from
damaging the instrument.
Clocks
Two elapsed time clocks are installed on
each instrument panel. The elapsed time
mechanism is started by pushing in on the
winding knob.
EMERGENCY EQUIPMENT
MASTER WARNING SYSTEM
An annunciator panel is located on the center
pedestal in each cockpit. Each panel con-
tains individual warning lights which indi-
cate malfunction or failures of equipment
and systems. Illumination of any individual
light also illuminates a red CAUTION light
on the upper portion of each instrument
panel. Once illuminated, the CAUTION
light can be extinguished (reset) by depress-
ing the light. The individual annunciator
panel light will remain illuminated. Another
malfunction will illuminate the CAUTION
light again. Warning lights are automati-
cally dimmed when the instrument panel
lights are on. The master warning system
does not include the fire warning and land-
ing gear unsafe lights. Power is furnished
by the essential dc bus.
NACELLE FIRE WARNING SYSTEM
A fire warning system is provided to detect
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 upper right side of each instru-
ment panel.
Nacelle FIRE Warning Lights
Left and right nacelle fire warning lights
are located on the upper right side of each
instrument panel. These lights illuminate
when nacelle temperatureat the turbine or
afterburner exceeds 1050�F + 50oF. They
are also illuminated for test�by depressing
the IND & LT TEST pushbutton switch. In-
dividual metal shields are provided which
can be pulled down over the lights to shade
them if necessary during illumination.
Power for the lights is furnished by the
No. 1 inverter.
STALL WARNING LIGHT
� A STALL WARNING light is located on the
annunciator panel in each cockpit and il-
luminates when the aircraft angle of attack
reaches +14 degrees and the nose landing
gear scissor switch is open. A steady tone
warning signal is also produced in the pilot's
earphone. Power for the stall warning light
is furnished by the essential dc bus.
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TA-12 SECTION I
LANDING GEAR SYSTEM
The tricycle-type landing gear and the main
wheelwell inboard doors are electrically
controlled and hydraulically actuated. The
main gear 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-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 lock-
ed. , Downlocks inside the actuating cylinders
hold the gear in place in the extended posi-
tion. L hydraulic pressure is on the gear
in the extended position as long as system
pressure is available. The landing gear
cylinders and doors are actuated in the cor-
rect order by two sequencing valves. Nor-
mal gear operation is by pressure from the
L hydraulic pump on the left engine. Should
pressure drop to 2200 psi during retraction,
the power source automatically changes to
the R hydraulic pump. R hydraulic pres-
sure will not, however, extend the gear in
the event of an L system failure; the manual
landing gear release must be used in that
case.
LANDING GEAR LEVER
A wheel-shaped landing gear lever is in-
stalled in the forward cockpit on the lower
left side of the instrument panel, just for-
ward of the throttle quadrant. The lever
has two labeled positions, UP and DOWN.
A locking mechanism is provided to prevent
the gear lever from being inadvertently
placed in the DOWN position. A pushbutton,
which extends upward from the top of the
lever, must be pressed forward in order to
release the lock mechanism. An override
button is installed just above the gear lever
to override the ground safety switch should
it become necessary 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 (forward cockpit) and the
GEAR NOT LOCKED light (aft cockpit) il-
luminate during cycling or when the gear is
in an unsafe condition. The aft cockpit has
a three-position guarded toggle switch lo-
cated on the lower left side of the instru-
ment panel. The switch is labeled UP and
DOWN. It is lock-wired in the off position
since it is to be used only for emergency
operation of the landing gear. The switch
will actuate the gear regardless of the posi-
tion of the landing gear lever in the forward
cockpit. Power for the circuit is furnished
by the essential dc bus.
Manual Landing Gear Release Handles
Handles, labeled GEAR RELEASE, for
lowering the gear when no L system hy-
draulic pressure is available are located
on the annunciator panels. When the gear
release handle is pulled, gear uplocks are
released in sequence and the gear falls and
locks down by force of gravity. The total
effective pull of the release cable attached
to the gear release handle is 9 inches, with
allowance for cable stretch and loosening
in the system the cable may be withdrawn
as much as 12 inches. Pulling the handle
out approximately 3 inches releases the
nose gear uplocks; continuing the pull for
the remaining 6 inches normal travel re-
leases the four main gear uplocks in the
sequence of right door first, aft locks be-
fore forward locks. If R hydraulic pres-
sure is available, the landing gear lever
must be put in the DOWN position before
pulling the gear release handle, or the
landing gear CONT circuit breaker must
be pulled; otherwise, �R system pressure
will retract the gear. After manual gear
extension, the gear may be retracted nor-
mally if L or R pressure becomes available.
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TA-12
LANDING GEAR SYSTEM
MANUAL LANDING GEAR
RELEASE HANDLE
(BOTH COCKPITS)
CROSSOVER VALVE
(PRESSURE)
ctttttttl
NOSE LANDING GEAR A D
ACTUATING CYLINDER
UL L
MAIN LANDING GEAR
ACTUAT. CYLINDER
II
LANDING GEAR
LEVER (FORWARD
COCKPIT)
LANDING GEAR SWITCH
UP (AFT COCKPIT)
DOWN
CROSSOVER VALVE
(RETURN)
a
ON
PRESSURE
SWITCH
-rr
1323
DOOR
SELECTOR
VALVE
0
07:17X
�CMLINCI
MA N LANDING GEAR
ACTUAT
CYLINDER
10[0151
001:13
aMit.
DOOR ACTUATING
CYLINDER (4 PLACES)
DOOR LATCH CYLINDER
.(4 PLACES)
UL
CABLE
ELECTRICAL CONNECTION
CHECK VALVE
RESTRICTOR VALVE Small arrow
indicates direction of restricted flow)
FLOW REGULATOR
RESTRICTOR VALVE. Restricted
flow in both directions)
MOM
COMO
R SYSTEM PRESSURE
R SYSTEM RETURN
L SYSTEM PRESSURE
L SYSTEM RETURN
Imo MLG DOORS CLOSED
COXEM0XC
MLG DOORS OPEN
ramming LANDING GEAR DOWN
rairaNzEzi LANDING GEAR UP
F201-45 (a)
1-58
Figure 1-24
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TA-12
SECTION I
Landing Gear Position Lights
Three green lights, located in the left side of
each instrument panel, indicate the down-
and-locked condition of the landing gear.
The location of each light corresponds to
the respective wheel it monitors. Power is
from the essential dc bus.
Landing Gear Warning Light and Audible
Warning
The warning light in the forward cockpit
gear handle illuminates red. When illum-
inated, it indicates at least one of the fol-
lowing conditions:
1. Gear is cycling.
2. Gear system is unsafe, though pro-
grammed UP or DOWN.
3. Gear is UP and power settings are be-
low minimum cruise.
An audible warning signal is produced in
the pilots headsets when the throttles are
retarded to less than minimum cruise set-
ting, the landing gear is not in the down-
and-locked position, and aircraft altitude
is below 10,000 (+500) feet. Power for the
light and audible warning circuit is furnished
by the essential dc bus.
Landing Gear Warning Cutout Button
The aural gear warning circuit may be dis-
armed by pressing the GR SIG REL push-
button switch which is located on the left
side of each instrument panel. The circuit
is rearmed when the throttles are advanced
to more than the minimum cruise setting.
Power is supplied from the essential dc bus.
Landing 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
streamers direct attention to their removal
before flight. Spare safety pins are pro-
vided in a box in the aft cockpit.
NOSEWHEEL STEERING SYSTEM
The nosewheel steering system provides
power steering for directional control when
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 pedals. �
L hydraulic system pressure from the nose
landing gear down line is routed to the steer-
ing system through a shutoff valve, con-
trolled by a nosewheel steering button on
each control-stick grip. Depressing the
nosewheel steering (NWS) button engages
nosewheel steering whenever the nosewheel
and rudder pedals are aligned. A holding
relay circuit maintains nosewheel steering
until the NWS button is depressed a second
time, when nose steering will be disengaged.
Steering is engaged at any time the NWS
button is held in and the nosewheel angle
and pedal position are matched. Nosewheel
steering radius is approximately 75 feet. A
mechanically 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 avail-
able and weight of the aircraft is
on any one gear. If the L system
pressure should drop below 1250
psi alternate nosewheel steering
may be obtained by placing the
brake switch to ALT STEER &
BRAKE position.
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SECTION I
TA-12
WARNING
The landing gear side load strength
is critical. Side loads during
takeoff, landing and ground oper-
ation must be kept to a minimum.
WHEEL BRAKE SYSTEM
The aircraft is equipped with artificial-feel
hydraulically operated power brakes. De-
pressing the rudder pedals actuates the 4
rotor brakes on each of the six main wheels.
The L hydraulic system furnishes brake
pressure with optional anti-skid operation.
The hydraulic pressure to the brakes is ap-
proximately 1200 psi. Should the L hy-
draulic system fail, alternate brakes are
available. The alternate brakes operate
from an independent system using R hy-
draulic pressure with no anti-skid provision.
A small accumulator is incorporated in the
normal brake system which should provide
up to five brake applications provided ac-
cumulator pressure has not been dumped by
selecting alternate brakes or the left hy-
draulic system has not been depleted by
actuation of anti-skid, leakage, or other
hydraulic malfunctions. Normal or anti-
skid 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 Switches
A three-position brake switch is located on
the left side of each instrument panel. When
in the NORM (center) position, brake pres-
sure from the L hydraulic system is avail-
able, but the anti-skid system is not oper-
ative. When in the ANTI-SKID (up) position,
the anti-skid system is operative whenever
the weight of the aircraft is on any one gear.
When in the ALT STEER & BRAKE (down)
position, the brakes, nosewheel steering
and air refueling system are powered by
the R hydraulic system if left system pres-
sure is below 1250 psi. When the aft cock-
pit switch is placed in the ANTI-SKID or
ALT STEER & BRAKE position, it is ca-
pable of overriding the forward cockpit
switch. Power for the circuit is furnished
by the essential dc bus.
WARNING
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 Lights
Illumination of the ANTI-SKID OUT indi-
cator light on each annunciator panel in-
dicates that the anti-skid system is inop-
erative. When the aircraft is on the ground,
the lights will be illuminated when either
cockpit switch is in the NORM or ALT
STEER & BRAKE position. The lights will
be off when either switch is in the ANTI-
SKID position, and the anti-skid control
box and wheel generators are operative. If
the fail-safe circuit within the anti-skid
control box is tripped the lights will illum-
inate and only power brakes will be avail-
able. The lights are off at all times when
the weight of the aircraft is off the gear.
DRAG CHUTE SYSTEM
The drag chute system is provided to re-
duce landing roll and aborted takeoff roll-
out distance. A ribbon-type parachute is
packed in a deployment bag and stowed in a
compartment in the upper aft end of the
fuselage. The chute rides free in the com-
partment and is snapped onto the airplane
in the initial stage of deployment. The
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TA-12 SECTION I
BRAKE SYSTEM
Lp
N2
FILLER
(VALVE)
OVERBOARD
DRAIN
BRAKE
RESERVOIRS
BRAKE
PEDAL
NORM ALT
NORM
MASTER
CYLINDERS
N2 PRESSURE
NORMAL
BRAKE
RELAY
VALVE
4111113
ICU
ALT
BRAKE
RELAY
VALVE
BRAKE
PEDAL
ALT
IIIMIN1111
L p
ANTI-SKID
SHUT OFF
VALVE
RELIEF VALVE
BRAKE DAMPER
III
L R
BRAKE BRAKE
RESTR I CTOR RESTR I CTOR
BRAKE SHUTTLE
VALVES
ANTI-SKID
GENERATORS
ANTI-SKID
SHUTOFF
VALVE
NITROGEN
CYLINDER
ALTERNATE
BRAKE
SHUTOFF
VALVE
RELIEF VALVE
ANTI
� SKID
NORMAL
a
ANTI-SKID
CONTROL
BOX
GASEOUS NITROGEN
�Eimin" L SYSTEM PRESSURE
msolgu L SYSTEM RETURN
mow MASTER CYLINDER SUPPLY
4.
ALT STEER
AND BRAKE
= BRAKE RELAY VALVE PRESS.
R SYSTEM PRESSURE (VALVE ENERGIZED)
moomml R SYSTEM RETURN
ELECTRICAL CONNECTION
F201-28(b)
Figure 1-25
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SECTION I
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�TA-12
chute mechanism incorporates a shear sec-
tion in the attachment yoke which ruptures
if the chute is deployed above the limit air-
speed. �Chute deployment is actuated elec-
trically from the forward cockpit by pulling
out and pushing in a drag chute handle, and
from the aft cockpit by operating a toggle
switch. The aft cockpit drag chute switch
has the capability of overriding the forward
cockpit drag chute handle. System power is
furnished by the essential dc bus.
Drag Chute Handle
A drag chute handle, labeled DRAG CHUTE,
is located on the upper left glare shield of
the forward cockpit. The handle is nor-
mally in the stowed (off) position, with the
handle horizontal. Pulling the handle out
to the limit of its travel activates a micro-
switch to deploy the drag chute. Rotating
the handle 90 degrees counterclockwise and
pushing in to the stop activates other micro-
switches to jettison the chute. Ground crew
personnel reset the handle to the neutral
position after flight.
Drag Chute Switch
A three-position drag chute toggle switch is
located on the upper left side of the aft cock-
pit instrument panel. The labeled switch
positions are CHUTE DEPLOY (up), off
(center), and JETT (down). The switch
functions are identical to those of the drag
chute handle.
AIR-CONDITIONING AND PRESSURIZATION
SYSTEM
Similar left and right air-conditioning
and pressurization systems utilize high
pressure ninth-stage compressor air from
each engine to pressurize and cool the cock-
pits and equipment compartments. System
shutoff valves allow compressor air to flow
when the engines are running and the sys-
tem switches are ON. Cooling is accom-
plished by ducting the bleed air through a
ram-air heat exchanger, primary and sec-
ondary fuel-air heat exchangers, and an
air-cycle refrigerator. Temperature of
the air supplied by each system is mod-
ulated by the positions of temperature con-
trol bypass valves located upstream from
the air-cycle refrigerators. The bypass
valves are positioned by control switches
located in the cockpits.
The left engine normally furnishes air for
the forward cockpit, ventilated flying suits,
inverters, and INS platform. The right
engine normally furnishes air for the aft
cockpit. A crossover system is provided
for emergency operation to supply right
engine system air to the forward cockpit
and equipment normally supplied by the left
engine system. High pressure canopy seal
and windshield defog air is furnished from
both right and left engine systems by ducts
connected downstream from the primary
fuel-air heat exchanger.
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
forward cockpit and 28,000 feet in the aft
cockpit.
TYPICAL COCKPIT PRESSURIZATION
SCHEDULE
Aircraft Alt
10,000 ft
20,000 ft
30,000 ft
35,000 ft & Up
Cockpit Alt
8,000 ft
16,000 ft
24,000 ft
26,000 ft
A crossover duct allows the pilot who has
control of the air-conditioning system to
divert aft cockpit air to the forward cockpit
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TA-12
SECTION III
AIR CONDITIONING
ON
-------
FORWARD COCKPIT
SYSTEM SWITCH
(BOTH COCKPITS)
E-BAY GROUND COOLING
(INEFFECTIVE UNLESS AFT
CANOPY IS CLOSED)
CROSSOVER
CHECK
VALVE
CROSSOVER VALVE (N.0.)
I. N. S. PLATFORM
COOLING CHECK VALVE
PRESSURE TEST
C_
SUIT
SUIT
P00 PER
CHECK VALVE
TO DISTRIB.
DUCTING
FORWARD
COCKPIT
SUIT FLOW
VALVE
0 DEFOG
� SWITCH
�� t INC
�)HOLD
OFF
DEFOG AIR
SUIT =1:011
P00 PER
TO DISTR.
CHECK VALVE
R.H. SYSTEM
IDENTICAL TO
HERE
GROUND
CONNECT.
AFT
COCKPIT
�
�
�
DEFOG
SWITCH
00)
INC
HOLD
OFF
�
�
CHECK
VALVE
INVERTER
COOLING
CANOPY
SEAL PRESSURE
FUEL-AIR HEAT
EXCHANGER
CHECK VALVE
I. N. S.
COOLING
PRESSURE
REGULATOR
L H. SYSTEM
SHOWN
COMPR.
F201-2.1(a)
ELECTRICAL (DAD CENTER
(L H. CHEEK)
I DEFOG AIR
PNEUMATIC
BYPASS
OVERBOARD FROM
FORWARD END
NLG BAY
DEFOG AIR
TEMP. CONTROL
BY PASS VALVES
1-1 FUEL-AIR
HEAT
EXCHANGER
RAM AIR
SYSTEM SHUTOFF VALVE
RAM AIR
HEAT
EXCHR
AUX. FUEL
PUMP
#3 #4
LH. ENGINE BLEED PORTS
Figure 1-26
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SECTION I TA - 12
in case of malfunction of the forward cockpit
system. The actuation of the crossover
system will not depressurize the aft cockpit
since the forward cockpit air exhausts into
the aft cockpit; however, a rise in temper-
ature will occur in the aft cockpit.
Forward Cockpit System Switches
The forward cockpit system three position
switches are installed on the upper left side
of each instrument panel. In the ON (left)
position the normally open system shutoff
valve is de-energized and the left system is
operative when the left engine is running.
In the OFF (center) position the shutoff valve
is energized closed, shutting off the air. In
the CROSSOVER (right) position, left system
air is shutoff and the normally open cross-
over valve closes, forcing right engine air
to the forward cockpit when the right engine
system is operating. The circuit is powered
from the dc essential bus.
Aft Cockpit System Switches
The aft cockpit system two position switches
are located on the upper left side of each
instrument panel. In the SYS ON (up) posi-
tion the right engine system's normally open
shutoff valve is de-energized so that right
engine air can flow to the aft cockpit. If
the forward cockpit system switch is in
CROSSOVER, this air will all be ducted to
the forward cockpit and will enter the aft
cockpit through the forward cockpit pressure
regulator valve. In the OFF position the
shutoff valve is energized and aft cockpit
system air is shutoff. The circuit is
powered from the essential dc bus.
Temperature Control Selector Switches
Two selector switches, one for each cock-
pit air installed on the upper left instrument
panels. 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 manual control positions. The
switches will normally be in the,AUTO posi-
tion; however, in case of a malfunction in
the automatic operation of the system, the
pilot can manually override the automatic
feature by moving the switch to either the
momentary COLD or WARM position. The
No. 1 inverter powers the cockpit temper-
ature control system.
Temperature Indicators and Monitor Switches
A temperature indicator and monitor switch
located on each upper left instrument panel
allows the pilots to monitor individual cock-
pit temperature conditions. The switches
are labeled FWD CKPT (left) and AFT CKPT
(right). Each pilot c,an monitor either for-
ward or aft cockpit air discharge temper-
ature by placing his switch in the desired
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 gage is provided
to allow anticipation of a tem-
perature' condition that might
eventually become too hot for
comfort. If the cocNit temper-
ature approaches 140 F, the suit
will not keep the pilot comfortable.
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TA-12
SECTION I
Temperature Control Knobs
Two temperature control rheostats, one
for each cockpit, are installed on the upper
left instrument panels. Arrows on the panel
adjacent to the knobs show the direction of
rotation necessary to increase temperature.
Generally, it is necessary to periodically
rotate the respective temperature control
rheostat toward the COLD (counterclockwise)
position to maintain a comfortable temper-
ature in the ventilated flying suits and keep
the temperature of the cockpits within tol-
erance. Electrical power for the cockpit
temperature control circuits is from the
No. 1 inverter.
Cabin Altimeters
A forward and aft cockpit pressure altitude
gage is installed on each left forward panel
and indicates either forward or aft cockpit
altitude as selected by the cabin altimeter
selector lever.
Cabin Altitude Selector Switches
A switch, labeled FWD CKPT in the up posi-
tion and AFT CKPT in the down position, is
installed on each left forward panel. Op-
erating the switch selects the respective
cockpit pressure altitude on the cabin alti-
tude gage.
Cockpit Depressurization (Dump) Switches
A guarded, two-position cockpit depressur-
ization switch, labeled PRESS DUMP, is
installed on the left side of each instrument
panel. Either pilot may depressurize (dump)
or repressurize both cockpits, but must first
obtain control of both cockpit air-conditioning
systems by use of the control transfer panel
(refer to Control Transfer Panels, Air-
Conditioning Switches and Transfer Lights,
this section). When control of the air-con-
ditioning is obtained, actuation of the PRESS
DUMP switch to the up position (guard up)
will depressurize both cockpits. When the
PRESS DUMP switch is moved to the down
position (guard down), the cockpits will re-
pressurize.
OXYGEN SYSTEM AND PERSONAL
EQUIPMENT
AIRCRAFT OXYGEN SYSTEM
The aircraft is equipped with two indepen-
dent, high-pressure, gaseous oxygen sys-
tems. Both systems supply each pilot, and
oxygen is consumed from the two systems
simultaneously. If one system fails, the
other system will continue to supply both
pilots, but with reduced duration. Each
system is supplied by one 875-cubic-inch,
1800-psi oxygen bottle. Both bottles are
located in the nosewheel well and are ser-
viced at the bottom of the right-hand chine.
As oxygen leaves the bottles the pressure is
reduced to 75 psi. ON-OFF levers for the
two systems are located on the oxygen con-
trol panels installed on the left consoles. A
dual system low pressure gage installed be-
tween the levers will read approximately 75
psi during normal operation. The needles
on the gage will fluctuate, indicating oxygen
flow when the pilot inhales. Oxygen quantity
is displayed on the dual indicating high pres-
sure gages located on the left side of the in-
strument panels just forward of the throttle
quadrants. The NO. 1 OXY LOW or the
NO. 2 OXY LOW lights on the annunciator
panels will illuminate when the respective
oxygen supply pressure decreases below
400 psi, or when the regulated pressure
drops to 58 + 3 psi.
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0
OVERBOARD DISCHARGE
VALVE
El�����������
(OXYGEN CYLINDER
SYS I
1800 AT 70�F
HIGH PRESSURE INDICATOR
(AFT COCKPIT)
PRESSURE SWITCH
(400 PSI OR LESS)
1800 PSI SYS 1
FILLER VALVE
�
�
�
II XIX
c u
ammiss�mi SYS 1
num. ON
BALANCE VALVE �
-n
cri 0 m
�
; PRESSURE NioNNE��91 . OFF OXYGEN
.0
REDUCER � � - �
I �
� N
�III � MI � PRESSURE SWITCH
P%) RELIEF VALVE
(120-140 PSI) � � s �
i LI �a � � (58 PSI OR LESS)
NI
� � �
�
� �Immium6 �
� 111-
.........1 .11-1;1111 .
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� s � � IN
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� � � � � I
� �
1 I I I II
� �
� I � I
OVERBOARD DISCHARGE � �� � � � � DISCONNECTS �
ii � �
s � � �
m � � SEAT OXYGEN �
a � I
VALVE � � � m �
� N � �
� � � � a
D�s����somm NEN � a � m a � UUI
� � � � � �
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. VENT LINE� � � � is VENT LINE
� . ii
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� � 75 PSI SYS 1 � in � � �
m � NI ia������������������� � � � �
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is a � � s
� � � !ii�mil�������siiiligoil�ial a
N IN � � �
� . 75 PSI SYS 1 � �
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� 11�����smausam�ma 111 ls������������������������� � I � ����������������� �
ismos�s������-���������^�����������������������414.11 61111."1.21111.111.11111.11111.1
U s
� 75 PSI SYS 2 a
1800 PSI SYS 2
LOW PRESSURE
WARNING LIGHTS
(AFT COCKPIT)
NO 1 OXYGEN LOW
NO 2 OXYGEN LOW
OXYGEN CONTROL PANEL
(AFT COCKPIT)
.L
� ME 101. MEN
1.
HIGH PRESSURE INDICATOR
(FORWARD COCKPIT)
LOW PRESSURE
WARNING LIGHTS
(FORWARD COCKPIT)
NO 1 OXYGEN LOW
NO 2 OXYGEN LOW
OXYGEN CYLINDER
SYS 2
1800 PSI AT 70�F
OXYGEN CONTROL PANEL
(FWD COCKPIT)
SYS 1 FF ^ OXYGEN OFF
SYS 21
ON ON
O
X I � �
� U. �
�
� � I
SEAT VENT
DISCONNECT
HIGH PRESSURE LINES (1800 PSI)
����I LOW PRESSURE LINES (75 PSI)
����..- ELECTRICAL LINES
VENT LINES
W31SAS N30AX0
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TA-12 SECTION I
EMERGENCY OXYGEN SYSTEM
Two independent emergency oxygen systems
are installed in the pilot's parachute pack.
Each system consists of three 20 cubic inch,
2100 psi cylinders attached to a common
manifold. These systems will supply oxygen
simultaneously during bailout or if the air-
craft oxygen system fails. An oxygen line
from each system is routed around both
sides of the pilot's waist and connects to the
suit controller valve. Check valves prevent
emergency flow when the aircraft systems
are supplying oxygen. When the emergency
system is actuated, check valves prevent
oxygen flow into the aircraft system. Emer-
gency oxygen flow pressure is slightly lower
than aircraft system pressure. Oxygen
duration of each emergency system is ap-
proximately 15 minutes.
Emergency Oxygen System Actuation
The emergency oxygen system may be ac-
tuated either manually by pulling the con-
vential green apple, or automatically by
the upward motion of the seat during ejec-
tion. The emergency oxygen system should
be actuated if the aircraft is not delivering
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 con-
ditions in the cockpit. The suit consists of
four layers, ventilation garment, bladder,
link net, and heat-reflective outer garment.
The ventilation garment layer allows ven-
tilation air to circulate between pilot's
underwear and the bladder layer. The
bladder provides an air-tight seal to hold
pressurized air in the suit. The link net is
a mesh which holds suit configuration in
conformance with the pilot's body. The
outer layer of heat-reflecting aluminized
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 provided by the
cockpit air-conditioning system. Temper-
ature of the ventilation air cannot be varied
except by changing cockpit air temperatures.
Ventilation airflow rate may be regulated 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, controlled by the pilot op-
erating the suit ventilation boost valve lever
which changes the air pressure of the in-
coming suit air. The aft cockpit has no
control, depending only on the valve setting
in the forward cockpit.
Suit Ventilation Boost Valve Lever
The suit ventilation boost valve lever, la-
beled SUIT VENTIL BOOST, is located in
the forward cockpit only, on the left console.
The lever positions are marked NORMAL
(aft) and EMERG (forward). Operating the
lever positions a butterfly valve in the cock-
pit air-conditioning air supply line in such
a way as to vary the pressure of the air
available to the suit system. Increased
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 suffi-
cient airflow to the suit. At IDLE engine
rpm the ventilation boost valve lever must
be kept at 2/3 of the way from NORMAL to
EMERG in order to provide sufficient air
for cooling the (INS platform, inverters and
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SECTION I
- TA-12
conditioning the suit when it is used. (When
the pressure suit is not worn the suit air
hose should be capped.) During takeoff and
normal flight the valve lever is kept in the
NORMAL position. If the pilot suffers dis-
comfort, such as might happen with a
gradual climb to an extreme altitude or dur-
ing low-rpm descents, the valve lever is
gradually moved toward the EMERG position
until a comfortable pressure and ventilation
condition is attained. The valve lever should
not be moved toward EMERG more than
necessary to provide pilot comfort; exces-
sive suit system pressure will unduly re-
duce 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 to keep internal suit pressure
at 3.5 psia (equivalent to pressure at 35,000
ft) in the event of cockpit depressurization.
A press-to-test button for each oxygen sys-
tem is installed on the controller valve,
which allows the pilot to check suit inflation.
Faceplate Heat Switches
Faceplate heat switches are installed on the
right console in each cockpit. Each switch
has four positions; OFF, LOW, MED, and
HIGH. Heat may be regulated to defog the
faceplate as required. Defogging is accom-
plished by the combination of faceplate heat
and oxygen flow.
HELMET
The helmet head area is divided into two
separate sections by a rubberized cloth face
seal. The front area between the faceplate
and the face seal receives oxygen from
either the aircraft or emergency oxygen
system through regulators built into the
helmet. Oxygen flows across the faceplate
from the inhalation valves inside the helmet
and accomplishes some faceplate defogging
before it is inhaled. The rear area re-
ceives ventilation 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 ventilation air from leaking for-
ward. An external crank on the helmet is
provided for adjusting the head band. But-
tons on each side of the helmet, when ac-
tuated, will lower the faceplate and visor.
The faceplate is opened by moving the but-
tons and dumping the pressure, allowing
the faceplate to be rotated upward. If the
aircraft or emergency oxygen supply to the
helmet is interrupted or exhausted, the re-
gulators in the helmet sense the drop in
pressure and the faceplate seal deflates,
allowing ambient air to enter the helmet so
the pilot will not suffocate.
GLOVES
Leather gloves attach to the suit at the
wrist rings. The inner liner of the glove
is similar to the suit inner liner and will
retain pressure. There is little or no
ventilation for the hands.
BOOTS
The sock or boot liner attaches to the suit
at the ankle by means of a zipper. The
boots are made of white leather to take ad-
vantage of heat reflection, and fit snugly
over the socks. A spur is attached to each
boot.
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TA-12 SECTION I
OXYGEN MASK AND REGULATOR
When permitted by appropriate regulations,
a substitute oxygen mask assembly may be
used in place of the pressure suit. The as-
sembly consists of a specially designed
A-13 oxygen mask, oxygen regulator, anti-
suffocation valve, and two oxygen personal
leads with connectors for both aircraft and
emergency oxygen systems. In the event
�the regulator should malfunction or the
oxygen supply be exhausted, the anti-suf-
focation valve, installed between the regu-
lator and the mask, will sense the drop in
oxygen pressure and allow ambient air to
enter the mask.
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-rear provides access to
the survival items stored inside. The kit
contains a two-way radio, smoke generator,
mirror, whistle, knife, matches, water,
food, first-aid kit, moccasins, and a com-
pass, all packed in a waterproof bag at-
tached to a 20-foot retention lanyard. If an
overwater flight is anticipated, a liferaft
may be stowed on top of the plastic bag and
attched to the lanyard. During ejection the
liferaft 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 liferaft. The
survival gear and liferaft remain attached
to the parachute harness by the retention
lanyard. During rapid abandonment of the
aircraft on the ground, the survival kit re-.
lease handle may be used to free the pilot
from the survival kit (including the lanyard)
without inflating the liferaft.
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
16,000 (+ 400) feet altitude, the drogue
chute is deployed first in order to stabilize
free fall of the pilot. The drogue is auto-
matically jettisoned at 15,000 (+ 400) feet
after an aneroid controlled opener deploys
the main chute. Below 15,000 feet the main
chute only deploys immediately. A manual
"D" ring 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 pilot's
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 manually actuate the
timer when bailout is made without using
the ejection seat.
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 fluo-
ride. A collapsible vision splitter is also
installed on the windshield center line to
minimize reflections.
DEFOG SYSTEM
The defog system delivers hot air from both
right and left air systems through check
valves to defog the windshields and canopies.
A plastic V-shaped air duct runs along the
lower edge of each windshield through which
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SECTION I TA-12
hot defog air is supplied when selected by a
defog switch that is located in both cockpits.
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 3-position defog switch is located on the
right console in each cockpit. When held in
the momentary DEFOG INCREASE (forward)
position the motor driven defog valve will
open. Time of travel to full open is approx-
imately 7-13 seconds. In the HOLD (center)
position the valve will stop at any partial
open position; in the OFF position the valve
will completely close. The circuit is pow-
ered by the essential dc bus.
WINDSHIELD RAIN REMOVAL SYSTEM
A rain removal system is provided for
clearing the forward windshield when op-
erating the aircraft in rain. It has a tank
that is pressurized by air and the tank is
connected to spray tubes located on each
side of the windshield center divider. A
pushbutton switch, located on the glare-
shield panel, is used to spray the rain re-
moval fluid onto the windshield. Power is
furnished from the essential dc bus.
CANOPIES
Each canopy consists of two high-tempera-
ture-resistant glass windows secured in a
reinforced titanium frame and hinged at the
aft end by two hinge pins. Operation of the
canopy is completely manual. Small holes
in each side of the canopy are provided as
lifting points from the outisde. No handles
are provided on the inside of the canopy. for
moving it up or down. A prop asembly
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. An air boost counter-
balancing system is provided to aid in the
manual opening and closing of the canopy.
Individual internal latching handles are in-
stalled below each right canopy sill, allow-
ing each canopy to be latched separately
from the inside. External fittings located
on the left side of the aircraft can be used
to operate the latches from the outside.
CAUTION
The canopy should be opened or
closed only when the aircraft is
stationary. Maximum taxi speed
with canopy open is 40 knots.
Gust or severe wind conditions
should be considered as a portion
of the 40-knot-limit taxi speed.
CANOPY SEAL
An inflatable rubber seal is installed in the
edge of each canopy frame. The seal seats
against the mating surfaces of the canopy
sill and windshield and provides sealing for
retaining cockpit pressurization.
Canopy Seal Pressurization Lever
A canopy seal pressurization lever, labeled
CANOPY SEAL PRESSURE, is located in
each cockpit above the forward right console.
The lever positions are ON (forward) and
OFF (aft). Moving the lever to the ON posi-
tion controls a valve to supply pressure to
the canopy seal.
CANOPY CONTROLS AND INDICATORS
Canopy Latch Handles
Canopy latch handles are located under the
right sill in each cockpit and rotate forward
to lock. Each sill trim is cut out to expose
the action of the locking lugs and pins as the
handle is rotated forward. A cam over-center
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SECTION I
CANOPIES AND CONTROLS
DETAIL A
1 CANOPY INTERNAL JETTISON HANDLES
2 CANOPY LATCH HOOKS
3 CANOPY LATCH HANDLES
4 CANOPY LATCH ROLLER BRACKETS
5 CANOPY LIFTING HOLES,
DETAIL B
6 CANOPY PROP ASSEMBLY AND UPLOCKS
7 CANOPY EXTERNAL LATCH CONTROLS
8 CANOPY EXTERNAL JETTISON HANDLE HIDDEN
9 CANOPY PROP (GROUND HANDLING)
966-5
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TA-12
action ensures that the handle will remain
only in the latched or unlatched position.
There are no canopy unsafe warning lights
installed in the aircraft.
Canopy External Latch Controls
Flush-mounted external latch fittings are
located on the left side of the aircraft, per-
mitting the canopies to be opened from the
outside. These controls accept a 1/2-inch
square bar extension. Once a canopy is un-
locked, it may be raised manually until the
prop locks it in the open position.
Canopy Internal Jettison Handles
A canopy jettison T-handle is located on the
left console wall adjacent to the left leg of
each pilot. The handle can be used to in-
dividually jettison the canopies without ini-
tiating the seat ejection system. Each
handle is held in the stowed position by a
safetywire and a ground safety pin. Cable
travel is approximately 6 inches.
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 both canopies simultaneously for
emergency entrance. Actuation of the jetti-
son handle jettisons the forward canopy im-
mediately and the aft canopy after a 1-sec-
ond delay. Sufficient cable length is pro-
vided to allow the operator to stand clear
of the fuselage during the jettison procedure.
Canopy Jettison Sequence
The canopy jettison system is designed to
unlatch and jettison each canopy individually
from the aircraft. Each system consists of
two initiators, which are independently ac-
tuated by either the ejection seat D-ring or
the canopy jettison handle; a canopy un-
latch 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 will fire
the unlatch thruster which unlocks the can-
opy. This thruster then activates the can-
opy removal thruster which jettisons the
canopy. During canopy jettisoning 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. Pulling the D-
ring jettisons the canopy as the initial step
in the ejection sequence.
EJECTION SEATS
Individual ejection seat systems utilize a
rocket-catapult, upward-ejection seat ca-
pable of safely ejecting the crewmember at
ground elevation provided that a level flight
path speed of 65 KIAS or greater is achieved
before ejecting. Each seat incorporates an
ejection ring, headrest, knee guards, auto-
matic foot retractors, automatic foot re-
tention separation devices, a pilot-seat sep-
aration device, a shoulder harness, an in-
ertia reel lock assembly, and an automatic-
opening seat belt. A speed sensor mounted
on the fuselage behind each seat automati-
cally selects one of two seat separation de-
lays, depending upon airspeed at ejection.
(See Seat Ejection System, this section).
Quick-disconnect fittings installed on the
seat rails and the floor of the aircraft per-
mit disconnecting the oxygen, ventilated suit,
and electrical lines.
Seat Vertical Adjustment Switches
The seats may be adjusted vertically by
means of an electric actuator mounted on
the lower end of the catapult. The switch
is located on the right side of the seat
bucket. Power for seat adjustment is fur-
nished by the essential dc bus.
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SECTION I
EJECTION SEATS (Both Cockpits)
8
2089
1 MANUAL CABLE CUTTER RING
2 HEADREST
3 SHOULDER HARNESS
4 LAP BELT
5 SHOULDER HARNESS INERTIA REEL LOCK LEVER
6 KNEE GUARDS
7 SEAT ADJUSTMENT SWITCH
8 EJECTION RING
9 EJECTION SEAT T HANDLE
10 FOOT RETRACTOR FITTINGS
10
F201-22(b)
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Shoulder Harness Inertia Reel Lock Levers
A shoulder harness inertia reel lock lever
installed on the left side of each 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 move-
ment. The reel also incorporates a locking
mechanism which will lock the shoulder
harness when a 2G to 3G force has been ex-
erted in a forward direction. When the reel
is locked in this manner, it will remain
locked until the lock lever is moved to the
LOCK position and then returned to the UN-
LOCK position.
Ejection (D) Rings
An ejection (D) ring, located on the front of
each seat bucket, is the primary control for
ejection. An ejection safety pin is installed
in the ejection ring housing bracket.
Ejection Seat 1-Handle
A secondary seat-ejection system is incor-
porated in each cockpit. The operating T-
handle for this ejection system is unlocked
and made accessible when the ejection D-
ring is pulled. If the seat fails to eject,
pulling the T-handle causes a separate ini-
tiator to fire the seat catapult and a 2-sec-
ond delay seat-separation and belt-opening
initiator.
WARNING
There is no safety interlock to
prevent actuating the secondary
seat ejection system with the
canopy in place.
Foot Spurs
Foot spurs (attached to the pilot's shoes)
are attached to each 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 the stowed position, the cables
to the foot spurs are reeled in, and the
pilot's feet are retracted into the foot rests
as part of the ejection sequence. The foot
cables are subsequently automatically se-
vered by a set of cutters during the ejection
sequence.
Manual Cable Cutter Rings
Each 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
cutter systems fail or rapid abandonment
of the aircraft is required on the ground.
PILOT-SEAT SEPARATION SYSTEM
Each ejection seat is provided with a pilot-
seat separation which operates in conjunc-
tion with the automatic seat belt release
system. A windup reel is mounted behind
the headrest, and a single nylon web is
routed from the reel halfway down the for-
ward 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, causes the pilot to be sep-
arated from the seat with a slingshot action.
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TA-12 SECTION I
AUTOMATIC SEAT BELTS
Each 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 normal ejection
sequence and requires no additional effort
on the part of the pilot.
SEAT BELT-PARACHUTE ATTACHMENTS
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 ap-
proximately 15,000 feet.
EJECTION SEQUENCE
Pulling the D-ring is the only action required
to initiate pilot ejection and results in firing
both the canopy jettison and ejection seat
systems.
Note
The ejection seat cannot fire until
the canopy jettison system has
fired. This design safety feature
Is necessary to prevent pilot ejec-
tion through the metal canopy.
All ejection actions occur automatically and
in a specific sequence. The D-ring cable
fires the ejection sequence initiator, ac-
tuating the canopy jettison system and the
leg-guard thruster. The leg-guard thruster
rotates the leg guards, retracts the pilot's
feet, activates the 2-seond delay 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 seat catapult initiator and
arms the speed sensor. (The 0.3-second
delay ensures that the canopy has separated
completely prior to seat ejection.) Gas
pressure from the catapult initiator fires
the seat rocket-catapult, the 4-second seat
separation delay initiator, and enters the
speed sensor.
Note
If airspeed is less than 265 KIAS,
the gas pressure passes through
the speed sensor and fires the 0.6-
second delay seat separation ini-
tiator. If airspeed is more than
300 KIAS, the pressure is blocked
by the speed sensor. Between
265 to 300 KIAS, seat separation
time will be 0.6 or 4 seconds,
depending on the tolerance of the
speed sensing unit.
Initial seat movement upward on the rails
disconnects normal oxygen, ventilated suit,
and electrical lines, and activates the emer-
gency oxygen supply. Either the 0.6-second
delay initiator or the 4-second delay ini-
tiator actuates the cable cutters, opens the
seat belt, and fires the seat separation sys-
tem.
Note
The 2-second delay cable cutter
backup initiator will actuate the
foot cable cutters and cut the
cables if they were not cut as a
result of the 0.6-second delay
initiator firing. The firing of the
4-second delay initiator will
again actuate the foot cable cutters,
cutting the cables if they were not
cut as a result of the 0.6-second
initiator or 2-second backup ini-
tiator firing.
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A static line, attached to the seat belt, is
pulled as the pilot separates from the seat
and activates the automatic parachute se-
quence.
EGRESS COORDINATION SYSTEM
An egress coordination system is installed
in the aircraft to supplement normal inter-
phone communication. With this system the
aircraft commander always has the capabil-
ity to issue and check compliance with a
bailout signal, regardless of which cockpit
he may be occupying. Power for the system
is furnished by the essential dc bus. See
Emergency Escape, Section III for additional
information.
Egress Lights and Switches
The forward cockpit lower right instrument
panel contains a guarded toggle switch, la-
beled BAILOUT (up), and two lights which
read BAILOUT (red) and AFT SEAT
EJECTED (amber) when illuminated. The
aft cockpit lower instrument panel contains
a guarded switch, labeled BAILOUT (up) and
a light which reads BAILOUT (red) when il-
luminated. Actuation of a BAILOUT switch
illuminates the BAILOUT light in the op-
posite cockpit. The AFT SEAT EJECTED
light is wired directly to a switch on the aft
cockpit ejection seat tracks and will illum-
inate whenever the aft seat is ejected.
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TA-12
SECTION II
NORMAL PROCEDURES
TABLE OF CONTENTS
Page
Page
Preparation For Flight
2-2
Cruise
2-16
Preflight Check
2-2
Descent
2-17
Starting Engines
2-9
Air Refueling Procedure
2-17
Before Taxiing
2-10
Before Landing
2-21
Taxiing
2-11
Landing
2-22
Before Takeoff
2-13
Go-Around
2-25
Takeoff
2-13
After Landing
2-25
After Takeoff Climb
2-16
Engine Shutdown
2-25
Climb
2-16
Abbreviated Checklist
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PREPARATION FOR FLIGHT
FLIGHT RESTRICTIONS
Refer to Section V for Operating Restric-
tions and Limitations.
FLIGHT PLANNING
Refer to the Appendix.
TAKEOFF AND LANDING DATA CARDS
Refer to Appendix for information neces-
sary to fill out Takeoff and Landing Data
Cards before each flight.
WEIGHT AND BALANCE
Refer to Section V for Weight and Balance
limitations. For detailed loading infor-
mation, refer to Handbook of Weight and
Balance Data. Before each flight, check AFT COCKPIT CHECK (Solo Flights Only)
takeoff and anticipated landing gross weights,
and Weight and Balance Clearance, Form 1. Lap belt shoulder harness and all
365F. personal leads - Secured.
PREFLIGHT CHECK
ENTRANCE
Ladder platform stands which overhang
the chine are used to gain entrance to the
cockpits. The canopies are unlatched ex-
ternally by rotating each external canopy
control clockwise with an L-shaped, 1/2-
inch-square bar. The canopies are manu-
ally raised to the fully open latched position.
AIRCRAFT STATUS
Refer to Form 781 for engineering, servic-
ing, and equipment status.
EXTERIOR INSPECTION
It is not practical for the pilot to perform
an exterior inspection while wearing a
pressure suit; therefore, the exterior in-
spection should be accomplished by other
qualified personnel.
BEFORE ENTERING COCKPIT
The following checks apply to both cockpits:
1. Manual cable cutter ring - Secured.
2. Ejection seat and canopy safety pins
installed - Check.
2. All circuit breakers - In.
LEFT CONSOLE
1. Emergency fuel control switches -
NORM.
2. Control transfer panel - Check.
3. UHF command radio - TR + G.
4. Oxygen supply lever - OFF.
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SECTION II
INSTRUMENT AND ANNUNCIATOR PANEL
1. Cockpit temperature rheostats -
Mid-range.
2. Cockpit temperature switches - AUTO.
3. Aft cockpit air system switch - ON.
4. Forward cockpit air system switch -
ON.
5. Landing and taxi lights switch - OFF.
6. Afterburner switches - OFF.
7. Brake switches - NORM.
8. Landing gear switch - OFF.
9. Pressure dump switch - NORM.
10. Drag chute handle - Neutral.
11. Forward bypass control - FWD CKPT.
12. Pitot heat - OFF.
13. Trim Power - ON.
14. Hydraulic reserve oil - OFF.
15. BUPD switch - OFF (guard down).
16. Pitch logic override switch - OFF
(guard down).
17. Yaw logic override switch - OFF'
(guard down).
18. Gear release handle - Stowed.
19. Air refuel switch - OFF.
20. Fuel dump switch - OFF.
21. Fuel transfer switch - OFF.
22. Emergency fuel shutoff - OFF (guard
down).
23. Fuel quantity selector switch -
TOTAL.
24. Inverter switches - OFF.
25. Generator switches - Neutral.
26. Battery switch - OFF.
RIGHT CONSOLE
1. Canopy seal pressure - ON.
2. SAS switches - All NORM.
3. SAS override control transfer
switch - FWD.
4. TACAN - ON and tuned.
5. ADF - As desired.
6. Faceplate heat switch - OFF.
7. Floodlight rheostat - OFF.
8. Instrument and Panel lights - OFF.
9. Defog switch - OFF.
10. Beacon'lights switch - OFF.
11. Flight recorder - OFF.
INTERIOR CHECK (Dual Flights)
For dual flights all items marked with an
asterisk must also be checked in the aft
cockpit.
*1. Throttles - OFF.
*2. Landing gear lever - DOWN.
*3. All circuit breakers - In.
*4. Foot retractors - Attach.
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CAUTION
Foot spurs must be attached or
removed while the cables are
fully retracted to prevent damage
to the cables.
*5. Accomplish personal equipment hook-
up. See Figure 2-1.
*6. Battery switch - Check.
a. Forward cockpit - EXT PWR.
b. Aft cockpit - OFF.
LEFT CONSOLE
*1. Emergency fuel control switches -
NORM.
2. IFF/SIF switches - STDBY (Proper
Mode and Code).
*3. Control transfer panel - As desired.
4. Suit ventilation boost valve lever -
Set at 2/3 of lever travel from NORM
to EMER.
*5. UHF command radio - T/R + G.
6. Radar beacon switch - ON.
*7. No. 1 and No. 2 oxygen systems - ON
(when using pressure suit). Check
system pressures.
8. Throttle friction lever - As desired.
*9. Aft bypass switch - CLOSED (mop).
INSTRUMENT AND ANNUNCIATOR PANEL
*1. Cockpit temperature rheostat - As
desired.
*2. Cockpit temperature indicator
switch - Check
a. Forward cockpit - FWD CKPT.
b. Aft cockpit - AFT CKPT.
*3. Cockpit temperature switch - AUTO.
*4. Aft cockpit air system switch - ON.
*5. Forward cockpit air system switch -
ON,
6. Periscope MIRROR SELECT handle -
Fully forward (projector).
*7. Landing and taxi lights switch - OFF.
*8. Afterburner switches - OFF.
Landing gear lights - Check green.
*10. Brake switch - Set.
a. Forward cockpit - ANTI SKID
b. Aft cockpit - NORM.
*11. Oxygen quantity indicator - Check.
*12. Cabin altimeter switch - Set.
a. Forward cockpit - FWD.
b. Aft cockpit - AFT.
*13. Pressure dump switch - NORM.
*14. Drag chute handle - Stowed.
*15. CIT gage - Pointers together and
indicating ambient temperature.
*16. TDI - Check for proper indication.
*17. Altimeter - Set.
*18. Clocks - Check.
*19. Forward bypass door indicator -
Check.
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SECTION II
PERSONAL EQUIPMENT HOOKUP
(Shirt sleeve flight)
'CAUTION
FOOT SPURS MUST BE
ATTACHED AND REMOVED
FROM SEAT BALL FITTING,
BY HAND. WHEN REMOVING
SPURS, THE BALL FMING
MUST BE GUIDED BY HAND
TO FULL RETRACTION
REASON - STAMPING TO
ENGAGE, AND KING
TO RELEASE BALL FITTING,
WILL DAMAGE THE RETURN
CABLE.
0 HOOKUP CHUTE
A -CHEST HOOK
B -2 LEG STRAPS;
(1 EACH LEG)
COHOOKUP SPURS
PUSH DOWN
TO LOCK
TIGHTEN STRAPS
A -CHEST STRAP
B -2 LEG STRAPS
C -2 SIDE STRAPS
0 PLUG IN OXYGEN HOSES (2 HOSES)
SHOULDER STRAPS
0 HOOKUP BELT AND SHOULDER
STRAPS AND PARACHUTE
LANYARD AND TIGHTEN
acir
0 HOOKUP HELMET ELECTRICAL
HELMET
MASK
ATTACHMENT
OXYGEN PANEL
()TURN ON OXYGEN AND
HOOKUP MASK
PARACHUTE LANYARD
PUSH DOWN
TO LOCK
F201-72(1)
Figure 2-1 (Sheet 1 of 3)
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SECTION II
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IA-IL
PERSONAL EQUIPMENT HOOKUP
PULL TO
ADJUST
0 LAP BELT
SECURE SHOULDER HARNESS STRAPS AND
PARACHUTE TIMER ARMING KEY. LOCK
BELT AND ADJUST
0
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 -rn HANDLE.
INSURE HANDLE IS SNAPPED SECURE
INTO HOUSING
Figure 2-1 (Sheet 2 of 3)
CHECK (TWO) PARACHUTE CANOPY ROCKET
JET RELEASES. INSURE ROLL BAR PIN
IS IN DOWN (LOCKED) POSITION. PULL ON
EACH RELEASE TO INSURE LOCK POSITION
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. 71 ONE AT A TIME. CHECK PRESSURE,
APPROXIMATELY 75 PSI AND FLUCTUATING
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
F201-72(2)
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TA-12
SECTION II
PERSONAL EQUIPMENT HOOKUP
0 HOOK UP SPURS
FOOT 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
DI S CONNECTED
OCOMMUNICATIONS (FACE HEAT AND RADIO)
CONNECT HELMET CHORD TO PARACHUTE
EXTENSION CHORD
OTURN FACE HEAT ON LOW (CONTROL ON
RIGHT HAND CONSOLE)
1 1
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 75 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
a PULL TO
'AD JUST
0 ADJUST KIT SEAT STRAPS; RIGHT AND LEFT SIDE
�
CONNECT EMERGENCY OXYGEN HOSES,
SLIDE KNURLED FITTING INTO PLACE,
INSERT SAFETY CLIP, PULL ON HOSE SLIGHTY
TO ASSURE OF LOCKED POSITION
NOTE
LEFT HOSE OVER HELMET HOLD DOWN STRAP
F201-72(3)
Figure 2-1 (Sheet 3 of 3)
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SECTION II
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TA-12
*20. Spike Controls - AUTO (Inopv.)
*21. Forward bypass controls - Check.
a. Forward cockpit - CLOSED.
b. Aft cockpit - FWD CKPT.
*22. Display MODE SELECT switch -
As desired.
*23. BEARING SELECT switch - TACAN.
*24. ATT/AP Select switch - As desired.
*25. Surface limiter handle - Pull out.
*26. Pitot heat switch - OFF.
*27. Trim power switch - ON.
*28. Hydraulic reserve oil switch - OFF.
*29. BUPD switch - OFF (guard down).
*30. Pitch logic override switch - OFF
(guard down).
*31. Yaw logic override switch - OFF
(guard down).
*32. Gear release handle - Stowed.
*33. Air refuel switch - OFF.
*34. Fuel dump switch - OFF (guard down).
*35. Fuel transfer switch - OFF (guard
down).
*36. Emergency fuel shutoff switches -
Fuel On (guard down).
37. Nitrogen indicator - Check.
RIGHT CONSOLE
1. Pitot-static source lever - NORMAL.
*2. Canopy seal pressure lever - OFF.
*3. SAS switches - Check
a. Forward cockpit channel engage
switches - ALL OFF.
b. Aft cockpit channel engage
switches - ALL NORM.
c. Aft cockpit SAS override panel
transfer switch - FWD.
4. Autopilot switches - OFF.
5. INS panel - As required.
*6. Defog switch - OFF.
*7. Faceplate heat - As desired.
*8. TACAN - ON and tuned.
*9. FRS TAKE CMD button - As required.
*10. FRS function selector switch - MAG.
*11. Floodlight switch - As desired.
*12. Instrument and panel lights - As
desired.
*13. Beacon lights switch - OFF.
*14. Flight recorder switch - OFF.
ELECTRICAL FUNCTION CHECK
1. Inverter switches - Check.
a. No. 1 and No. 3 switches
NORM.
b. No. 2 switch - ON.
*2. IND & LT TEST switch - Press.
a. N2 quantity indicators should
decrease to zero. Nitrogen
quantity warning lights on an-
nuciator panel should come on
at 1-liter point.
b. Fuel tank boost pump lights
should illuminate.
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SECTION II
c. All warning and FIRE lights should
illuminate.
d. Landing gear unsafe warning
should be heard.
3. Crossfeed and manual boost pump
switches - Press (check lights ON).
4. Pump release switch - PUMP REL,
then release.
5. Tank boost pumps - Check 1, 2, and
6 tank lights ON (automatic sequencing).
6. Cross'feed 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. UHF and IFF/SIF - Check.
9. IFF/SIF - As required.
STARTING ENGINES
Either engine may be started first. Nor-
mally, the left engine is the first to be
started.
CAUTION
Before starting an engine, verify
that wheels are firmly chocked.
There is no parking brake and
brakes are inoperable until hy-
draulic pressure is available.
Determine that intake and ex-
haust areas are clear of person-
nel and ground equipment. Ground
personnel using interphone equip-
ment will be in position to observe
exhaust nozzle and nacelle in-
spection panels during start.
El
Do not move either control
stick until at least 1500 psi
hydraulic pressure can be
maintained on the A or B
system gages; otherwise, a
control system inspection
will be necessary.
Check with INS crew prior to start-
ing engines.
2. Fuel low pressure lights - OFF.
3. Ground air supply - Request ON.
4. Engine start switch - GND at 15 per-
cent rpm.
5. Throttle - IDLE above 16percent rpm.
6. Fuel flow - Check.
7. Engine start switch - Release when
EGT rises.
8. EGT - Check (400�C maximum).
9. Ground air - Request OFF at 40 per-
cent rpm.
10. Idle rpm - Adjust to 60-64percent rpm.
11. Engine and hydraulic pressure instru-
ments - Check normal.
a. Fuel flow - Check.
b. EGT - Check.
c. Oil pressure indicator - Check.
CAUTION
Discontinue start if oil pressure
rise is not observed within 60
seconds from start of rotation.
d. Hydraulic system pressures
Check.
12. Other engine - START, using same
procedure.
INS mission only.
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CAUTION
If throttle is inadvertently re-
tarded to OFF during the start
do not advance in an attempt to
restart engine. In case of false
start use Clearing Engine pro-
cedure, this section. After-
burner duct must be visually
checked and unburned fuel re-
moved prior to attempting
another start.
. Never apply ground air supply
to engine starter when engine
rpm exceeds 40 percent rpm.
CLEARING ENGINE
When a false start occurs, trapped fuel
and fuel vapor may be removed from engine
by using following procedure:
1. Throttle - OFF.
2. Start switch - OFF (release).
3. Continue cranking engine or request
ground crew to apply starter air
supply at 40 percent rpm or below.
4. Ground air. - Signal OFF after 15
seconds.
CAUTION
Never reapply air to starter when
engine rpm exceeds 40 percent
rpm.
BEFORE TAXIING
1. Emergency fuel system - Check.
a. Throttle - IDLE.
b. Emergency fuel control switch -
EMER.
c. Tachometer - Check for stable
engine operation. RPM may
increase, decrease, or show
no change.
d. Emergency fuel control switch -
NORMAL.
2. Generator switches - RESET above
60 percent rpm after check with INS
crew.
3. Battery switch - BAT (within 3
seconds).
4. Generator-out lights - Check out.
Note
If the generator-out warning
lights fail to extinguish, return
the battery switch to the EXT
PWR position and repeat steps
2 and 3.
[5j INS DEST/FIX switch - VARIABLE
DEST.
.1 INS MODE switch - NAY. (Check
F-7.1
r9
with INS crew prior to actuating
switch.) Depress STORE pushbutton
and check HSI heading marker for
10-deg right and DTG for 122 nrni.
INS indications - Report destination
coordinates, distance-to-go, and
ground speed when slewing is com-
pleted.
INS DEST/FIX switch - VARIABLE
FIX and STORE. Check INS FIX
REJECT light on.
INS DEST/FIX switch - VARIABLE
DEST and STORE. Check INS FIX
REJECT light off.
Li10. INS umbilical cord - Disconnected
(confirmed by INS crew).
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11. External power - Signal for dis-
connect.
12. Inlet air forward bypass - Ground
crew confirm open.
13. SAS engage switches - ALL ON.
14. SAS recycle lights - Press (lights
should go out).
15. SAS light test switch - Press.
(All lights should illuminate.)
16. Autopilot pitch and roll engage
switches - ON,
17. Autopilot disengage switch (control
stick) - Press. Check that auto-
pilot disengages.
18. SAS engage switches - OFF. Pitch
and yaw A and B lights illuminate,
both MON lights must stay out.
19. Surface trim - Check operation and
set to zero. Confirm that direction
of movement corresponds with indi-
cation.
20. Control system - Check for correct
direction of movement. Individually
check each axis in both directions
and request ground personnel to
verify proper deflection of control
surfaces.
21. Seat and canopy safety pins - Re-
moved and stowed.
*22. Canopy - Close and lock.
*23. Canopy seal pressure lever - ON.
CAUTION
The canopy should be opened
or closed only when the air-
craft is completely stopped.
Maximum taxi speed with
canopy open is approximately
40 knots. Gusts or severe
wind conditions should be
considered as a portion of the
40-knot limit taxi speed.
24. Rear view periscope - Check.
25. Taxi clearance - Obtain from
control tower.
26. Chocks and downlock pins - Signal
for removal. Observe crewchief
for clearance to taxi.
27. Nosewheel steering - Engage.
Check operation of nosewheel
steering.
TAXIING
1. Brakes - Check.
WARNING
Do not select alternate brakes
with both L and R hydraulic
systems operative.
*2. Flight instruments - Check.
(Check turn-and-slip indicator for
turn needle deflection in the direc-
tion of turn while taxiing and ball
free in race).
*3. Navigation equipment - Check oper-
ation. Check ADF, TACAN and INS.
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TURNING DIAGRAM
/30� MAX
96.2 FT
76.7 FT
96.3 FT
56 0 FT
75.2 FT
CENTER OF TURN
NOTE: 151.9 FT MINIMUM RUNWAY WIDTH REQUIRED FOR 180-DEGREE TURN
(MAIN GEAR WHEELS ON EDGE OF RUNWAY AT START OF TURN).
F201-61(a)
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Figure 2-2
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TA-12
SECTION II
CAUTION
All taxiing and turns should be
accomplished at slow speeds 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 instruments - Check at
Military thrust (run up one engine
at a time).
2. SAS channel engage switches - All
ON.
3. SAS recycle lights - Press, if neces-
sary (lights should go out).
4. Surface trim indicators - Check for
zero setting.
5. Tanks 1, 2, and 6 - Check ON.
INS - Check and fix as required.
At designated runway position, se-
lect correct STORED FIX position
and fix. Check INS FIX REJECT
light off. Reset DEST/FIX briefed
initial destination position, and
�store. Check distance to go after
slewing completed, then reset
DEST/FIX to STORED AUTO if
desired.
*7. Compasses - Check and synchronize
FRS and check INS if applicable and
return display mode select switch to
desired position. Check standby
compass against runway heading.
8. Pitot heat - ON,
*9. All warning lights - OUT.
*10. Shoulder harness - LOCK.
11. Beacon light switch - ON (if re-
quired).
12. Flight controls - Cycle and check
hydraulic pressure.
13. Suit ventilation boost valve lever -
NORMAL just before taking runway
and applying power.
TAKEOFF
1. Brakes - HOLD.
2. Elapsed time clock - Start.
3. Nosewheel steering - Engaged.
4. Throttles - Advance.
5. Brakes - Release at 80 percent rpm.
CAUTION
The tires may skid if the brakes
are held at high thrust.
6. Engine instruments - Check at
Military thrust.
7. Afterburner switches - ON (simul-
taneously). Afterburners should
light within 5 seconds, indicated by
a noticeable increase in thrust.
Faulty afterburner operation can
be detected by EGT and fuel-flow
indications.
WARNING
Monitor nosewheel steering
because the afterburners may
not light simultaneously.
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TAKEOFF
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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
USE NOSEWHEEL STEERING AS NECESSARY
FOR DIRECTIONAL CONTROL
THROTTLES - ADVANCE TO TAKEOFF THRUST
ENGINE INSTRUMENTS - CHECK AT
MAXIMUM THRUST
AFTERBURNER SWITCHES - ON
THROTTLES - ADVANCE TO
MILITARY THRUST
BRAKES - RELEASE
THROTTLES ADVANCE TO 80 PERCENT
NOSEWHEEL STEERING - ENGAGE
BRAKES - HOLD
F201-3(a)
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SECTION II
8. Engine instruments - Recheck at
Maximum 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
indication of improper engine
performance during power ad-
vancement, 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. Refer to per-
formance data, Appendix I, for
takeoff information. Directional
control can be maintained with
nosewheel steering up to nose-
wheel liftoff speed.
9. Acceleration - Check. Check indi-
cated airspeed against computed
acceleration check speed at selec-
ted acceleration check distance.
Refer to performance data, Ap-
pendix I, for takeoff information.
10. Rotation - Begin at computed air-
speed approximately three seconds
before reaching takeoff speed. Ap-
ply smooth, constant back pressure
so that required control deflection
and rotation to takeoff attitude oc-
curs at takeoff speed. Refer to
Appendix I for rotation and takeoff
speeds.
Note
Use indicated airspeed during
takeoff and climb until proper
climb schedule speed is indi-
cated 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 nosewheel steering is no
longer available. Rudder pressure must
be held to counteract the cross wind
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
problems 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.
CAUTION
Avoid abrupt rotation since
this can impose an excessive
load on the tires and cause
blowouts.
In general, the tires are more critical
during takeoff than landing because of the
higher groundspeeds and gross weights
involved. Wing lift quickly relieves the
gear load as the nose is raised. Start
rotation at computed rotation speed, approxi-
mately 3 seconds before reaching the
scheduled takeoff airspeed. Premature
nosewheel lift-off should be avoided because
the unnecessary drag will extend the ground
run. Delayed rotation also extends the
ground run and can result in excessive tire
speeds.
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TA-12
AFTER-TAKEOFF CLIMB
When definitely airborne:
Landing gear lever - UP. Accel-
erate to correct climb speed.
Note
The landing gear will retract in
approximately 12 seconds. Ob-
serve landing gear limit speed
while gear is extended.
WARNING
Single engine operation is criti-
cal immediately after takeoff.
Increasing airspeed and de-
creasing angle of attack has
greater benefits than gaining
altitude at a maximum rate.
2. Engine instruments - Check.
3. Surface limiter handle - Push in at
Mach 0.5.
*4. Altimeters - Set to 29.92 in. Hg
prior to reaching FL 180.
5. IFF/SIF - As briefed.
CLIMB
The aircraft accelerates rapidly to climb
speed. A definite rotation and pitch change
is required to establish the climb. With
maximum afterburning, the climb angle is
about 20 degrees at low altitude. Begin the
rotation sufficiently in advance of reaching
climb speed to avoid exceeding the recom-
mended airspeed schedule. The recom-
mended climb speed schedule with after-
burning is 350 KEAS to approximately
26,000 feet and a constant Mach 0.9 above
26,000 feet. The recommended Military
(non-afterburning) climb speed is a constant
300 KEAS. Refer to Appendix I for climb
performance.
CRUISE
Observe limitations of Section V.
Center-of-gravity control is important
for optimum cruise performance. Fuel
load distribution and the automatic tank
sequencing provide a forward center-of-
gravity for takeoff and climb. During
cruise, the automatic sequencing provides
an aft center of gravity to minimize elevon
deflection and resulting trim drag. Supple-
mental manual control of fuel usage is also
possible, but should only be used if auto-
matic fuel tank sequencing malfunctions.
SUPERSONIC ACCELERATION
Maximum afterburning is required for
all supersonic accelerations. Variations
in outside air temperature have a pro-
nounced effect on supersonic acceleration
capabilities. Advantage of cold tempera-
tures should be taken to obtain the best
acceleration. One of three procedures is
recommended, depending on temperature.
Climbing Acceleration
A 400-KEAS climbing acceleration is not
recommended if the temperature is warmer
than 5o above standard at altitudes above
20,000 feet. When this procedure is used,
accelerate to 370 KEAS and rotate to climb
attitude. Rotation should not be delayed
because it is possible to inadvertently accel-
erate to 425 KEAS or faster. Establish
400 KEAS and climb at this speed. Mach
number will increase with altitude. The
rate of climb and acceleration will be slow
between 25,000 and 30,000 feet; approxi-
mately Mach 1.0 to 1.1.
Note
The bypasses should be opened
manually at Mach 1.35. If the
doors are not opened, duct "buzz"
will occur at approximately Mach
1.4.
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SECTION II
Diving Acceleration
A diving acceleration from 40,000 feet is
recommended when the temperature is warm-
er than standard at low altitudes and cold
above 35,000 to 40,000 feet. Starting from
approximately Mach 1.0 and 40,000 feet,
make a shallow dive to Mach 1.2 at 35,000
feet. The acceleration will be slow between
Mach 1.1 and 1.2. Pull out and start climb-
ing from approximately 35,000 feet at a
constant 400 knots EAS.
Note
Do not make an abrupt pull-
out because this will increase
load factor and bleed off Mach
number.
Manually open the bypass doors at Mach
1.35.
Level Acceleration
Level accelerations can be made at 40,000
feet if desired. Accelerate at constant
altitude to 400 knots EAS and then climb at
this speed. The acceleration will be notice-
ably slower in the range from Mach 1.1 to
1.2. Manually open the bypass doors at
Mach 1.35.
DESCENT
Descent performance charts are shown in
Appendix I. The descent fuel consumption
should be minimized to obtain maximum
flight duration with the J-75 engines.
The inlet air bypass doors can be opened
to act as thrust spoilers and increase the
rate of descent.
1. Defog switch - As desired.
2. Landing gear - As desired below
speed limit.
Note
The landing gear may be low-
ered to increase the rate of
descent.
3. IFF/SIF - As briefed.
4. Altimeter - Set to station pressure
when passing through flight level
180.
Note
Use pitot-static instruments for
airspeed and altitude data during
all operations below flight level
180 except in climbs.
5. Bypass doors - OPEN (if desired).
AIR REFUELING PROCEDURES
Either of two methods of handling power
during refueling may be used. Whenever
the initial fuel quantity remaining is over
approximately 15,000 pounds it is possible
to use minimum afterburning on one engine
and less than Military thrust on the other.
This allows refueling to be accomplished
at a constant altitude of approximately
32,000 feet, using the non-afterburning
engine for thrust control. Normally or
when at light weight, the initial contact
should be made using non-afterburning power
settings. One afterburner should then be
lighted after temporarily disconnecting when
the aircraft becomes power limited at
Military thrust. The conventional procedure
of completing refueling without use of an
afterburner can also be used; however, a
toboggan to approximately 25,000 feet will
be necessary as the tanks are filled.
Prior to air refueling, stabilize and trim
at refueling speed for contact. Observe the
tanker for director light signals and
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SECTION II TA-12
maneuver as directed by the lights. A suc-
cessful connection is confirmed by a mild
jolt to the aircraft, steady illumination of
the director light panel and the extinguishing
of the READY light. Slight maneuvering
may be necessary at this point to illumin-
ate the azimuth and elevation neutral lights
during fuel transfer. Contact can be main-
tained between the aircraft and tanker
during a turn or in a descent. No adverse
flight characteristics are present due to
tanker downwash. After the disconnect
occurs, separation is made down and to the
rear of the tanker.
PRIOR TO REFUELING
Accomplish the following prior to refueling:
1. Air refuel switch - READY.
Note
Amplifier requires up to approxi-
mately five minutes for warmup.
2. Forward transfer switch - FWD
TRANS. (2000 lbs to tank 1).
3. Fuel quantity indicator selector -
TOTAL. Monitor total fuel quantity.
4. Seat - Lower.
NORMAL REFUELING
Normal refueling is accomplished as follows:
1. Establish contact.
After contact is made:
2. READY light - Check out.
3. Total fuel quantity - Monitor.
When refueling is complete:
4. Control stick disconnect - Press.
5. Air refuel switch - OFF.
(When probe is clear of receptacle.)
6. Tanks 1, 2, 6 - Check ON.
In case L hydraulic pressure is lost, R
pressure may be utilized for refueling by
moving the brake switch to ALT STEER &
BRAKE position.
CAUTION
Do not leave the brake switch in
the ALT STEER & BRAKE position
after refueling.
ALTERNATE REFUELING
When in observation position after rendez-
vous with tanker.
5. READY light - Push on (green) if
The boom may be latched in the refueling
receptacle manually as an alternate pro-
cedure by using the following procedure:
necessary.
1. Air refuel switch - MANUAL.
Check READY light on.
6.
Forward transfer switch - OFF.
2. Control stick disconnect - Press and
7.
Stabilize in pre-contact position.
hold.
8.
Beacon light switch - OFF.
When nozzle has bottomed in the receptacle:
9.
Observe tanker director lights illumi-
nated and boom in ready for contact
position.
3. Control stick disconnect - Release.
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TA-12
SECTION II
AIR REFUELING BOOM ENVELOPE LIMITS
UP ELEVATION LIMIT
20�
6 FEET EXTENDED
INNER LIMIT .
18 FEET EXTENDED
OUTER LIMIT
LEFT AZIMUTH
LI MIT
EXTENDED OUTER
LIMIT
Y
100
EXTENDED
INNER LIMIT
RI GHT AZIMUTH
LIMIT
DOWN ELEVATION LIMIT
F2O1-31
Figure 2-4
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SECTION II
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TA-12
AIR REFUELING DIRECTOR LIGHTS
.�NC�
1010.
lb 01'
I 1,
I
u.. 1-- �
Li- t I.: . ". 1 I
1...
LA.. .." ....! .
eCr os. Li
..... .
CI � cv ,,, L...
II
����i' ����1
�O�
,111
I
11
CENTERED
APPROACHING FORWARD LIMIT
210�
21.5�
- 23.50
24.50
26.0�
30.00 �
34.00
- 3550
- 37.00
3. 50
40.00
APPROACHING AFT LIMIT
COLOR CODE
/ RED
M- GREEN
P201 ..32()
2-20
Figure 2-5
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TA-12
SECTION II
CAUTION
If the disconnect trigger is re-
leased before the nozzle is in the
bottom of the receptacle, it is
possible for the nozzle to damage
nozzle latches, preventing any
further refueling.
4. Fuel quantity - Monitor TOTAL fuel.
When refueling is complete:
5. Control stick disconnect - Press.
CAUTION
The automatic limit disconnect
system is inoperative. All dis-
connects must be initiated by the
receiver aircraft, as the tanker
operator is unable to release the
nozzle latches during manual
boom latching.
6. Air refuel switch - OFF. READY
light out.
Note
If a malfunction occurs which
prevents disconnecting the boom,
place the Air Refuel switch in
the MANUAL position, depress
the IFR DISC trigger. If dis-
connect is not accomplished,
proceed with brute force pullout
by retarding throttles.
BEFORE LANDING
1. Fuel transfer switch - FWD TRANS
(if required).
Note
When tank 5 or 6 contains fuel,
transfer 1000 to 3000 pounds
forward to obtain slight nose-
up pitch trim.
2. Surface limiter handle - Pull out
(released) at Mach 0.5.
3. Periscope MIRROR SELECT handle -
Fully forward.
4. Hydraulic pressures - Check.
5. Fuel transfer switch - OFF.
*6. Brake switches - Set.
a. Forward cockpit - As required.
b. Aft cockpit - NORM.
*7, Shoulder harness - Manually locked.
*8. Faceplate - Open.
*9. Oxygen - OFF.
10. Traffic pattern entry - 275 to 350
KIAS, 1500 feet above field eleva-
tion.
11. Downwind - 250 KIAS, 1500 feet
above field elevation.
12. Landing gear lever - DOWN (check
gear warning lights).
Note
Normal gear extension time is
approximately 16 seconds. Ob-
serve gear limit speed with
gear extended.
13. Base leg .220 to 230 KIAS, descending.
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SECTION II
TA-12
14. Final approach - Maintain best final
approach speed, minimum 165
KIAS.
Note
165 KIAS is best final approach
speed with 5000 pounds fuel re-
maining. Increase approach
and landing speeds 1 knot for
each additional 1000 pounds of
fuel above 5000 pounds remaining.
. See Figure 2-3 for a typical landing
pattern.
15. Landing and taxi lights switch As required.
LANDING
NORMAL LANDING
Refer to the Appendix for landing ground
roll distances. If airspeed becomes ex-
cessively low, a high sink rate will develop,
resulting in a hard landing. During the
flare, throttles are moved to IDLE and
touchdown is made at approximately 10 de-
gree pitch angle (nose approximately on the
horizon).
1. Throttles - Retard to IDT.F in flare.
CAUTION
Allow throttle to follow quadrant
curvature so that the hidden
ledge at the IDLE position will
prevent inadvertent engine cut-
off.
2. Touchdown speed - As required.
3. Hold nosewheel in air.
CAUTION
Fuselage angle must not exceed
14 degrees to avoid scraping the
tail.
4. Drag chute handle - Pull out to
deploy. Chute deployment requires
approximately 3 seconds.
5. Lower nosewheel at 110 KIAS.
6. Engage nosewheel steering for di-
rectional control. Steering will not
engage until rudder pedals align
with no position (straight
ahead) and weight of aircraft is on
any one gear.
7. Brakes - Apply after chute deploys.
Moderate braking may be used prior
to chute deployment.
CAUTION
If the chute does not deploy, ob-
serve the brake energy limit speeds
in Section V.
8. Drag chute handle - Rotate counter-
clockwise and push in to jettison
chute.
Note
The drag chute should be jet-
tisoned at 55 KIAS unless the
crosswind component exceeds
12 knots.
CAUTION
If the chute is not jettisoned,
the elevons should not be moved
during taxiing as the shroud
lines may jam between the in-
board elevons and the fuselage
and cause structural damage.
AFT COCKPIT LANDING TECHNIQUE
Fly a normal traffic pattern. After rollout
onto final approach establish an approach
angle sufficiently steep to permit full view
of runway threshold.
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ZLI.O�2900 SILO/L10Z :aseaia Joi pancuddV
NOTE
NORMAL FINAL APPROACH SPEED IS 165 KIAS PLUS
ONE KNOT PER 1000 LB OVER 5000 LB FUEL REMAINING.
LANDING SPEED IS 20 KIAS BELOW FINAL APPROACH
SPEED. SPEED FOR MINIMUM LANDING ROLL IS
10 KNOTS LESS THAN FOR NORMAL PROCEDURE.
REDUCE AIRSPEED TO 250 KIAS.
LOWER LANDING GEAR AND CHECK
INDICATORS. MAINTAIN 1500 FEET
ABOVE FIELD ELEVATION
REDUCE AIRSPEED TO
230 KIAS.
Nttb
ENTER TRAFFIC PATTERN AT
AIRSPEED 275-350 KIAS
ALTITUDE 1500 FEET ABOVE
FIELD ELEVATION
ADJUST AIRSPEED AS REQUIRED
(165 KIAS MINIMUM)
MAINTAIN 275-350 KIAS
1500 FEET ABOVE Fl ELT)-
ELEVATION
LEVEL TURN AT 1500 FEET
ABOVE FIELD ELEVATION
4a NORMAL TOUCHDOWN AT 145 KIAS./i
RETARD THROTTLES TO IDLE.
DEPLOY DRAG CHUTE. ENGAGE
NOSEWHEEL STEERING AFTER
NOSEWHEEL IS ON THE GROUND
N11311Vd ON
ZLI.O�2900 SILO/L10Z :aseaia Joi pancuddV
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SECTION II TA-12
CROSSWIND LANDING
The traffic pattern for a crosswind landing
should be normal, making proper allow-
ances for velocity and direction of the cross-
wind. Proper runway alignment on final
approach can be maintained by crabbing or
dropping one wing; however, a combination
of the two is recommended just prior to
flare. Remove crab before touchdown,
using wing low technique to prevent side
drift. Reduce sink rate to a minimum to
accomplish smooth touchdown. At increas-
ed crosswind components, sink rate must
be minimized due to increase of side loads
imposed on the landing gear. In severe
crosswinds the nose should be lowered and
nosewheel steering engaged prior to drag
chute deployment.
LANDING ON SLIPPERY RUNWAYS
Wet Runways
Use normal technique. Landing roll will
increase due to reduction in available
braking force. Braking effectiveness is
increased if ANTI-SKID is off.
WARNING
Wet runway braking tests have
not been completed. Preliminary
tests indicate that the aircraft will
plane with heavy water conditions
on the runway. With this con-
dition, directional control in a
crosswind may be difficult.
Icy Runways
Landing on an icy runway is the same as
landing on a wet runway except that braking
effectiveness is further reduced.
CROSSWIND COMPONENT CHART
60
50
40
30
20
10
0
0 10 20
NOTE CROSSWIND COMPONENT-KNOTS
FOR CROSSWIND COMPONENT ENTER CHART WITH
MAXIMUM REPORTED GUST VELOCITY
Figure 2-7
MINIMUM ROLL LANDING
30
35
F201-75
a. Make touchdown near the approach
end of the runway at minimum air-
speed. This is essential for a suc-
cessful short field landing.
b. Deploy the drag chute as quickly as
possible after touchdown. Lower
the nosewheel while the chute is
deploying.
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SECTION II
c. Apply optimum braking immediate-
ly after chute deployment. Moderate
braking may be used prior to chute
deployment.
d. Move throttles to IDLE during flare
or immediately after touchdown.
e. Move right engine throttle to OFF
after touchdown.
Note
Retarding both throttles to OFF
further reduces thrust, but elim-
inates nosewheel steering and
braking. If the brakes are burn-
ed out at the end of the runway,
and speed will permit a safe turn
off, the nosewheel steering sys-
tem will "save" the landing.
The throttle technique is dependent on the
pilot judgement of the particular field con-
ditions.
WARNING
Engine shutdown will result in
loss of hydraulic actuating pres-
sure for the following systems:
a. Right engine shutdown -
Alternate brakes and nose-
wheel steering.
b. Left engine shutdown - Normal
and anti-skid brakes and
nosewheel steering.
GO-AROUND
A go-around may be initiated anytime during
the approach, or during landing roll when
sufficient runway remains for takeoff.
1. Drag chute - Jettison, if deployed.
2. Throttles - MILITARY thrust,
MAXIMUM thrust if required.
3. Landing gear lever - UP after posi-
tive climb angle established.
4. Trim - As necessary.
AFTER LANDING
1. Pitot heat - OFF.
2. SAS channel engage switches - OFF
(before taxiing).
*3. Lighting switches - As required.
4. Suit ventilation boost valve lever -
Set at 2/3 of lever travel from
NORMAL to EMERG.
ENGINE SHUTDOWN
1. Wheel chocks - Installed.
F
INS - As briefed.
CAUTION
INS must be off prior to open-
ing canopies to prevent possi-
bility of excessive temperatures
of INS components.
*3. Canopy seal pressure levers - OFF.
*4. Canopy - Open.
Note
In event of engine fire during
shutdown, the engine can be
motored with fuel OFF to blow
out fire if starter unit is con-
nected. Refer to Section III.
*5. Recorders - OFF.
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TA-12
GO AROUND
NOTE
The excess thrust available to perform a go-around varies
with airspeed, gross weight, airplane configuration, field
elevation and ambient temperature. As extremes of these
variables are approached, the ability to perform a successful
go-around with military thrust decreases, thus requiring
afterburning thrust. Refer to appendix for charts showing
variation in performance to be expected with changes in
these operating conditions.
NOTE
A MINIMUM OF 1000 LBS. OF FUEL IS
REQUIRED FOR A GO-AROUND WITH A
NORMAL PATTERN
� THROTTLES - MILITARY THRUST
(MAXIMUM THRUST IF NECESSARY)
� LANDING GEAR LEVER-UP (AFTER
DESCENT IS CHECKED)
� TRIM - AS NECESSARY
F201-64
2-26
Figure 2-8
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TA-].Z
SECTION II
*6. All appropriate electrical switches -
OFF.
7. All inverter switches - OFF.
8. Generator switches - TRIP (mo-
mentary).
9. Throttles - OFF.
10. Battery switch - OFF.
11. Seat and canopy safety pins -
Installed.
STRANGE FIELD PROCEDURES - AS BRIEFED.
ABBREVIATED CHECKLIST
Normal and emergency procedures ab-
breviated checklists are furnished separately.
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TA-12
SECTION III
EMERGENCY PROCEDURES
TABLE OF CONTENTS
Page
Page
Introduction
3-2
Emergency Entrance
3-19
Propulsion System Failure
3-2
Ditching
3-20
During Takeoff
Immediately After Takeoff
3-5
3-5
Fuel System Failure
3-20
Nozzle Failures
3-6
Fuel Dumping Procedure
3-21
Afterburner
Air Inlet
3-6
3-7
Electrical Power System
3-22
Compressor Stall
3-7
Hydraulic Power System
3-23
Engine Flameout
Air Start Procedures
3-7
3-8
Flight Control System
3-24
Glide Distance
3-8
SAS Emergency Operation
3-25
Engine Oil Pressure
Engine Fuel Control
3-8
3-11
Landing Gear System Emergency
Operation
3-31
Fire
3-11
Wheel Brake System Failure
3-32
Elimination of Smoke & Fumes
3-12
Air Data Computer
3-32
Emergency Escape
3-13
Pitot-Static System
3-33
Takeoff & Landing Emergencies
3-14
Air Conditioning and Pressurization
Failures
3-33
Abbreviated Checklist
3-34
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SECTION III TA-12
INTRODUCTION
This section includes recommended pro-
cedures to be used in the event of emergency
or abnormal operating conditions. The pro-
cedures ensure safety of the pilots and air-
craft under most situations until a safe land-
ing can be made or other appropriate action
taken. Multiple emergencies, adverse
weather, and other peculiar conditions may
require modification of procedures; there-
fore, it is essential that the pilot determine
the correct course of action by using com-
mon sense and sound judgement. Critical
steps of procedures are presented in cap-
ital letters. These steps should be me-
morized so that they may be performed im-
mediately without reference to checklists.
PROPULSION SYSTEM FAILURE
By definition, propulsion system failure may
be total or confined to the components con-
sidered as part of the propulsion system,
including the main engine, afterburner, in-
let, nozzle, tailpipe, fuel controls lubri-
cation and ignition systems.
Complete Engine Failure
Complete engine failure is mechanical
failure within the engine characterized by
extreme vibration, seizure, or explosion.
Other symptoms may be a sharp drop in
thrust, rpm, or EGT. Engine should be
shutdown immediately after positively de-
termining which engine has completely
failed. If complete engine failure occurs,
it probably will not permit normal wind-
milling operation. Shut off the fuel. An
airstart should not be attempted since this
can result in fire or explosion. Land as
soon as possible.
Engine Mechanical Failure
Engine mechanical failure is an engine or
engine accessory failure which requires
pre-cautionary shutdown to avoid or delay
complete engine failure. Mechanical failure
situations include uncontrollable oil tem-
perature, EGT, or RPM and abnormal oil
pressure, fuel flow, or vibration. Normal
windmilling speeds can be expected. A
landing should be made as soon as possible,
after precautionary shutdown.
Engine Flameout
Engine flameout is characterized by a loss
of thrust and a drop in EGT and rpm.
Flameout can result from interruption of
fuel supply, component malfunction, or
compressor stall. Immediate airstarts
may be possible provided the attempt is
made before compressor rpm has appre-
ciably decreased; the higher the rpm, the
quicker and more consistent will be the
airstart. In the event an engine flames out
at high Mach number, an immediate air
start can be attempted after the flamed-out
engine is positively identified. It may be
difficult to immediately determine the
flamed out engine. The pilot should cross-
check the turn-and-slip indicator, EGT,
fuel flow and the rpm, as an aid in positively
determining the failed engine. Probability
of a successful airstart is greater if at
least 7 psi CIP is attained. Airstarts have
been made while in roughness but restor-
ation of stable inlet conditions increases
the probability of success. While it is ex-
peditious to use crossfeed to assure ade-
quate fuel supply during an air start attempt,
crossfeed should not be left on after the
start is obtained. Turn an additional tank
on to the side where flow interruption is
suspected before crossfeed is discontinued.
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TA-12 SECTION III
Afterburner Flameout
Afterburner flameout may occur as the re-
sult of engine stall. Stall conditions must
be corrected before attempting afterburner
relight. Reducing altitude, increasing air-
speed, or increasing engine thrust will im-
prove the afterburner lighting character-
istics, especially at high altitude. If these
measures fail, the afterburner igniter re-
cycling period should be varied by allowing
shorter or longer intervals between attempts
to obtain an afterburner light. A compari-
son of duct pressures may reveal that a
difference in inlet efficiency was respon-
sible for the flameout. Afterburner opera-
tion cannot be sustained at levels of turbine
discharge pressure below 10 inches of mer-
cury absolute; however, this corresponds to
altitudes of 59,000 feet at 0.8 Mach number
and 70,000 feet at 1.8 Mach number.
Compressor Stall
Compressor stall is usually indicated by
compressor pulsations and afterburner
flameout may be expected. Other indica-
tions are loss of thrust, rapid reduction or
fluctuation of rpm, and failure of rpm to
increase during acceleration. Compressor
stall may be caused by abrupt or erratic
throttle movement, failure of the nozzle to
open as soon as the afterburner starts to
operate, or unstable inlet conditions.
Single Engine Flight Characteristics
The aircraft design is such that no flight
system is dependent on a specific engine;
thus, the loss of an engine will not result
in subsequent loss of all hydraulic or elec-
trical systems.
The engines are located outboard on the
wing, away from the direct influence of the
fuselage air flow to obtain optimum
performance from the inlet ducts during
normal operation. If an engine fails at low
speed just after takeoff, the large amount
of asymmetric thrust may require full rud-
der and a mild bank toward the good engine
for control. Minimum single engine di-
rectional control speeds are shown on the
chart in this section. A chart showing max-
imum weights for single engine climbout is
included in the performance data appendix.
Acceleration to climb speed and climb to
pattern attitude can be accomplished with
maximum thrust on the operating engine.
Full rudder trim can be used to assist in
control. Pitch trim changes while dumping
fuel can be expected due to shifting center
of gravity as the tanks empty. Directional
trim is quite sensitive to changes in air-
speed and power settings during pattern op-
eration.
At high speed, engine failure or flameout
could cause large amounts of yaw at high
rates. The yaw SAS has a large degree of
authority to prevent this. Temporary thrust
reduction on the good engine (minimum
afterburning) helps to counteract the asym-
metric thrust condition, and followup rudder
action is necessary. If large yaw angles
develop, inlet duct airflow disturbances
may cause the other engine to stall or flame-
out.
In the event of engine flameout at high Mach
number, an immediate airstart can be at-
temped after the flamed out engine is posi-
tively identified. This may be difficult.
The pilot should endeavor to cross check the
turn and slip indicator, EGT, fuel flow, and
rpm in order to positively determine the
failed engine. If encountered, intensity of
inlet roughness increases with Mach number.
If a start is unsuccessful, or if engine
failure has occurred, a descent to inter-
mediate altitudes will be necessary. The
bypass doors should be open on the wind-
milling engine. Descent range can be ex-
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TA-12
MINIMUM SINGLE ENGINE CONTROL SPEEDS
5
4
3
1
4
3
2
1
0
120
140 160 180
INDICATED AIRSPEED - KNOTS
FULL AFTERBURNING
WITH WATER INJECTION
Full rudder deflection
4500 ft. altitude
Takeoff angle of attack
200
220
Rosemount pitot-static
120
140 160 180
INDICATED AIRSPEED - KNOTS
FULL AFTERBURNING
WITH WATER INJECTION
Full rudder deflection
4500 ft. altitude
Takeoff angle of attack
200
220
F201-26
3-4
Figure 3-1
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TA-12
tended by decelerating with Military thrust
on the good engine. Either a slight bank or
yaw with full rudder trim or a bank of 8 to
10 degrees with minimum rudder trim and
moderate yaw can be used for cruise.
Fuel management during protracted engine-
out operation should be directed toward
maintaining optimum center-of-gravity con-
ditions, making all of the fuel available to
the operating engine and, when necessary,
continuing the fuel cooling of necessary sys-
tems. Improper cg conditions will be indi-
cated by abnormal pitch trim requirements.
The crossfeed valve should be opened after
tanks 5 and 6 are emptied by right engine
consumption, or if this is the failed engine,
by successive forward transfer operations.
This accomplishes the dual purpose of main-
taining cg and using all available fuel. Fuel
cooling is continued automatically when the
inoperative engine is windrnilling unless its
emergency fuel shutoff switch is actuated.
Crossfeed should never be used during for-
ward transfer when fuel remains in tanks 5
or 6. If it were, most of the fuel transferred
would come from the operating tank(s) of
group 2, 4, or 3 and only a small forward
cg shift would result.
Double Engine Failure
The possibility of a double engine failure is
greater at high speeds because it is possible
for the second engine to flameout as a re-
sult of the yaw angles induced by the first
engine failure. In this case, first restart
the engine that flamed out due to the yaw
maneuver.
If a double engine failure occurs at ex-
tremely low altitude and sufficient airspeed
is available, the aircraft should be zoomed
to exchange airspeed for an increase in
altitude. This will allow more time for ac-
complishing emergency procedures. At-
tempt an air start immediately and repeat
as many times as possible.
THRUST FAILURE DURING TAKEOFF,
TAKEOFF REFUSED
If either the acceleration check speed is
marginal or the thrust of either engine de-
cays or fails and conditions permit:
1. ABORT
Use abort procedure given in this sec-
tion for Takeoff and Landing Emer-
gencies.
AFTERBURNER FLAMEOUT DURING TAKEOFF,
TAKEOFF CONTINUED
If an afterburner fails before leaving the
ground, and a decision is made to continue:
1. AFTERBURNER SWITCH - OFF.
After 5 seconds:
Z. AFTERBURNER SWITCH - ON.
If afterburner does not light:
3. AFTERBURNER SWITCH - OFF.
4. Trim - As necessary.
ENGINE FAILURE IMMEDIATELY AFTER
TAKEOFF
If an engine fails immediately after takeoff
and the decision is made to continue, main-
tain maximum thrust on the operating en-
gine. Normal takeoff speeds are equal to
or faster than minimum directional control
speed. Lateral and directional control of
the aircraft can be maintained when air-
speed remains above the minimum single
engine directional control speed. However,
the ability to maintain altitude and accele-
rate or climb depends on weight, drag,
altitude, airspeed, and temperature. Re-
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TA-12
fer to Appendix for takeoff climb capability
data. When at heavy weight for the existing
air temperature, dumping fuel may reduce
the gross weight sufficiently to remain air-
borne:
1. THROTTLES - MAXIMUM THRUST.
2. CONTINUE STRAIGHT AHEAD.
3. LANDING GEAR LEVER - UP.
4. Fuel dump switch - NORM (if necessary).
5. Throttle - Failed engine OFF.
If not mechanical failure:
6. ATTEMPT AIR START (Refer to Air
Start Procedure this section).
For obvious mechanical failure:
7. Emergency fuel shutoff switch - FUEL
OFF.
DOUBLE ENGINE FAILURE IMMEDIATELY AFTER
TAKEOFF
If a double engine failure occurs, proceed
as follows:
1. IF GEAR IS DOWN AND CONDITIONS
PERMIT - LAND STRAIGHT AHEAD.
2. IF GEAR RETRACTION HAS BEEN
INITIATED OR CONDITIONS DICTATE-
EJECT.
WARNING
Decay of engine RPM will result
in rapid loss of A & B hydraulic
system pressure and subsequent
loss of aircraft control.
AFTERBURNER NOZZLE FAILURE DURING
FLIGHT
Nozzle Failed Closed
Normally, the nozzles open before the
afterburners light. If the afterburner noz-
zle either fails to open or closes during
afterburner operation, there may be a com-
pressor stall, a rapid increase in exhaust
gas temperature, or a decrease in rpm.
If one of these conditions occur:
1. Afterburner switch - OFF.
Nozzle Failed Open
If the nozzle fails open, a maximum of ap-
proximately 60% of Military thrust remains
available without afterburning. The after-
burner can be turned on if necessary and
full afterburner thrust will be available.
AFTERBURNER FLAMEOUT
In the event of a flameout of the afterburner:
1. Afterburner switch - OFF. Wait at
least 5 seconds.
2. Afterburner switch - ON.
If afterburner fails to light:
3. Afterburner switch - OFF.
AFTERBURNER CUTOFF FAILURE
In the event of an electrical failure, or
failure of the afterburner actuator motor
during afterburner operation, the after-
burner may be turned off as follows:
1. Throttle - Retard below afterburner
detent.
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TA-12 SECTION III
1. Inlet air bypass switch - OPEN.
WARNING
2. BYp DR MAN OPEN light on - Check.
A pronounced loss of thrust will
result when the throttle is retarded
below the afterburner detent position.
When the above cutoff procedure has been
used to terminate afterburning after elec-
trical failure, a relight cannot be obtained
until electrical power has been restored.
However, the full range of non-afterburning
thrust is available. A relight cannot be ob-
tained if a complete failure of the after-
burner motor actuator has occurred, or if
the relay in the afterburner electrical con-
trol circuit has failed in the afterburning
position.
To relight the afterburner, proceed as
follows:
1. Afterburner switch - OFF for at least
5 seconds.
Z. A/B CONT circuit breaker - Check.
Reset if necessary.
3. Throttle - Advance to afterburner range.
4. Afterburner switch - ON.
AIR INLET CONTROL SYSTEM
MALFUNCTIONS
A malfunction of the air inlet bypass system
can be caused by hydraulic power loss or
mechanical failure. System malfunctions
and action available to the pilot are as fol-
lows:
BYPASS DR NOT OPEN Light On
Illumination of this warning light indicates
that the bypass doors are closed with land-
ing gear extended.
BYP MAN OPEN Light Not On With Bypass
OPEN Selected
If the inlet air bypass doors remain closed
when OPEN is selected avoid using high
thrust settings at low speeds before landing.
COMPRESSOR STALL OR UNSTABLE INLET
1. INLET AIR BYPASS SWITCHES -
OPEN.
2. AFTERBURNER SWITCHES - OFF.
3. Airspeed - Adjust toward 350 KEAS.
4. Throttles - Adjust to clear stall.
5. Restart engine if flamed out or shut
down.
ENGINE FLAMEOUT PROCEDURE
1. AFTERBURNER SWITCHES - AS
REQUIRED.
2. AIRSPEED - ADJUST TOWARD 300
KEAS OR MACH 0.8.
If not obvious mechanical failure:
3. ATTEMPT AIR START.
If obvious mechanical failure or air start
is unsuccessful:
4. Failed engine generator - TRIP.
5. Crossfeed - As required.
6. Land as soon as possible.
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DOUBLE ENGINE FLAMEOUT
When altitude permits:
1. Use appropriate steps of Engine Flame-
out or Air Start procedures, as appli-
cable.
When altitude is critical or engines will not
start:
2. EJECT.
AIR\ START PROCEDURES
The cause of a flameout should be deter-
mined and corrected prior to attempting an
air start. The estimated air start envelope
for a windmilling engine is shown in this
section. The recommended procedure for
air start is as follows:
1. THROTTLE - OFF (affected engine).
2. INLET BYPASS SWITCH - OPEN.
3. CROSSFEED - ON.
4. THROTTLE - IDLE.
5. ENGINE START SWITCH - AIR (hold).
If no evidence of start within 15 seconds:
6. THROTTLE - OFF.
7. EMERGENCY FUEL CONTROL
SWITCH - EMER.
8. THROTTLE - ADVANCE TO 800-900
PPH FUEL FLOW.
9. ENGINE START SWITCH - AIR (hold).
After start:
10. Throttle and cockpit switches - As
required.
Note
Emergency fuel control switch may
be returned to NORM position after
start unless flameout was caused
by normal fuel control failure.
GLIDE DISTANCE WITH BOTH ENGINES
INOPERATIVE
The glide distance chart shows zero-wind
glide distances with both engines wind-
milling. The glide speed is 0.8 Mach num-
ber above 30,000 feet and 300 knots EAS
below 30,000 feet. This airspeed will pro-
vide near-maximum glide distance capa-
bility and sufficient engine speed to maintain
hydraulic pressure.
With both engines out, the ac generators
furnish rated electrical power at windmill-
ing speeds above 2800 rpm. The emer-
gency battery provides SAS operation at
lower windmilling speeds. There is suf-
ficient hydraulic flow to operate the control
surfaces at satisfactory rates above 3000
rpm and operation at reduced rates is
available to a windmilling speed of approxi-
mately 1500 rpm.
LANDING WITH BOTH ENGINES
INOPERATIVE
Landing with both engines inoperative
should not be attempted.
ENGINE OIL SYSTEM FAILURE
Failure of the engine oil system is indicated
by the oil pressure gage. If an oil system
malfunction causes oil starvation of the
engine bearings, the result will be pro-
gressive bearing failure, engine roughness,
oil seal failure, loss of oil, and subsequent
engine seizure. Bearing failure due to oil
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SECTION III
ESTIMATED AIRSTART ENVELOPE (Wind milling Engine)
55
ALTITUDE - 1000 FEET
50
45
40
35
30
STARTING
ENVELOPE
.4 .5
.6 .7
Figure 3-2
.8
1.0
F201-25
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DEAD ENGINE GLIDE DISTANCE
70
60
50
40
ALTITUDE - 1000 FEET
30
20
10
0
TWO ENGINES
I
AT WINDMILLING
I
RPM
GEAR UP SPEED:
0.8 MACH ABOVE
300 KEAS BELOW
85,000 LBS.
ZERO WIND
30,000 FT.
30,000 FT.
GROSS WEIGHT AND UNDER
14111�11\
GEAR DOWN
0.65 MACH ABOVE
230 KEAS BELOW
GEAR DOWN
SPEED:
30,000 FT.
30,000 FT.
DISTANCE = 1/2 GEAR UP
DISTANCE
\
0
20
ao 60
GLIDE DISTANCE - NAUTICAL MILES
80
F201-24 .
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starvation is generally characterized by a If engine is operating:
rapidly increasing vibration. If this is noted,
accompanied by a pressure loss on the gage, 1. Throttle - OFF.
do the following:
2. Have ground crew continue cranking
1. Throttle - OFF. engine.
2. Land as soon as possible.
ENGINE FUEL CONTROL FAILURE
When a failure of the normal fuel control is
suspected, do the following:
CAUTION
Never apply ground air supply to
engine starter when engine rpm
exceeds 3490.
3. Emergency fuel shutoff switch -
1. Throttle - Adjust for smooth switchover. FUEL OFF.
2. Emergency fuel control switch - EMER. 4. Battery switch - OFF.
After fuel control malfunction, do not re- 5. Abandon aircraft.
turn the emergency fuel control selector
switch from EMER to NORM for the dur-
ation of the flight. To do so might result ENGINE FIRE DURING TAKEOFF, TAKEOFF
in an engine flameout. Descents should be REFUSED
made with the throttle advanced to provide
no less than 1500 pph fuel flow per engine If either FIRE warning light illuminates
to prevent excessive fuel temperatures in before leaving the ground, do the following:
the fuel control. Careful throttle movement
and close monitoring of EGT is necessary 1. ABORT.
as the emergency fuel control only senses Use abort procedure given in this
compressor inlet pressures and fuel flow is section for Takeoff and Landing Emer-
controlled by the throttle. When practicing gencies.
emergency procedures with a properly
functioning normal fuel system, the trans- 2. THROTTLE - OFF (AFFECTED
fer back to NORM should only be made after ENGINE).
first retarding the throttle to 7860 rpm or
less to avoid excessive pressure surge in 3. EMER FUEL SHUTOFF SWITCH -
the engine fuel system. FUEL OFF.
FIRE
ENGINE FIRE DURING GROUND START
If there is evidence of fire during ground
start, keep the engine rotating until the fire
is extinguished. Apply chemicals from out-
side the engine as a last resort.
4. Shut down other engine after stopping.
5. Seat pin - Insert if time permits.
6. Abandon aircraft.
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ENGINE FIRE DURING FLIGHT
Illumination of a nacelle FIRE warning light
indicates a nacelle compartment temperature
above approximately 1050 F. An immediate
check for abnormal EGT, trailing smoke, or
any other indication of the presense of fire
should be made. In case of doubt, assume
that a fire does exist and proceed as follows:
. THROTTLE - RETARD BELOW AFTER-
BURNER DETENT (AFFECTED
ENGINE).
If light remains on:
2. THROTTLE - IDLE ABOVE MINIMUM
CONTROL SPEED.
If light still remains on:
3. THROTTLE - OFF.
4. EMER FUEL SHUTOFF SWITCH -
FUEL OFF.
5. CHECK FOR OTHER INDICATIONS OF
FIRE.
If fire confirmed:
6. EJECT.
If no fire:
7. Land as soon as possible.
ENGINE FIRE AFTER SHUTDOWN
Use applicable steps of Engine Fire During
Ground Start procedure, this section.
ELECTRICAL FIRE
The pilot's ability to detect an electrical
fire is somewhat limited when wearing a
pressure suit because he is not exposed to
the characteristic odor. He must depend on
visual detection of smoke in the cockpit.
The method of fighting an electrical fire is
different from the customary procedure in
that it is not desirable to turn off all elec-
trical power simultaneously. Such action
would turn off the SAS and fuel boost pumps;
however, the battery and one generator may
be turned off with no adverse effect on es-
sential systems.
1. Both generators should not be turned
off simultaneously unless absolutely
necessary.
2. Turn off all non-essential electrical
systems.
3. Turn electrical equipment back on in-
dividually in an attempt to isolate mal-
function.
4. Land as soon as possible.
ELIMINATION OF SMOKE AND FUMES
The pilot cannot detect fumes when wearing
a pressure suit. The helmet oxygen system
is independent of the cockpit and suit air
supply system. Smoke can be eliminated
promptly by dumping cabin pressure unless
the smoke enters from the cockpit air con-
ditioning system. With the cockpits pres-
surized, shutting off the cockpit air systems
will not depressurize the cockpits; however,
there will be a minimum of ventilating air
flow. If the smoke is introduced by the
forward cockpit air supply system, switch
the cockpit system to CROSSOVER. The
defog system should be off at all times when
not required.
WARNING
When pressure is dumped, cockpit
depressurization will occur at an
extremely rapid rate and the pilots
will be dependent on their suit
pressure for altitude protection.
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SECTION III
EMERGENCY ESCAPE
Escape from the aircraft in flight should be
made with the ejection seat. The following
is a summary of ejection expectations:
a. At sea level, wind blast exerts only
minor forces on the body up to 525
knots; appreciable forces from 525 to
600 knots. The aircraft limit airspeed
Is below these speeds.
b. Ejection at 65 KIAS and above during
takeoff roll results in successful chute
deployment.
c. The free fall froth high altitude down to
15,000 feet with drogue chute stabili-
zation will result in stabilized descent
in the quickest manner.
CAUTION
Flight with oxygen mask and re-
gulator are restricted to below
FL 500 and below 420 KEAS because
of wind blast forces anticipated in
the event of ejection. Before actual
ejection airspeed should be reduced
to subsonic and as slow as conditions
permit.
BEFORE EJECTION, IF TIME AND CONDITIONS
PERMIT
1. Altitude - Reduce so that the pressure
suit is not essential to survival.
2. Airspeed - Reduce to subsonic and as
slow as conditions permit.
3. Head aircraft toward unpopulated area.
4. Transmit location and intentions to
nearest radio facility.
5. IFF - EMERGENCY position.
EJECTION
A minimum risk ejection can be performed
at any height when airspeed is above 65
KIAS with wings level while in level or
climbing flight; however, accomplish bail-
out above 2000 feet when feasible.
1. If possible, aircraft commander will
notify crewmember of decision to eject.
2. Actuate BAIL OUT light switch.
3. Observe forward seat ejection (or AFT
SEAT EJECTED light illuminate if air-
craft commander in forward cockpit).
WARNING
If forward seat ejects first, do not
eject from aft cockpit or jettison
canopy until forward seat is ob-
served to go. Ejection of pilot
may be determined visually, by
noting rocket blast, by feeling
aircraft shake, or by hearing seat
ejection system fire.
4. Lower visor.
5. GREEN APPLE - Pull.
Pull green apple if at altitude.
6. EJECTION RING - PULL.
If seat fails to eject after a normal delay,
do the following:
7. JETTISON CANOPY. Operate canopy
jettison handle. If canopy does not
jettison, attempt to blow off by open-
into slipstream with canopy latch
handle.
8. EJECTION SEAT T-HANDLE - PULL.
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WARNING
Do not pull ejection seat T-handle
until canopy is gone. There is no
safety interlock to prevent ejection
by T-handle when canopy is in place.
In the event the ejection seat still fails to
eject, continue as follows:
9. Slow aircraft to between 250 and 300
KEAS.
0. FOOT SPURS - RELEASE.
1. PERSONAL LEADS - DISCONNECT.
2. TRIM AIRCRAFT FULL NOSEDOWN,
HOLD STICK NEUTRAL.
3. ROLL INVERTED, LEAN FORWARD.
4. SIMULTANEOUSLY RELEASE LAP
BELT AND CONTROL STICK.
5. AFTER CLEAR, PULL PARACHUTE
ARMING LANYARD (RED KNOB) ON
PARACHUTE HARNESS.
6. Survival kit release handle - Pull after
parachute opens to reduce landing
impact.
EMERGENCY EXIT ON THE GROUND
To exit on the ground in an emergency, pro-
ceed as follows:
1. Ejection seat safety pin - Install if
time permits.
2. Lap belt and shoulder harness
3. Personal leads - Disconnect.
- Release.
4. Parachute harness attachments - Re-
lease.
5. Foot spurs - Release manually. (Use
cable cutter if unable to release spurs.)
6. Canopy - Unlatch or jettison as appli-
cable.
7. Abandon aircraft.
TAKEOFF AND LANDING EMERGENCIES
ABORT
The abort procedure assumes that a deci-
sion to abort will be made before rotation
speed is reached. Aborts from above ro-
tation speed are not prohibited, but the
risks associated with aborting from such
a high initial speed at takeoff weight must
be balanced against those of continuing a
takeoff when making the decision. In gen-
eral, after rotation speed is reached, the
most reasonable course of action is to con-
tinue rather than abort unless the emer-
gency is such that the aircraft can not fly.
Engine Management
Both throttles should be retarded to IDLE
and the brakes applied with the nose down
as soon as the decision to abort is made.
Reaction time and residual thrust will usu-
ally cause airspeed to continue increasing
until engine rpm begins to decrease. The
planned rotation speed may be exceeded as
a result; however, the nosewheel should be
kept on the runway to take advantage of
nosewheel steering in combination with
rudder control. Shutdown of one engine will
shorten the stopping distance, but shutdown
is not necessary unless the drag chute does
not operate properly. In the event of chute
failure, shutdown the right engine after
both are idling, or complete the shutdown of
a failed or flamed out engine.
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WARNING
Wait until rpm and EGT show that
both engines are idling or that one
engine is failing before selecting
the engine to shutdown. Loss of
both engines may result in loss of
hydraulic pressure for braking.
Aircraft Attitude, With Decision to Abort
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Brake Switch
Unless rotation has been initiated, keep the
nosewheel on the runway. If a nose up atti-
tude has been established but can be checked
short of lift-off, lower the nose immediately
if below 140 KIAS. If above 190 KIAS, use
an angle of attack of 10o to 12o while decel-
erating toward 190 KIAS. Lower the nose
as 140 KIAS is approached. Energize the
brakes simultaneously with nosewheel con-
tact.
When rotation is well advanced, the aircraft
may accelerate beyond takeoff speed and
liftoff before rotation can be checked. In
this case hold the aircraft off sufficiently to
regain control and then touchdown without
sideslip if possible. Fly the aircraft back
to the runway, attempting to regain the
center. Lower the nose as 190 is approached.
Chute Deployment
The drag chute requires 4 to 5 seconds for
deployment after drag chute actuation. It is
permissible to actuate the deploy handle
while decelerating in anticipation of reach-
ing 190 KIAS; however, premature deploy-
ment can result in destruction of the chute.
Actuation of the chute system so as to reach
190 KIAS simultaneously with loading of the
chute is not recommended unless the risk
is justified by a very marginal distance re-
maining situation. It is better to actuate
the drag chute switch at or below 190 KIAS
while decelerating.
The normal ANTI-SKID ON brake switch
setting provides nosewheel steering and
braking power from the L hydraulic system
and anti-skid protection. It is not necessary
to change the switch setting unless the left
hydraulic pressure has failed or anti-skid
off is desired. Selection of ANTI-SKID OFF
or ALT STEER & BRAKE causes the ANTI-
SKID OUT warning light on the annunciator
panel to illuminate.
ABORT PROCEDURE
WARNING
. Do not unfasten the lap belt or
shoulder harness until the air-
craft has come to a stop.
. The landing gear should be left
In the extended position.
1. THROTTLES - IDLE.
Retard both throttles to IDLE. Do not at-
tempt to shut down either engine immedi-
ately unless failure to do so would vitally
endanger the aircraft.
2. NOSEWHEEL STEERING - ENGAGE.
3. BRAKES - OPTIMUM BRAKING.
For dry runway: use moderate to
heavy brake pressure.
For wet runway: use light to
moderate brake pressure.
4. DRAG CHUTE HANDLE - PULL.
The limit airspeed for drag chute de-
ployment is 190 KIAS.
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5. BRAKE SWITCH - As required.
Set the brake switch to ALT STEER
& BRAKE when the L hydraulic
system is below normal pressure
due to system or left engine failure.
CAUTION
Selection of ALT STEER gz BRAKE
changes the source of brake pres-
sure from the L to the R hydraulic
system and disables the anti-skid
system.
6. Shut down one engine (if necessary).
Shut down of one engine is considered
necessary in the event of drag chute
failure.
If drag chute fails to deploy, use
DRAG CHUTE FAILURE Procedure,
this section.
Shut down the right engine if both
engines are idling or if the right
engine has failed.
Shut down the left engine if it has failed.
WARNING
Positively identify the failed engine
before attempting engine shutdown.
DRAG CHUTE FAILURE
If the drag chute should fail to deploy and
stopping distance is critical, proceed as
follows:
Dry Runway
1. LOWER NOSE IMMEDIATELY.
2. NOSEWHEEL STEERING - ENGAGE.
3. BRAKES - AS REQUIRED UP TO MAX-
IMUM BRAKING.
4. RIGHT ENGINE THROTTLE - OFF, IF
REQUIRED.
5. HOLD AS MUCH UP ELEVON AS POS-
SIBLE AND STILL KEEP THE NOSE-
WHEEL ON RUNWAY FOR DIREC-
TIONAL CONTROL.
Wet Or Icy Runway
1. LOWER NOSE.
a. LANDING - AT 110 KIAS.
b. ABORT - IMMEDIATELY AT 190
KIAS.
2. NOSEWHEEL STEERING - ENGAGE.
3. BRAKES SWITCH - NORM.
4. BRAKES - MAXIMUM PRESSURE.
5. RIGHT ENGINE THROTTLE - OFF.
6. HOLD AS MUCH UP ELEVON AS POS-
SIBLE, BUT KEEP THE NOSEWHEEL
ON THE RUNWAY FOR DIRECTIONAL
CONTROL.
Note
This wet or icy runway technique
will probably blow the tires early in
the landing roll; however, direc-
tional control can still be maintained
and the blown tires will remain on
the wheels. Additional pedal pres-
sure will be required as each tire
blows. Maximum wing aerodynamic
braking is more effective than wheel
braking on a wet or icy runway until
the nose is lowered but the nose up
attitude must not be held to a point
that the nosewheel will slam onto
the runway. Use of maximum
possible up elevon after the nose
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is lowered while keeping the nose-
wheel on the runway provides aero-
dynamic drag and additional down
load on the main wheels.
SINGLE ENGINE LANDING
A single engine landing is basically the
same as a normal landing, except that the
pattern may be entered at any point and is
expanded to avoid steep turns. Airspeed
is maintained above normal on final ap-
proach. The outstanding difference from
normal landings is the noticeable change in
directional trim with power changes. The
most marked trim change will occur as the
throttle is retarded during flare. This is
reduced by setting the rudder trim to neu-
tral on trim indicator after final approach
is established. Directional heading is
maintained by rudder pressure until thrust
is smoothly reduced during the flare. The
landing gear may be lowered after lining up
on final approach if the L hydraulic system
is operating; however, at least 90 seconds
must be allowed for emergency gear ex-
tension if the L hydraulic system is inop-
erative.
1. Fuel - DUMP and TRANSFER as re-
quired.
2. Review hydraulic services available.
3. If left engine has failed, brake switch-
ALT STEER & BRAKE.
4. Inoperative engine SAS pitch and yaw
switches - OFF.
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. SAS roll switches - Both OFF.
. Operative engine SAS roll switch - ON.
. Landing gear lever - DOWN.
8. Establish steeper than normal final ap-
proach.
9. Maintain 200 K1AS minimum until land-
ing assured.
Note
If it is necessary to land with more
than 35,000 pounds of fuel remain-
ing increase minimum approach
speed 1 knot for each additional
1000 pounds.
10. Rudder trim - Neutral.
Note
Partial afterburning thrust may
be required during final approach.
WARNING
If the throttle is retarded below
the afterburner detent or the after-
burner switch is turned off while in
the partial afterburning range, a
significant loss of thrust will result.
.1. When landing is assured - Retard
throttle.
.2. Make normal touchdown.
SINGLE ENGINE GO-AROUND
Make go-around decision as soon as possible
on final and prior to flare.
1. Throttle - As required.
2. Afterburner switch - ON, if required.
3. Continue approach until go-around as-
sured.
4. Landing gear lever - UP, when de-
scent is checked.
5. Trim - As necessary.
6. Accelerate to 250 KIAS climb speed.
SIMULATED SINGLE ENGINE LANDING
Directional trim changes will be more pro-
nounced during an actual single engine
situation with one engine windmilling.
1. Retard one throttle to IDLE.
2. Follow Single Engine Landing procedure.
FORCED LANDING
At least one engine must be operating if a
forced landing is to be attempted. All
forced landings should be made with the
landing gear extended regardless of terrain.
High airspeed or nose-high angle of impact
during landings with gear retracted causes
the aircraft to "slap" the ground on impact,
subjecting the pilot to possible spinal injury.
It is recommended that a gear-up landing
not be attempted with this aircraft; EJECT
instead. If a forced landing is necessary,
proceed as follows:
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SECTION III
1. Fuel dump switch - NORM, dump as re-
quired. Terminate fuel dumping at
least 30 seconds prior to contact. Refer
to Fuel Dump, this section.
2. Fuel transfer - As required.
3. Landing gear lever - DOWN.
4. Shoulder harness - Lock.
5. Canopy - Jettison during approach if
desired.
CAUTION
If crash and rescue personnel are
Immediately available it may be
preferable to retain the canopy
until aircraft stops to reduce pos-
sibility of burns in case fire occurs
during landing.
6. Throttles - OFF at touchdown.
7. Drag chute handle - Pull out to deploy
chute.
8. Emergency fuel shutoff switches -
FUEL OFF.
9. Battery and generator switches - OFF.
10. Canopy - Manually open or jettison if
not accomplished during approach.
11. Abandon aircraft as soon as possible.
LANDING GEAR UNSAFE INDICATION
A landing gear unsafe indication could be
caused by low L hydraulic system pressure
or malfunction within the landing gear ex-
tension or indication system. If the L hy-
draulic system has failed, refer to Landing
Gear System Emergency Operation (Ex-
tension) procedure, this section. Upon de-
tecing an unsafe indication, proceed as
follows:
1. Landing gear circuit breakers and
lights - Check.
2. L hydraulic pressure - Check.
3. Landing gear lever - Recycle to DOWN
position; repeat if necessary.
If landing gear still indicates unsafe:
4. Landing gear position - Determine by
reference to tower or other aircraft.
If the landing gear appears down and locked:
5. Make normal approach and land on side
of the runway away from suspected
unsafe gear and observe the following
precautions.
a. Inertia reel lock lever - LOCK.
b. Hold weight off unsafe gear as long
as possible. If nose gear indicates
unsafe hold off then lower smoothly
at approximately 110 KIAS.
c. Allow aircraft to roll to a stop
straight ahead and do not attempt
to taxi or shut down engine until
landing gear ground safety pins
are installed.
6. If any gear remains fully retracted
refer to Landing Gear System Emer-
gency Operation (Extension) procedure,
this section.
7. If all gear extended, but not fully, refer
to Partial Gear Landing procedure,
this section.
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Note
. Increasing airspeed may assist in
locking a partially extended nose
gear.
. Yawing aircraft may assist in
locking a partially extended main
landing gear.
PARTIAL-GEAR LANDING
A landing with the nose gear retracted or
with all gear up should not be attempted. A
landing with the nose gear and one main
gear locked down is not recommended.
Under ideal circumstances, a landing with
the nose gear extended and both main wheels
retracted may be possible. If this config-
uration can be accomplished, base a de-
cision to land or eject on whether or not
other factors are favorable. An unob-
structed runout surface adjacent to the run-
way is desirable. A dry lakebed landing
might be preferable. Wind velocity and di-
rection are important in selection of the
landing heading.
If a decision is made to land, conventional
final approach and landing speeds and atti-
tudes are recommended. This will result
in the tail touching while the nose is at less
than normal height. An attempt to hold the
aircraft off by using a higher pitch angle is
not recommended because of the greater
possibility of high impact loads as the nose
gear slaps down. An empty tank 1 condition
is desired.
1. Accomplish nose-gear-only configu-
ration if necessary, as follows:
a. Landing gear CONT circuit breaker-
Push in.
b. Landing gear lever - Up.
c. Landing gear CONT circuit
breaker - Pull.
Note
Nosewheel steering will not be
available.
d. Manual landing gear release
handle - Pull to release nose gear
only (first lock releases nose gear).
Check nose gear down light - ON.
2. Do not transfer fuel forward.
3. Fuel dump switch - NORM, if necessary
to reduce weight.
4. Battery switch - OFF.
5. Inertia reel lock lever - LOCK.
6. Canopy jettison handle - Pull, if de-
sired.
Note
If the canopy is not jettisoned
prior to landing, it should not
be unlocked until the aircraft
has stopped.
7. Make normal approach and landing.
8. Drag chute handle - Pull out to deploy
chute.
9. Use rudders for directional control.
10. Throttles - OFF, when directional con-
trol is no longer possible.
11. Abandon aircraft as quickly as possible.
MAIN OR NOSE GEAR TIRE FAILURE ON
TAKEOFF
If takeoff is continued after a tire has failed
or is suspected to have failed, do not retract
the landing gear until the tire has been vi-
sually checked by another aircraft or a con-
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TA-12
SECTION III
trol tower. If no fire is evident and the
gear is to be retracted, move the brake
switch to the NORM position and apply brakes
prior to raising the gear. Stopping wheel �
rotation will prevent tire fragmentation
damage in the wheel well. (With the brake
switch in the ANTI-SKID position the brakes
cannot be applied when the weight of the air-
craft is off the main gear.) After the gear
has been retracted return brake switch to
the ANTI-SKID position. The most likely
place to blow a main gear tire is during the
latter portion of the takeof run. The follow-
ing procedure is recommended when a main
or nose gear tire fails during takeoff run:
1. IF SPEED AND RUNWAY PERMIT -
ABORT. Refer to abort procedure.
2. If rotation speed has been reached,
continue takeoff. (Do not retract the
gear until a visual check is made.)
MAIN GEAR FLAT TIRE LANDING
Plan the landing for minimum gross weight
with touchdown to be made on the side of
the runway away from the flat tire. It is
possible that only one or two of the three
tires has failed. If only one tire has failed,
little danger exists when landing at low
weight because two tires have sufficient
strength to support the aircraft.
WARNING
Maintain IDLE rpm until fire-
fighting equipment arrives.
Engine shutdown allows fuel to
vent in the vicinity of the wheel
brake area, thus creating a fire
hazard.
NOSE GEAR FLAT TIRE LANDING
If it is necessary to land with a flat nose-
wheel tire or tires, proceed as follows
after making a normal touchdown:
1. Drag chute handle - Pull out to deploy
chute.
2.
Nose gear - Hold off.
Hold the nosewheel off as long as
practicable (approximately 110 KIA.S)
and then lower gently to runway.
3. Brakes - Use differential braking to
maintain directional control.
EMERGENCY ENTRANCE
The procedure to be used by rescue per-
sonnel when assisting a disabled pilot from
the aircraft following a crash landing is as
follows:
1.
Touchdown on good tires.
a.
2.
Drag chute handle - Pull out to deploy
chute as soon as possible.
3.
Nosewheel - Lower.
b.
4.
Nosewheel steering switch - Engage.
c.
5.
Hold weight off flat tire.
6.
Brakes - Use differential braking to
maintain directional control.
d.
If aircraft is on fire or if external
latch control cannot be operated, jetti-
son canopies by pulling canopy jettison
T-handle on left chine.
Shut off oxygen supply at oxygen control
panel.
Open helmet faceplate before discon-
necting oxygen line.
Ensure that seat will not fire acciden-
tally - Safety the ejection ring.
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e. Release lap belt and shoulder harness.
I. Disconnect pilot's personal leads and
emergency oxygen hose.
g. Disconnect parachute attachments.
h. Pull manual cable cutter ring.
i. Remove pilot gently to avoid aggra-
vating possible internal injuries.
DITCHING
Ditching should not be attempted. All emer-
gency survival equipment is carried by the
pilot; consequently, there is nothing to be
gained by riding the airplane down. Ejec-
tion is the best course of action when the
alternative is ditching.
FUEL SYSTEM FAILURE
INCORRECT FUEL SEQUENCING
Incorrect automatic fuel sequencing is in-
dicated primarily by the fuel boost pump
lights. (A switch may illuminate out of
normal sequence, or fail to illuminate on
schedule.) The remedy for this is to se-
quence manually until either automatic se-
quencing resumes or a landing is made. It
is possible that faulty fuel sequencing may
manifest itself by secondary indications,
such as a fuel low level light coming on
prematurely, or an abnormal adjustment
required in pitch trim (due to cg change by
faulty fuel distribution). Note that forward
cg requires increased power to maintain
speed and altitude. If normal sequencing
does not resume, and manual sequencing is
either inconvenient or impossible, turn
crossfeed on or transfer fuel to ensure that
any available fuel will get to the engines.
CAUTION
Do not permit a fuel boost pump
to continue operating in an empty
fuel tank or the boost pump will
be damaged.
FUEL PRESSURE LOW WARNING
If one or both low fuel pressure lights il-
luminate, proceed as follows:
1. Crossfeed - Press ON.
2. Tanks containing fuel - Press ON.
3. Analyze difficult and attempt to re-
store normal sequencing.
4. Crossfeed - Press OFF.
If normal operation can not be restored:
5. Land as soon as possible.
FUEL BOOST PUMP FAILURE
Loss of all boost pumps can only result
from multiple failures. It would be indi-
cated by illumination of both low fuel pres-
sure lights. If this occurs during takeoff,
fuel tank pressurization will supply suffi-
cient fuel to the engine driven pumps to
maintain engine operation. The takeoff
should be aborted if speed and runway
length permit.
WARNING
Fuel cannot be dumped with com-
plete boost pump failure. Use
caution if heavy weight landing is
required.
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SECTION LTI
Partial boost pump failure may not be indi-
cated by the low fuel pressure lights. Im-
proper fuel sequencing and center of gravity
shift may be the first indication. Proceed
as directed in INCORRECT FUEL SE-
QUENCING. Crossfeed may be required;
however, more fuel will tend to feed from
forward tanks that have boost pumps op-
erating when crossfeed is on. This could
cause an aft cg shift which might be haz-
ardous when operating with a failed pitch
SAS.
FUEL TANK PRESSURIZATION FAILURE
Fuel tank pressurization failure is indicated
by the tank pressure gage and warning light
and by the liquid nitrogen quantity indicator.
The liquid nitrogen quantity low warning
light on the annunciator panel indicates im-
pending failure. No corrective action is
possible after both liquid nitrogen systems
are depleted except to limit rates of descent.
In descent, the fuel tank suction relief valve
in the nosewheel well opens when slightly
negative tank pressures occur. Rates of
descent should be limited so that tank pres-
sure does not become less than -1/2 psi.
This is the minimum tank pressure limit
and is based on structural capabilities of
the fuselage tanks.
FUEL DUMPING PROCEDURE
Fuel dumping provides a means of rapidly
reducing the aircraft weight in an emergency.
All tanks containing fuel except tank I will
empty in the normal fuel tank usage se-
quence: Tank 1 fuel cannot be dumped
since the boost pumps in this tank are
stopped when the fuel dump switch is moved
to either the NORM or EMER position.
When in the NORM position, dumping will
continue until the remaining fuel in tank 3
reaches 5000 pounds. Dumping will then
terminate and the boost pumps in tank 1 will
automatically start. When in the EMER
position, dumping will continue until all fuel
except any remaining in tank 1 is dumped.
To avoid fuel pressure fluctuations, the
boost pumps in tank 1 must be started be-
fore tank 3 completely empties, by moving
the fuel dump switch to the OFF position.
The boost pumps in tank 1 may also be
started by pressing the tank 1 boost pump
switch; however, this will terminate dump-
ing. To increase the dump rate, manually
select boost pumps for all tanks containing
fuel, except tank 1. The R FUEL PRESS
LOW warning light may illuminate at low
engine speeds.
NORMAL FUEL DUMPING
Accomplish normal fuel dumping as follows:
1. Fuel dump switch - NORM.
2. Fuel quantity - Alternately monitor
total fuel and tank 3 fuel.
3. Fuel dump switch - OFF when 5000
pounds remain in tank 3.
EMERGENCY FUEL DUMPING
If the fuel level in tank 3 has prematurely
reached the 5000 pound level and dumping is
required (excessive fuel in tanks 4, 5 or 6)
proceed as follows:
1. Fuel dump switch - EMER.
2. Tanks 4, 5 or 6 containing fuel - Press
on.
3. Forward transfer switch - FWD TRANS
(if required).
4. Fuel quantity - Alternately monitor
tanks 1 and 3.
When tank 1 quantity reads 3000 pounds:
5. Forward transfer switch - OFF.
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When required amount of fuel remains:
6. Fuel dump switch - OFF.
FORWARD FUEL TRANSFER AND FUEL DUMPING
PROCEDURE
Forward fuel transfer and fuel dumping may
be accomplished simultaneously as follows:
1. Fuel dump switch - NORM.
2. Forward transfer switch - FWD TRANS.
3. Fuel quantity - Alternately monitor
tanks 1 and 3.
When tank 1 fuel quantity reads 3000 pounds:
4. Forward transfer switch - OFF.
5. Fuel dump switch - OFF when 5000
pounds remain in tank 3.
FUEL QUANTITY LOW WARNING
If the fuel-quantity-low warning light comes
on with appreciably more than 5000 pounds
of TOTAL fuel indicated on the quantity
gage, determine total fuel from the individ-
ual tank quantities. Monitor tank 3 quantity
and land as soon as practicable. Quantity
indications are affected by pitch attitude and
longitudinal acceleration. Total quantity in-
dication is also affected by the fuel distri-
bution in the individual tanks.
If the fuel quantity low warning light does
not come on with less than 5000 pounds of
TOTAL fuel indicated on the quantity gage,
test warning light and land as soon as pos-
sible.
ELECTRICAL POWER SYSTEM FAILURE
SINGLE AC GENERATOR FAILURE
Failure of one ac generator will be indicated
by illumination of the L GENERATOR OUT
or the R GENERATOR OUT warning light.
One generator in normal operation is suffi-
cient to carry the entire aircraft electrical
load. In the event of generator failure, pro-
ceed as follows:
1. Generator switch - RESET then release.
If the generator fault was momentary the
generator will be reconnected to the system
and the warning light will extinguish.
If the light remains on:
2. Generator switch - TRIP.
3. Land as soon as possible.
DOUBLE AC GENERATOR FAILURE
If both ac generators fail, the dc monitored
bus will be dead. The only souce of power
will be the battery which will automatically
power the dc essential bus if the battery
switch is on. With reduced usage the bat-
tery will last approximately 30 minutes with
all inverters on. Some dc systems which
may not always be essential for flight can-
not be turned off by the pilot unless the cir-
cuit breakers are pulled. These are diffi-
cult to reach when wearing a pressure suit.
The UHF radio should be off except when
absolutely necessary because its large
power requirement will deplete battery
power rapidly. With complete generator
failure, fuel boost pumps are inoperative.
Proceed as follows:
1. Battery switch - Check ON.
2. Generator switches - RESET.
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SECTION III
3. If both generators do not reset -
Land as soon as possible; conserve
battery power if both generators are
inoperative.
TRANSFORMER-RECTIFIER FAI LURE
One transformer-rectifier will supply the
normal electrical demands. Fixed-fre-
quency ac power systems will continue to
operate normally. A double failure of the
transformer-rectifers removes power from
the dc monitored bus. The battery will op-
erate the dc essential bus. Conserve bat-
tery power and land as soon as possible.
INVERTER FAI LURE
The No. 2 inverter may be used to supply
either the No. 1 or No. 3 inverter bus in
addition to supplying its own bus. If both
No. 1 and No. 3 inverters should fail, and
the respective switches are placed in the
EMER position, only No. 1 and No. 2 in-
verter buses will be powered by the opera-
tive No. 2 inverter. Failure of No. 2 in-
verter will make the standby attitude gyro
and B SAS channels in pitch, yaw, and roll
inoperative. If the No. 1 or No. 2 inverter
fails, proceed as follows:
1. A & B SAS roll channel switches - OFF.
2. No. 2 inverter switch - Check ON.
Check NO. 2 INVERTER OUT light not il-
luminated:
3. Failed inverter switch - EMER.
Check that NO. 1 INVERTER OUT or NO. 3
(as applicable) INVERTER OUT light extin-
guishes.
4. A & B SAS roll channel switches -
ON.
CAUTION
Both SAS roll channels should be
disengaged prior to turning ON a
normal or emergency inverter,
or switching inverter loads in
flight.
5. Illuminated SAS recycle lights - Press.
HYDRAULIC POWER SYSTEM FAILURE
With both engines out, the hydraulic pumps
provide sufficient flow for satisfactory flight
control system operation at windmill speeds
above 3000 rpm. Reduced control system
capability is available down to a windrn.illing
speed of approximately 1500 rpm. With one
engine windmilling, all primary and most
utility services are supplied by the operating
engine hydraulic systems. The windmilling
engine utility system pressure and flow may
be sufficient to supply service until the en-
gine is almost stopped.
PRIMARY HYDRAULIC SYSTEM FAI LURE
The loss of either A or B hydraulic system
will be indicated by the warning light on
both annunciator panels, the master caution
light, and the hydraulic pressure gage. Re-
duce speed to less than 350 KEAS if either
A or B system fails.
Disengagement of the failed hydraulic sys-
tem SAS channels is necessary to maintain
full yaw and roll damping capability. As a
hydraulic system failure is not sensed by
the SAS equipment, it is necessary to double
the SAS signal gain of the operating channel
to give the equivalent control response in
yaw and roll. Airspeed reduction with a
single hydraulic system is a precautionary
procedure which allows for the reduction
In available hinge moment capability. DiB -
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SECTION III
TA-12
engagement of the failed system SAS pitch
channel is not mandatory, but it may be
more desirable to disengage all three chan-
nels than only the yaw and roll switches.
Monitor all system operations closely and
attempt to determine if a complete failure
is imminent. Be prepared for ejection
prior to complete failure.
UTILITY HYDRAULIC SYSTEM FAILURE
The loss of L or R hydraulic system will be
indicated by the hydraulic pressure gage.
If the pressure of the L system falls below
2200 psi, crossover for gear retraction is
automatic. The manual release must be
used to lower the gear. Items which are af-
fected by the L hydraulic system are normal
brakes, UHF antenna retraction, nosewheel
steering, forward cockpit air-conditioning,
aerial refueling system, and the left inlet
control actuator. Items which are affected
by the R hydraulic system are the aft cock-
pit air-conditioning, right inlet control
actuator, alternate steer and brake and air
refueling system.
A OR B HYDRAULIC SYSTEM FAILURE
1. Reduce speed to less than 350 KEAS.
CAUTION
Do not exceed 350 KEAS with either
an A or B hydraulic system inop-
erative. If either system fails
above this speed, reduce speed as
soon as possible. Flight control
responsiveness will be reduced
during single hydraulic system op-
eration at high KEAS and Mach
numbers, and flight under these
conditions should be held to a
minimum.
2. Affected SAS yaw and pitch switches -
OFF.
3. SAS roll switches - OFF.
4. Operative roll channel switch - ON.
Note
When one roll SAS channel is dis-
engaged or turned off, the sim-
plified logic circuit will disen-
gage the other roll channel. The
desired roll channel switch must
be turned OFF and then reengaged
to regain single channel roll SAS
operation.
5. Reserve hydraulic oil switch - A or B
(whichever system is operative).
A AND B HYDRAULIC SYSTEMS
FAILURE
1. EJECT.
WARNING
If both the A and B hydraulic sys-
tems have failed all flight controls
will be inoperative.
FLIGHT CONTROL SYSTEM FAILURE
With both engines out, the ac generators
furnish rated electrical power at windmill
speeds above 2800 rpm. The emergency
battery provides SAS operation lower wind-
mill speeds. There is sufficient hydraulic
flow to operate the control surfaces at sat-
isfactory rates above 3000 rpm and op-
eration at reduced rates is available to a
windrnilling speed of approximately 1500
rpm.
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Note
During single engine operation a
windmilling engine may not develop
sufficient system hydraulic pres-
sure to maintain operation of its
associated SAS servo channels. To
avoid nuisance disengagement of
SAS channels, turn off all three
SAS channel switches for the wind-
milling engine hydraulic system
when lower than normal pressure
is indicated. Pitch and Yaw SAS
damping will continue on one chan-
nel. The operative engine SAS
roll channel must be cycled OFF
then ON to maintain damping in the
Roll axis.
FLIGHT CONTROL SYSTEM EMERGENCY
OPERATION
If either the A or B hydraulic system fails,
the control forces will not change. Either
system will operate the control surfaces,
but at a slower rate and with some reduction
in control responsiveness at high KEAS and
Mach numbers.
If control difficulties are encountered:
1. Check A and B hydraulic system pres-
sures. If either A or B hydraulic sys-
tem has failed proceed as directed for
A and B hydraulic system failure this
section.
2. Disengage autopilot and check control.
3. Check SAS warning lights. If SAS
failure has occurred, proceed as di-
rected under SAS Emergency Operation
this section.
SAS EMERGENCY OPERATION
SAS emergency operating procedures and
the applicable flight limitations should be
used whenever there has been a channel
disengagement or a reduction in SAS effec-
tiveness. Disengagements may result from
failures of any of the following systems or
components: SAS gyro or electronics cir-
cuitry, flight control servos, or electrical
power supply. Disengagement or loss of
effectiveness may occur as a result of com-
plete or partial loss of A or B System hy-
draulic power. Disengagement of any
channel is indicated by illumination of the
master caution light, the SAS CHANNEL
OUT light on the annunciator panel, and
one or more of the recycle indicator lights
on the SAS control panel.
When a malfunction is indicated in any SAS
axis, initiate the following preliminary
actions:
a. A & B hydraulic system pressures -
Check normal. If hydraulic system
failure is indicated, follow A and/ ,
or B Hydraulic System Failure pro-
cedure, this section.
b. INVERTER OUT Warning Lights -
Check.
If illuminated, use Inverter Failure
procedure, this section.
c. Proceed to appropriate Pitch and
Yaw axis or Roll Axis Failure pro-
cedures, this section.
A single failure or sequence of failures in
the pitch and yaw axes which leaves one A
or B channel operating in each of these axes
does not change the aircraft flight charac-
teristics. However, some undesirable
cross-coupling in the pitch and yaw axes
may result from failure of one roll channel.
Characteristics which change as a result of
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SECTION III TA-12
failures affecting both the A and B channel
servos in an axis are described as second
condition failures with the appropriate pro-
cedures. Refer to the SAS Warning Lights
charts which illustrate the probable causes
of failure indications, remaining capabilites,
procedures, and limits which apply after
channel disengagement.
Pitch and Yaw Axis Failures
A "first" condition failure exists after at-
tempts to extinguish one or more recyle
lights are ineffective and either an A or B
channel is operating (light Off) in each of
the pitch and yaw axes. A "first" condition
failure exists with a single A, B, or M
channel light illuminated or in some cases
after simultaneous or progressive illumi-
ation of two or more of these lights, as il-
lustrated by the SAS Warning Lights Chart.
Note
Consider that no failure exists
when all pitch and yaw recycle
lights have been extinguished, re-
gardless of previous combinations
of illumination, if normal operation
of the recycle lights is verified by
depressing the SAS Lights Test
button.
Flight may be continued without restriction
when a first condition failure exists except
that maximum airspeed is limited to 350
KEAS in the case of combined channel fail-
ures due to low hydraulic system pressure.
A "second" condition failure is defined as
existing whenever the A and B recycle lights
in one axis remain illuminated after attempts
to extinguish them are ineffective. When a
"second" condition failure exists, flight
speed is restricted to Mach 2.8 and 350
KEAS. Transfer fuel as required to obtain
either 2o nose up trim or 3000 pounds in
tank 1.
Note
Each instance of recycle light illum-
ination presents a new situation and,
if the light(s) can not be extinguished,
the condition must be determined as
being a "first" or "second" condition
of failure in accordance with the
definitions provided above.
Logic override procedures are usable after
a "second" condition failure when the se-
quence of light illumination indicates that
a channel with operative servos is available.
Refer to After Second Failures, SAS Warn-
ing Lights Chart. When use of logic over-
ride is effective, flight characteristics are
the same as with SAS fully operational.
However as a precaution against subsequent
hardover failure signals, the autopilot must
not be engaged in that channel and second
condition failure limits apply.
WARNING
If logic override is recommended,
use it only in the channels specified
and only after decelerating to sec-
ond condition failure limit speeds
in order to prevent excessive
structural loads which could result
from a hardover failure at higher
speeds.
Neither logic override nor BUPD
operation should be attempted with
either channel known to have a
failed servo.
BUPD plus logic override procedures are
available after a "second" condition failure
in the pitch axis. The BUPD is optimized
for operation at air refueling speeds, and
It should not be operated above 330 KEAS
or 0.85 Mach. It may or may not improve
flight characteristics at other flight con-
ditions.
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SECTION III
SAS FAILURE WARNING LIGHTS CHART
PITCH OR YAW RECYCLE LIGHTS ON
INDICATIONS AFTER FIRST FAILURE
PITCH A
OR
YAW
BO
Light( s 1 on
1 t 0 2nd M
0
0
0
0
0
00
0
0
0
0
0
0
0
0
0
C.)
0
0
0
0
0
0
0
SEQUENCE
OF
ILLUMINATION
FIRST
A
servo
B
servo
M
gyro
A
gyro
B
gyro
A
servo
B
servo
M
gyro
M
gyro
SECOND
,
,'N
M or A
gyro
M or B
gyro
A
servo
B
servo
CHANNELS
REMAINING
OPERABLE
B
A
A
and
B
B
A
B
A
B
A
ACTION: First try to
press light(s) off
No further action when first failure lights stay on
then I* A or B light Is off
LIMITS
NONE
INDICATIONS AFTER SECOND FAILURE
P ITCH A
OR
YAW
B
0 Ist 0 2nd M
0
0
0
0
0
0
0
s 0
0
0
0
0
0
0
0
0
0
0
0
0
0
SEQUENCE
OF
ILLUMINATION
FIRST
M
gym
A
gym
B
gym
A
servo
B
servo
9?
servo
servo
SECOND
A or B
gyro
B or M
gyro or
B servo
A or M
gyro or
A servo
B
gyro
A
Wm
B
servo
A
servo
FUNCTIONS
OPERABLE
A or B
Channel
A servo's
possibly
B
channel
B servo
possibly
A
channel
B
servo
A
servo
NONE
ACTION
First try then 10
to press
lights off
If A and B lights
stay on
Note: Use of
Logic Override is
not mandatory
pitch
Try
A
or yaw:
Override
B
I A
If pitch
Try BUPD
plus
override
- -
NO ACTION
or
B
Unless subsonic
then
BUPD
B A
I
first
plus
pitch override
A
I B
If Yaw
No Action
UNUSABLE
pitch,
_ or
yaw SAS
To use pitch or yaw Logic Override:
A and B Channels off. Select A or B override.
Beep Channel switch ON.
To use BUPD:
A and B channels OFF. BUPD - ON
Select A or B Override. Beep one Channel on.
Channel off If no improvement.
Do not use
Logic Override
or
BUPD
LIMITS
Mach 2.8 and 350 KEAS maximum.
Fuel transfer is necessary for 2� noseup trim up to 4000 lb.
With override - No autopilot that axis
With BUPD - Mach 0.85 and 330 KEAS
F2O1-40(2)
Figure 3-4 (Sheet 1 of 2)
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SAS FAILURE WARNING LIGHTS CHART
COMBINATIONS OF PITCH, ROLL AND YAW DISENGAGE LIGHTS
INVERTER OUT AND A OR B HYDRAULIC LOW
WARNING LIGHTS ON
INDICATION
ELECTRICAL FAILURE
HYDRAULIC FAILURE
P ITCH YAW
A A
ROLL
B B
M M
00
0
00
a
00
8
00
� i e
00
0
00
o of
� �
�
� �
� �
��
0
00
00
00
0
��
00
No lights
on but
operation
poor
CHECK
INV 3
INV 1
INV 2
ANY TWO
A System
B System
BOTH
A and B
ACTION
1 Check circuit
breakers
a Inv 3 GB-
SAS pitch-
yaw mon
b Ess DC bus-
SAS M
2 Inverter Switch
3 Press recycle lights
4 Recycle roll channel
if light is on
5 Do not use logic
1 Check circuit
breakers
a Inv 1 GB-
SAS yaw A
b Ess DC bus-
SAS A
- EMER
off
switch
override
1 Check circuit
breakers
a Inv 2 GB-
SAS yaw B
b Ess DC bus-
SAS B
NOTE:
INV 2 load
cannot be
xfrtl to
EMER
1 Check circuit
breakers
a Inv 1,2,3
b Ess DC bus-
SAS A,B,M
NOTE:
M Channels will
be inoperative
Channel off if pressure is low
With normal pressure:
1 Cycle roll channel switch
2 Press recycle lights off
NOTE: Any combination of A, B, and/or
roll lights may occur
0 Lights may illuminate simultaneously
or progressively
LIMITS
NONE
2nd Failure 350 KEAS maximum
Figure 3-4 (Sheet 2 of 2)
If logic override procedures are not effec-
tive or possible after a second condition
failure in the yaw axis, tests at high Mach
numbers indicate that neutral to slightly
positive stability exists but that there is
little damping of yaw oscillations after they
commence.
1. Illuminated recycle light(s) - Depress
and release.
If light( s) stays on or reilluminates, no further
action is required unless a second condition
failure exists.
2. Channel switch - OFF.
If another failure should occur in the same
axis:
3. Illuminated recycle lights(s) -Depress
and release.
F201 -40(1)(c)
4. If lights do not extinguish with second
condition failure - Comply with limits.
If SAS lights indicate a good servo is avail-
able:
5. A or B logic override - Engage as indi-
cated by servo availability.
a. Pitch or yaw logic override switch-
A or B position depending on fail-
ure analysis.
Note
Refer to SAS Warning Light
Chart.
b. Appropriate A or B channel switch-
Beep ON.
Recycle light should extinguish.
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TA-12 SECTION III
c. If control does not improve -
Channel switch - OFF.
d. Logic override switch - OFF.
For pitch axis second condition failure;
when speed is below 330 KEAS and 0.85
Mach:
6. BUPD - Engage as required.
a. Pitch SAS A and B channel switches-
OFF.
b. BUPD switch - ON.
c. Pitch logic override switch - A or
B position as indicated by servo
availability.
d. Appropriate A or B pitch SAS
channel - Beep ON. Recycle light
should extinguish.
e. If control does not improve -
Channel switch OFF.
f. Logic override switch - OFF.
g. BUPD switch - OFF.
h. Depending on failure analysis this
procedure may be repeated using
other SAS channel if indicated.
Roll Axis Failures
Illumination of the roll channel disengage
light shows that both roll channels and the
roll autopilot are disengaged. When there
is no apparent fault in the hydraulic sys-
tems or electrical power supply which
would cause disengagement, check for a
transient disengagement as follows:
1. A or B channel switch - OFF, then ON.
A transient or intermittent fault existed if
the light remains off. If the light does not
extinguish, or reilluminates while man-
euvering, a first condition failure exists in
the roll mode.
For a first failure:
2. A and B channel switches - OFF.
Unless the failure can be associated with a
specific hydraulic or electrical power sup-
ply, regain the use of one channel by the
following arbitrary step sequence:
3. A Channel switch - ON.
Note
. Be prepared to move the switch to
OFF immediately if a hardover
signal results, indicating that the
failed channel was inadvertently
selected.
. Operation with only one roll chan-
nel engaged results in overriding
of logic circuitry. There is no
automatic protection against in-
advertent selection of a failed
channel, or against subsequent
failure of a properly operating
channel which has been engaged.
If a hard-over signal is obtained on engage-
ment or during subsequent operation, or if
no improvement is noted in flight character-
istics:
4. A Channel switch - OFF.
5. B Channel switch - ON.
Note
Be prepared to disengage the
channel immediately if a hard-
over signal results.
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TA-12
If a hard-over signal is obtained on engage-
ment or during subsequent operation, or if
no improvement is noted in flight character-
istics, a dual or second condition failure
exists.
For a second condition failure:
6. Roll channel switches - Both OFF.
Some undesirable cross-coupling may occur
during single roll SAS channel operation.
This appears as small amplitude oscillations
in the pitch and yaw axes, as the elevons on
only one side of the aircraft respond to roll
signals during single channel operation and
compensation for the asymmetric roll sig-
nals is provided by pitch and yaw axis con-
trol operation.
Scheduled activity may be continued for the
remainder of the flight with a single roll
SAS channel operating. The roll autopilot
may be engaged and the automatic navigation
feature of the INS used as desired.
Notes
Operation with both roll channels
disengaged is permitted if cross-
coupling about the pitch and/or
yaw axes prevents precise air-
craft control with one roll channel
engaged.
In the event of single engine failure
at low speed, or during single
engine landing, failure of one roll
SAS channel and simultaneous
automatic disengagement of the
other roll channel may occur due
to loss of hydraulic power from
the windmilling engine.
To avoid changes in control characteristics
at a critical time during single engine land-
ings, either make the approach with both
roll SAS channels disengaged or with the
roll channel which is powered by the inop-
erative engine disengaged.
A second roll SAS channel failure while at
high speed will probably be indicated by ab-
normal pitch transients and small roll trans-
ients without illumination of either pitch or
roll SAS indicator lights. The symptoms
may be difficult to attribute to roll channel
failure. When pitch transients occur with
one roll channel engaged, disengage both
roll SAS channels and check for control im-
provement. If no improvement is noted,
the single roll channel may be reengaged
if desired.
Failure or intentional disengagement of both
roll SAS channels is expected to increase
pilot fatigue, reduce mission effectiveness,
and will disable the roll autopilot; however,
no hazard to safety should result and there
are no flight restrictions on continued oper-
ation.
TRIM FAILURES
Pitch, yaw or roll trim malfunctions maybe
of the inoperative type or the runaway type.
Runaway trim failures in pitch may occur
at slow speed (0.15 /sec change in elevon
deflection) if due to autopilot/Mach trim
motor operation or at fast speed if due to
manual trim motor operation (1.5 /sec
change in elevon deflection). A low speed
runaway type of malfunction will be apparent
by the need for constant manual pitch trim-
ming. The runaway yaw trim rate if ap-
proximately 1.5o per second trim change.
The roll trim rate is approximately 1� /sec.
Runaway yaw trim will be accompanied by
rudder pedal deflections as the surfaces
move. Runaway pitch or roll trim will not
be accompanied by stick movement due to
surface movement.
In the event trim runaway failure is suspected,
proceed as follows:
1. TRIM POWER SWITCH - OFF.
If circumstances permit:
Z. Reduce speed to below 350 KEAS and
2.5 Mach.
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SECTION III
With runaway nose up pitch trim:
3. Transfer fuel forward to reduce forward
stick force requirement.
WARNING
Do not transfer fuel if nose down
pitch trim has occurred.
When initial speed is above Mach 2, decreas-
ing Mach normally requires increasing nose
up pitch trim.
When time and conditions permit:
4. Autopilot - ON. Check for control im-
provement.
5. Affected trim circuit breakers - Pull.
Note
Both A & C phase circuit breakers
must be pulled on the suspected
circuit.
Trim Malfunctions:
a. If runaway slow speed pitch trim
Pull auto pitch trim circuit
breakers.
b. If runaway high speed pitch trim
Pull manual pitch trim circuit
breakers.
c. If inoperative manual pitch trim
Pull the Mach trim dc circuit
breaker.
Note
If Mach trim dc circuit breaker is
pulled, the normal Mach trim
speed stability augmentation in the
transonic region will -..tY.; be avail-
able.
d. If runaway roll or yaw trim - Pull
roll or yaw circuit breakers.
6. Trim power switch - ON.
With manual pitch trim inoperative and
auto trim available, engagement of the
pitch autopilot will gradually correct an out
of pitch trim condition. This will relieve
the pilot of a need for maintaining stick de-
flection to maintain attitude. The pitch
autopilot can also be used when the auto
trim motor is inoperative, but automatic
pitch trim synchronization will not be avail-
able.
CAUTION
Disengagement of the pitch auto-
pilot when not in trim may be ac-
companied by a considerable
transient.
If the trim malfunction is a runaway in the
roll axis, right or left stick deflection will
be required for the rest of the flight but
stick force will not be more than normally
required for the same amount of deflection.
If the malfunction was a runaway in the
yaw axis, rudder pedal force will be re-
quired to maintain neutral rudder pedal
position.
LANDING GEAR SYSTEM EMERGENCY
OPERATION
RETRACTION
There is no emergency system for retracting
gear in flight; the gear lever is the only con-
trol retracting the gear. If the gear lever
cannot be moved to the UP position after
takeoff, do the following:
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1. Ground retract button - Depress and
hold.
2. Landing gear lever - UP.
This procedure overrides a solenoid, nor-
mally actuated by the landing gear switch,
and permits the landing gear lever to be
moved.
EXTENSION
The manual landing gear release handle un-
locks the landing gear uplocks and allows
the landing gear to fall free to the down-
and-locked position. If L or R hydraulic
system pressure is available the landing
gear lever must be placed in the DOWN
position or the landing gear CONT circuit
breaker must be pulled to permit emergency
extension. Approximately 90 seconds is
required for emergency gear extension. The
manual landing gear release handle must be
pulled approximately 9 inches for actuation
of all gear uplocks. If it is not pulled all
the way out one or more gear may fail to
extend. If the L hydraulic system has failed,
or normal gear extension is unsuccessful,
proceed as follows:
1. Landing gear lever - DOWN.
2. Manual landing gear release handle -
Pull.
3. Verify gear down and locked.
If landing gear remains retracted:
5. Aft cockpit landing gear switch - DOWN.
If gear still retracted:
6. Landing gear CONT circuit breaker -
Pull.
7. Repeat steps as necessary.
Note
When the landing gear CONT cir-
cuit breaker is pulled nosewheel
steering will be inoperative.
WHEEL BRAKE SYSTEM FAILURE
Without antiskid brakes operating, proper
braking technique is required to prevent a
skid. A skid is hard to detect in this air-
craft because of its size, weight, and land-
ing gear geometry. At high speed a skid
will usually blow the tires before corrective
action can be taken. Proper braking tech-
nique is achieved by applying a steady, con-
stantly increasing pedal pressure as air-
craft speed decreases.
BRAKE SYSTEM EMERGENCY OPERATION
If normal braking is not effective, or L hy-
draulic pressure is not available and R hy-
draulic pressure is, proceed as follows:
1. Brake switch - ALT STEER & BRAKE.
If both engines are shut down during ground
roll, the brake switch should be left in the
ANTISKID or NORM position and steady
pedal pressure applied until the aircraft
comes to a complete stop.
AIR DATA COMPUTER FAILURE
If malfunction or failure of the air data
computer (ADC) is suspected, proceed as
follows:
1. Cross-check TDI instrument against
pitot- static -operated air speed and
altimeter.
If cross-check shows TDI to be inaccurate:
2. Revert to use of pitot-static-operated
instruments for aircraft control.
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SECTION III
3. Mach trim - Pull MACH TRIM circuit
breaker.
4. Autopilot - OFF.
PITOT-STATIC SYSTEM FAILURE
Under some conditions the two normal ADC
and pitot-static operated systems may be-
come inaccurate or inoperative from a com-
mon malfunction. Failure of the pitot heater
may simultaneously affect both normal sys-
tems in icing conditions. The pitot probe
could also be plugged by a foreign body. If
both normal systems fail, the pilot should
proceed as follows:
1. Attempt to restore operation by se-
lecting alternate source.
2. Maintain aircraft control by use of atti-
tude and power indicating instruments.
3. Request escort aircraft for letdown and
landing.
AIR CONDITIONING AND PRESSURIZATION
FAI LURES
LEFT ENGINE OR FORWARD COCKPIT SYSTEM
INOPERATIVE
At any time the left engine is shut down:
1. Forward cockpit system switch -
CROSSOVER.
FORWARD COCKPIT AND VENTILATED SUIT
OVERTEMPERATURE
1. Defog switch - OFF.
2. Cockpit temperature indicator - Check.
If temperature indication is too high:
3. Forward cockpit auto temperature
rheostat - Rotate toward COLD.
Note
The hot and cold valves are motor-
operated and travel from full hot
to full cold in approximately 7 to
13 seconds.
If auto temperature control is not effective
and forward cockpit temperature remains
too high:
4. Forward cockpit temperature control
switch - Hold in manual COLD.
Note
In this position the motor-driven
valves take 12 to 24 seconds to
travel from full hot to full cold.
If no decrease in temperature occurs in
30 seconds:
5. Forward cockpit system switch -
CROSSOVER.
WARNING
Aft cockpit system switch must be
ON.
If suit temperature cannot be controlled by
the above steps:
6. Suit flow valves - OFF.
7. Reduce altitude and speed.
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AFT COCKPIT OVERTEMPERATURE
If the aft cockpit temperature indication is
too high, proceed as follows:
1. Aft cockpit auto temperature rheostat -
Rotate toward COLD.
Note
The above step should be accom-
plished in increments as there will
be a lag in the temperature indi-
cation.
If auto temperature control is not effective
and aft cockpit temperature remains too
high:
2. Aft cockpit temperature control switch-
Hold on manual COLD
Note
The manual cold valve will take
from 12 to 24 seconds to travel to
FULL COLD. �
COCKPIT DEPRESSURIZATION
Cockpit depressurization above approxi-
mately 35,000 feet will be indicated by pres-
sure suit inflation. If suit inflates, proceed
as follows:
1. Cockpit altitude - Check.
2. Canopy seal levers - Check ON.
3. Cockpit pressure dump switch - Check
OFF.
WARNING
During this time, the pilots will
be depending on the pressure suit
only for altitude protection.
If cockpits still do not repressurize:
4. Suit ventilation boost lever - EMER.
5. Descend as soon as possible.
ABBREVIATED CHECKLIST
The emergency abbreviated checklist is
furnished separately.
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TA-12
SECTION IV
AUXILIARY EQUIPMENT
TABLE OF CONTENTS
Control Transfer Panels
4-2
Inertial Navigation System
4-16
Communications and Navigation
Periscope
4-31
Equipment
4-4
Rear View Periscope
4-36
Transponder (1FF)
4-11
Lighting Equipment
4-36
Interphone Control Panels
4-13
Flight Recorder
4-37
Flight Reference System
4-14
Autopilot System
4-37
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CONTROL TRANSFER PANELS
Control transfer panels, located on the left
console in each cockpit, allow either pilot
to assume control of the TACAN and UHF
equipment, the fuel boost pumps, the fuel
quantity gage, and the air-conditioning sys-
tem. (See figure 4-1.)
FORWARD COCKPIT CONTROL TRANSFER PANEL
The forward cockpit control transfer panel
has seven switches and seven transfer lights.
The ADF switch and transfer light is inop-
erative.
Communications Equipment Switches and Transfer
Lights
Two 2-position toggle switches, labeled
TACAN/INSTR, UHF, and two transfer
lights, located on the upper left portion of
the panel, permits the pilot to assume con-
trol of the TACAN, flight instruments, and
UHF equipment. Control is obtained by
moving the respective toggle switch fore or
aft. Control transfer is made when the
transfer light illuminates.
Fuel Switches and Transfer Lights
Two 2-position toggle switches, labeled
FUEL CONT and FUEL QTY, and two trans-
fer lights, located on the lower left portion
of the panel, allow the pilot to assume
manual control of the fuel boost pumps and
also to obtain readings on the fuel quantity
gage. Control of the fuel boost pump panel
and the fuel quantity gage is obtained by
moving the respective switch either fore or
aft. Control transfer is made when the
transfer light illuminates.
Air-Conditioning Switches and Transfer Lights
Two 2-position toggle switches, labeled
AIR COND CONT TRANS, and
lights, labeled FWD and AFT,
on the right side of the panel.
labeled FWD CKPT and NORM
control switch. The unlabeled
control transfer switch.
two transfer
are located
The switch
is the mode
switch is the
When the mode control switch is in the
FWD CKPT position one cockpit has control
over the air-conditioning of both cockpits
and one transfer light (FWD or AFT) will
be illuminated, indicating which cockpit has
control. Moving the control transfer switch
in either cockpit to the alternate position
will transfer control to the other cockpit.
When the mode control switch is in the
NORM position either the aft cockpit has
control of air-conditioning for both cockpits
(AFT light on) or each cockpit has control
over its own air-conditioning (both lights
out). Moving the control transfer switch to
the alternate position reverses these
functions.
The FWD light will not light in either cock-
pit if the mode control switch is in the
NORM position.
The following chart summarizes the func-
tion of the cockpit air-conditioning system:
Transfer
Lights
� FWD
O AFT
Control Function
-� �._ _��
Fwd ckpt has control of air-
cond of both ckpts. Mode con-
trol sw in FWD CKPT position.
O FWD
� AFT
Aft ckpt has control of both
ckpts. Mode control sw in
FWD CKPT or NORM position.
O FWD
O AFT
Aft and fwd ckpts have control
of their own air-cond. Mode
control sw in NORM position.
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SECTION IV
CONTROL TRANSFER PANELS
FORWARD COCKPIT
1 TRANSFER LIGHT
2 ADF CONTROL TRANSFER SWITCH
3 UHF CONTROL TRANSFER SWITCH
4 AIR CONDITIONING CONTROL TRANSFER SWITCH
5 AIR CONDITIONING MODE CONTROL SWITCH
6 FUEL QUANTITY GAGE CONTROL TRANSFER SWITCH
7 FUEL BOOST PUMP CONTROL TRANSFER SWITCH
8 TACAN AND FLIGHT INSTRUMENT CONTROL
TRANSFER SWITCH
T IRAN
TACAN INSTR DF UHF
CONT TRANS
FUEL CONT FUEL QTY
6
NT TRAN
TACAN/INSTR OF HF
CONT TRANS
FUEL CONT FUEL QTY
4
F201 -35(a)
Figure 41
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SECTION IV.
TA-12
AFT COCKPIT CONTROL TRANSFER PANEL
The aft cockpit control transfer panel is
identical in appearance and function with the
forward control transfer panel except that
there is no air-conditioning mode control
switch. The ADF switch and transfer light
are inoperative.
COMMUNICATIONS AND NAVIGATION
EQUIPMENT
UHF COMMAND RADIO, AN/ARC-51
The AN/ARC-51 UHF command radio pro-
vides two-way communications on 1750 dif-
ferent frequencies extending from 225.0
through 399.9 megacycles. Any of these
frequencies may be selected manually; how-
ever, the radio is preset on the ground to
the 18 frequencies most commonly used
during normal operation. In addition to the
main receiver, the set utilizes a second
guard receiver which can cover a frequency
range between 238.0 and 248.0 megacycles,
but which is normally pretuned to 243.0
megacycles. Power for the set is furnished
by the essential dc bus. Refer to Control
Transfer Panels, this section, for further
information.
AN/ARC-51 Control Panels
A control panel is installed on the left con-
sole in each cockpit. The panels contain a
function switch, rotary channel selector
switch, volume control and four manual
tuning knobs.
Channel Selector Switch
A rotary channel selector switch labeled
CHAN permits selection of any one of 18
preset channels, the guard (G) frequency
channel or the manually (M) set frequency
channel.
Function Switch
The function switch has four positions la-
beled OFF, T/R (transmit-receive), T/R
+ G (transmit-receive + guard) and ADF
(inoperative). In the T/R position both the
receiver and transmitter are tuned to the
preset or manually selected channel. When
the switch is in the T/R + G position, the
radio will receive signals simultaneously
from the main and guard channels of the
receiver.
Frequency Selector Knobs
Four frequency selector knobs permit man-
ual selection of any of the 1750 frequencies
for transmit-receive operation. The man-
ual frequency windows indicate a direct
reading in megacycles and tenths of a mega-
cycle.
Volume Control
Audio level may be increased by rotating
the volume (VOL) control knob clockwise.
UHF Antenna
L hydraulic system pressure extends and
retracts the UHF antenna. The antenna will
extend from its stowed position in the right
chine when the function selector switch is
moved to an operating position. The antenna
will retract when the function selector switch
is moved to OFF. The antenna is spring-
loaded and extends if L hydraulic pressure
is lost.
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SECTION IV
UHF COMMAND RADIO CONTROL PANEL (Both Cockpits)
1 CHANNEL SELECTOR SWITCH
2 MANUAL FREQUENCY SELECTOR KNOBS
3 FUNCTION SWITCH
4 VOLUME CONTROL
F201-36 (a)
Figure 4-2
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TA-12
UHF Command Radio Operating Procedures
1. Obtain control of set with the UHF
switch on the control transfer panel and
check transfer light illuminated.
2. Function switch - As desired.
3. Channel selector switch - As desired.
4. To select a frequency other than one of
the preset channels:
a. Channel selector switch - M.
b. Manual tuning knobs - Position to
set desired frequency. A digital
readout of the selected frequency
will be shown in the windows at the
top of the control panel.
5. To transmit as well as receive on
guard channel:
a. Channel selector switch - G.
b. Function switch - T/13 or T/R + G.
ADF RECEIVER
The ADF radio receiver is an automatic or
manual direction finder and a low and broad-
cast range aural receiver. The equipment
consists of a radio receiver, control panel
(aft cockpit only), flush sense antenna, flush
fixed loop antenna, and HSI. The receiver
covers a frequency range of 0.19 to 1.75
megacycles in three bands. Power for the
equipment is furnished by the essential dc
bus and the No. 1 26V instrument trans-
former.
ADF Control Panel
The ADP control panel is installed on the
right console of the aft cockpit. The controls
are described below.
Function Switch
; -
The function switch is the larger of the two
concentric knobs on the inboard side of the
panel. The labeled positions are OFF,
ADF, ANT, and LOOP. In the ADF position
the equipment functions as an automatic di-
rection finder with a continuous indication
of the bearing to the radio station, shown on
the HSI. In this position also, the sense and
loop antennas are connected to the receiver.
In the ANT position, received signals are
obtained only from the sense antenna, and
the equipment functions as a conventional
aural radio receiver. In the LOOP position
received signals are obtained only from the
loop antenna and the equipment functions as
a manual direction finder to enable the pilot
to determine the bearing to the radio station
by aural null procedures.
Band Selector Switch
The band selector switch is the larger of
the concentric knobs in the outboard side
of the control panel and is used to select
the desired frequency band. The correct
frequency scale will also appear in the
frequency indicator window for the band se-
lected as follows:
Band Frequency
190 - 400 KC
Coverage
FAA low frequency
band
400- - 840 KC
International distress
frequency and lower
standard broadcast
band
840 - 1750 KC
Upper standard broad-
cast band
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SECTION IV
ADF CONTROL PANEL (Aft Cockpit Only)
AFT COCKPIT
(ONLY)
LOOP CONTROL
2 BFO SWITCH
3 FREQUENCY INDICATOR WINDOW
4 TUNE-FOR-MAX INDICATOR
5 BAND SELECTOR SWITCH
6 TUNING CONTROL
7 GAIN CONTROL
8 FUNCTION SWITCH
F201-38(b)
Figure 4-3
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SECTION IV TA-12
Tuning Control
The tuning control is the smaller of the out-
board concentric knobs and tunes the re-
ceiver within the frequency band selected.
The tuned frequency is indicated on the scale
of the frequency indicator. The control is
also rotated slightly for maximum reading
on the tuning indicator.
Loop Control
The control labeled LOOP is used to ac-
complish the electrical equivalent of ro-
tating the loop antenna. The control is la-
beled L and R and the left or right rotation
effect will be apparent in the headset and
the tuning indicator. The speed of the ro-
tating effect may be slowed by turning the
loop control approximately half way to the
L or R labeled position.
Gain Control
The gain control is the smaller of the in-
board concentric knobs and is provided to
adjust the receiver audio level.
BFO Switch
The BFO switch provides a beat frequency
oscillation to aid in tuning the receiver or
to receive coded transmissions.
Operating the ADF Receiver as a Conventional
Radio Receiver
1. Function switch - ANT.
2. Band selector switch - Select desired
band.
3. Tuning control - Rotate to desired fre-
quency.
4. Volume - Adjust as desired.
5. The BFO switch can be used to tune in
continuous-wave signals or to zero-
beat modulated signals.
Operating the ADF Receiver as an Automatic
Direction Finder
1. Tune receiver as above and positively
identify the station.
2. Function switch - ADF.
3. Tuning control - Tune for maximum
reading on tuning meter.
4. HSI bearing select switch - ADF.
5. Read bearing to station on HSI bearing
marker.
Operating the ADF Receiver as a Manual
Direction Finder (Aural Null)
1. Tune receiver as above and positively
identify the station.
2. Tuning control - Tune for maximum
reading on tuning meter.
3. Function switch - LOOP.
4. Loop control - Turn to R or L as
necessary to acquire null.
5. HSI bearing select switch - ADF.
6. Read bearing to station on HSI bearing
marker.
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TA-12 SECTION IV
FLIGHT INSTRUMENT CONTROL PANEL
MODE SELECT BEARING
NA
MAG
JACAN SELECT
TACAN
FLIG'HT INSTRUMENT CONTROL PANEL
A flight instrument control panel is centrally
located at the bottom of the instrument panel
in each cockpit. There are three selector
switches on each panel, labeled MODE SE-
LECT, BEARING SELECT and ATT/AP
SELECT. The flight instruments in both
cockpits will have identical indications at
all times and only one cockpit has control.
Control is transferred from one cockpit to
the other by using the TACAN/INSTR control
transfer switch on the left console in either
cockpit.
Mode Select Switch
The MODE SELECT switch affects the HSI
indication only. This is a rotary switch
with three positions; NAY, MAG, and
TACAN. When the switch is in NAY or MAG
position the heading marker indicates the
command steering course from the NAY
system, the course arrow is servoed to the
lubber line and the course deviation bar is
4411110p
AFT COCKPIT
Figure 4-4
centered. When the switch is in TACAN
position the heading marker is manually
set, the course arrow is manually set to
the desired tacan course and the course de-
viation bar is operative. When this switch
is in NAY position, the compass card indi-
cates true heading, and in MAG or TACAN
position the compass card indicates mag-
netic heading.
Bearing Select Switch
F26-76
The bearing select switch has two positions
labeled TACAN and ADF, and is used to
select the source for the HSI bearing pointer
indication, regardless of mode select
switch position.
ATT/AP Select Switch
The ATT/AP select switch has two positions
labeled FRS and INS and is used to select
the reference source for the attitude indi-
cator and the autopilot. The INS position
must be selected for AUTO-NAY operation.
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SECTION IV
TA-12
TACAN CONTROL PANELS
TACAN SYSTEM, AN/ARN -52
AFT CKPT
F201-74
Figure 4-5
The Tacan System provides continuous in-
dications of bearing and slant distance to a
selected surface beacon and range only to
another aircraft containing the necessary
transponder equipment. The system trans-
mits interrogation pulses which trigger re-
sponding pulses from the selected ground
station or aircraft. Slant distance to the
station or aircraft is computed from the
elapsed time. Both bearing and distance
are visually displayed on the Horizontal
Situation Indicator which is located on each
instrument panel. The system is capable of
operation on any one of 126 channels and has
a range of about 300 nautical miles. The
transmitting frequency range is 1025 to 1150
megacycles. Frequency ranges for recep-
tion are; low band normal, 926-1024 mega-
cycles, air to air 1088-1150 megacycles,
high band normal, 1151-1213 megacycles,
air to air 1025-1087 megacycles. Power
for the set is furnished by the essential dc
bus.
ANARN-52 Control Panel
A control panel is installed on the right
console in each cockpit. The panel contains
a channel selector switch, mode selector
switch and a volume control. The cockpit
having control is determined by use of the
control transfer switches.
Channel Selector Switch
A channel selector is used to select any one
of 126 available channels. Channel selec-
tion is accomplished by setting the desired
channel number in the window using the
concentric knobs. The outer knob selects
the first two digits and the inner knob se-
lects the third digit of a desired channel.
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SECTION IV
Volume Control Knob
Audio level of the Tacan station identifica-
tion signals is increased by rotating the
volume (VOL) control clockwise.
Mode Selector Switch
The function selector switch has four posi-
tions.
OFF - The set is de-energized.
BEG - The set is energized and presents
bearing and course information on the HSI.
T/R - Same as the REC position and also
presents range in nautical miles to a Tacan
station.
A/A - Presents slant range only in nautical
miles to another cooperating AN/ARN-52.
Operation of the Tacan System
1. Obtain tacan control on the control
transfer panel.
2. Display MODE SELECT switch - TACAN.
3. BEARING SELECT switch - TACAN.
4. TACAN mode selector switch - REC.
(Allow 90 seconds for warmup.)
5. Channel selector switch - Desired
channel.
6. Verify station identification.
7. Observe bearing pointer and to-from
indicator on HSI.
8. Tacan mode selector switch - T/R or
A/A.
9. Observe range to station or aircraft on
HSI.
TRANSPONDER (IFF) - 914-X-1
The 914-X-1 transponder provides recep-
tion, detection, decoding, encoding and
transmission of signals in the IFF Mark X
(SIF) system and has a locally installed
MODE X discrete operating function. The
transponder will also recognize a Mode 4
interrogation; however, the set will not de-
code or encode a reply without accessory
equipment. Any one of numerous coded re-
plies available for Modes 1, Mode 3 or X
can be selected by rotating the appropriate
selector switches on the panel. The set is
capable of transmitting an emergency reply
regardless of the interrogation mode. A
provision is also incorporated to identify
position of the aircraft. Power for the set
is furnished by the essential dc bus. Addi-
tion of the Mode X capability deletes the
Mode 2 function from the transponder. Con-
trols are provided in the forward cockpit
only.
TRANSPONDER (1FF) CONTROL PANEL
The transponder control panel is installed
on the upper left console of the forward
cockpit. The panel contains two code se-
lectors for Mode 1 and Mode 3/X codes,
Mode 1 and Mode 3 toggle switches, an
I/P switch, IFF power selector switch and
an emergency switch bar.
Power Switch
The IFF power switch has three positions:
Off, LO, and ON. When the switch is
placed at LO, only local (strong) interro-
gations are recognized and answered. With
the switch in the ON position, there is full
sensitivity for recognition and reply. The
IFF power switch activates Mode X when in
the ON or LO position. Response to Mode
1 and Mode 3 interrogations is dependent
on the position of the Mode 1 and 3 toggle
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IFF/SIF CONTROL PANEL
FWD COCKPIT
Figure 4-6
switches. When the Emergency switch bar
is up, the power switch is forced to the ON
position. A 30 second time delay is incor-
porated in the power switching before the
equipment is operative.
Mode Switches
Two two-position mode switches, one for
Mode 1 and one for Mode 3, control trans-
mission of Mode 1 and Mode 3 replies.
Correctly coded interrogations will be an-
swered when a mode has been made active
by selecting the IN position. When a Mode
1 or Mode 3 switch is in the OUT position,
that mode is not active and does not trans-
mit upon interrogation except in Emergency.
Mode X is active at all times when the power
switch is in the ON or LO position and is
not affected by the Mode 1 or Mode 3 toggle
switch position.
Code Selectors
F201-13(b)
Two rotating type code selectors are pro-
vided. The code selector for Mode 1, con-
sists of two rotary digital/indicating switches.
The first digit window will indicate 0, 1, 2,
3, 4, 5, 6, or 7. The second digit window
will indicate 0, 1, 2, or 3. The Mode 3/X
code selector will indicate 0, 1, 2, 3, 4, 5,
6, or 7 for each digital window. The mode
3 code selection also controls the Mode X
code transmission.
Emergency Switch Bar
The emergency switch bar, when placed in
the EMERGENCY up position, operates two
toggle switches that controls emergency
response and also pushes the IFF power
switch to the ON position if it is in the off
or LO position. When the emergency bar
is in the up position an emergency indicat-
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SECTION IV
ing pulse group (code 7700) is transmitted
on Mode X each time an interrogation is
made on Mode X. Mode 1 and 3 are also
turned on by the emergency bar irrespec-
tive of the position of the Mode 1 and 3 In-
Out switches. In the EMERGENCY position
Mode 1 will respond on the code selected
but Mode 3 will respond on code 7700 irre-
spective of code selected.
Note
The ground radar scope indication
from this transponder is coded in
a different manner than the normal
AN/APX-46 transponder.
Identification of Position (I/P) Switch
The identification-of-position (I/P) switch
is used to control transmission of 1/P pulse
groups. The switch has three positions;
MIC, OUT and a spring-loaded T/P position.
When the switch is momentarily in the 1/P
position, the 1/P timer is energized for 30
seconds. If an interrogation is recognized
on any active mode within this 30 second
period, 1/P replies will be made. When the
switch is in the OUT position, transmission
of the I/P pulse groups is withheld. The
MIC position is inoperative at present.
OPERATION OF THE IFF SYSTEM
1. Power switch - ON or LO.
2. Emergency bar - Down.
3. Mode 1 and Mode 31N-OUT switches -
As required.
Note
Mode X operation is\ continuous
when the power switch is in the LO
or ON position. For secure IFF
operation, both the Mode 1 and
Mode 3 toggle switches must be in
the OUT position.
4. I/P switch - As required.
5. Code selectors - As required.
To make an emergency response to Mode 1,
Mode 3 and Mode X interrogations;
6. Emergency bar - Push up.
INTERPHONE CONTROL PANELS (AN/AIC-10)
An AN/AIC-10 interphone control panel for
each cockpit is installed on a shelf behind
a lower hatch under the aft cockpit. Each
panel contains a call button and a NORMAL-
AUX LISTEN switch. No ON-OFF switch is
provided and the equipment is operative
whenever the essential dc bus is energized.
A remote volume control is located on the
right console in each cockpit.
Call Button
The call button is inoperative.
Normal-Aux Listen Switch
The NOR MAL-AUX LISTEN switch has two
positions, NORMAL and AUX-LISTEN. The
normal position allows all audio signals to
pass through the AN/A1C-10 amplifier. Se-
lecting the AUX LISTEN position bypasses
the amplifier, and audio intensity must be
adjusted with the individual equipment
volume control. The switch is safety-wired
in the NORMAL position.
Transmitter-Interphone Control Switch
A momentary contact, center-off slide
switch on the control stick grip permits the
microphone circuit to be connected to the
UHF transmitter (TRANS position, up) or to
the interphone circuit (1NPH position, down).
An interphone' jackbox, connected to the
common interphone circuit, is mounted in
the load center bay for ground crew use.
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Throttle Microphone Button
A microphone button is provided on the in-
board throttle in each cockpit for use during
taxi, takeoff, and landing when the pilot's
left hand must be on the throttles. These
are pushbutton switches which must be held
for radio transmission.
FLIGHT REFERENCE SYSTEM
The flight reference system supplies infor-
mation for indication and control of aircraft
heading and attitude. The system consists
of a flight reference platform, induction
compass transmitter, heading and attitude
couplers, a control panel in each cockpit
and the rotating compass card of each HSI.
The two modes of operation, magnetic slaved
mode and directional gyro mode, provide ac-
curate directional reference for all latitudes.
The directional gyro mode is the more re-
liable at latitudes near the magnetic poles,
since the magnetic slaved mode is subject
to severe magnetic distortion near the poles.
When in the magnetic slaved mode, the sys-
tem is basically a gyro stabilized compass
slaved to the induction compass transmitter.
This mode provides magnetic heading with-
out northerly turning error or oscillations.
The directional gyro mode may be used at
all latitudes, but is most useful when the
magnetic field is weak or distorted or when
navigating in the polar regions. When in
the directional gyro mode, the system is
free of magnetic influence and operates as
a directional gyro, indicating an arbitrary
gyro heading as selected by the pilot. In
directional gyro mode, with the proper lati-
tude selection made on the control panel,
the gryro is made to precess the correct
amount required to overcome gyro drift at
the selected latitude. In either mode, head-
ing information is furnished to the autopilot
and HSI provided that the ATT/AP SELECT
switch is in the FRS position, and the display
MODE SELECT switch is in the MAC posi-
tion.
MANUAL FAST SLAVING
Before Takeoff
The normal slaving rate of the system is
about 1 1/2o per minute. When the compass
system is energized before takeoff, the gyro
may be as much as 180 from the proper
heading. About 1 1/2 hours would be re-
quired to slave to the correct heading at
normal slaving rates. Manual fast slaving
is provided by actuating the set heading�
switch, which increases the rate to 720
per minute. This corrects a 180o error
in 15 seconds.
In-Flight
Normally, if the compass is properly slaved
before takeoff, no in-flight manual fast
slaving is required unless free directional
gyro operation is selected. When operating
as a free gyro, the desired heading can be
established by using the set heading switch.
Note
The autopilot must be turned OFF
during manual fast slaving when
the FRS is being used as a heading
reference.
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FLIGHT REFERENCE SYSTEM (FRS) CONTROL PANEL
F20.1-73
Figure 4-7
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FRS CONTROL PANELS
A control panel is installed on the right
console in each cockpit. The panel con-
tains a function selector switch, set heading
switch, latitude selector knob, synchroni-
zation indicator, malfunction indicator and
hemisphere selector switch, and take com-
mand button.
Latitude Selector Knob and Indicator
The latitude selector knob may be rotated
to select and display the desired latitude in
degrees and tenths of degrees in the indi-
cator window. The knob is operable in the
DG mode only and selects the latitude in
which the airplane is operating. When in
DG operation, with the operating latitude
selected, the directional gyro will be cor-
rected for apparent drift due to the earth's
rotation.
Note
The proper corrections will be
made only if the hemisphere selec-
tor switch is indicating the correct
hemisphere.
Take Command Button
A combination button and light on the con-
sole panel provides for transfer of control
by depressing the button and observing.the
light.
Malfunction Indicator
The malfunction indicator monitors the
power supply plus other prime system func-
tions. Any deviation from normal operation
that would cause the system to render er-
roneous information will cause the indicator
to display 3 white triangles.
Hemisphere Selector Switch
The hemisphere selector switch is used to
select the hemisphere in which the aircraft
is operating.
FRS SYSTEM OPERATION
1. ATT/AP switch - FRS.
2. Hemisphere selector switch - As
required.
3. Functions selector switch - As
desired.
4. Latitude selector knob - Set to proper
latitude when operating on free gyro.
5. Set heading switch - Fast slave com-
pass card of HSI to proper heading.
6. Synchronization indicator - Center
needle when operating as a magneti-
cally slaved system.
INERTIAL NAVIGATION SYSTEM (INS)
The inertial navigation system is self-
contained and operates in all modes without
the use of electromagnetic radiation or ex-
ternal references. The system consists of
a gyro-stabilized platform, platform elec-
tronics, coupler and power supply, repeater
and converter power supply, control panels,
and distance-to-go, groundspeed, and a di-
rection indicator.
In operation the system displays present
position, groundspeed and the direction and
distance to go to any of 42 preselected posi-
tions as continuous readouts. When operated
in autopilot AUTO NAV, and INS STORED
AUTO mode, the aircraft will be steered
automatically to each point in the flight plan
sequentially, with no pilot action required.
If the flight plan is being flown in sequence
in the STORED AUTO mode, the destination
select light will illuminate if the destination
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TA-12 SECTION IV
displayed on the destination select panel does
not agree with the destination towards which
the aircraft is flying. This light is extin-
guished when the pilot sets the selector panel
to the number of the stored destination being
approached.
The destination select panel provides se-
lection of destination numbered 0 through
41. The first 27 preselected positions are
assigned to preplanned mission destinations,
fix points, targets, rendezvous points, or
other points occurring sequentially during
the mission. The computer computes and
stores the great-circle courses between
each pair of these numerical points, and the
aircraft will adhere to these great circle
courses. Turns from one course to another
will be made with bank angle optimized
with a maximum bank of 30 degrees) for
the groundspeed and heading change required.
The heading marker of the HSI will point to-
ward the optimum path to follow to place the
aircraft on the next course. If the pilot
switches to a subsequent destination in
STORED MAN before completing the route
segment he is on, the turn will be made in
accordance with computer program direc-
tions.
Positions 27 to 41 provide ADF type steering
for courses to these points and are not meant
to be used in the STORED AUTO mode.
These positions are available for alternate
destinations or may be used to employ an
alternate flight path to a position included
in the first 27. A sufficient number of alter-
nate destinations is available to provide
adequate coverage throughout the mission.
Duplication of any of the first 27 positions
in this group provides a steering indication
on the HSI heading marker, resembling that
of ADF navigation, i. e., the pointer points
directly to the next destination within a 45
degrees needle deflection.
The basic reference of the inertial naviga-
tion system is provided by three single-axis
accelerometers mounted at right angles to
each other on a gyro-stabilized platform.
The platform employs three floated inte-
grating gyros, also mounted at right angles.
The platform is initially aligned with a co-
ordinate reference frame, represented by
a plane tangent to the surface of the earth
and oriented to any convenient azimuth at
the point of origin. The platform stable
element is isolated from the airframe
through a system of three gimbals which
provides 360 degrees freedom of rotation
in yaw and roll, and pitch angles of + 60
degrees. All platform outputs are changed
to digital form before entering the computer.
In normal operation the platform also pro-
vides attitude outputs in analog form through
resolvers and synchros to the autopilot and
the attitude indicator. Conversion of pre-
sent position to latitude and longitude read-
out is accomplished continuously by the
digital computer when in operational mode.
Cooling air, necessary to the system, is
supplied by the aircraft air-conditioning
and pressurization system. A self-contained
heating system is incorporated in the plat-
form to ensure that gyros and precision
sensing components are maintained at tem-
perature within an optimum operating range.
The system is powered by the No. 3 in-
verter, the LH generator, and the monitored
dc bus.
Note
Accuracy of INS information will.
be, slightly degraded if pressure
altitude data supplied by the air
data computer is lost or is in-
accurate.
The INS is controlled from two control
panels, the navigation panel and the des-
tination select panel. See figure 4-
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TA-12
NAVIGATION CONTROL PANEL
The navigation control panel, located on the
right console, consists of a DEST/FIX se-
lector switch, STORE pushbutton, MODE
selector switch, FIX ADS knob, two sets of
geographic coordinate digital readout win-
dows, labeled PRESENT POSITION and
DEST/FIX POSITION, a VARIABLE INPUT
indicator labeled LAT and LONG, with
thumb-wheels for manual insertion of geo-
graphic coordinates and a switch for selec-
tion of N or S latitude. The controls and
indicators are as follows:
Mode Selector Switch
The INS MODE selector switch is a rotary
switch with five positions and is labeled
OFF, RST, ALGN, NAY and FRS. The FRS
position is inoperative.
Note
During flight the INS MODE se-
lector switch must not be switched
to any position other than NAV,
otherwise the INS will be deacti-
vated and will not function until
the switch is moved through OFF,
RST, and ALGN positions in con-
junction with the ground operating
equipment and normal INS pre-
flight procedure.
CAUTION
Do not move the INS MODE se-
lector switch from the OFF posi-
tion in flight, if the INS has not
been cycled from OFF to the NAY
mode prior to flight, otherwise,
the INS system will be damaged.
RST Mode
The RST (reset) mode is used only on the
ground during INS preflight when the plat-
form has reached operating temperature.
It permits the GOE operator to check cor-
rect power switchover from ground to
aircraft power, start the gyro spin motors,
and make the computer ready for use.
ALGN Mode
The INS must be completely warmed up,
stabilized, and aligned to a coordinate re-
ference frame before it can be operated.
This is necessary to minimize the drift of
the stable reference platform once it is
aligned to the coordinate reference frame.
The complete warmup and alignment pro-
cedure at normal ambient conditions takes
about 1 and 1/2 hours. During this period
the destination loading operation is accom-
plished, normally by use of a punched tape.
However, the coordinates of the present
location and 42 destinations or targets may
be set in manually by the VARIABLE INPUT
thumbwheels and N-S selector and entered
Into the computer memory by pushing the
STORE or DEST FIX pushbutton for each
position. After a period of gyro stabili-
zation, the platform is torqued to the co-
ordinate reference frame and the gyros are
drift-trimmed. The two transverse hori-
zontal accelerometers are used to sense
the local vertical and their outputs are used
in the servo loops that torque the platform
and measure the amount of gyro drift. The
presence of output signals from each ac-
celerometer indicates that the platform is
not level in that axis. While level align-
ment of the platform is being accomplished
automatically, platform azimuth is aligned
with a selected reference which is trans-
ferred to the platform by the ground op-
erator. The platform is drift-trimmed at
the reference points thus established, and
the drift reduced to certain preestablished
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TA-12 SECTION IV
INS PANEL AND INDICATORS
FWD COCKPIT
HORIZONTAL SITUATION
IND ICATOR
VIEW A
RIGHT CONSOLE PANEL
DISTANCE TO GOIGROUND
SPEED INDICATOR
DESTINATION
SELECT PANEL
F201-33(c)
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SECTION IV TA-12
rates before the system can be operated.
The MODE selector switch has a detent be-
tween NAV and ALGN positions and cannot
be moved either way between these two posi-
tions until it is first depressed.
NAV Mode
Switching to the NAV mode permits the GOE
to be disconnected, and places the platform
in the operational mode. The gyros are es-
sentially memory devices that memorize
the coordinate frame established. The sys-
tem operates using these memorized co-
ordinates to perform the navigation problem,
and the accelerometers measure translations
of the platform caused by movement of the
aircraft. The accelerometer outputs are
integrated once to provide velocity on each
axis, and a second time to establish their
displacement from the point of origin. These
displacements (distances flown) are trans-
lated into geographical position coordinates
by the computer. In addition to indicating
position coordinates to the pilot, this posi-
tion is also used to torque the platform to
the local vertical and azimuth as the air-
craft changes position. The coordinate
frame thus rotates about the earth to main-
tain its orientation on a plane tangent to the
surface of the earth at the position of the
aircraft.
FRS Mode
The FRS mode position is inoperative. The
reference source for the HSI rotating com-
pass card is selected with the display
MODE SELECT switch on the flight instru-
ment control panel.
WARNING
If the INS should fail, the DISPLAY
MODE selector switch should be
moved to the MAG or TACAN
mode without delay in order to
retain a heading indication on
the HSI display.
DEST/FIX Switch
The DEST/FIX switch is a five-position ro-
tary selector switch with positions as fol-
lows:
STORED
AUTO, FIX, MAN
VARIABLE
FIX, DEST
STORED AUTO. The INS will automatically
sequence consecutively through the 42 pre-
stored destinations as each is reached when
the switch is in the STORED AUTO position.
STORED FIX. To use a prestored des-
tination as a fix point, the switch is set to
the STORED FIX position, the destination
select panel is set to the desired destination
number, and the STORE or DEST FIX push-
button is depressed when the fix point
crosses the horizontal line on the periscope
screen.
STORED MAN. To select any of the 42 pre-
stored coordinate positions as a destination,
out of the automatic consecutive sequence,
the switch is set to the STORED MAN
(manual) position, the destination select
panel is set to the desired destination num-
ber, and the STORE or DEST FIX pushbutton
is depressed.
VARIABLE FIX. To use a variable (un-
stored) fix point as a point of reference,
the switch is set to the VARIABLE FIX
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TA-12 SECTION IV
position, the VARIABLE INPUT thumbwheels
are set to the desired coordinates, and the
STORE or DEST FIX pushbutton is depressed
when the fix point crosses the horizontal line
on the periscope screen.
VARIABLE DEST. To select a variable
(unstored) destination, the switch is set to
the VARIABLE DEST position, the VARI-
ABLE INPUT thumbwheels and N-S selectors
are set to the desired coordinates, and the
STORE or DEST FIX pushbutton is de-
pressed.
FIX ADJ Knob
The fix-adjust knob, labeled FIX ADJ, con-
trols a flight cursor on the periscope and
is used to update the INS by means of vi-
sual fixes on known coordinate points. It
is not necessary to fly directly over the fix
point to obtain useful data. Viewing the fix
point on the screen, the pilot positions the
cursor with the FIX ADJ knob to coincide
with the fix point as it crosses the hori-
zontal reference line on the display. Refer
to discussion of fix-taking for further in-
formation.
STORE Pushbutton
The STORE pushbutton is used to store in
the computer memory, either selected des-
tination information or position information
which has been selected by the VARIABLE
INPUT thumbwheels and N-S selector. It
also initiates the computations required to
navigate to the coordinates selected.
Note
The DEST/FIX pushbutton on the
destination select panel is identical
in function to the STORE button on
the navigation panel. They may be
used interchangeably. Do not push
either button unless a course
change or fix is desired.
N-S Hemisphere Selector
The N-S selector switch may be placed in
either the N or S position, depending upon
which hemisphere the desired destination
or fix is located. This switch is used only
in conjunction with the variable input thumb-
wheels to manually insert a destination or
fix point in flight.
VARIABLE INPUT Indicator
The VARIABLE INPUT indicator has thumb-
wheels that are used to manually insert any
desired reference coordinates in to the sys-
tem, thus giving the pilot added flexibility
of operation in flight It is good practice to
put the DEST/FIX switch in the VARIABLE
DEST or VARIABLE FIX position prior to
setting the coordinates in the indicator. To
insert variable destination coordinates into
the system, select VARIABLE DEST with
the DEST/FIX switch, then insert the de-
sired destination coordinates with the VAR-
IABLE INPUT thumbwheels; select desired
hemisphere with the N-S selector and de-
press the STORE or DEST FIX pushbutton.
The DEST/FIX POSITION indicator will read
out the new coordinates immediately after
the STORE or DEST FIX button is depressed,
and the INS will navigate the aircraft to the
new destination using ADF type steering.
Variable update fix is inserted in the com-
puter in the same way as a destination, ex-
cept that VARIABLE FIX is selected with
the DEST/FIX switch.
PRESENT POSITION Indicator
The PRESENT POSITION indicator is set at
the geographical coordinates of the flight
origin site prior to takeoff. In flight it con-
tinuously indicates the coordinates of the
aircraft position as computed by the INS.
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DEST/FIX POSITION Indicator
The DEST/FIX POSITION indicator normally
displays the latitude and longitude coordi-
nates of the destination to which the INS is
navigating. This display may be the coordi-
nates of any selected destination from the
42 prestored positions or the coordinates of
any selected variable destination. This co-
ordinate display normally changes at such
times as the computer calculates a new
course to a newly selected destination. For
STORED MAN or VARIABLE DEST modes,
this change will occur when the DEST FIX
or the STORE pushbutton is depressed, For
sequential or out of sequence destination
selections in STORED AUTO mode, the des-
tination coordinate display will change co-
incident with roll out to the new destination
course. The minutes counter portion of the
latitude display may also change whenever
a fix is taken. When either a STORED FIX
or VARIABLE FIX is taken, the calculated
correction (in nautical miles) is displayed
on the latitude minutes display, without
changing longitude, or the degrees portion
of latitude on the DEST/FIX POSITION indi-
cator. The portion of the latitude display
used for the fix distance indication is blocked
off in white on the indicator (see Figure 4-9).
The calculated fix correction is displayed
up to a maximum value of 59 nautical miles
whether position is updated or whether the
fix is rejected. The calculated fix correc-
tion will continue to be displayed until
another fix is taken or until a new destin-
ation is selected and displayed. When a
new destination is selected, the latitude
minutes counters will revert to a display
of destination latitude until such time as
another fix is taken.
DESTINATION SELECT PANEL
The destination select panel, labeled NAV,
is located on the instrument panel. The
panel has a two-place digital counter, con:-
trolled by thumbwheels, and a self-illumi-
nated pushbutton switch which reads out
DEST FIX when lighted. The number of a
stored destination or fix (0 through 41) may
be set on the counter manually and inserted
into the INS computer by depressing either
the DEST FIX or the STORE pushbutton
when the DEST FIX switch is in the STORED
MAN or STORED FIX position.
Note
Positions 42 through 49 can be
displayed, but are inoperative.
Except when flying out of sequence in the
STORED AUTO mode the DEST FIX push-
button illuminates when the destination
number on the panel and the destination ap-
proached by the aircraft are not the same.
When they are again the same (thumbwheels
must be rotated), the light will go out. In
all modes the light will come on when pilot
action is required. When the DEST/FIX
switch is placed in either STORED or VARI-
ABLE FIX, the light will come on. When
the STORE or DEST/FIX pushbutton is de-
pressed the light will go out. In any mode,
in which anew destination is selected by
depressing the STORE or DEST/FIX push-
button, the light will go out when the sys-
tem accepts the new destination. In
STORED MAN, the light will come on if a
destination is passed by 15 miles without
selecting a new destination.
DISTANCE-TO-GO AND GROUNDSPEED
INDICATOR
A distance-to-go and groundspeed indicator
is installed on the instrument panel. Di-
gital indicators display the distance be-
tween the aircraft position and the destin-
ation, and the groundspeed, in units of 1
nautical mile and knots, respectively.
When a new destination is selected either
automatically or manually the indicator
will change to show the new distance-to-go.
The distance-to-go indication will decrease
toward zero while approaching the destin-
4-22
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
TA-12 SECTION IV
ation, then increase after passing the des-
tination if flight is continued on the same
course. Distance-to-go will not read zero
at destination if the computed cross-course
distance is greater than 1/2 nautical mile,
since readout resolution is to the nearest
nautical mile.
HORIZONTAL SITUATION INDICATOR (HSI)
The INS computes true heading and steering
information which can be displayed by the
HSI on each instrument panel. The rotating
compass card of each HSI receives the true
heading signals when the controlling display
mode selector switch is in the NAY position.
When the display mode selector switch is in
the MAG or TACAN position the compass
card is driven by the FRS signals to indicate
magnetic heading although the INS system
is still generating true heading information.
The heading marker is driven by the steer-
ing signal from the INS when NAY or MAG
modes are selected, and is manually set
when TACAN mode is selected. The bear-
ing pointer points to either an ADF or
tacan station, whichever is selected with
the controlling BEARING SELECT switch,
regardless of display MODE SELECT switch
position.
Note
The aircraft will automatically fly
the course computed by the INS and
selected by the pilot only if the auto-
pilot is in the AUTO NAY mode.
COURSE SELECTION
The INS is capable of providing steering
information to any selected destination when
the path from source to destination is
greater than 30 nautical miles but less than
21,500 nautical miles (from 1/2 degree to
179 degrees of great circle arc). The se-
quence in which courses are provided de-
pends upon the position of the DEST/FIX
switch on the navigation control panel. In
STORED AUTO position, course directions
will be provided to stored destinations auto-
matically in their numerical sequence;
however, an out of sequence deviation can
be made in STORED AUTO by selecting the
desired out of sequence destination number
on the destination select panel and depress-
ing either the DEST FIX or STORE push-
button. After the out of sequence deviation,
other destinations will then continue to be
automatically selected in numerical se-
quence. In the STORED MAN or VARIABLE
DEST positions, steering directions to in-
dividual destinations are supplied after each
destination is selected by depressing either
the DEST FIX or STORE pushbutton. For
STORED AUTO or STORED MAN modes,
the steering information provided by the
computer is a great circle flight path only
if the destination selected is one of the first
27 sets of stored coordinates (00 through 26).
ADF type steering will be commanded for
STORED destination selections numbered
27 or greater and for all VARIABLE DEST
mode selections. In STORED MAN mode,
the computed course starting point is de-
termined as follows:
a. The position of the current desti-
nation is selected by the computer
as the starting point for the new
course if the aircraft computed
position is within 100 miles of
this point when the STORE button
is depressed.
b. The computed position of the air-
craft is selected by the computer
as the starting point for the new
course if the distance to go is
more than 100 miles from the
current destination.
4-23
Approved for Release: 2017/07/25 C06230172
Approved for Release: 2017/07/25 C06230172
SECTION IV TA-12
INS STEERING CHARACTERISTICS
DISTANCE TO GO FOR START OF TURN-AUTO NOW STEERING
DISTANCE TO GO-NAUTICAL MILES
17
1
150
140
130
120
110
100
90
80
70
60
50
40
30
20
10
0
I +---+--
120�
_4--
--441-CL *UST INATION B
FLIGHT
TRACK
..-�
NOTE
D.T.G.
1
FOR START
I
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I
TURN
TO
GREAT
CIRCLE
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NEW COURSE
GROUND
CHANGE
IS A FUNCTION OF
SPEED AND COURSE
SCHEDULE
DESTINAT
ON A
...--
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NEAR PATH
STEERING
I
VARIABLE
i
110�
\----
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