1. s
F~
rs
------------SECURITY INFORMATION - RESTRICTED------------=-
Fs
AN 01-60JLC-1D
SAFETY OF FLIGHT SUPPLEMENT
FLIGHT HANDBOOK
USAF SERIES
F-86D
AIRCRAFT
This publication supplements AN 01-60JLC-1. Reference to this supplement will be mode on title
page of the basic handbook by personnel responsible for maintaining the publication in current status.
NOTE COMMANDING OFFICERS ARE RESPONSIBLE FOi
HINGING THtS SUPPLEMENT TO THE ATIENTION OF ALL AF
PERSONNEL CLEAHD FOR OPERATION OF SUBJECT AIRCRAFT.
PUBLISHED UNDER AUTHORITY OF THE SECRETARY OF THE AIR FORCE
AND THE CHIEF OF THE BUREAU OF AERONAUTICS
NOTICE: This document contains informatioo affecting the national defense of the United States within the
meaning of the Espionage Laws, Title 18 U.S.C., Sections 793 and 794. Its transmission or the revelation of
its contents in an7 manner to an unauthorized person is prohibited b7 law.
__________.*___________ Fs
1. PURPOSE.
This supplement describes changes to the engine shutdown pro-
cedure for all F-86D aircraft. These changes will tend to reduce the
number of shroud rlng replacements.
2. INSTRUCTIONS.
!· Prior to engine shutdown, advance the throttle to 60 per cent
rpm and manually open the tailpipe variable-area nozzle. Operate
the engine under this condition for approximately 2 minutes.
b. Upon retarding the throttle, place the variable-nozzle switch
in tlie NORMAL posltlon.
AF-WP-0-13 JUL 53 7,200
15 JULY 1953 Fs
F5
1

2. SECURITY INFORMATION - RESTRICTE AN 01-60JLC-1
~LIGHT HANDBOOK
SMAMA-Aug M - 7355 RESTRICTED
10 JANUARY 1953
REVISED 25 JUNE 1953
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3. SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
Reproduction for non-military use of the information or illustrations contained in th is publi-
ca tion is not permitted without specific approval of the issuing service (BuAer or USAF )
The policy for use of Classified Publications is establ ished for the A ir Force in AFR 205-1
and for the Navy in Navy Regulations, Article 1509.
r--------------LIST OF REVISED PAGES ISSUED--------------
INSERT LATEST REVISED PAGES. DESTROY SUPERSEDED PAGES.
A
NOTE : The porrion of the text affected by the current revision is indicated by a vertical line in the outer margins of the page.
Page
No.
Date of Late;t
Revision
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N o.
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* The a5terisk indicates pages revised, added or deleted by the current revision.
ADDITIONAL COPIES OF THIS PUI.LICATION MAY I.E OBTAINED AS FOUOWS:
USAF ACTIVJTJES.- ln accordance with T echnical Order No. 00-5-2.
NAVY ACTIVITIES.- Submit request to nearest supply point listed below, using form NavAer-140: NASO, Philadelphia, Pa.:
NAS. Alameda, Calif.; NAS, Jacksonville, Fla. ; NAS, Norfolk, Va.; NAS, San Diego, Calif.; NAS, Seattle, Wash. : ASD.
NSC. Guam.
For listing o f avai lable material a nd details of distribution see N aval Aeronautics Publications Index NavAer 00· 500.
RESTRICTED Revised ·
25 June 1953
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4. SECURITY INFORMATION - RESTRICTED
AN 01-60JLC-1
1'Jl81E o, co•r••rs
StdNII I osc•TIOI . 1
SectlN H IOllW NOCfDUlfS 47
S.CtlH HI El&GBICY PIOCBIUllES 6S
SectiH IV DESCRIPTION AND OPEIATION
Of AUXILIARY EQUIPMENT 81
SectiH V OPERATING LIIAITA1IONS 107
SectiH VI FLIGHT CHAIAaEIISTICS 113
SectiH VII SYSTEMS OPEIATION 123
SectiH VIII CIEW DUTIES (NOT APPUCABU)
Secti11 IX AU-WEATHER OPERATION 13S
Appeacllx I OPERATING DATA 141
Alphahetkal l1dex 163
itevised 1S April 1953 RESTRICTED i
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5. ii
• •
SECURITY INFORMATION- RESTRICTED
AN 01-60JLC-1
.lust
'
a lllo111ent
hanCJ on to it J
• • •
As a source of technically accurate information, the Air Force and the manufacturer have
joined forces to compile this book in a manner intended to give you data in a form best-suited for
your usage. Descriptions, flight characteristics, techniques, and specific procedures are designed to
enable you to fly this airplane in the safest and most efficient manner. REMEMBER, many accidents
and fatalities could have been prevented had the pilot used his Flight Handbook to its full intent.
The information and instructions are based upon engineering and service data plus observations
of experienced Air Force pilots. This book, prepared in a new style identified by the full-page cover
illustration, is easy to read and amply illustrated to better enable you to assimilate the increased
amount of data necessary to operate the complex airplanes of today. You should READ AND
STUDY this book to gain a sound knowledge of the airplane and always KEEP IT HANDY as a
reference manual co answer specific questions.
Take advantage of the opportunity the Air Force gives you to secure a copy of this handbook
for your own personal use as outlined in T .O. 00-5-2. As frequent revisions and Immediate Atten-
tion Technical Orders keep the information in this handbook up to date throughout the life of the
airplane, it is up to YOU co see that your personal copy is kept current. Prompt distribution of all
applicable publications can be ensured by checking with your Base Inspector to make sure that
your requirements are entered in the Publication Requirement Table (T.O. 00-3-1) and the Auto-
m:itic Distribution List.
This handbook is divided into nine sections, an appendix, and an index as follows:
Section I, DESCRIPTION-a detailed discussion of the airplane and the equipment and systems
(including all emergency equipment which is not part of the auxiliary equipment) which are
essential to flight.
Section II, NORMAL PROCEDURES-opera'ting instructions arranged in proper sequence from
the time the airplane is approached until it is left parked on the ramp.
Section III, EMERGENCY PROCEDURES-concise procedures to be followed in meeting any
emergency that could reasonably be expected (except those of the auxiliary equipment).
Section IV, DESCRIPTION AND OPERATION OF AUXILIARY EQUIPMENT-description of,
and normal and emergency operating instructions for, all equipment not essential for flying
the airplane.
Section V, OPERATING LIMITATIONS-all limitations and restrictions that must be observed
during operation.
Section VI, FLIGHT CHARACTERISTICS-a discussion of flight characteristics, the advantageous
a·s well as the dangerous, that are peculiar to the airplane as based on~extensive flight tests.
Section VII, SYSTEMS OPERATION- a supplementary discussion of special characteristics and
factors involved in operation of some of the airplane systems under various conditions.
Section VIII, CREW DUTIES-omitted as not applicable for a single-place airplane.
RESTRICTED Revised 25 June 1953
SCANNED HV AlTfAI.OGS.COH
I 1

6. SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
Section I
l
-·-----
· - ·
·
··
-·
-----·
DESCRIPTION
THE AIRPLANE.
The North American F-86D is a single-place, all-
weather fighter-interceptor. Principal recognition fea-
tures are its swept-back low wing and empennage and
the radome, similar in appearance to a propeller spin-
ner, on the nose above the air intake duct. The engine
is an axial-flow, turbojet unit equipped with an after-
burner. Other noteworthy features include electronic
engine control, leading edge wing slats, fuselage speed
brakes, and the combination of elevators and horizontal
stabilizer into a single surface known as the controllable
horizontal tail. Designed primarily as a high-altitude
fighter-interceptor, the airplane is equipped with the
latest and most advanced equipment (a large part of
which is electronic) to accurately and effectively launch
the rockets with which the airplane is armed. The
rocket-firing control system computer tracks the target
and tells the pilot the course to follow. This system
requires precision flying in order to operate with a
minimum amount of error, as it will more accurately
compute the true rocket interception course if abrupt
maneuvers, high G, rapid rates of roll or pitch, or steep
angles of bank are avoided. For better utilization of the
full capabilities of the rocket-firing control system and
for prolonged service life of the equipment, smooth,
coordinated precision flying should be practiced. How-
ever, it must be emphasized that whenever evasive ac-
tion is necessary, high-G turns and abrupt maneuvers
section
within the structural limitations of the airplane are
permissible.
AIRPLANE DIMENSIONS.
Over-all dimensions of the airplane are as follows:
Wing span ............................................37.1 feet
Length ....................................................40.3 feet
Height ....................................................15.0 feet
AIRPLANE GROSS WEIGHT.
The normal take-off gross weight of the airplane, in-
cluding full internal fuel and ammunition, is approxi-
mately 18,250 pounds.
ARMAMENT.
Armament consists of twenty-four 2.75-inch, folding-
fin rockets. The rockets are suspended in a package,
located in the under portion of the fuselage, which
extends for firing and then retracts when the firing is
completed.
MAIN DIFFERENCES TABLE.
The main differences between the F-86D Airplane and
the other models of the F-86 Series are shown in figure
1-4.
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7. Section I
2
SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
GENERA& AIIIIANGEMENT
Figure 1-2 (Sheet 1 of 2)
RESTRICTED
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IAtOGS.COH

8. Revised 15 April 1953
SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
Figure 1-2 (Sheet 2 of 2)
RESTRICTED
SCANNEI) HY AlrIAtOGS.COH
Section I
3

9. Section I SECURITY INFORMATION-RESTRICTEO
AN 01-60JLC-1
810CIC NUMBER DESIGIIJl1'1011
... AfSM55 - - -491
.. :Aff0.ffl"'1.-517
... Af50-518 "1,. ..553
... Af-554 "'1. -576,
... Af-704 "'1. -734
... AFSl-2944 tJI,. -3131
... Af51-5857 "1,. -5944
.. . Af51-5945 ~ -6144
... AFSl-6145 d,,,. -6262,
... AFSl..8274 d,,,. -8505
... AF52-3598 dJUI -4304,
Figure 1-3
ENGINE.
The Model ) 47-GE-l 7 turbojet engine is characterized
by an afterburner and a variable-area jet nozzle. The
engine has a rated sea-level static thrust of approxi-
mately 5425 pounds at Military Power and 7350 pounds
at Maximum Power (with afterburner operating). In
engine operation, air enters the intake duct below the
nose of the airplane and passes through an axial-flow
compressor, where it is compressed progressively in 12
stages. The compressed air then flows to the combus-
tion chambers, where atomized fuel is injecced and
combustion occurs. From the combustion chambers, the
hot exhaust gas passes through a turbine and out the tail
pipe in gradually expanding form to provide the high-
velocity jet and reaction thrust. The turbine, which
is rotated by the exhaust gas passing through it, is
directly connected to, and drives, the compressor. When
maximum performance for short periods is desired,
the exhaust gas may be reheated by the injeccion of
fuel into the tail pipe aft of the turbine. The burning
of this injected fuel, providing additional thrust, is
called afterburning. An automatically controlled, vari-
able-area nozzle in the end of the tail pipe provides
correct nozzle conditions for optimum performance,
with or without afterburner operating.
ELECTRONIC ENGINE CONTROL.
The airplane is equipped with an electronic engine
control, the main advantage of which is the extremely
rapid response of the engine to changes in throttle set-
ting. In addition, starting and operating technique is
considerably simplified. Flame-outs and compressor
stalls due to excessively rapid throttle movement are
eliminated because the controls will always respond at
the correct speed, regardless of how fast the throttle is
moved. Control of the engine is tied in with the control
of the variable-area nozzle, thereby automatically main-
taining correct tail-pipe temperature. When the throttle
is moved, an electronically amplified signal energizes
the main and afterburner fuel control valves and the
variable-area nozzle. Refer to Section VII for further
information on the electronic engine control and its
characteristics.
If the electrical inverters do not function
properly when external power is applied for
a start, the electronic power control will be
inoperative, and, a hot start may result. In-
verter warning lights should be checked and
all circuit breakers pushed in when external
power has been applied before an engine start
is attempted. The starting circuit will not be
energized if a-c power is not available from
the inverters.
4 RESTRICTED Revised 25 June 1953
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10. SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
Section I
ENGINE FUEL CONTROL SYSTEM.
Engine fuel control is provided by a main fuel control
valve and an emergency fuel regulator. (See figure
1-5.) Two engine-driven pumps, a booster pump and
a dual pump, supply fuel to the main control valve,
which meters the amount of fuel required by the en-
gine. The engine-driven booster pump provides suf-
ficient fuel boost pressure in case of failure of the tank.-
mounted electrical booster pumps. Excess fuel is by-
passed through the open emergency fuel regulator back
to the booster fuel pump inlet. Metered fuel from the
control valve passes through a throttle-actuated engine
fuel stopcock to the engine combustion chambers. The
setting of the control valve is determined electronically
by throttle position, airspeed, tail-pipe temperature,
engine rpm, altitude, and outside air temperature. The
control valve automatically regulates fuel flow to the
engine during starting; the correct fuel flow is pro-
vided for ignition and then for acceleration to the
flM
speed previously selected by the throttle. In case of
failure of the main control valve, the throttle-controlled
emergency fuel regulator takes over engine fuel control
when the emergency fuel system switch is positioned at
TAKE OFF or ON. The emergency regulator limits the
amount of by-passed fuel in accordance with throttle
setting, airspeed, and altitude. In normal operation, the I
emergency regulator is held in a full by-pass position by
an energized holdout solenoid. The solenoid may be de-
energized by positioning the emergency fuel system
switch to TAKE OFF or ON, enabling the throttle to me-
chanically control the emergency regulator. With the
switch at TAKE OFF (stand-by) during 100% rpm opera-
tion, a drop in fuel pressure causes the emergency regu-
lator to take over control of fuel flow at a pressure
slightly below normal (depending on outside air tem-
perature). The emergency regulator, at a position deter-
mined by the throttle, forces fuel around the main
control valve to the combustion chambers. When con-
trol by the emergency regulator is selected, the holdout
ENGINE J47-GE-17
ENGINE AIR
INTAKE DUCT
CANOPY
SURFACE CONTROLS
ENGINE CONTROL
AUTOMATIC PILOT
ARTIFICIAL FEEL
Revised 25 June 1953
.....
...
NO
NO
BELOW IIAD0Mf
IN NOSE
HINGED AT HAI
COMPLmLY HYDRAULIC
IHEVHSIIIU CONffOL
ELECTIONIC
YES
YES
E-3 FIRE CONTROL•
E-4 FIRE CONTROLt
N-9-1 (STANO-BY)
Figure 1-4
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IAtOGS.COH
NO
-
5

11. Section I SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
6
ALCOHOL
TANK
DEICING
SHUTOFF
VALVE
NOTE:
ELECTRIC
PUMP
TIMER
FROM
FUEL SUPPLY
FUEL FILTER DEICING SYSTEM INSTALLED
ON F-86D-20 AND F-86D-25 AIRPLANES AND
F-86D-30 AIRPLANES AFSl-5945 THRU -6014
DUAL
FUEL PUMP
WARNING LIGHT
ENGINE-DRIVEN 1...,;.._.::,.._..;._;..,::...1
DUAL
FUEL PUMP
MAIN FUEL
CONTROL VALVE
(CLOSES WHEN
ENERGIZED BY
EMERGENCY
SYSTEM SWITCH)
MAIN
ELECTRONIC
AMPLIFIER
OVERSPEED
GOVERNOR
I
I
I
I
I
t PRESSURE
GAGE
..._____
*F-860.-1 THRU F-860.-10, AND F-860-40
AND SUBSEQUENT AIRPLANES
tF-860.-1 THRU F-86D-35 AIRPLANES
EMERGENCY
FUEL SYSTEM
TEST BUTTON
MAIN SYSTEM
BY-PASS VALVE
(REGULATES WHEN
ENERGIZED, OPENS
WHEN NOT ENER-
GIZED)
ENGINE MASTER
SWITCH
,,-------
ENGINE
FUEL
STOPCOCK
'I
I
I
EMERGENCY
FUEL"SYSTEM
SWITCH
ENGINE FUEL
MAIN FUB. PLOW
MAIN FUEL IV.PASS
GOVHNID FUEL IV.PASS
EMBGENCY FUEL FLOW
flLTHDIICEAI.COHOL
EMERGENCY
FUEL SYSTEM
INDICATOR
LIGHT
EMERGENCY..._--1'"-'"--'.......,.
FUEL REGULATOR
(OPENS TO BY-PASS
WHEN ENERGIZED)
.,. ------
*FLOWMETER
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
/
165-48-22F
Figure 1-5 (Sheet 1 of 2)
RESTRICIEO
SCANNEI) HY AlrIAtOGS.COH
Revised 25 June 1953

12. SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
Section I
CONTROL SYSTEM
FROM ENGINE
COMPRESSOR
AFTERBURNER FUEL
AFTERBURNER BY-PASS
PUSSURIZED AIR
EI.ECfflCAL CONNECTIONS
MECHANICAL CONNECTIONS
FROM
FUEL SUPPLY
TURBINE-DRIVEN
AFTERBURNER
FUEL PUMP
TO FUEL
BOOSTER PUMPS
(HIGH-SPEED
OPERATION)
SPRING-LOADED
AFTERBURNER
FUEL SHUTOFF
VALVE
AFTERBURNER
SHUTOFF
SWITCH
AFTERBURNER
ELECTRONIC
AMPLIFIER
ENGINE
• FILTER
AFTERBURNER
FUEL CONTROL
VALVE
COMIIUSTION
NOZZLES
Figure 1-5 (Sheet 2 of 2)
Revised 15 April 1953
165-48-58
solenoid is de-energized, allowing the emergency regu-
lator to take over engine fuel control. Simultaneously,
a main system by-pass valve is opened and the main
fuel control valve is closed to ensure that the emergency
regulator will control fuel flow. A complete loss of
electrical power to the primary bus, or movement of the
battery-starter and generator switches to OFF during
engine operation will also cause the holdout solenoid
to be de-energized. This will allow the emergency fuel
system to take over fuel control, regardless of the posi-
tion of the emergency fuel system switch and without
illumination of the emergency fuel system indicator
light. The emergency regulator is set to deliver nor-
mally a slightly lower fuel pressure than the main fuel
control valve. The emergency regulator does not pro-
vide automatic fuel control in starts, accelerations, and
decelerations; therefore, any rapid engine acceleration
with the emergency regulator operating would supply
too much fuel to the engine, causing flame-out, stall,
or excess temperatures. A mechanical overspeed gov-
ernor is installed downstream of the main control valve
and emergency regulator. This limits maximum fuel
flow to prevent excessive overspeeding of the engine.
THROTTLE ( POWER CONTROL).
The throttle quadrant on the left console is labeled
"CLOSED" and "OPEN." Immediately adjacent to the
throttle, the markings are "START IDLE," "MILI-
TARY," and "AFTERBURNER." There are three stops
on the quadrant: one at the CLOSED position, one at
START IDLE to prevent shutting off the fuel supply inad-
vertently, and one at MILITARY. Outboard movement
of the throttle, which is spring-loaded inboard, allows
the stops to be by-passed. Throttle movement during
engine starting has been simplified; the throttle is
simply moved outboard and forward to the desired
power setting. With the engine master switch at ON,
the initial outboard movement of the throttle actuates
a microswitch that energizes the ignition system; the
subsequent movement of the throttle to START IDLE
opens the engine fuel stopcock. Ignition is automati-
cally cut off by means of a starter cutout relay when
engine speed reaches approximately 25 % rpm. Since
the mixture in the combustion chambers will burn con-
tinuously after once being ignited, ignition is required
only during the starting procedure. When the engine
is running, movement of the throttle in the normal
operating range changes engine rpm by electronically
changing the setting of the main fuel control valve.
The throttle setting determines the position of the
variable-area nozzle which maintains optimum exhaust
temperatures. When the throttle is moved past the stop
at MILITARY, compressor air is supplied for afterburner
fuel pump turbine operation, the afterburner fuel
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13. Section I SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
SPEED BRAKE
SWITCH
ENGINE
CONTROL
LOCKUP
BUTION
,,---M ICROPHONE
BUTION
1'HROffLE GRIP CONJ'R01S
!65-43-15B
Figure 1-6
I
shutoff valve is opened, and the fuel tank booster
pumps begin operating at high speed. In the after-
burner operating range, the afterburner fuel control
valve is controlled by movement of the throttle. After-
burning continues until the throttle is retarded from
the AFTERBURNER range. The throttle grip (figure 1-6)
contains an engine control lockup button (red) marked
"CAGE," a microphone button, and a speed brake
switch.
ENGINE MASTER SWITCH.
An engine master switch is located on the engine con-
trol panel (figure 1-24) on the left instrument subpanel.
Turning the master switch to the ON position opens the
main fuel shutoff valve, starts the fuel booster pumps,
and completes circuits to the throttle-actuated micro-
switch, which provides ignition during starting. The
circuit to the air start switch is also energized. The en-
gine master switch is powered from the primary bus.
EMERGENCY FUEL SYSTEM SWITCH.
A means of selecting the proper engine fuel control is
provided by a switch on the engine control panel. (See
figure 1-24.) The switch has three positions: ON, TAKE
OFF, and NORM and is powered by the primary bus.
The TAKE OFF position is used only during take-off and
for initial climb to safe altitude. When the switch is at
TAKE OFF at 100% rpm, fuel flow to the engine is elec-
tronically controlled by the main control valve, with
the emergency regulator serving as a stand-by. With the
switch in this position, the emergency fuel system indi-
cator light is on to serve as a warning that rapid throttle
movements should not be made.
If it becomes necessary to reduce power, the
emergency fuel system switch should be re-
turned to NORM before throttle is readvanced,
to avoid undesirable power surges, possible
compressor stall, or overtemperature condi-
tion.
If the main fuel system pressure drops while the switch
is at TAKE OFF, the emergency fuel regulator automati-
cally takes over fuel control. However, if the main fuel
control valve again becomes operative, it will serve as
the primary control and the emergency fuel regulator
will return to stand-by. When the emergency fuel
system switch is at NORM, the main control valve con-
trols fuel flow without the emergency regulator serving
as a stand-by. The switch should normally be in this
position for all flight conditions except take-off and
initial climb. When the switch is moved to ON, the
emergency regulator takes over fuel control, and the
main system by-pass valve is opened.
• Accelerations must be made very cautiously
while the emergency fuel system switch is
ON; otherwise, compressor stall or flame-out
will result.
• If loss of electrical power to the primary
bus occurs, or if the battery-starter and gen-
erator switches are moved to OFF during
engine operation, the emergency fuel sys-
tem will take over fuel control, regardless
of the position of the emergency fuel system
switch and without illuminating the emer-
gency fuel system indicator light. Thus, all
subsequent throttle movements must be
made cautiously to prevent compressor stall
or flame-out.
EMERGENCY FUEL SYSTEM TEST BUTTON.
Operation of the emergency fuel system may be checked
by means of a test button to the right of the emergency
fuel system switch on the engine control panel. (See
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14. SECURITY INFORMATION-RESTRICTED
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figure 1-24.) The emergency fuel system switch must
be at TAKE OFF during the test. When the test button is
pr~ssed, completing the circuit from the primary bus,
the main fuel control system is disabled by electrical
closing of the main fuel control valve. A decrease in
I engine rpm to a lower stabilized value (figure 2-4) will
indicate that the emergency fuel regulator has taken
over fuel control. When the test button is released, the
main fuel control valve is allowed to regain fuel con-
trol, and the emergency fuel regulator is again at
stand-by.
EMERGENCY FUEL SYSTEM INDICATOR
LIGHT.
An emergency fuel system indicator light, labeled
"WARNING," is located on the engine control panel.
(See figure 1-24.) The light will come on when the
emergency fuel system switch is moved to ON or TAKE
OFF. Illumination of the light serves as a warning that
the emergency regulator is in control or in stand-by.
In either case, any accelerations must be made very
cautiously. The light is powered from the primary bus.
DUAL FUEL PUMP WARNING LIGHT.
A warning light on the left console (17, figure 1-20;
16, figure 1-21) comes on when either element of the
engine-driven dual fuel pump has failed. Although
failure of one pump element will result in negligible
change in engine performance under most operating
conditions, failure of both elements will cause engine
failure. Consequently, illumination of the warning
light serves as an indication that flight should be dis-
continued as soon as practical. The warning light op-
erates on power from the primary bus when the engine
master switch is ON.
ELECTRONIC ENGINE CONTROL LOCKUP SYSTEM.
An engine control lockup system is installed in this
airplane to provide protection against possible loss of
power during take-off should the electronic engine con-
trol be deprived of a-c power. The lockup system is
used in conjunction with the electronic engine control.
One of two types of engine control lockup is used:
manual lockup, or automatic lockup. With the manual
system, the red button, marked "CAGE," on the throttle
handgrip is depressed by the pilot to manually lock the
~ main and afterburner fuel control valves and the vari-
able afterburner nozzle in position during take-off and
initial climb. When the engine is stabilized for take-off
and the emergency fuel system switch has been moved
to TAKE OFF, a momentary depression of ..he lockup
button will lock the electronic engine controls in that
position. When a safe altitude is reached and before
throttle is retarded, moving the emergency fuel system
switch from the TAKE OFF position will remove the
lockup. In an emergency, manual lockup can also be
removed by rapidly retarding the throttle below the
half-throttle •position. Automatic lockup differs from
manual lockup in that the electronic engine control
system maintains full engine control until an a-c power
supply fa_ilure is actually encountered. Whenever the
a-c power supply to the electronic engine control falls
belo~pproximately 92 volts, the main and afterburner
fuel control valves and the variab
-le afterburner nozzle
are automatically locked in positiqn. This locked power
setting is then maintained until approximately 15 sec-
onds after the a-c power supply rises above approxi-
mately 98 volts. The lockup is then automatically re-
moved and the electronic engine control resumes op-
eration. The 15-second interval following inverter
change-over is required so that the electronic engine
control components can be rewarmed after a main in-
verter failure. If the a-c power supply is not re-
established above approximately 98 volts, the engine
controls will stay in automatic lockup until manually
released. Manual release from automatic lockup is ac-
complished by converting to operation on the emer-
gency fuel system. Power settings should not be changed
when engine controls are in manual or automatic
lockup, since engine overtemperature or surging may
result when lockup is removed. During an engine
start, the automatic lockup system is made ineffective
when the battery-starter switch is momentarily posi-
tioned at STARTER. After the engine is operating, posi-
tioning the emergency fuel system switch to TAKE OFF
will again make the automatic lockup system operable.
It will then remain operable during normal operation
with the emergency fuel system switch at the TAKE
OFF or NORM position. If the engine control lockup
system is manual, a lockup button is installed on the
throttle handgrip; if automatic, no lockup button is
provided. On airplanes equipped with an automatic
lockup system, a lockup indicator light, labeled "ENG.
CONT. LOCKUP," is provided to indicate when en-
gine controls are in lockup.
ENGINE CONTROL LOCKUP BUTTON.
On airplanes with a manual lockup system, a contact
button marked "CAGE" is provided on the grip of the
throttle (figure 1-6) to enable the pilot manually to
lock the main and afterburner fuel control valves and
the variable afterburner nozzle in position during take-
off and initial climb. When the engine is stabilized for
take-off, a momentary depression of the lockup button
will lock the electronic engine control in that position.
The lockup button is powered from the primary bus .
when the emergency fuel system switch is at TAKE OFF
and the throttle is advanced well forward of the mid-
position.
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15. Section I SECURITY INFORMATION-RESTRICTED
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With the electronic engine control locked, it
is possible under certain conditions, after
take-off and during climb, to exceed the 100%
rpm limit. If the 100% rpm limit is exceeded
and is accompanied by an exhaust temperature
below 705°C, increasing the angle of climb
will reduce rpm to within operating limits.
The lockup condition of the electronic engine
control should be removed during climb to
altitude whenever exhaust temperature in-
creases to approximately 705°C.
ENGINE CONTROL LOCKUP INDICATOR
LIGHT.
An amber engine control lockup indicator light (17,
figure 1-18; 16, figure 1-19), labeled "ENG. CONT.
LOCKUP," is provided on airplanes equipped with an
automatic lockup system. The light is powered by the
primary bus. Whenever a-c power loss occurs, causing
the electronic engine controls to be locked in position,
this light will illuminate. This light will remain illumi-
nated until a-c power is restored and the electronic en-
gine controls have been automatically released from
automatic lockup. Consequently, during operation on
the main fuel control system, this light serves as a warn-
ing that power settings should not be changed as long as
the light remains on. If it is necessary to operate on the
emergency fuel system, engine operation on the emer-
gency fuel system will have no effect on the illumination
of the light. The light will remain illuminated as long
as a-c power is not available and the electronic engine
controls remain in lockup.
Switching the emergency fuel system switch
from the ON position to the NORM position
should not be attempted while the lockup in-
dicator light is illuminated.
ENGINE AFTERBURNER SYSTEM.
For maximum power for short periods, fuel discharged
from combustion nozzles is burned in the tail pipe. In
this manner, the engine exhaust gas is reheated and
provides additional thrust. Fuel to the afterburner com-
bustion nozzles is controlled by an afterburner fuel con-
trol valve. (See figure 1-5.) Throttle position, altitude,
airspeed, engine rpm, and tail-pipe temperature elec-
tronically determine the afterburner fuel control valve
setting; excess fuel is by-passed through the valve back
to the engine-driven booster pump inlet. An after-
burner fuel pump, driven by engine compressor air,
supplies fuel to the afterburner fuel control valve.
Afterburning can be obtained by further advancing the
throttle past the MILITARY Stop to the AFTERBURNER
10
range. This action causes an afterburner fuel shutoff
valve and afterburner fuel pump air shutoff valve to be
opened. The fuel shutoff valve allows fuel to flow to
the afterburner control valve, and the fuel pump air
shutoff valve supplies compressor air to the afterburner
fuel pump turbine, which begins operating. At the same
time, the fuel booster and dual pumps in the forward
fuselage tank begin operating at high speed to increase
the rate of fuel flow. As afterburner fuel flow is in-
creased, the variable nozzle is automatically opened to
prevent excessive exhaust temperatures. When the vari-
able nozzle is fully open, afterburner fuel flow is limited
as required to prevent excessive exhaust temperatures,
regardless of throttle position. The afterburner is shut
off when the throttle is retarded from the AFTERBURNER
range.
Extensive use of t/fter/Jurner mt/terially re-
duces flij//t range and operational time.
AFTERBURNER SHUTOFF SWITCH.
In case of failure of electronic engine control, the after-
burner may be shut off by means of an afterburner
shutoff switch on the engine control panel. (See figure
1-24.) When the switch is moved from the guarded
NORMAL position to OFF, the afterburner fuel pump air
shutoff valve closes to stop the afterburner fuel pump,
the afterburner fuel shutoff valve closes, and the fuel
tank booster pumps return to low-speed operation. The
shutoff switch is powered from the primary bus.
VARIABLE-AREA NOZZLE.
Two clamshell-type doors form the variable nozzle at
the end of the tail pipe. The electrically actuated nozzle
is electronically controlled by throttle position and
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16. I
(
l
~:
II
SECURITY INFORMATION-RESTRICTED
AN 01-60JLC-1
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figure 1-7
exhaust temperature to maintain correct nozzle tem-
peratures for optimum performance. The nozzle may
also be manually actuated. When the variable-area
nozzle is full open, its area is approximately 160 percent
that of a fixed nozzle; when variable-area nozzle is fully
closed, its area is 95 percent that of a fixed nozzle. When
automatically controlled, the nozzle remains in a con-
stant, partially open position up to approximately 75%
rpm and then grad~ally closes as rpm is increased up to
Military Power. At Maximum Power (afterburner),
the nozzle opens and closes automatically to prevent
excessive exhaust temperatures. When the nozzle is
fully open, or if it jams, afterburner fuel is automati-
cally limited in order to maintain correct exhaust tem-
perature. The nozzle is held partially open below 75%
rpm for two reasons: to permit rapid engine accelera-
tion, and to maintain reduced thrust at high engine rpm
for landing. In the event of a go-around, an increase
in engine rpm above 75% will cause the nozzle to close,
resulting in the necessary rapid thrust increase. To
provide more rapid accelerations from low rpm, the
nozzle is automatically held in a partially open position
until engine speed is stabilized and is then rapidly and
automatically readjusted to the correct position.
VARIABLE-NOZZLE SWITCH.
For automatic operation of the variable nozzle through
the electronic engine control, a variable-nozzle switch
on the engine control panel (figure 1-24) is placed
in the guarded NORMAL position. When the switch is
turned OFF, automatic control is cut off and, if the
switch is left at OFF during normal operation, engine
performance will be limited. If the automatic controls
fail or are to be overridden, the nozzle may be actuated
by movement of the switch from OFF to either of the
spring-loaded positions, OPEN or CLOSE. The nozzle will
move more slowly when actuated by means of the switch
than when automatically controlled. The switch is pow-
ered from the primary bus. However, when switch is at
NORMAL, a-c power must be available for automatic
operation of the variable nozzle.
While the nozzle is being closed by means of
the switch, exhaust temperature gage must be
carefully watched to prevent excessive exhaust
/~
temperature.
~
~---
· '.s::::.. --:,a=~=~
{!t,""4#
II ex!Hlust temperature of 10()0° Cis fe(Jtl,ed,
s11cl1 overtemperatur, operation rtf(lirts engine
removal and overlJaul.
VARIABLE-NOZZLE POSITION INDICATOR.
A position indicator ( 15, figure 1-20; 15, figure 1-21) for
the variable nozzle is located on the left console, aft
of the throttle on F-86D-1, F-86D-5, and F-86D-10 Air-
planes. On F-86D-15 and subsequent airplanes, the in-
dicator (19, figure 1-19), is located in the lower right
corner of the main instrument panel. The indicator dial
is calibrated in quarters from minimum to maximum
nozzle area. Power for operation of the position indi-
cator is supplied from the primary bus.
IGNITION SYSTEM.
The ignition system, providing ignition through two
spark plugs, functions only during normal ground
starts and windmilling air starts. For normal ground
starts, with the engine master switch ON and the
battery-starter switch in the STARTER position, the in-
itial outboard movement of the throttle (to pass the
START IDLE stop) actuates a microswitch that supplies
d-c power to the ignition vibrator units. This d-c power
is converted by the ignition vibrator unit into high-
voltage alternating current and is fed to the spark plugs
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17. Section I SECURITY INFORMATION-RESTRICTED
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WARNING
WITH AIR-START SWITCH IN
AIRSTART
MOST ELECTRICALLY POWERED EQUIPMENT IS TURNED OFF
EQUIPMENT POWERED BY
SECONDARY BUS
MONITORED NO. 1 BUS
MONITORED NO. 2 BUS
MONITORED NO. 3 BUS
A-C BUS
WHICH INCLUDES
RADIO EQUIPMENT
WING FLAP CONTROLS
LANDING GEAR CONTROLS•
LANDING GEAR INDICATORS•
•f-86D-1 AND F-86D-S AIRPLANES
165 -43-12A
Figure 1-8
in the No. 3 and No. 7 burners. After ignition occurs
and the engine has accelerated to approximately 25%
rpm, the ignition system is de-energized by the action
of the automatic starter drop-out relay. During wind-
milling air starts, the main ignition system is energized
through the AIR START (up) position of the air start
switch and the throttle microswitch as in normal ground
starts. After an air start, it is necessary to return the
air start switch to NORM (down) to de-energize the main
I
ignition system. Afterburner ignition is accomplished
by spontaneous combustion, the temperature of the
gases from the engine being utilized to ignite the fuel
discharged from the afterburner combustion nozzles.
ENGINE MASTER SWITCH.
Refer to ENGINE FUEL CONTROL SYSTEM in this
section.
AIR START SWITCH.
To provide ignition for restarting the engine in flight,
a two-position switch is installed on the engine control
panel. (See figure 1-24.) When the switch is moved
frc:,m the guarded NORM (down) position to the AIR
START (up) position, current is supplied for engine igni-
tion.* On all subsequent airplanes, the throttle also
must be moved outboard to complete the ignition cir-
cuit. At the same time, an inverter for the fuel pressure
gage is energized and all electrical equipment powered
by the secondary bus, monitored bus No. 1, monitored
bus No. 2, monitored bus No. 3, and the a-c bus is auto-
matically turned off to temporarily decrease the load
on the battery. The air start switch is powered from the
primary bus. (See figure 1-11.)
When the air start switch is positioned at
AIR START (up), all controls to the main fuel
control valve are cut off in order to conserve
battery power. Therefore, the emergency fuel
system switch must be ON during air starts.
When the air start switch is returned to the NORM
(down) position, after the start is completed and the
generators are again operating, ignition shuts off and
electrical equipment