Created: 6/15/1968

OCR scan of the original document, errors are possible


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TDC changes norma, operation procedure Tor Dcacen; and Engine Shutdown.



The Abbreviated Checklist will be changed and replacing puges

NOTE : The technical data Information furnished herein in Intended to be used aa INTERIM data .only. It will be replaced and superseded at the time of issue of the next revision to the flight manual.



This TDC transmits revised pages which replace and supersede previou sly furnished pages for the Flight Manual dated Incorporation of previously furnished TDC's providesperformance which includes:

Revised Military Climb performance atRDC Atmosphere).

Revised Normal Climb performance at variousfor6 ARDC and "Mean Tropic" Atmospheres.

Revised Cruise performance at various temperatures.

Revised Cruise Profiles covering: Long Range Cruise* High Altitude Cruiseeilincr Cruise

Additional descriptive and operating information has been incorporated including Emergency forward transfer, updated engine time. EGT limits, additional tire limitsew drag chute deploy limits. The Pilot's Abbreviated Checklist will be revised and issued to conform.


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No. I Tropical Atmosphere Climb Performance Superseded

by TDC 3


No.Climb and Cruise

RDC Atmosphere t, "MEAN TROPIC" Atmosphere)

No. 4A Limits l> EGT

No.Cruise Plight


No.and Cruise

No.Chate Deploy

No.Printed Change Dated

No. 10 Deployment to

No.Operation for Descent tt Engine

No. 12 Climb Performance

No. 13 Printed Change Dated



DC No.58



This TDC transmits revised pages which supersede previously furnished pages for the Flight Manual dated All previously issued TDC'a are incorporated. In addition, this TDC includes:

Deployment to ARCP data.

presentation of normal climb performance

presentation of cruise performance for long rangealtitudeRDC and "MEAN

single engine descent data for various speeds,for6 ARDC and "Mean Tropic" atmospheres.

descriptive material.

Previously issued checklist changes conform with procedures supplied In this manual.




n Normal Emergency Procedures




SystemsIX All Weather

Appendix: Performance





table of contents



And Afterburner

Gear System

Inlet Syifam

Steering System

Supply System

Brake System

Refueling System

Chute System

Power Supply System

Conditioning and

Power Supply System

System .

Control System

Systems and Personal

Flight Control System

Augmentation System

Static System

Dota Computer




elta wing, single placepowered by two axial flow bleed bypass turbojet engines with afterburners. The aircraft Is built by the Lockheed-California Company and is designed to operate at very high altitudes and at high supersonic speeds. Some notable features of the aircraft are very thin delta wings, twin canted rudders mounted on the top of the engine nacelles,ronounced fuselage chine extending from the nose to the leading edge of the wing. The propulsion system uees movable spikes to vary inlet geometry. The surface control! are elevons and rudders, operated byactuators with artificial pilotfeel. lngle-polnt pressure refueling system is installed for ground and in-flight refueling. rag chute is providedoll.

The overall aircraft dimensionsollows:


Length ft.

Height {to topstabilizer)

Tread {MLG

aircraft gross weight

The ramp gross weighto of these aircraft may vary fromb.b.0 gallons of fuel. This is based an zero fuel weights0 lb.0uel density5 lb. per gallon, and varying equipment loading configurations.




3 15

12 14 16 17




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DE-ic1nc WARNING LIGHT MASTER caution ucmt ltikCTfR

pijiscopi viiwinc

OTNiVl' inut

IUa OtRIOWIN' warning 31


vertical spud indicatorinlet

elapsed riMi

warning lights

triplelay indicator igniter purge SfllTDi



(urgency run sMuron n





28 29



ins asT and stuer panel







trim swhcmes







cutout BUTTON







See the weight and balanceT.O.or information regarding specific aircraft and equipment configurations.


Thrust Is supplied by two Pratt and Whitneyleed bypass turbojet engines with afterburners. The Interim maximum afterburning static thrust rating of each engine0 pounds at sea level with standard day conditions. The engines are designed for continuous maximum thrustat the Mgh compressor Inletassociated with high Mach number and high altitude operation. There is no time limit on maximum thrust operation. The engineingle rotor, nineressure ratio compressorompressor bleed bypass cycle for high Mach number operation. The bypassbleeds air from the fourth stage of the compressor, and six external tubes duct the air around the rear stages of thesection aad the turbine. The air reentors the turbine exhaust ahead of the afterburner and Is used for increased thrust augmentation. When the engine goes into bypass operation, the afterburner fuel control resets to furnish additional fuel to the afterburner. The transition to bypass operation Is scheduled by the main fuel controlunction of compressor Inlet temperature (CIT) and engine speed. The transition normally occursIT of150 , correspondingach number range.

Engine speed on the ground, or at low Mach numbers, varies with throttle movement from IDLEosition slightly below MILITARY thrust. Between this throttle

position and the maximum afterburning thrust position the main fuel controlengine speedunction of CIT and modulates the variable area exhaust nozzle to maintain approximately constant rpm. Throttle movement lit the afterburning range varies the afterburner fuel flow,position and thrust. At high Machand constant inlet conditions, the engine speed Is essentially constant for all throttle positions down to and Including IDLE. ixed throttle position, the engine speed will vary according to schedule when Machand CIT change.

The enginewo stage turbine. discharge air cools the first stage and Is then returned to the exhaust gas stream. Turbine discharge temperatures are monitored by indications of exhaust gas temperatures. hemical Ignition system Is used lo Ignite the low vapor pressure fuel. eparate engine driven hydraulic system, using fuel as hydraulic fluid,tlie exhaust nozzle, chemical ignition system dump, compressor bypass and starting bleed systems. The main fuel pump, engine hydraulic pump andare driven by the main engine The afterburner fuel pump la powered by an air turbine driven by compressorair. The ac generator, aircraft hydraulic pumps and fuel circulating pump are locatedemote gearbox driven by the engine power takeoff pad through agearbox.


The engine thrust ratings are defined by the power lever position at the main fuel control. The power lever Is mechanically linked to the throttle,elationshipthrottle position and thrust ratings.



Maximum Rotcd Thrust

Maximum raced thrust ie obtained inby placing the throttle against the quadrant forward stop.

Minimum Afterburning Thrust

MINIMUM afterburning thrust iswith the throttle just forward of the quadrant MILITARY thrust detent. Ignition is automatically actuated when the throttle Is advanced past the detent and afterburner fuel flow Is terminated when the throttle is retarded aft of the detent. Afterburning fuel flow and thrust areby moving the throttle between the detent and the quadrant forward stop. afterburning isf maximum afterburning thrust at sea level andt high altitude. The basic engine operates at MILITARY rated thrust during all afterburning operation.

Militory Roied Thrust

MILITARY rated thrust is the maximum non-afterburning thrust and is obtained by placing the throttle at the MILITARY thrust detent on the quadrant.


IDLEhrottle position for minimum thrust operation. It Is not an engine rating. Minimum thrust is always obtained when the throttle is at the IDLE stop on the


There is no distinct throttle position for starting. Starting is accomplished bythe throttle from OFF to the IDLEas the proper engine speed Is reached. This directs fuel to the engine burners by actuation of the windmill bypass valve and actuates the chemical ignition system.


The aft stop on the quadrant is the engine OFF throttle position. In this position, the windmill bypass valve cuts off fuel to the burners and bypasses it back to the aircraft system. This provides engine oil, fuel pump and fuel hydraulic pump cooling when an engine Is windmilllng at high Mach number.


Engine fuel system components Include the engine driven fuel pump, main fuel control, windmill bypass valve and variable area fuel noazles in tlie main burner section.

Main Fuel Pu-np

The engine driven main fuel pumpwo stage unit. The first stage consistsingle centrifugal pump actingoost stage. The second stage consists of two parallel g'ear type pumps with discharge check valves. The parallel pump and check valve arrangement permits one pump toin the event the other fails. The pump discharge pressure is determined by theand metering function of the main fuel control. The maximum dischargeIssi. elief valve Is provided in the second stageto prevent excessive fuel system pressure.

Main Fuel Control

The main fuel control meters main burner fuel flow, controls the bleed bypass and

start blend valves and controls exhaust nozzle modulation. Thrust is re (tula toda function o( throttle position, compressor inlet air temperature, main burner pressure and engine speed. The bypass and start bleed valve positions are controlled unction of engine speed biased by CIT. For steady atatc inlet conditions at high Mach number. Ihe control providesonstant engine speed at all throttledown to and Including IDLE. On the ground and at lower Mach numbers, engine speed varies with throttle position from slightly below MILITARY down to IDLE. Afterburner operation is always atrated engine speed and EGT. The fuel control is providedilottrimmer for EOT regulation. There is no emergency fuel control system.

Windmill Bypati and Dump Volve

The windmill bypass and dump valve directs fuel to the engine burners for normalor bypasses fuel to the recirculation system for accessory, engine component and engine oil cooling during windmllling operation. The valve is actuated byfrom the main fuel control. The valve also opens to dram the fuel manifold when the engine is shut down.

Fuel Nonlei

The engine has eight can-annular typechambers with forty-eight variable area dual orifice fuel nozzles In clusters of six nozzles per burner. The nozzlesixed area primary metering orificeariable area secondary metering orifice, dischargingommon opening. The secondary orifice opensunction oforifice pressure drop.


The derlcrhinent system provide" protection against severe turbine over-temperature during high altitude operation. When EGTr more with the system armed, the fuel:alr ratio In the engine burner cans is reduced, or derlched, below normal values. This is accomplishedolenoid operated valve and orifice which bypasses metered engine fuel from the fuel oil cooler to the afterburner fuel pump inlet. The solenoid valve la actuatedignal from ths EGT gages reached. Once actuated) it remains open until the syetem la turned off. Two warning lights are provided to indicate when tha valve Is open. Power for the derich circuits isfrom the essential dc bus.

Fuel DorJchmenr Arming Switch

A two position fuel derichment arminglocated on the left side of the In the ABM (up) position thocircuits are armed and thesolenoid valve will openand remain open If. In the OFF positionsolenoid valve is closedsystem can not provide derichmentis furnished dram the essential

Fuel Detichmonl Wornlng Uflhtt

The fuel derichment warning lights, located on the left and upper center of thepanel. Illuminate and remain on while the derichment solenoid valve Is open. The lights will be extinguished when the arming switchlaced in the OFF position and will remain extinguished when the arming switch La reset to ABM if both EGT* are.



In the event of derichment theewitch muat be placed in the OFF position prior to relighting the afterburner to prevent engine fipced suppression and subsequent Inlet unstart. It engine flameout Is experienced with inlet unstart the arming switch should also be placed to OFF prior to relighting the engine.

Derichment at sea level willhrust loss of approximately i% If In maximum afterburning orf at Military. oss in thrustpm speed suppression will occur during cruiae with maximum afterburning.


Afterburner fuel system components Include the centrifugal afterburner fuel pump,fuel control and spray bars.

Afterburner Fuel Pump

The afterburner fuel pumpigh speed, single stage centrifugal pump. The pump is driven by an air turbine which leby engine compressor discharge air. The compressor discharge air supplyutterfly valve in response to the demand of the afterburner fuel control. The turbine is protected from overspeed by an aero-dynamic speed limiting airvcnturl.

Afterburner Fuel Control

The afterburner fuel controlydro-mechanical fuel control which schedules metered fuel flowunction of throttle position, main burner pressure andinlet temperature. Fuel flow is

meteredredetermined schedule to provide fuel Into both zones of thespray bars simultaneously. The controleset mechanism which increases the afterburner fuel flow when the bypass valves open and decreases the fuel flow when the valves close.


Each engine Is provideduelsystem for actuation of theexhaust nozzle and the start andbleed valves. Engine hydraulicpressure is also required to dump the unused chemical Ignition fluid. Pressure is suppliedigh temperature, engine driven, variable delivery, piston type pump. The pump maintains systemup0 pslaximum flow ofpm for transient requirements. Engine fuel is supplied to the pump from the main fuel pump boost stage. Some high pressure fuel Is diverted from the hydraulic system to cool the non-afterburningline and the windmill bypass valve discharge line. This fuel le returned to the aircraft system. Low pressure fuel from the hydraulic pump case Is returned to the main fuel pump boost stage. Hydraulic system loop cooling Is provided by ths compensating fuel supplied from the main fuel pump.

Exhouit Nozzle Actuation

The exhaust nozzle control and actuation system Is composed of four actuators to move the exhaust nozzle, and an exhaust nozzle control modulating the hydraulic pressure to the actuators in response to engine speed signals from the main fuel control. The exhaust nozzle control Is mounted on the aft portion of theressure regulator Is containedeparate unit located near the exhaust nozzle control.


start bleed and bypass valve actuation

section i

and Bypass Bleed Valve Actuation

The bypass bleed control and actuation system consists of four two-posltlonto move the bleed valves,ilot valve to establish the pressure to the actuators. The pilot valve controls the bleed valve position In response to asignal from the main fuel control. Blend valve position is scheduled within the main fuel controlunction of engine speed and compressor inlet temperature. The starting bleed control and actuationis similar to the bypass bleed system except that three actuators are used and the pilot valve controls starting bleed valve position in response to the main fuel pump boost stage pressure rise.


Tha variable area, iris type, exhaust nossle la comprised of segments operatedam and roller mechanism and four hydraulic actuators. The actuators arc operated by fuel hydraulic system pressure. Thenozzle Is enclosedixed contour, convergent-divergent ejector nozzle to which free floating trailing edge flaps are attached. In flight, the inlet cowl bleed suppliesairflow between the engine andfor cooling. During ground operation, suck In doors in the aft nacelle area provide cooling air. Free floating doors around the nacelle, just forward of the ejector, supply tertiary air to the ejector nozzle at subsonic Mach numbers. The tertiary doors and trailing edge flaps open and close withinternal nozzle pressure, whichunction of Mach number and engine thrust.

proximate percentage of open position. dot markings above and belowndosition marks are for calibration purposes. The indicators are remotelyby electrical transducers located near the exhaust nozzles. Each transducer Is cooled by fuel and is operated by the afterburner nozzle feedback link. Power for the Indicators is supplied by thenverter.


engine and reduction gear box areby an engine contained, "hotlosed systom. The oil is cooled bythrough an engine fuel-oil cooler. The oil tank la mounted on the lower right side of the engine compressor case andsable capacityals. Total tank capacityals. The oil Is gravity fed to the main oil pump which forces the oililter and the fuel-oil cooler. The filter is equippedypass Incase of clogging. From the fuel-oil cooler the oil Is distributed to the engine bearings and gears. Oil screens are Installed at thejets for additional protection. Scavenge pumps return the oil to the tank where it Is deaerated. The main oil pump normally maintains an oil pressure oforessure regulating valve keeps flow and pressure relatively constant at all flight conditions. Because of the highof the oil, It is diluted with trlchloro-ethlene at lower temperatures and special cold weather shut down procedures may be required.

Main Fuel-Oil Cooler

Nozzle Petition Indicator

Each engine Is providedozzleindicator located on the right side of the Instrument panel. The indicators are markedond Indicate the ap-

This unit provides oil cooling by using engine fuel to absorb tha heat. The oil temperature Is controlled by fuel flow through the cooler. ypass valve Isto bypass fuel around the cooler when the fuel flow Is greater than the cooler flow capacity of0 pounds per hour.



purge switch Off

FRO" FUEl hyd pjwp

mainburkr ignition main ignition si gnu DUMP signal dump signal drain

run cooling in

cowpkesso* discharge pressure


ajb ignition line

turbine disch press

ignition signal



Quantity Low light*

An Indicator light lor each engines oilis located on the lower instrumentpanel. The lights areIL QTY LOW and illuminate when the respective engine oil quantity is reduced5 gals, power is furnished by thedc bus.

Engine Oil Te-rperoture Light

LIL TEMP lights are installed on the annunciator panel. These lights will Illuminate when respective engine oil inlet temperature is leas than +or greater than

Re-rote Gear Bo- Oil System

The remote gear box contains anwet sump lubricating system with its own oil supply and pressure pump. There Is no scavenge pump. It Is vented to the engine breather system through the remote gear box drive shaft. The oil Is coaled by circulation through the remote gear box fuel-oil heat exchanger.


Trlethylborane (TEB) Is used for ignition of main burner and afterburner fuel. Special handling procedures are required because TEBill burn spontaneously upon exposure to air above. The TEB is contained4 pint) storage tank pressurized with nitrogen. Theprovide* inartlng and operatingtoetered quantity of TEB to either the main burner or afterburner section. Operation is in responseuel pressure signal from the appropriate Actuation Lb automatic with throttle movement. echanical counter for each engine, located ait of the throttles, indicate TEB shots remaining. inimum of 16

injections can be made with one full tank of TEB. The TEB tank is engine mounted and Is cooled by main burner fuel to maintain the TEB temperature within safe limits. If the TEB vapor pressurealeupture disc Is provided tothe vaporized TEB and tank nitrogen through the afterburner section. N'o pilot Indication of TEB tank discharge is The engine la also equipped with catalytic igniters installed on thelameholders to provide ImprovedIgnition system reliability andcapability. Turbine exhaustmuat be aboveoatisfactory afterburner 'light'* by the catalytic Igniters.

Igniter Purge Switch

A lift-lock toggle switch labeled IGNITER PURGE la installed on the upper right side of the Instrument panel. When the switch Is pulled out and held In the upolenoid operated valve supplies fuelsystem pressure to the chemical ignition system dump valve. Thia allows the remaining TEB to be dumped into the afterburner section; while the engine is running. .Approximatelyeconds Is Electrical power Is furnished by the essential dc bus.


Both electrical power and engine fuel hydraulic preaaure are necessary to purge the TEB If the engine is not rotating the system will not normally dump.


Do not actuate the Igniter Purge switch unless the engine isinpm range to prevent damage to the afterburner flame holders.



A starter cart is provided for ground starts. This may beelf-contained gas engine cart or multiple air turbine cart. The output drive gear of cither cart coiuiectstarter gear on the main gear box at the bottom of the engine. There are no aircraft controls for this system. It is turned on and off by tha ground crew in response to signals from the pilot. Air starts are made by wlndmllllng the engine.


Two throttle levers, one for each engine, are locateduadrant on the left forward console. The right throttle is mechanically linked to the right engine main fuel control and the loft throttle to the left enginefuel control with parallelogram type linkages designed to compensate for heat expansion. The afterburner and main fuel controls are Interconnectedlosed loop cable. The throttle quadrant I* labeled OFF, IDLE and AFTERBURNER. When the throttles are moved forward from OFF to IDLE) they dropidden ledge to the IDLE position. This ledge preventsengine cutoff when the throttles are retarded to IDLE. When retarding the throttles from IDLE to OFF they must be lifted in order to clear the IDLE stop ledge. Forward throttle movement from IDLEILITARY atop controls thethrust range of the engine. The throttles must be slightly raised to clear the stop before additional forwardof the throttle can actuate theIgnition. The AFTER BURNER range extends from the Military atop to the quadrant forward stop. The right throttle knobadio transmissionswitch. Throttle quadrants are marked to indicate 82 power lever angle (PLA) for assistance In determining the cruise power position.

Tlroltle Friction Lever

The throttles are prevented from creepingriction lever located on the Inboard side of the throttle quadrant. When the lever la full aft, the throttles are freo to rnuvo. Moving the lever forward as the INCREASE FRICTION label indicates,increases the amount of friction to hold the throttles in the desired position.

TEB Remaining Counter!

A mechanical TEB remaining counter for each engine Is located aft of each throttle. The counters are spring wound and set torior to engine start. Eachhrottle le moved forward from OFF to IDLE or MILITARYhe counter will reduce one number.

Exhouit Got Temperofuro Trim Switchei

Individual exhaust gas temperature trim switches for each engine are located on the lower loft side of tho Instrument panel. The switches are spring loaded, momentary contact, three position switchesenter OFF position. When held in the(up)emote trimmotor on the engine fuel control Isto slightly increase main burnor fuel flow and turbine inlet temperature. The trim motorsuel flow or EGT travel range ofate of changeer second. When held in the(down) position, the trim motor reduces the fuel flow and decreasesinlet temperature. Aa increase or decrease in turbine inlet temperature will be reflected on the respective exhaust gas temperature gage. These switches are the only provision for main engine control when the throttles are in the afterburning range. They have no effect on rpm when the nozzle is modulating to provide the scheduledspeed. Power for the trim motors is furnished by the respective ac generator


ENGINE INSTRUMENTS Exhaust Gas Temperature Gages

Two exhaust gas temperature gages, one for each engine, are mounted on the right side of the instrument panel. They are calibrated In degrees centigradend Indicate the temperature sensed bydischarge thermocouples. The four digit windows at the top of the gages indicate the exhaust gas temperature to the nearest degree. An OFF window at the bottom of each dial when visible indicates instrument power failure. mall red light on the dial will light when EGT reaches This will activate the respective derichmentif armed. The indicating systempower from thenverter.

brated up0 rpm and the subpointer makes one complete revolution for0 rpm. The tachometers are self-energized and operate independently of the aircraft electrical system.

Lnyire Oil'

An oil pressure gage Is provided for each engine on the right side of the Instrument panel. The gages indicate output pressure of the respective engine oil pump in pounds per square inch. The gages are calibratedal in incrementssi. Power for the gages is furnished by thenverter bus throughvolt auto-transformer.

Fuel Flow Indicators

A fuel flow indicator for each engine is mounted on the instrument panel andtotal fuel flow (engine and afterburner) in pounds per hour. The dial is calibrated0 pound per hour Increments0 pph. The five digit canter window indicates the fuel flow to theph. The indicator Is not compensated for return flow and indicates total fuel flow to engine,and heat sink system. ositiveis normal during windmill operation and the Indicator will read high when the mixer and temperature control valve iscooling loop fuel back to tankuring descent after high speed cruise both high and low fuel flows and flow oscillations may be indicated. Power for the Indicators is supplied by thenverter.


A tachometer for each engine Is mounted on the right side of the instrument panel. The tachometers indicate engine speed inper minute. The main pointer Is cali-

Compressor Inlet Te-npprolure Gooe

A dual Indicating compressor inletgage is mounted on the upper right side of the Instrument panel. It is calibrated In SO incrementsnd 10 Increments. The needles indicate the airforward of the first compressor stage of each nacelle. The system uses platinum realatance sensors and power is furnished by thenverter.

Cot*pressor Inler Ai- St.itjc Pressure Goge

A dual Indicating compressor inlet air static pressure gage located on the upper center of the Instrument panel, measures absolute pressure at the engine compressor inlet. The gage Is calibrated In one psl increments and has marked red rangessi andosireen radial marksl. hite striped third pointer on the CIP gage indicates pressure to be expected when tbe Inlets are operating normally if overEAS. The L>abeled pointers indicate actual inlet static pressures. Power Is furnished from thenverter.


The air inlets for each nacelle arc canted inboard and down to align with the localpattern. The inlet system consists of the cowloving Spike to help provide optimum Internal airflowa spike porous centerbody bleed and an internal cowl shock trap bleed for Internal shock wave position and boundary layer flow control, forward and aft bypass doors for control of airflow In the Inlet and to the engine, cowl exhaust louvers,controls, sensors, actuators and Suck-In doors are alsoIn the aft nacelle area for ground cooling. Nacelle cooling air is provided in flight by air from the cowl shock trap bleed and aft bypass. Normally, the spike and forward bypass are operated automatically by the air inlet control system. InletIs controlled so that the proper amount of air is supplied to the engine and, atairspeeds, the positions of shock waves ahead of the inlet and in the Inlet throat are controlled so as to provide maximumram pressure recovery at the engine face. Controls are provided in the cockpit for incremental control of the aft bypass for those conditions where additional bypass area Is required oreduction in forward bypass flow ia desired. Manual controls are provided to override thespike and forward bypass control


The spike is locked In the forward position for ground operation and flight0 feet. It is unlocked above this altitude and is programmed during automatic operation tonch off the forward stop at. Above, the spikeso as to increase the nacelle Inlet area and decrease the area of the throat or narrowest portion of the duct. Spikeis scheduled primarilyunction of

M;ich numbersensed by the Rosemount boom pitot static ports with biasing for angle of attack and yaw angle. The spike moves aft approximatelynches during transition between. The inlet control alsohock expulsion sensor (SES) and restart feature which can operate automatically when speeds for inlet scheduling are reached. It is effective above approximately. If an inlet becomes unstable and expels the internal shock, the shock expulsion sensor for that inlet overrides the automatic spike andbypass schedule. It causes thebypass to open fully and the spike to move forward as much asnches. Spike retraction is started5 second* after the expulsion is sensed and, when schedule position is reached, thebypass is returned to automatic The SES reference pressure is CIP, and the system is triggeredomentary decrease of CIPr more. Thisharacteristic CIP indication of inlet unstart occurrence. However, it may also operateesult of pressureif CIP decreases rapidly below the previous normal condition duringstalls. The SES feature does nota manually operated spike or forward bypass control. Manual operation of aswitch overrides the SES operation for that inlet. Spike and forward bypass door position changes may be observed during SES operation on the spike andbypass position Indicators. Local pitch attitude and yaw angle are sensedressure probe mounted on the Rosemount pitot boom. The spike porous centerbody bleeds boundary layer air from the inlet throat to prevent flow separation. This air is ducted overboard through the supporting struts and nacelle louver*. The spikes can be fully controlled by use of cockpit controls when hydraulic pressure is available. Emergency Spike forward


provide pneumatic pressure to move and lock the spikes forward in the event of hydraulic system failure.


The forward bypass provides an exhaust for Inlet air which la not required by the engine, and controls the inlet dlffuaer pressure so as to properly position the inlet shock. It consistsotating basket which opens duct exhaust portshort distance aft of the inlet throat. When the landing gear is down, the forward bypass doors are held open by an electrical override signalanding gear door switch. The switch Is positioned to allow manual or automatic control of the bypass when the landing gear retracts. In automatic operation, thebypass remains closedow supersonic speed is reached, then Itin accordance with air Inlet control system Mach and pressure schedules. The Inlet usually "starts" at, that is, the inlet shock Is positioned near the cowl shock trap bleed In the Inlet throat area. As speed Is increased, ths forward bypass schedules as required to maintain the Inlet shock at the throat position.

The forward bypass position is controlled by the ratio of inlet duct static pressureeference total pressure. The Inlet duct static pressure Is sensed by taps located aft of the shock trap bleed.

Tho reference total pressure Is sensed by two external probes one located on the lower Inboard side of the nacelle and the other at the top of the nacelle. The forward bypass control also senses an unstartesult of the sudden decrease in pressure at the engine face and controls the Inletimed sequence. The minimum Mach number at which the automaticactuates varies with the Intensity of the unetart but is generally In the vicinity of. An overriding switch holds the forward bypass closed at speeds lower than.


The aft bypass consistsing ofperipheral openings allowing amass flow off that available from the forward bypass. The ring is located just forward of the engine face. Those openings allow excess inlet air to be bypassed around the engine. The bypassed air joins cowl shock trap bleed air and passes between the outside of the engine and afterburner and the inside of tbe nacelle. This flow augments the exhaust gas in the ejector area. Each aft bypass ring is positionedydraulic actuator which is powered by theydraulic system and is controlled by the cockpit switch. The bypass Is held closed during takeoff and landing by an electrical signal from the nose gear downlock. It is also closed during subsonic operation. Position in flight is aet manually inwith determined Mach number and engine operating requirements.







section i

inlet control system

The air Inletomputer which utilises electricallypneumatic pressure signal* toschedule and reposition the spikes and forward bypass. The computer also servesalibrated path for the manual spike and manual forward bypass control. Major components for each inlet control are the computer, pressure transducer, angle transducer and two pressure ratio The spike and forward bypassconsist of four rheostat type knobs and two Inlet restart switches and anspike switch. Aft bypass control is by means of two rotary type switchesabove the throttle. Three annunciator panel lights are pertinent to the inlet control system.

Nine different pressures are sensed forcontrol. The Rosemount airspeed boom provides pilot total and static pressures to the pitot pressure transducer. The pitch and yaw attitude probe on the left side of the boom provides angle of attack and yaw angle pressures for conversion to electricalby the attitude transducer. At each nacelle local pitot pressure and two inlet duct static pressures are sensed to enable two sensors within the pressure ratioto convert pressure ratios tosignalsirect forwardcontrol,ause an automaticfollowing shock expulsion. Somefunctions are also accomplished within the pressure transducer. Most of theoutputs of the pilot pressureattitude transducer, and bothratio transducers are transmitted to the computer. The computer alsoignal from the main landing gear doors to assure that the forward bypass will be open whenever the main gear Is down.

Spike ConrroU

pike control* are located on the lower left side of the instrument panel. The controls are labeled AUTO, FWD, aad have labeled mark*ach numbers. Intermediate marksach increment* allow the knobs to be positioned manually atach number. In the detented AUTO position, spike position Is scheduled automatically by the Inlet control system. In the detented FWD position, thewill move to the full forward position. The Mach numbered positions are used In manual operation. Use of settingsto aircraft flight Mach number moves the spike aft to the correct position for proper Inlet performance. The spike control also biases the forward bypassunction of control knob position when the bypass Is being manually controlled. The forward bypass position Indicator andcontrol knob will not be In agreement by the amount of bias. Control power for the left spike is from thenverter and for the right spike thenverter.

Forward Bypass Controls

The L. &t'R BYPASS control* are located juat Inboard of the spike control*. ontrol Is turned full counterclockwise to the detented AUTO position, operation of the respective forward bypass iscontrolled by the Inlet computer. Ai the control is turned clockwise the first da-tented position will position the forward by-pas* to the full open. As the controlurned further clockwise the forward bypass will Incrementally move towards the closed position and will be fully closed In the full clockwise position. Markingsn increments ofercent allow theto be positioned at any percentage of full open. Power for the circuits is from the essential dc bus andndnverters.




Manual operation of the forward bypass la permissible with the spike operating on its automatic schedule; however, when the spike is operated manually, the forward bypass must be operated manually or the bypass will open fully and will not schedule.

Iris' Raito'f Swilchsi

osltion toggle switches are located on the left side of the Instrument panel. witches are labeled BESTABTWD DOOR OPEN (center) and OFFn the RESTART position the spike andcontrol settings are overridden, the forward bypass is opened and the spike la moved forward. In the center FWD DOOR OPEN position the forward door Is moved to/or held open but the spike positionto its control knob. In the OFFboth the spike and forward bypass are controlled by their respective controls. Power for the restart circuit is supplied by the essential dc bus.

Einergeoty Spike Switch

Aswitch, labeled

EMER SPIKE, Is provided below thepanel. The switch Is guarded In the center OFF position. After the guard is Opened the switch may be positioned inositions as necessary. In the eventydraulic failure, the one shot emergency pneumatic bottle in thenacelle Is activated to drive and lock the spike in the full forward position. Power for the emergency spike circuit Is from the essential dc bus.

Inlet AflSwitchc* ond Indlcotof Ughn

The inlet aft bypass switches and Indicator lights are located above the throttle They are four-position rotary type switches equipped with concentric lever handles. The switch positions from lop to bottom are labeled%% Left and right amber lights, located near the switch levers illuminate to indicate when an aft bypass position and the switch setting do notight should Illuminate each time ita switch la moved, then extinguish as ths bypass reaches the required position.econds Is required for the aft bypass to move from full closed to full open. The aft bypass actuator controlare powered by the essential dc bus.

Spike Position Indlcofor

A dual spike position indicator Is located on the lower right side of the instrument panel. abeled pointers indicate the position oftthe respective spike in Inches aft of the forward position. It Is calibrated in Inchesoithndnch labeling, power is furnished from thanverter for the left spike and thenverter for the right spike.


A dual forward bypass position indicator is located on the lower right side of thepanel. abeled pointers indicate the opening of the respective tor-ward bypassncrements. Labeled positions0PEN. Power le furnished from thenverter for the left bypass and thenverter for the right bypass.


Manual Inlet Indicator Light


annunciator panel MANUAL INLET light, when Illuminated, indicates that one or more of the four rotary spike and forward bypass controls is not in the AUTO position or that an inlet restart switch is not in the OFF position. Power for the light is furnished by the essential dc bus.

There arc sLx Individual fuel tanks,from forward to aft as tanks,,nd 6. Interconnecting plumbing and electrically driven boost pumps are utilised for fuel feed, transfer, and dumping. Other components of tbe system include pumpnitrogen Inerttng, scavenging, pres-auriaation andingle-pointreceptacle,uel quantitysystem. In addition to furnishing fuel to the engines, automatic fuel manage-


ment provides center of gravity and trim drag control. The fuel le also used to cool cockpit air, engine oil, accessory driveoil, and hydraulic fluid by means of the fuel heat sink system.


The Integral, Internally sealed, fuel tanks are contained in the fuselage and wing root. The tanks are Interconnected by right and left fuel manifoldsingle vent line. Submerged boost pumps supply fuel through the manifolds and transfer fuel for c. g. Forward transfer is accomplished by manual control of the right manifold. Aft transfer Is accomplished automatically through the left manifold. uel dump valve is installed in each fuel manifold. Normal sequence of tank usage is controlled by float switches to automatically maintain anc. g. for cruise. The left engine Is normally sequenced from tanks,ndhe right engine is sequenced from tanks,nd 4. Normal automatic tankla aa follows;

flight refueling. Ground refueling laby use of an in-flight refueling probe specially modified toand operated locking device so that refueling may be done without hydraulic power. Fuel from the receptacle flowi through themanifold to each tank. The useifferent size orifice for each tank allows all tanks to be filled simultaneously InIS minutesoasleofsl. Dual shutoff valves In each tank terminate refueling flow when the tank Is full. eluding fitting Is installed on the right fuel feed manifold in the lower right side of tank 3. ndhich feed the left manifold, are defueled bythe crossfeed valve.


Any fuel Inust be balancedike amount of fuel In the other tanks during ground fueling or defucltng to prevent the aircraft fromdown on the tail.


ank 4

Li Engine

and 6

ank 4

The fuel manifolds can be connected bythe crossfeed switch. Thisotor operated valve between the fuel

manifolda and ts mainly used during single

engine operation.


A single point refueling receptacle installed on top of the fuselage aft of the airbay io used for both ground and In -



Sixteen single stage centrifugal ac powered boost pumps are used to supply the fuel manifolds. ndhich normally feed both engines, are equipped with four pumps and tanksave two pumps each. Either pumpair Isof supplying fuel to Its manifoldate sufficient for normal engine operation In the eventailure of the other pump. The pumps In each tank may be operated out of the normal sequence by use of thetank boost pump control switches located on the right side of the Instrument panel. These switches supplement auto-


fuel system


matic tank sequencingank falls Co feed in the proper sequence. Ilnecessary tothe pump release switch tothe manually actuated pumps when the tank ia empty. Normally, eachhich are protectedommon float switch) is protectedloat switch that deactivates the pump when the tank is empty. One ol tlie float switches En each tank illuminates the yellow tank empty light contained In the respective boost pump tank switch. For example, the float switch for theump In tankis used to Indicate thata empty ands off. (Theightgreen whenrc on. Whens on and insequencing, the green light may not Indicate the status of otherumps whose operation Is affected by automatic features of the ullage and refuelinghe boost pumps that feed tlie left hand manifold are normally powered from the loft generator bus and the pumps that feed the right hand manifold are normally powered from the right gonerator bus. circuit breakers for each pump are located in the compartment behind theand are not accessible in flight.

Emergency Fuel Shuloff Swltchei

A guarded fuel shutoff switch for each engine is installed on the lower right side of the instrument panel. Each switch is guarded in the down (fuel on) position. Fuel is shut off in the OFF (up) position. witchotor operated valve in the respective engine feed line to operate. Motor power is supplied from the corresponding ac generator bus.

Fuel Boost Pump Switchoi ond Indicator lights

Six pushbutton type fuel boost pump switches with green and yellow indicator lights are installedertical row on the right side of the Instrument panel. These switches are provided for manual control of the fuel boost pumps.


Manual operation supplements, but does not terminate the normal automatic fuel tank sequencing.

The switches have an electrical hold and ball arrangement that allows manualof only one tank of tank groupnd one tank of tank groupt the same time. This feature is intended to prevent more than eight boost pumps from operating simultaneously.


It Is possible to operate more than eight boost pumps at onceombination of automaticand manual actuation; however, this condition will not overload the electrical system except when operatingingle generator.

et of boost pumps is actuated, either automatically orreen light will Illuminate In the pushbutton. ank Isellow EMPTY light In the pushbutton illuminates. When depressec the boost pump switch will hold downuntil released by the pump release switch. Power for the boost pump switch circuit and lights Is furnished by thedc bus.

Pump Release Switch

A momentary pump release switch ison the Instrument panel below the fuel boost pump switches. The switch has two positions, PUMP BEL (up) and NORM (down). When placed in the momentary PUMP REL position, any boost pump switch that has been depressed during

manual boost pump selection will be released and automatic sequencing of the fuel tanks is continued. Power for the circuit is furnished by the essential dc bus.


A manually selected boost pump should be releasedank indicates empty so that the pumps In that tank will be shutoff;damage to the pump may occur.

CrcH'occ Switch

A pushbutton type crossfeed switch iaabove the boost pump switches on the instrument panel. When depressed, ita green light In the switch,otor operated valve between the left and right fuel manifolds, allowing operating boost pumps to pressurize both fuel The switch must be depressed atime to terminate crossfeeding. Power for the circuit is furnished by the essential dc bus.

Fuel Tronsfer Switch

A guarded three-position fuel transfer switch is located on the right side of thepanel. The switch io guarded In the OFF position. When the guard is raised and the switch is moved to the FWD TRANS position, the pumps inrealve at the forward end of the right fuel manifold opens intof fuel manifold pressure is abovesi and fuel will transfer forward through the right side fuel manifold as long as automatic or manual pump sequencing continues. will be automatically terminated by closing of the forward transfer valve when theuel level0 pounds.oost pumps will remain inactivated until either tank 4 has bs remaining or the transfer switch is moved to the OFF (down) position. umps will also start when theump switch is pressed. The forward transfer valve is not closed by manual selection ofut right side boost pump pressure

makes forward transfer ineffective. The lift-lock forward transfer switch may also be pulled out and placed in theRANS position. In this position,umps are inactivated, the right side pumps inre turned on, ands turned off if operative. The transfer is only from tankhich prevents the accumulation of hot fuel innd puts the warmer fuel intohere It will be used immediately after an air refueling.


Forward transfer should bebefore refueling is started to restore normal tank sequencing.

Transfer is automatically terminated when the0 pound float switch operates, and theumps remain off until eitherounds remaining or the transfer switch is moved to the OFF Power for the transfer controlis furnished by the essential dc bus.

Those aircraft1 are modified to replace theorward Transfer position with an EMER forward transfer position on these aircraft. When the lift-loc switch is pulled out and replaced in the EMER position,umps areand the0 lb stop transfer float switches inre replaced byloat switches. This allows forward transfer to continue untils full.


The EmK^osT?e used only in case of an. emergency.

Fuel Dump Switch

Aosition lift-lock fuel dump switch is located on the right side of thepanel. The switch is guarded in the OFF (down) position. In the DUMP (center) position dual type solenoid dump valves in each manifold are opened and tlie pumps inre inactivated unlessmanually. If fuel pressure is abovesl, all other tanks dump In normal usage Sequence untils down0 pound remaining level. Dumping nor-


maliy stops at this point and. If fueltn lankheumps will start unless the forward transfer switch Is in either the FWD TRANS orRANS position. The switch knob must be pulled out to put the switch through the DUMP position either to the EMER or OFF position. In the EMER position,0 pound stop dump switch ins bypassed and fuel dumping will continue from all tanks except tank 1. Ifs to be completely dumped,hould be pressed on beforempties In order to avoid fuel pressure fluctuation asmpties. Power for the circuit is furnished by the essential dc bus.


Emergency fuel dumping must be terminated by placing the dump switch to DUMP or OFF. All fuel can be dumped with EMER dump on andelected manually.

Fuel Quantity Selector Switch and Quantity Indicotor

A fuel quantity Indicatorotary seven-position fuel quantity selector switch Ison the lower right side of thepanel. Positions on the selector switch are marked for TOTAL and each of the six tanks positions. The switch Isto the indivicj.il tank or TOTALto select the desired reading on the fuel quantity Indicator. The dial Is calibrated0 pound Increments from zero0 pounds. The Indicator igitalwindow Indicating to theounds. Power for the circuit is furnished by thenverter.

Fuel Quantity Low Light

A FUEL QTY LOW light on the annunciator panel will Illuminate when total fuelIn0 pounds or leas. Power for the light Is furnished by thedc bus.

Fuel Protiure Low Warning Llghti

Fuel pressure warning lights,UEL PRESS LOW are located on the annunciator panel. Illuminationthat engine fuel inlet pressure has fallen belowel. The light Is extinguished by restoring fuelabove approximatelysl. Power Is furnished by the essential dc bus.


U Is possibleuel 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.

rtd-Tonk Lights Switch

And tank lights switch Is Installed below the boost pump switches on thepanel. The switch has twoup and down (spring loaded down) and Is used to test the operation of the liquid nitrogen indicators, nitrogen systemlight, derichment light and fuel boost pump lights. When the switch Is moved to the up position, the liquid nitrogen Indications will move down-scale toward aero andTY LOW annunciator light, fuel boost pump lights and derichment light will Illuminate. Power for the circuit Is furnished by the essential dc bus.

FUEL SYSTEM Pressurization


Thr fuel preasurlzatlon system consist* of two Dewar flasks, located In the notewheel well) and associated valves and plumbing to the fuel tanks and Indicators. These flasks are equipped with automatic ac powered heaters and contain liquid nitrogen. The forward flask containsiters and the aft flaskiters. They supplygas to the fuel tanks5 psl above ambient pressure. This lnerts the ullage space above the fuel and will produce some fuel flow to tbe engine-driven pump in case of boost pump fail-are. The liquidfrom the bottom of the flasks is routed through submerged heat exchangers ino ensure that the nitrogen has become gaseous. The nitrogen gas is then ported to the common vent line and to the top of all tanks.

The venting system consistsommon vent line through all tanks with two vent valves in each tank except theank which has only one vent valve and the open forward end of the vent line. The forward vent valve Ins equippedelief valve to relieve tank pressuresi,loat valve that closes the vent valve when the tank la full. The float shutoff la provided to keep fuel fromthe vent line. The aft vent valve is similar to the forward except It has novalve. The common vent line tees Into two lines innd both go through the rear bulkhead. In the tail cone area thereelief valve In each line with the left valve set to relieve pressuresl above ambient pressure. In tbe event of failure of thia valve, the right valve will rellvve prcsauro55 psl. uction relief line and valve connects to the common vent line Innd terminatesell mouth fitting in the aft end of the nosewheel well.

Two valves are provided in the vent system to prevent fuel from surging forward In the vent line during aircraft deceleration. heck valve prevents fuel that Is coming forward fromrom going farthor than tank 5. ython valve located Inrevents fuel coming fromrom going any farther than tank 3. This float actuated valve doses the vent when fuel is moving forward in the vent line and diverts It Into tank 3. Acceleration presents no problem of fuel shift between tanks.

Liquid Nitrogen Quonrlry Indlcotor

A dual liquid nitrogen quantity indicator I* Installed on tho right side of the Instrument panel. The Indicator displays the quantity of liquid nitrogen remaining in each of th* two dewar flasks. The Indicator Is markediter incrementsower for tho Indicator Is furnished by thenverter bus.

N2 QuontUy Low Light

An Indicator lightTY LOWprovided on th* annunciator paneL The light will Illuminate when either hand on the liquid nitrogen quantity gageiter remaining. Power for the light isby the essential dc bus.

Fuel TankIndicator

A fuel tank pressure indicator Is Installed on the right side of the Instrument panel. The gage indicate* the pressure existing In theuel tank, and la marked fromon incrementsound per square Inch. Power for the indicator Is furnished byvolt Instrument transformer.

section i

fuel heat sink system







Tonk Pressure Low Light

A TANK PRESSURE LOW warning light Is located on the annunciator panel and will il. luxninate when the tank pressure reduces50 psl. Power for the light isby the essential dc bus.

that has space for It. During single engine operation with the inoperative engine throttle in OFF, actuation of the fuel cross feed valve also allows the hot recirculated fuel from the windmllllng engine toand mix with the cooling loop andfuel for the operating engine.


Fuel from the fuel manifolds Is usedooling medium for the air conditioningthe aircraft hydraulic fluid, and the engine and remote gear box oil. Circulated fuel from the engine fuel hydraulic system is also used to cool the TEB tank and the control lines which actuate the afterburner nozzle. Engine oil Is cooled by main engine fuel flow through an oil cooler, locatedthe main fuel control and the windmill bypass valve. This fuel Is then directed to the main burner section. The other cooling is accomplished by fuel circulation through several cooling loops. Hot fuel returning from the remote gear box heat exchanger, the primary and secondary air conditioning heat exchangers, the hydraulic fluid heatthe spike heat exchanger and the exhaust nozzle actuators is circulatedixing valve and temperature limiting valve (smart valve) and returned to the main engine and afterburner fuel if the mixed fuel temperature Isall of the hot fuel will be burned by the operating engine and afterburner. Ii the temperature of the mixed cooling ^oop and incoming engine fuelthe smart valve starts to close and aof the cooling loop fuel Is prevented from mixing with the Incoming englnoressure operated valve routes the hot fuel to tank 4. The smart valve Isclosednd all cooling loop fuel is returned to tank 4. Ifs full, the hot fuel will be diverted to the next tank


The aircraft Is equipped with an airsystem capable of receiving fuellow rate of0 pounds per minuteoom type tanker aircraft. The system consistsoom receptacle, doors,hydraulicignal amplifier and control switches and indicator light. power for the system Is normally supplied fromydraulic system. Ifydraulic system is Inoperative the refuel system can operateydraulic pressure by selecting alternate steering and brakes. Electrical power Is supplied by the essential dc bus.

Air Refuel Switch

An air refuel switch is Installed on the right side of tho Instrument panel. The switch has three positions! READY, OFF and MANUAL. When the switch Is placed in the READY (up) position hydraulicopen the refueling doors, the boom latches are armed, the receptacle lightsand the green READY light The receptacle door* are opened by spring action If hydraulic pressure Is not available. In the MANUAL (down) position the latching dog* In the roceptacle aro closed. They may be opened by holding the disconnect (trigger) switch on the control stick until the boomeated. When the disconnect switch Is released the latches

will close and hold the boom. The latches will open to release the boom when theswitch is depressed. This position is used in the eventecond time delay Is incorporated to prevent nozzle damage if the manualis selected during refueling contact.

Air Refuel Reset Switch ond Indicator Light*

A square dual indicator light and resetlabeled IFR PUSH TO RESET, isat the top left side of the Instrument panel. The top half is labeled READY and will illuminate green when the air refuel switch Is In the READY or MANUAL,and the refueling receptacle Is open and ready to accept the refueling boom. The light will extinguish after the boom Is If the boom disconnects from the fueling receptacle the lower half of the switch will illuminate amber and show DISC. If the air refuel switch Is In the READY position the light button le then pressed to reset the system amplifier for another If the air refuel switch is In the MANUAL, position the READY light will be illuminated and manual engagement andare controlled by the disconnect switch on the control stick. Power for the switch and light is supplied by the essential dc bus.

Ditconfect Switch

A momentary contact trigger type switch is installed on the forward side of the control stick. Depressing the trigger switch will normallyisconnect. Theswitch is also depressed to open the receptacle latches when the air refuel switch is in the MANUAL, position. Releasing the disconnect switch will close the latches.


A disconnect may he accomplished in four ways:

Automatically, if boom envelope limits are exceeded (except when usingboom latching).

Automatically, when manifold pressure*si.

Manually, by the boom operator.

Manually, by actuating the disconnect switch on the control stick.

Pilot Director Lights (Tanker)

Pilot director lights are located on theof the tanker fuselage between the nose gear and the main gear. They consist of two rows of lights; the left row for elevation and the right row for boom telescoping. The elevation lights consist of five colored panels with strip green, triangular green and triangular red colore andnd U, for down and up respectively. Background lights arebehind the panels. The colored panels are illuminated by lights that are controlled by boom elevation during contact. The colored panels that Indicate boomare not Illuminated by background lights. An illuminated white panel between each colored panel serveseference. Theor aftor forward are visible at the ends of the boom telescoping panel. The Air Refueling Director Lights Profile) shows the panelat various boom nosale positions within the boom envelope. There are no lights to Indicate azimuth;ellow line Is visible on the tanker tothe centered position. When contact is made, the panels automatically reflect the correction the pilot must make toposition.



Threeolt ac power laby two engine driven generator* rated atoVA depending on the Installation. Each generatoreparate ac busmpere transformer rectifier. Output of the transformer rectifiers is paralleled andvolt ac power to an essential dc busonitored dc bus andystem ofVA In the eventingle generatorus transfer and protectionconnects the two generator buses. amp hour batteries are furnished toemergency power to the essential dc bus In the event of complete power failuremaller battery provides emergency power to the DNS and thenverter.


Each engine drives an ac generator through Its remote gear box tohase power. There are no constant speed drive units, so the ac frequency varies directly with engine rpm; however, the frequency is essentially constant at scheduled engine speed during climb and cruise. When the output of either generator dropsps, it Is automatically tripped and the other generatorprovides power through the bussystem. Generator cutout occurs at an engine speed of0 rpm. Conventional switches are provided for manual control of the generators.


The aircraft is equipped with twofor connecting ac and dc external power sources to the aircraft electrical system. These receptacle* are located in the noac-wheel well. When external power isto the aircraft and the power switch Is In the EXT PWB position, the acare automatically disconnected from their respective bust* and the busespower from the ground power unit. External dc poweraralleled with the dc output of the two aircraft transformer rectifiers. External dc power and Inverter cooling air must be connected In order for the external ac power to be available.


Electrical power for the essential and monitored dc buses Is normally supplied by th* paralleled output ofmprectifier* which are poweredby the ac buses. The twompere-hour emorgency batteries arcto supply the essential dc bus with powerimited time when bothrectifiers or both generators are Inoperative.


Fixed frequency ac power is supplied byA solid state air cooledhese inverters, located In the cheeks of th* nosewheel well, are controlled byswitches and powered by the essential dc bus. Thenverter is alsoto the INS battery whenever the INS mode switch is on. Normally the No.ndnverters furnish power to their respective buses. Thenverter Is normally off. Inverter power distribution Is so arranged that thenverter bus andvolt Instrument transformer powers most of the flight and engine Thenverter bus furnishes ac power for the INS. In the event offailure or other electrical system malfunction, any one of the three Inverteray be operated from then-




oxycw quantityi law uc ciab irvtR

j ml OtAJOrrr ihoicaim stuck* switch * INVinTB SWITCHES

GLNERATOB switches

capwitch I BArrfHV switch


cajin aitiwtt*

CAtlft altINTit* StUCTO"

H hah and WARNING LIGHTS ttst potion


verier power supply. Cmtain relatedla transferred from the No.ndnverter by operation of the autopilotswitch to maintain the proper power phase relationships. The AN/ABCHF radio has Its own rotary inverter aupply. Refer to Electrical Power Distributionthis section.


Circuit B'sokeri

The cockpit circuit breaker panels areon the right and left consoles andthe annunciator panel. The circuit breakers are push to reset, pull out type breakers for certain ac and dc circuits as listed on the electrical power distribution chart, Circuit breaker panels which are not accessible during flight, but which should be Inspected before flight, are located In the air conditioning bay (Ju"tof the refueling receptacle) andload center (left hand side of nose-wheel well).

Gsnerotor Switches

A switch for each generator Is located on the right aide of the Instrument panel and is powered by the essential dc bus. Each switch has three positions; GEN BESETPD? (down) and center (neutral). The switches are spring loaded to the center neutral position. Holding the switch In the GEN RESET (up) position will return thegenerator to normal operation if It has been removed from the bus for anyother than complete generator failure. In the TRIP (down) position, th* generator output will be removed from the generator bus and the auto bus transfer system will aupply that bu* from the other generator If itperating.


The generators must be reset and connected to the bus after thearc started and before the ac ground power Is removed.

Power Switch

A three-position battery-external power switch Is located on the right side of the instrument panel. When In flight or on the ground with ground power disconnected, placing the power switch in the BAT (up) position causes the emergency batteries to supply power to the essential dc bus. In the EXT PWB (down) position, the external power sources furnish power for thesystems. In the center OFF position, external ac power Is disconnected but power from the dc external receptacle willto supply the essential and monitored dc buses and dc power will not beby moving the power switch from the EXT PWB to OFF positions.

Inverter Switches

Switches for No.nd No. 3are located on the right side of the instrument panel below the generator switches. In the NOBM (up) position, the respective inverter is energized andpower to its individual bus. In the OFF (center) position therom the essential dc bus. In the EMERG (down) position the No. 4Is activated and connected to that inverter bus. In the event of multiplefailure, the lowest numberedswitch that is placed In the EMERG position receives power from the No. 4 Under thiaigher numbered Inverter can not receive power even if Its Inverter switch Is In the EMERG

position.nverter also may receive dc power from the small INS battery If the INS mode switch is not In the OFF position.

Generotor Out Indicotoi Lightt

ENERATOR OUT Indicator lights, located on the annunciatorumlnateenerator is not furnishing power to its ac bus.

TiofMlormef-gectifier Out Indicator Liohti

FHR-HECT OUT indicator lights, located on the annunciator panel,to Indicate that the respective transformer-rectifier is not furnishing power to the dc buses.

Inverter Out Indicator Light*

Three INVERTER OUT Indicator lights are located on the annunciator panel. Whenthe numbered light indicates that the respective Inverter bus voltage is too low. An inverter switch must be placed in the OFF position to disconnect thatfrom the bus. isconnected inverter is switched to the EMERG position, thenverter is activated and willpower to the respective inverter bus and tha light will be extinguishedower numbered Inverter switch has already boon turned to EMERG.

Emergency Battery On Indicator Light

The EMER BAT ON light located on thepanel illuminates when thebatteries aro furnishing power to the essential dc bus.


Four separate hydraulic systems areon the aircraft, each with Its own pressurized reservoir and engine-driven pump. The pump* fornd L, system are driven from the left engine remote gear box andystem pumps arefrom the right engine remote gear box. Hydraulic fluid is cooled by fuel-oil ox-changers, using the aircraft fuel supply as the cooling agent. ydraulic systems provide power for operating the flight controls. ystemspower for all other hydraulicof the aircraft. Under normalconditions, the systems areof one another. The L, hydraulic system provides hydraulic power to the left engine air Inlet control, the landing gear (Including uplocks and doorormal brakes. In-flight refueling door, UHF retractable antenna, and normal nose-wheel steering. ydraulic system provides hydraulic power to the right air Inlet control and also to the alternate brakes, nosewheel steering, refueling door and landing gear (emergency retraction only) whenydraulic system has failed. Whenydraulic systempower to the brakes, the anti-skid feature is Inoperative.

Hydroullc System Preiiuret Gogei

Two dual Indicating hydraulic gages areon the lower center portion of tha Instrument panel. The right hand gagehydraulic pressure offlight controls) systems, and the left hand gage indicates hydraulic pressure ofystems. The gages are calibratedel Increments0 pel. Pressure indication on the gages isby means of remote transmitters in the individual systems. Twcnty-slx volt ac power is furnished by the instrumentand thenverter.








quantify CAcr

hydroullc wornlng light*

Six hydraulic warning Lights are located on the annunciator panel. YD PRESS LOW Ughis will Illuminate when the pressure In the respective system0 psi. YD LOW light will illuminate when the quantity is lessallons. YD LOW light will illuminate when the respective reservoir quantity Is less8 gallons. Power for the lights Is furnished by the essential dc bus.

minute. Indown) position thevalves toystem are opened and the reserve fluid will supplyystem. Power for the valves Is furnished by the essential dc bus.


Reserve hydraulic fluid Is to be usee only to supply theystem In the event of malfunction of the other system.

System Quantity Gage

a quadruple hydraulic fluid quantityInstalled on the right side of thepanel. nd It concentric needles are on the left side of the gage andoncentric needles are on the right aide of the gage. The dials are marked in gallons. Power is furnished from thec instrument transformer.


A reserve oil supply fornd bsystems Is contained Inreserve tank mounted in theuel tank. The reserve hydraulic oil isby gravity flow and nitrogenthrough solenoid operated nhutoff valves to eitherydraulic system.

Hydroullc Reserve Oil Switch

The hydraulic reserve oil switch ia mounted on the left aide of the annunciator panel. Ithree position switch, guarded In the center OFF position. Inup) position, solenoid operated shut off valves are opened toydraulic system suction and tank vent lines. This allows the reservefluid to supplyystemneeded up toallon per

flight control system

The cockpit flight controls consist of acontrol stick and rudder pedals. The delta wing configuration utlllr.es elevons Instead of separate aileron and elevator control surfaces. The elevons, movingin the same direction, function as elevators and when moving In oppositefunction as ailerons. Eachconsists of an Inboard and outboard panel with the Inboard panel located between the fuselage and the nacelle and thepanel outboard of tho nacelle. Both panels on one side functioningle unit with the servo input to the outboard elevon connected directly to the inboard elevon surface. The dual canted rudders are full moving, one piece, pivoting surfacesmall 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 aa follows.:


operated mechanical stops areIn tho cockpit mechanism to limit the surface movement at high Elevon travel In roll Is limited to 7 up, 7 down and rudder travel Is limited to 10 right,left. An additional stop la installed In each rudder servo package to limit thetravel. These stops arc electrically controlled and hydraullcally operated by separate electrical and hydraulic systems. If no electrical power ia available, thewill be limited to approximately 10 I* and P. travel. If electrical power leto one stop, that rudder only will have the full 20 ravel available. The rudder cable must be stretched to obtain this travel,oticeable Increase in rudder pedal force.


Cable systems are utilised to transfermovements from the control stick and rudder pedals to the flight control The pitch and roll axis cableare duplicated from the cockpit to the mixing mechanism In the aft fuselage. The rudder system has two separate closed loop single cable systems, one to each Cable tension regulators and slack absorbers are incorporated in the cable systems,


Flight control trim Is accomplished bythe control surfaces through the use of electrical trim actuators. The roll and pitch trim actuators are locatedof the feel springs so that edck position remain" neutral, irrespective of the amount of trim. The trim actuator and feel spring location is combined In themechanism and yaw trim Is reflected by rudder pedal position.

Travel limits of the trim system aredown up In up and down (each side) In roll, andleft toright In yaw. Trim position Indicators are provided for each axis. Trim rates are aa follows:

j Roll

Total Diff.

Total Dtff.

Automatic pitch trimeparate, slow speed motor for autopilot synchronisation. The automatic pitch trim rate5 /sec maximumsec minimum. Trim power Is normally furnished by the Rbus.


Primary control for the rudders consists of conventional rudder pedals mechanically connected by cables, bell cranks and push-rod" to hydraulic control valves at thehydraulic actuators. The rudder pedals are released for adjustment by pullingandle labeled PEDAL ADJ locatedthe annunciator panel. Wheel brakes are controlled conventionally by toe action on the rudder pedals; refer to Wheel Brake System, this section. Rudder pedalalao controls noaewheel steering; refer to Nosewheel Steering System, this section. The pedals are hinged to foldand upward, providing foot space on the cockpit floor.






control stick grip


The useull power Irreversible control system for Actuation of the surfaces prevents air loads and resulting "feel" from reaching the cockpit controls. Therefore, feel springs are installed In each of the pitch, roll and yaw axis mechanisms to provide an artificial sense of control feel. The springs apply loads to the pilot controls Into the degree of control deflection.


The control stick Is mechanically connectedorque tubo, push rods and bell cranks to tbe dual cable system which operates the roll and pitch quadrants In the aft fuselage tall cone. Mechanical push rod linkages mix the control movemonts and position dual hydraulic control valves. Theseystem hydraulicto the inboard elevon actuating

Push rods, bell cranks and torque tubes transfer inboard elevon deflection tothe outboard dual hydraulic control valves. These valves directystem hydraulic pressure to the outboard elevon actuating cylinders. ush rod followup system closes off the flow offluid to the actuators when theelevon deflection Is obtained. on the control stick grip is apitch and yaw trim switch, ancontrol stickoaewheel steeringicrophone switch for both Interphone and radioombination autopilot disconnect andrefueling disconnect switcham override pushbutton.

Control Stick Command Switch (CSC) Refer to Autopilot System, Section IV.

Pitch ond Yow Trim Switch

Pitch and yaw trim control Is providedpring-loaded, four position thumbswitch installed on the control stick gripenter OFF position. The switch positions are LEFT, RIGHT, NOSE UP and NOSE DOWN. The switch controls trim motors powered by the right generator bus through theolt ac trim actuator transformer and trim power bus.


The trim power switch must be in the ON position before the pitch, roll and yaw trim switches will operate.

Lateral movement of the switch to the left corrects for right yaw and lateralto the right corrects for left yaw. Forwardmovement of the switch produces down elevon operation of the trim motors and actuators (aircraft nose down). Aft movement moves tho elevens up (aircraft nose up).

Trim Power Switch

A trim power ON-OFF switch is located on the annunciator panel. It enables the pilot. If necessary, to disconnect power to all trim motors quickly as the main trim power ac circuit breaker la not available to the pilot. To prevent Inadvertentthe switch must first be pulled outit can be moved from th* ON to the OFF position. In the ONhase ac power from the right generator bus is applied to the primary side of the trim actuator transformer. Individual 28

achases of the Manual Pitch, Auto Pitch, Boll and Yaw trim circuits are located on the right

Roll Ttlm Switch

A three-poaitlon roll trim switch is located just forward of the throttle quadrant. The switch positions are Indicatedleft)right) arrows. The Switch Is spring-loaded to the center off position. When the switch Is held In thaosition, the roll trim motor actuates to move the right ele-vons up and the left elevons down. Actuation of the switch to the I. position moves the right elevons down and left elevons up. volt ac power is furnished from the trim power bus.


A three-position rudder synchronization switch is installed just forward of the throttle quadrant. The switch positions are indicatedright) arrows. It is eprlngloaded to the center off position. In the L,ositions the switch provides electrical power to the right rudder trim motor which moves the right rudder to agree with the position of the left. Rudder synchronisation is obtainedeedles on the yaw trim gage. volt ac power is furnished by the trim power bus.

Roll, Pitch ond Yow Trim Indicator*

Separate roll, pitch and yaw trim Indicators are located on tho left side of the instrument panel. The roll trim indicatorouble ended needle and displays the amount of roll trim fromtodifferential. The pitch trim Indicator displays the amount of pitch trim from 5 nose down to 10 nose up,onlynose down andnose up trim Is available. The yaw trim Indicator displays the amount of yaw trim from 10 left to 10 right for both rudders. Rudder synchronisation Is obtainedeedles on the yaw trim gage. volt ac power for theIs furnished by the Instrumentand thenverter.

Surfoce Umlter_Control Hoodie

andle, labeled SURF LIMIT RELEASE, le located on the left side of the annunciator panel. When tho handle le turnedounterclockwise and released, thestops in the roll and yaw axis of the cockpit control system are activated. This action also opens an electrical switch whicholenoid operated valve in each rudder servo package and activates the servo package rudder stops. When the handle la pulled out and rotated 90 clockwise, the mechanical stops In the cockpit are released and the solenoid Is energized, releasing the servo package stops. Power for the rudder limitingis furnished by ths essential dc bus.

Surfoce Limltar Indicator Light

When speed exceeds, an Indicator light on tho annunciator panel will Illuminate until the surface llmlter handle is released. If the speed Is belownd thellralters are on, the Indicator light will Illuminate until the surface llmlter handle ia pulled out. Power for the lights Is furnished by the essential dc bus.


The automatic flight control system includes stability augmentation, autopilot, and air data systems, plus additional subsystems furnishing attitude and navigational course Inputs for the autopilot. The air data ays-

Cam furnishes signals to tha autopilot and lnertlal navigation systems. Th* stability augmentation system supplies signals to the hydraulic servos that operate the control surface*. The lnertlal navigation system supplies attitude and navigational course Inputs for the autopilot. Heading andreference signals for the autopilot are also supplied by the Flight Reference The autopilot moves the aircraftservos through the SAS. Forinformation on the autopilot andnavigation systems, refer to Section IV.


The three axis stabilityombination of electronic andequipment which augments thestability of the aircraft. It is designed for optimum performance at the basiccruise speed and altitude, but also provides improved stability for In-flight refueling, landing and takeoff. The SAS Is part of the aircraft's basic control system and is normally used for all flight

Dual electronic channels are provided for all axeshird monitor channel Isfor both the pitch and yaw axis. Logic circuits compare the functioning of each pitch and yaw channel and automaticallyailed channel. The pilot Is also providedisual warningailed channel.

In the roll axis, each channel control* the elevons on only one side of the aircraft. The pilot mayingle channel If Reliabilityrovided through dual hydraulic and inverter supplies. Each active channel In each axis Is powered by separate supplies so that the two halve* of each system are operated Independently. eparate gyro system is provided for each channel In each axle. The design is ouch that no single failure except overheatingomplete gyro package can cause loss of all channels In one axis. Even If thisIt Is unlikely lhat all of the gyros In the package would fall simultaneously. The SAS system compare"lectronic systems andalfunctioninghannel. Automatic gainIs applied to the remaining channels so that control response remains thealfunctioning electronic* channel isby Illumination ofight.

STABILITY AUGMENTATION PITCH AXIS The pitch axis SAS consists of twoactivehirdhannel. The two Independent activerovide the desired control through two pairs of tandem servos. There is one pair of servos on each sld* of the aircraft. The servos are in aerie* with the autopilot and the pilot'* control movement*. Damping signals to the elevon* do not move the control stick. hannel drives one servo on the left aide of the aircraft and one on the right side. hannelydraulic systemhannel usesydraulic system. This avoid* loss of both channel* in case of failure of eitherydraulic systems. The sensors for the pitch axix are rate gyro* located In tank No.he gyros provide signals In proportion to the rate of pitch attitude change oflagged" pitch rate gain ia programmed Into the pitch SAS electronic circuits. This pitch rate signal changeover may bo folt aa an abrupt pitch transienturn while climbing or descending throughoot level. Refer to Section VII, Pitch Axisdue to Lagged Pitch Rate Switching. Phasing of the gyro signals is such tbat aa angular pitch motion produces elevon movement to oppose and restrict attitude change. The system will take corrective action rapidly in the eventust disturbance. Pilot inputs are also opposed; however, the elevon motion produced by theesigned to aid the pilot In avoiding ovcrcontrol and improve tbe handling qualities of the aircraft.














The logic circuit Is able toAS failure In either the electronics or the alfunction Is isolated, the failed active channel will disengage and the system continues In operationingle channel. Malfunctioning and dlsengagelng of channels is Indicated by indicator lights. The pitch axis canaximum elevon surface travel up down. Dual or single channel operation produces the same corrective action of the elevon surface. Powerhannel is fromhase ofnverter bus. Powerhannel Is fromhase of No. 2 Monitor channel power Is fromhase of thenverter. Each power source Is protected by Individual circuit breakers in the cockpit.


The yaw axis of the SAS is very similar to the pitch axis, using twohannelsonitor channel. There is one pair of hydraulic servos for each rudder, each pair mountedhifHetree arrangement. Damping signals to thedo not move the rudder pedals. hannel drives one servo on each side of the aircraft. ydraulicIs connectedhannel and the Bsystemhannel. The rate gyro sensors for the three channels are identical to the pitch rate gyros, except for the physical orientation to sense yawing motions. Hi Pass" filter circuit Isto allow passage of normal short term damping signals, bul will stop the signalseliberate turn la mado. ateral accelerometer sensor Is also used In each channel of the yaw axis. Thisprovides an input for high gain lateral acceleration function toore rapid rudder response during engine failure conditions. However, this function willthe pilot when he is purposely trying to sideslip.

The logic circuit is idsntleal to the pitch axis snd functions In the same manner. The yaw axis canaximum rudder travel of 8 left toright. Corrective

surface motion is the same regardless of one or two channel operation due togain doubling if only one channel is operative. Powerhannel Is fromhase of thehannel fromhase of thenverter and the monitor channel fromhase of thenverter. The circuitry from each power source is protected by Individual circuit breakers.


Roll axis reliability requirements are not as severe as pitch and yaw; therefore, less complicated circuitry and components are used. The roll axis has two independent channels, each operating the elevons on one side of the aircraft. hannel positions the left elevon surfaces and operatss fromydraulic system. hannel positions the right elevon surfaces and operates fromydraulic system. There is nochannel. Each channel can be operated Individually. Although the system gain is the same as two channel operation, roll control Is not symmetrical. Coupling into the yaw and pitch axas is possible, but the systems operating In those axes minimize undesirable aircraft motion. Maximum elevon travel In the roll axis is 2 upown (eachotal of 4 with both systems operating. Powerhannel ishase of thenverterhannelhase of thenverter.


The SAS control panel on the right console contains six channel switches,hannels of the pitch, roll and yaw axis. The panel alsoress-to-test switch and six Indicator lights for thend MON channels In the pitch and yaw axis. Three guarded switches for the backup pitch damper, pitch logic override and yaw logic override are located on the right side of the annunciator panel. oll channel disengage light is located between the roll channel switches. Individual circuit breakers are located on both right and left consoles.

Chonnel Switches

There are tlx toggle twitches located on the SAS control panel. Therene pair for each axis; pitch, roll and yaw. The forward switch of each pairhannel and the rear switchhannel. The switches have two positions; ON (forward) and OFF (aft). When electrical power Is on the aircraft and the channel switches are OFF, the SAS electronlca are powered, but the channel servos are not engaged into the control system. Moving the switches to the ON position engages the SAS servosthe recycle light is extinguished. If the recycle light is not extinguished It must be depressed for engagement.

Recycle Indicator llohrt

Six Indicator lights are located on the SAS control panel adjacent to the pitch and yaw channel engage switches. One light Isfor eachnd MON channel in the pitch and yaw axes. When the channel switch is on and the light is not illuminated, the channel is functioning properly. If the light is illuminated, it indicates that the channel has disengaged and the light may be pressed to recycle the channel. In the event the failure was momentary, this will reengage the channel. If the light relllum-Inates, the channel is malfunctioning, but It Is not necessary to turn the channelswitch off because the light indicates that automatic disengagement has occurred.


The lighted recycle indicator light should be pressed down firmly and released. ontrol surfacewill occurardover servo exists in that channel. Refer to Section IE.

The six recycle lights will be Illuminated when electrical power Is applied to the The channel switches musr be on and the recycle lights must be pressed to engage the channel electronics to the servos When engaged and operating, the channel lights will be out.

Roll Chonr>el Disengage light

A single roll channel disengage light Isbetween the two roll channel switches. When illuminated it indicates (hat both roll channels have disengaged. Disengagement results when the roll servo channel outputs differ by more than an amount equivalent surface deflection. When operatingingle roll channel the light will not be illuminated and disengagement In the eventailure Is not provided. The switch must be ON for the active channel and OFF for the malfunctioning channel.

Light Test Switch

A pushbutton light test switch Is located In the center of the SAS control panel. Press' Ing the button illuminates all SAS lights for test.

Backup Pitch Damper Switch

A guarded BUPD switch is located on the right side of the annunciator panel. It is guarded in tho OFF position. It is used In case the SAS pitch channels are unusable due to electronic malfunctions orof the pitch gyro package. In the ON position the backup gyro, located In the electronic compartment, supplies pitch rate signals through an independentchannel to eitherervos. The pitch logic override switch must be used to selectervo operation.


The primary purpose of Ihe BUPD Is to provide an emergency system for pitch stability augmentationrefueling and landing approach. The system Is optimised for use at light weight, aft center of gravity and subsonic speeds. It is notas an emergency backupduring cruise. Refer to Section IU. Emergency Procedures.

SAS Pitch Logic Override Switch

A guarded, throe-position SAS pitch logic switch is located on the right side of the annunciator panel. It Is OFF in the center guarded position and the logic circuit is Placing the switch inup) position deletes the logic circuit andhannel operation. Indown)the logic circuit Is deleted and Bla selected. The switch must be placed in eitherosition when the BUPD la used. This selects operation of eitherervos.


The override switch Is only used as an emorgency procedure. Refer to Section HI.

SAS Yaw Logic Override Switch

A guarded, three-position SAS yaw logic switch is located on the right side of the annunciator panel below the pitch logic override switch. It Is guarded In the OFF position. up) position deletes the logic circuit andhannel operation.down) position deletes the logicandhannel operation.


The override switch Is only used as an emergency procedure. Refer to Section HI.

pitot-static systems

The pltot-statlc system supplies the total and static pressure necessary to operate the basic flight Instruments and air data system components. The pressures arc sensed by an electrically heated probe mounted on the nose of the aircraft. The probe and forward nose also serves as an antenna for the high frequency radio. The pitot orifice of the probe Is divided Inside the head to provide two separate pressure sources. It also has two circumferential sets of four static pressure ports each. One pitot and the aft set of static portspressure signals to the air dataand inlet air control systems. The other set of pickups supply pitot and static pressure directly to the speed sensors on the ejection seats, the altimeter, the rate of climb and airspeed indicators. An offset head on the left side of the probe provides yaw and pitch pressure signals to the inlet spike controls and to the stall warning light sensor.

The heating elements of the probe areby the pitot heat switch located on the left side of the annunciator panel. Power la furnished by the left ac generator bus.

An alternate heated pitot static source is available from the Flight Recorder System. Refer to Flight Recorder, Section TV.




Pilot-Heat Switch ond Irtdlcotor Light

A two-position toggle switch Is located on the left side of the annunciator panel. In the ON (up) position ac power Is applied to the heating elements of the pltot-statlc probe. The probe Is grounded toanner which permits the HF radio to be operated while pitot heat is on. In the OFF (down) position ac power lafrom the probe heating elements.

The circuitry also incorporates an altitude switchITOT HEAT light located on the annunciator panel. The pitot heat light will be on when the switch is In the ONand the altitude Is0 feet, and aleo when the switch is in the OFFand the altitude is0 feet. The light will be OFF if when0 feet and pitot heat Is ON, and when0 feet with the switch in the OFF


The air data computer performs twocomputation and display. Tbe total and static pressures from tbe pltot-statlc probe are converted to electrical signals required for the pilot's triple displaycompressor inlet pressure Indicator system, the automatic flight control and In-ertial navigation systems. The ports which supply pressure to the air data computer are separate from those that furnishto the basic flight instruments. failure of the air data computersource will not leave the pilot without the altitude, vertical velocity or airspeed Information. The air data computerpltot-static pressures Intorotary shaft positions which are equivalent to pressure altitude and dynamic

pressure. These shaft positionsechanical analog computer made up of came, gears and differentials to drive the output functions. Outputs of the air data computer and the usingare listed below:




Equivalent Airspeed


Display Indicator


Inlet Pressure Indicator !


Mach Bate


Dynamic Pressure



Navigator Computer

Power for the air data computer Is furnished either by thernverteron the position of the autopilotswitch.

Triple Dhploy Indicator;

A triple display Indicator Is located on the Instrument panel to provide digital displays of airspeed, altitude, and Mach number as computed by the Air Data Computer. The altitude indication range of the TDI la from

00 foot. eel the first digit in dropped,0 feeteet pressure altitude, thelimit of the ADC signal to lhe The Mach number display capability range of each instrumentthe minimum indication at staticnormally rangesach number and the maximum indication would beormallyinstrument. This range corresponds to the range of signals which the ADC isof providing. The TDI displaysin knot* equivalent airspeed (KEAS) within an instrument capabilityEAS; however, the minimumIs normallyEAS towith the minimum ADC signal The maximum signal provided by the ADC results in an airspeed indication which decreasesEAS at sea levelEAS0 feet and, and then decreases further at high altitudes to show the KEAS corresponding tond the existing pressure An off flag appears on the face of the instrument If the ADC loses power. Power for the Instrument la from thernverter.


. Indications of the triple display indicator and the basic pltot-static flight instruments should be periodically cross checked to confirm proper system operation. Refer to figureppendix i.

. Tho triple display indicator is primarily used for aircraftabovend toproper airspeed controlcllmba to FL ISO. Basic pitot-statlc operated flightshall be used in the landing pattern, during takeoff untilclimb schedule is established on the TDI, and during allor actual instrument flight below.

. If KEAS Indications oscillatetwo values on the high ond of the range, it is an indication that the Indicator limit is being approached.


For information regarding Instruments tha: are an Integral partarticular system, refer to applicable paragraphs in this section and Section IV.

Ainpeod-Moch Meter

A combination airspeed and Mach meter operating directly from pitot-statlcis located In the flight instrument group. Thispecial instrument with airspeed and Mach number rangeswith aircraft performance Mach number and Indicatedare road simultaneously on theand outer index respectively. imit airspeed needle (white barred) shows the airspeed limit of the aircraft. The actual airspeed limit is in equivalent airspeed; however, the needle varies with altitude to read the indicated airspeed that converts to equivalent airspeed.


A sensitive pressure altimeter Is located on the Instrument panel. In addition to0 footoot pointers, it also0 foot pointer. This pointer extendi to the edge of the dialriangular marker at Us extremity. The center discutout through which yellow and blacc warning stripes appear at altitudes0 feet. The barometric pressure Bcale isutout at the right side andetnob located at the lower left aide of tho Instrument.

Altilode Indicoro. (MM-3)

The attitude indicator is located In the baaic flight instrument group on the instrument panel. It provides constant visual Indication of note and wing position in relation to the earth's surface. Attitude indications are presentedpherical graduatedeferenceank pointer,arked outer ring. orizontal line is formed on the spherical dial by tho meetingray, upper climb sectionlack lower dive section. The instrument shows attitude In climb or dive up toegrees.


At approximatelyegrees climb or dive, the attitude indicator will flip but will not tumble. egreo flip In roll will be very rapid and the instrument will accurately indicate pitch and roll attitudesthereafter. Some small Inaccuracies may developeries of maneuvers beyond theegree climb or dive attitude. These Inaccuracies willbe cancelled out at the erec-

tion rate of .8 per minute.

eference line remains fixed with the marked outer ring and represents the aircraft In miniature. The spherical dial moves up or down, or the whole spherical dial assembly rotates within the instrument case behindeference line and outer ring to indicate aircraft attitudes. As the dial assembly rotates, the bank pointer moves with It to indicate degrees of bank on tho outer ring. The outer ring indicates-bank. The spherical dial and pointer are capable ofegrees of roll with the aircraft. Pitch attitude of tho aircraft is indicated by the position of the horizon line in relation to the miniature aircraft. itch adjustment knob on (lie lower right side la used to change the position of the spherical dial as desired. During initial gyro erection, and when power Is off or is insufficient to keep

the gyroarning OFF flag appears at the bottom of tho Indicator, The autopilot and attitude reference selector switch In used to select pitch and rollsignals from either the INS or FRS stable platforms.


To avoid gross pitch attitude errors the pitch adjustment knob of the attitude indicator should be adjusted to align the index marks before theand attitude reference selector switch is changed in flight.


Toossible malfunction of the attitude indicator, anaccuracy check should be made by comparing It against the standby attitude indicator and other basic flight instruments.

The system is powered by thendnverter depending on the position of the autopilot selector switch.

Standby Attitude indicofor

The standby attitude Indicator located on the lower loft side of the Instrument panelthe pilot with an Independent attitude reference. Itphere 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 of2 degrees and roll anglesogrees. The bank angle scale is marked on the lower periphery. itch reference adjustment knob is provided on the lower right corner of the Instrument for positioning the reference bar aaast orect pushbutton Is providedmall panel above the throttles.




1 OXY tOt*


2 OXY low



HEAT high





hyd PRESS ion







XT MR-RE ct out







hyd LOW



iiv3 LCiY




Do not hold fast erect button for more thaneconds to prcvont overheating of fast erect motor.


This Instrument has its own self-contained gyro and is not dependent on anothersource. The OFF fiag will be visiblo whenever power to the indicator is interrupted. Power is provided byhase of thenverter.

Verticol Velocity Indi

A vertical velocity indicator is located on the instrument panel and shows the rate of change oi altitude in fcot per minute. Changes in pressure due to changes inare.sensed by the static system and transmitted to the indicator. Depending on the Instrument installed the Instrument la capable of Indicating vertical speed0 feet per minuteeet per minute. An over-pressure diaphragm and valve prevent excessive rates of climb or descent from damaging the Instrument.

Twin and Slip Indicator

A turn and slip Indicator is installed on the instrument panel. The Indicator isforwo or four minute turn. The indicator is powered by the essential dc bus. An additional larger slip indicator is mounted on the upper center instrument panel beneath the CIP indicator.








klgurc 9


An elapsed time clock la located on thepanel. It contains an elapsed time mechanism that Is started and stopped by pushing the winding knob. A clock Is also Installed in the panel. The second hand is started and stopped by the small button on the upper right corner. The third hand serves0 minute recorder.

Nocellc Fin; Warning Lights

Left and right nacelle FIRE warning lights located on the top right side of thepanel. Illuminate when nacelleat the turbine or at theexceeds* Flip down glare shields are provided for night flying. Power for the circuit is furnished by thenverter.


An annunciator panel Is mounted on the lower instrument panel. The panel contains Individual warning lights that Indicateor failures of equipment and Illumination of any Individual light also illuminates an amber master caution light on the upper portion of the Instrument panel. Once Illuminated, the master caution light can be extlngulahed (reset) by depressing the light. The individualpanel light will remain illuminated. Another malfunction again illuminates the master caution light. Warning lights are automatically dimmed when the Instrument panel lights are on. The master warning system does not Include the fire warning and landing gear unsafe lights. Power is furnished by the essential dc bus.


A fire warning system detects and indicates the presenceire in the engineot spot anywhere along the length of the detection circuit will Illuminate the light of that particular nacelle. The lights areon the pilot's Instrument panel above the respective column of Instrumentsto each engine.


A STALL WARNING light Is located on the annunciator panel which Is Illuminated when the aircraft angle of attack reaches nd the nose landing gear scissor switch is open. Pressurenlets on the pitch and yaw probe are senseditchunit to actuate this light. teady tone warning signal Is alao produced in the pilot's earphone. Power for the stalllight is furnished by the essential dc bus.


The tricycle 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 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 la no hydraulic pressure on the gear when It ts up and locked. Down locks Inside the actuating cylinders hold the gear In place In the extended position. HydraulicIs also on the gear in the extended positionystem pressure le The landing gear cylinders and doors are actuated in the proper order by twovalves. Normal gear operation is

powered byydraulic pump on the left engine. Should pressure dropsl during retraction, the power source automatically becomesydraulicydraulic pressure will not extend the gear In tho event ofystem failure and the manual landing gear release must be used. Normal gear extension timeeconds.


A wheel shaped landing gear lever ison the lowar left side of thepanel Just forward of the throttle Th* lever has twoP and DOWN. ocking mechanism Is provided to prevent the gear lever from beingplaced in the DOWN position. utton which extends upward from the top of the lever mult be pressed forward in order to release the lock mechanism. Anbutton Is Installed just above the gear lever and may be used to override the ground safety switch should it becometo raise the gear when the weight of the aircraft Is on the landing gear. Once energised, tike gear lever must be recycled to the DOWN position In order to bring the ground safety switch back Into theed light installed in the transparent wheel illuminates during cycling, or when the gear la in an unsafe condition. Power for the circuiturnished by the essential dc bu*.

Monual Landing Geor Release Handle

A manual landing gear release handleGEAR RELEASE Is installed on the annunciator panel. Ifydraulichas failedydraulic pressure is available, tlie landing gear lever must be in tho DOWN position or the landing gear CONT circuit breaker must be pulled out before pulling tho CEAR RELEASE handle. Otherwise,ystem will retract the

gear. The gear extends by gravity force.nches of pull on the handle Is required since the uplocks are released at different positions along the cable length. The nose gear uplock Is released first followed by the right gear then the left. Gear retraction Is possible after being lowered by the manual gear release handle,ydraulic system pressure is available.

Gear ond Worning Light Test Button

A gear and warning light pushbutton switch Is located on the left forward panel. When depressed It illuminates the landing gear lever red light, all annunciator panel lights, the right and left nacelle fire warning lights, and actuates the gear warning tone In th* headset. It Is also used to test the three green landing gear position lights when

Landing Gear Position Lights

Three green lights, located on the left side of the Instrument panel Indicate the down and locked condition of the landing gear. The location of each,light correspond* to th* respective wheel It monitors. Power is from the essential dc bus.

Lending Gear Worning Light ond Audible Warning

The red landing gear warning light In the landing gear lever handle when illuminated Indicates;

Gear la cycling.

Gear system is not locked In the UP or DOWN position.

Gear la UP and throttle settings areMILITARY and altitude la0 feet.

A pulsed tone warning signal la alioin the pilot's earphones when the throttles are retarded belowhe distance between the IDLE and MIX, throttle settings, the landing gear Is not In the down and locked position and aircraft altitude is0eet. Power for the light and pulsed tone warning Is furnished by the essential dc bus.

Lending Geor Warning Cutoot Button

The audio gear warning signal can beby pressing the GR SIG BELswitch on the instrument panel. The circuit le reactivated when the throttles are advanced above tho minimum cruise setting. Power is supplied by the essential dc bus.

Lond Geor Ground Safely Piru

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. Warningdirect attention to their removal before flight. An additional set of ground safety pins is providedontainer behind the scat.


A landing gear strut damper system lato control gear "walking" during brake operation. The system la sensitive to less thanhange in fore and aft acceleration. The damping ia controlledonitoring valve whichIncreases or decreases the brake pressure as required. Hydraulic pressure for the damper system Is provided byystem.


The nosewheel steering system provides power steering for directional control when the aircraft weight Is on any one gear. The nosewheel is steerableegrees cither side of center. Steering Is accomplishedydraulic steer-damper unit controlledable system by the rudderydraulic system pressure from the nose landing gear down line Is routed to the steei Ing systemhutoff valve, which Is controlled by the nosewheel steering (NWS) button on the control stick grip. Steering Is engaged by depressing the NWS button and matching pedal position with nosewheel angle. olding relay circuit allows the NWS button to be released after it Is once depressed and steering will stay engaged. It Is disengaged when the NWS button is again pressed and releasod. Steering is engaged at any time the NWS button Is held depressed. Nosewheelradius Is approximatelyeet. Aoperated centering camcenters the nosewheel when It Power for the system Is furnished by the essential dc bus.


Nosewheel steering is operable only ii essential dc bus power is available and weight of the aircraft Is on any one gear. Ifystem pressure should dropsi alternate nosewheel steering may be obtained by placing the brake switch to ALTRAKE position.


The landing gear side load strength Is critical. Side load during takeoff, landing and ground operation must be keptinimum.









The aircraft le equipped with artificial feel hydraullcally operated power brakes. De -pressing the rudder pedals actuates the four rotor brakes on each of the alx main wheels. ydraulic system furnishes brake pressure with optional antiskid The hydraulic pressure to the brakes Is0 pal. Shouldydraulic system fall, alternate brakes are available. The alternate brakes operate from an Independent system using Hpressure with no antiskid provision.

A small accumulator Is incorporated in the normal brake system which should provide up to five brake applicationsydraulic failure provided accumulator pressure has not been dumped by selecting alternate brakes. Certain types ofsystem failures suchroken line could deplete the system Quid. Normal or antiskid brakes are usable If left hydraulic pressure la steady and0 psi. brakes are used if left hydraulic system pressure is below this pressure.

Broke Switch

A three-position brake switch is located on the left side of the Instrument panel. In the NORM (center) position, brake pressure fromydraulic system Is available, but the antiskid system le not operative. In the ANTISKID (up) position, the antiskid system is operative. In the ALT STEER It BRAKE (down) position, the brakes, NWS and air refueling system are powered byydraulic system If left systemia0 psi. Power for theIs furnished by the essential dc bus.


Do not switch to alternate brakes unless normal left hydraulicIs unavailable or normal brakea are inoperative. Pressure may be trapped In the brakea after the pedals are released, causing grabbing or locking.

Antl-tktd CM Indicator Light

Illumination of the ANTI-SKID OUTlight on the annunciator panelthat the anti-skid system le When the aircraft is on the ground, the light will be Illuminated when the brake switch la in the NORM or ALTRAKE poaltlon. The light will be off when the switch is in the ANTI-SKID position, If the anti-skid control box and wheelare operative. If the fall safewithin the anti-skid control box Is tripped and power from the essential dc bus la on the system, the light will Tho lightff at all times when the weight of the aircraft Is not on the gear.


The drag chute system is provided tolanding roll and aborted takeoff roll out distance. foot ribbon typeIs packedeployment bag and stowed In the upper aft end of the fuselage. It ridos free in the compartment and Is locked onto the airplane at the initial stage of Ita deployment action. The neck of the drag chute linkreakaway section to protect against aircraft structural damage If the chute Is deployed at toopeed. The chute deployment is actuated electrically and power is furnished by the essential dc bus.



Prog Chute Hondlo

Tho drag chuto doploy and Jettison handle ta located on the left edge at the Instrument glare shield. When pulled the handlemicro awitchoa which deploys the drag chute. When turnedegreesand pushed tn, the drag chute is jettisoned. Power for the circuit Is furnished by the essential dc bus.


Similar left and right hand air conditioning and pressurixatlon systems utilize high pressure ninth stage compressor air from each engine to pressurise and cool theand equipment compartments. System shutoff valves allow compressor air to flow when the engines are running and the system switches are ON. Cooling Is accomplished by ducting the bleed airam air heat exchanger, primary and secondary fuel/air heat exchangers, and through an air cycle refrigerator. Temperature of the air supplied by each system is modulated by temperature control bypass valves located upstream from the air cycle refrigerators. The bypass valves are positioned by control switches located In the cockpit.

A water separator is Installed tn each air conditioning system downstream of the air-cycle refrigeration units. Below an altitude of0ressure switch In the automatic temperature control circuit limits the minimum outletof the air from the air-cycletoo prevent freezing of water in the separator. Using the manualcontrols will allow lowerair to come from the refrigerator but icing of the water separator may occur if humidity is high. 0 feet the altitude pressure switch opens the water separator bypass valve and air does not flow through the separator.

The left engine normally furnishes air for the cockpit, nose compartment, ventilated flying suit, inverters and INS platform. The right engine normally furnishes air toay where it mixes with cockpitair for ventilation ofay, and the aft equipmentixed orifice restrictionuctInto two outlets provideortion of the right system air to flow to the upper part of the cockpit. rossover system is provided to supply right engine system air to the cockpit and equipment normally supplied by the left engine system. The operation of the crossover system will not depressurizeay since the cockpit air exhausts Intoay;ise In temperature will occur inay. High pressure canopy and hatch seal air and windshield defog air is furnished from both right and left engine systems by ducts connected downstream from the primary fuel/air heat exchangers.


When the aircraft is at high altitude, the pressurlzatlon systemsonstant altitude of0 feet In the cockpit and nose0 feet Inay. '

Cabin Pteaure Schedule Switch

The cockpit pressure schedule switchwo position toggle switch labeled CABIN PRESS located on the lower center of the Instrument panel. In the NORMAL, (down) position, the cockpitay pressuri-zatlon systems provide the normal pressure schedule and will maintain constant altitudes00 feet when the aircraft is0 feet. In0 feet (up) position, the cockpit pressure Is regulatedsi maximum differential and will0 foot cockpit altitude up to


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ret. 0 foot position leforduring subsonic low altitude ferry flights but Is not restricted for use as desired during climbs, descents and high altitude cruise. ate control lawhich limits the pressure change0 ft/mln when changing schedules.


During descents from high altitude, only the normal cockpit pressure schedule will provide optimum cockpit cooling.0 foot schedule will not cool the cockpit in descents as well as the normal schedule due to Increased turbine back pressure.

Cockpit Air Switch

The cockpit air switchhree position switch with labeled positions of NORM (left) OFF (center) and EMER (right). In the NORM position the left system shutoff valve Is deenerglaed to open and the left engine system furnishes air to the cockpit. In the OFF position the left shutoff valve isto closed, shutlng off the normal cockpit air. In the EMER position leftair is shutoff, the crossover valve in the right system Is energised closed and the right system shutoff valve Is deener-glzed to open and right system air iato the cockpit. The circuit Is powered by the essential dc bus.


In the EMER positionay system switch OFF position Isand right systom air must be shut off by moving the cockpit air switch to the NORM position.

'em Switch

ay system switch has two positions and Is located on the upper left side of the Instrument panel. In the ON (up) position the right engine system's shutoff valvo la deenorglzed to open so that right engine air can flow toav. If the cockpit air switch is In the crossover or EMER position this air will be ducted to dieand will enteray through the cockpit regulator valvlng. In the OFFthe shutoff valve is energised to offay system air Is shutoff If the cockpit air switch is in NORM position. TheIs powered by the essential dc bus.

Temper ot Lire Control Selector Switches

Two selector switches, one for the cockpit and one foray and/or emergency cockpit air, are Installed on the upper left instrument panel. Each switch has four positions; AUTOOLD (downARM (down right) and HOLDhe switches are spring loaded to HOLD from the COLD and WARM positions. The switches will normally be in the AUTOhowever, the pilot can manuallythe automatic feature by moving the switch to either the momentary COLD or WARM position. Tho manual COLD control will provide colder air. If required, than the automatic control. Thenverter powers the cockpit temperature control system. Thenverter powersay and/or emergency cockpit aircontrol system.

Terpp--rotuie Indicator Selector Switch

A temperature indicator selector switch located on the upper left Instrument panel allows the pilot to monitor cockpitay temperature. Cockpit temperature iswhen the switch is placed In the CK.PT (left) positionay temperature when











the switch IsinAY (right) position. Power for the indicator laby the essential dc bus.


Upoint, the Insulation and ventilation of the pressure suit will keep the pilot comfortableockpit environment that Is too warm. The temperatureis provided so as to allow anticipationemperature condition that might eventually become too hot for comfort. If the cockpit temperaturehe suit will not keep the pilot comfortable.

T en-pet at utt Control Rheostats

Two temperature control rheostats, one for the cockpit and one foray and/or emergency cockpit air are Installed on the upper left instrument panel. Arrowsthe direction of rotation necessary to Increase temperature. Generally, It la necessary to periodically rotate thetemperature control rheostatthe COLD position to maintain atemperature In the ventilated flying suit and keepay temperature In tolerance. Electrical power for the cockpit temperature control circuits is from thenverter. ay and/or cockpit emergency air control is powered by thenverter.

Pressure Altitude Gope

A cockpitay pressure altitude gage is located on the left forward panel andeither cockpitay altitude as selected by theay selector.

Altitude SelectotLever

This switch type lever is located on the left forward panel. It is labeled CABIN ALT in the up positionAY ALT In the down position and selects the respective pressure altitude to be indicated on the gage.

Depressor! rationSwitch

A two position lift-lock depressurlsatlon switch labeled PRESS DUMP and PRESS NORM is located on the upper leftpanel. When the switch is pulled out and moved to the PRESS DUMP position, both the cockpitay will be deprea-surlzed by the opening of the safety valves. When moved to the PRESS NORM position the safety valves will close and the cockpitay will repressurize.


Depressuriaation and repressur-lzatlon will occur at an extremely rapid rata.

Nose Hotoh Seol Shutoff Lever

A nose hatch seal shutoff lever, located on the forward right side of the cockpit,the noae hatch seal shutoff valve. It is normally In the ON position to allow canopy seal pressure to inflate the nose hatch seal. In the OFF position the nose hatch seal is Isolated from the canopy seal system. This prevents the deflation of the cockpit canopy seal in the event of excessive nose hatch seal leakage.

Nose Air Shutoff Hondle

A nose airandle Is located at the bottom of the annunciator panel. It lain the locked ON position. Tho handle is turned counterclockwise to unlock and then pulled out to shut off airflow to the pressurized nose compartment.


oxygen system and personal equipment

The aircraft Is equipped with dual liquid oxygen system*. Two liquid oxygenlocated In the right aide of the noee-wheel wellapacity of ten6 gallons) each. The liquid oxygen flow*, by gravity. Into the pressure buildup coil and vaporize* because of exposure to ambient temperature surrounding the coil*. The gas flows through the pressure closing portion ol the pressure control valve and theand gas ports of the fill valve and then back Into the top of the container where it collects and developsigher pressure. This cycle contlnuos until the systempressure Issi) at which time tha pressure closing valve closes and slope tlie flow of llqutd oxygen through the pressure buildup coll. The liquidwill now flow through the check valve and out the converter supply port to theheat exchanger. During periods of shut down system pressure will continue to rise because of normal liquid boll off. The Increase In pressure la sensed at theopening valve. Atal this valve opens dumping the gas back into the The pressure will continue to slowly rise, due to boll off, until It reaches reflief valve opening pressuresi. The excess pressure is ventedthrough the relief valve. Two ON-OFF levers for the two systems are located on the oxygen control Installed on the left console. The noodles on the pressure gage will fluctuate, indicating oxygen flow when the pilot inhales. Liquid oxygen Is warmed and converted to gas for breathingeat exchanger which consists of additional length of tubing In the supply line. Tbe low pressure gage on the oxygen control panelormal pressure0 psl.

Liquid Oxygen Quantity Gogs

The liquid oxygen quantity gage is located on the left side of the Instrument panel. It la calibratediter Increments The quantity gageouble needle type and indicates the quantity of

liquid oxygen remaining in therystems. Whened OFF Indicator at the bottom of the gage Indicate tlie gage Is not receiving power from thenverter.

Indicator Teit Switch

A red test button labeled IND TEST Ison the left *lde of the Instrument panel. When this button Is pressed thequantity gage needles will reduce As the oxygen needles approachiter mark the OXY LOW warning light will illuminate. When the button iathe gage needles will resume their original position. The CIT and apike and forward bypass position indicators are also tested by this button.

Oxygen lor. ifxiicotlng ug-'i

Two oxygen low warning lights are located on the pilot's annunciator panel. The lights are labeledXY LOW andXY LOW. Each light will Illuminate when oxygen pressure dropsal oriter or less remains In the system.

emergency oxygen system

Two independent emergency oxygen systems are Installed In the pilot's parachute pack. Each system consists5 cubic0 pel cylinder. The systems will supply oxygen simultaneously during bailout and when tho aircraft oxygen systems fail. An oxygen line is routed around each side of tha pilot's waist and connects to the suit controller valve. Emergency oxygen flow pressure la alLghtly lower than aircraft system pressure. Oxygen duration of each emergency system Is approximatelyinutes,



The emergency oxygen system la actuated either manually by pulling the conventional green apple, or automatically by the upward motion of the coat during ejection. The emergency oxygen syatem ahould beIf the aircraft systems are notthe deaired amount of oxygen or hypoxia or noxloul fumea are auapected.


A full preeaure ault ia provided which is capable of furnishing the pilotafe environment regardless of pressurein the cockpit. The suit consists of four layers, ventilation manifold, bladder, link net, and heat-reflective outer garment. The ventilation manifold layer allows vent air to circulate between the pilot'sand tho bladder layer. The bladder provides an air-tight seal to holdair In the suit. The link netesh which holds suit configuration inwith the pilot's body. The outer layer of heat-reflecting cloth provides some protectionot environment. Air prcssuro to the suit is regulateduit controller valve, located on the front of the suit just above the waist.

Pressure Suit Ventilation Air

Air for ault ventilation ia provded by the cockpit air-conditioning system. of the ventilation air cannot be varied except by changing cockpit Inlet air Ventilation airflow rate mayuit flow control valve installed at the hose connection point on the suit. Ventilation air and exhaled breathing air are exhausted from the suit.

Suit Ventilation Boost Valve lever

The suit ventilation boost valve lever,SUIT VENTIL BOOST. Is located on the left console. The lever is marked NORMAL, (aft) and EMERG (forward). the leverutterfly valve in the cockpit air-conditioning air supply line inay as to vary the pressure of the air available to the suit system. 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 thr NOHMAL position (used when engine rpm Is high) the cockpit air-conditioning line requires no constriction to provide sufficient airflow to the suit. At IDLE engine rpm the ventilation boost valve lever must bo keptf the way fromto EMERG In order to provideair for conditioning the suit andthe INS platform and Inverters Inay. During takeoff and normal flight the valve lever is kept in the NORMAL If the pilot suffers discomfort, such as might happenradual climb lo an extreme altitude or during low-rpm descents, the valve lever is gradually moved toward the EMERG positionomfortable pressure'and ventilation condition is The valve lever should not be moved toward EMERG more than necessary to provide pilot comfort; excessive suit system pressure, will unduly reduce the available refrigeration.

Suit Controller Volvo

AU four aircraft and emergency oxygenlines enter the controller valve at the front waist of the pressure suit. Thevalveensor thatairflow and oxygen to keep Internal

suit pressuresi (equivalent0 ft) in the event of cockpitization. resa-to-test button for each oxygen system Is Installed on thevalve, which allows the pilot to check suit inflation.


Leather gloves fasten onto the milt at the wrist rings. The inner liner of the glove in similar to the suit inner liner and will retain pressure.

Plate Hear Switch

A face plate heat switch is Installed on the right console of the cockpit. The switch has four positions: OFF, LOW, MED and HIGH. Heat may be regulated to defog the face plate as required. Dafogglng is accomplished by the combination of face plate heat andflow. The face plate heater circuit is powered by the essential dc bus.


The helmet head area is divided into two separate sectionsubberised cloth face seal. The front area between the face and the face seal receives oxygen from either the aircraft or emergency oxygen system through regulators built into the helmet. Oxygen flows across the face plate from the Inhalation valves inside the helmet andsome face plate defogglngit Is Inhaled. The rear area receives vent air for helmet Interior temperature regulation. The face Seal ia not positive; however, the pressure of the oxygen in the front area la slightly higher to prevent vent air from leaking forward. An external crank on the helmet la provided for head band adjustment. Buttons on each aide of the helmet, when actuated, will lower the face plate and visor. The face plate Is opened by moving the buttons and dumping the pressure, allowing the face plate to be rotated upward. If the aircraft oroxygen supply to the helmet Isor exhausted, the regulators In the helmet sense tha drop In pressure and the face plate seal deflates, allowing ambient air to enter tho helmet so the pilot will not suffocate.


The sock or boot liner fastens onto the suit at the thigh by meansipper. The boots are made of white leather for heatand fit snuggly over the aocks. pur that fastens to the seat Is attached to each boot.


When permitted by appropriateubstitute oxygen mask assembly may be used In placeressure suit for flights at low or Intermediate altitudes. Theconsistspecially designed oxygen maskxygen regulator, anti-suffocation valve and two oxygenleads with connectors for bothand emergency oxygen systems. In the event that the regulator shouldor the oxygen supply is exhausted, an antl-suffocation valve Installed between the regulator and the mask will sense the drop in oxygen pressure and allow ambient air to enter the mask to prevent suffocation.


A reinforced flberglas survival kit container fits Into the seat bucket and attaches to the parachute by anap attachments on eachoor oo the top provides access to the survival items stored inside. The kitstandard survival Items such aa radio,mirror, whistle, knife, matches, rations, water, compass and first aid kit. Various additional items depending on the terrain and soason may be

provided. The kit le peckedater proof bag attached0 foot retention lanyard. If an overwater flight laa life raft may be atowed on top of the plastic bag and attached to the lanyard. During ejection tho life raft inflating device la armed. Following ejection, the survival kit release handle should be pulled before reaching the ground. This action separates the survival gear from the pilot and Inflates the life raft. The survival gear and life raft remain attached to the parachute harness by the retention lanyard. apid abandonment of the aircraft on the ground, the survival kit release handle may be used to free the pilot of the survival kit {including the lanyard) without Inflating the life rait.


A special parachute5 foot canopy la used. The large canopyormal descent rate with the bulky personalrequired for high altitude flight. mall diameter, ribbon type stabilizing drogue chute is also provided. 0 feet altitude, the drogue chute Isfirst in order to stabilise free fall of the pilot. The drogue is automatically jettisoned0) feet after an aneroid controlled opener deploys the main chute. 0 feet the main chute only deploys Immediately. andle is also available for opening the main chute. The chute pack is equipped with conventional quick release bucklea. The emergency oxygen bottles are located between the chute canopy and the pilot's back. ombination hand squeezed bulb and manually operated pressure relief valve located adjacent to the suit controller is used to adjust cushion pressure ased knob located on the left harness strap is connected to the parachute timer arming cable and is used to actuate the timer when bailout is marie.


The windshield is composed of two glass assemblies secured and sealedhaped titanium frame. The glass surfaces are coated with low reflective magnesium fluoride. ollapsible vision splitter is also Installed on the windshield center line to minimize reflections.


The windshield defog system delivers hot air from both right and left air systems through check valves to defog the windshield and canopy. hapod air duct rune along the lower edge of the windshield. Hot defog air is supplied through this duct when selectedwitch that Is located on the upper left console. The air is directed to the windshielderies of holes on the upper surface of the duct. Holes are also provided at the aft ends of the duct to direct air toward the canopy glass.

Defog Switch

A three position defog switch is located al tha forward end of the upper left console. When held in the momentary DEFOG(forward) position the motor driven defog valve will open. Time of travel to full open Isoeconds. In the HOLD (center) position the valve will atop at any desired partial open position; in the OFF position the valve will completely close. The circuit Is powered by the essential dc bus.


Hot air la ducted from the L.roeaurl-zation aupply downstream of the fuel air heat exchanger and upstream of theregulator and air cycle refrigerator,











eries of orifice located on the left side of the outside center windshield support. The system Includes left and right solenoid shutoff valves controlledwitch in the cockpit. Power Is furnished by the essen-tall dc bus.

Windshield Deice Switch ond lodicotor Light

osltlon windshield delce switch is located on the upper left instrument panel. In the OFF (right) position the shutoff valves are closed and no delcing air is supplied. InN (center) position the hot air is furnished by the rightlow Is available for delcing. Int ON (left) position bothhut-off valves are opened and full flow Isto the windshield orifice*. Power for the switch and lights Is furnished by the dc essential bus.


onsiderable amount of air is used when operating the delcing system InN position. This may reduce the cockpitay air supply when operating In the lower ranges of engine rpm.

. The delcer Indicator light, located above the switch, will be Illuminated at any time the deice switch is not In tho OFF position.


A rain removal system ia provided for clearing the windshield when operating the aircraft in rain. Itank that Isby air from the windshield deicer System and the tank is connectedpray tube located on the left side of thecenter divider. ushbutton switch, located on the upper instrument panel, la used to spray the rain removal fluid onto the left windshield. Power is furnished by the essential dc bu*.


Do not apply rain repellentry windshield as prolonged obscuration may result.


The canopy consists of two highresistant glass windows securedeinforced titanium frame which is hinged at the aft end of two hinge pin*. Operation of the canopy is completely manual. Small holes In each side of the canopy areas lifting points from the outside. N. handles are provided on the Inside of the canopy for moving It up or down. rop assembly locks the canopy In the full open position. The canopy la secured in the dosed and locked positionour hook interconnected latching mechanism. Aboost counterbalancing system Is provided to aid in the manual opening and closing of the canopy. This nitrogen is also used* to force water into the map case when the destruct system Is actuated.


Actuation of the destruct system tends to deplete the nitrogen boost counterbalance system andthe manual force needed to open the canopy. Canopymay be necessary for rapid egress.

An Internal latching handle Is installedtho right canopy sill, allowing the canosy to be latched from the inside. An external fitting located on the left side of the alrcrsdb can be used to operate the latches from the outside.



The canopy should be opened or cloned only when the aircraft is completely stopped. Maximum taxi speed with the canopy open isnots. Gusts or severe windshould be consideredortion of thenot limit.

Conopy Latch Handle

A canopy Latch handle Is located under the right sill In the cockpit and rotates forward to lock. The sill trim Is cutout to expose the action of the locking lugs and pins as the handle is rotated forward. am over center action allows the handle to remain only in the latched or unlatched position. No canopy unsafe warning light Is provided.

ernal Letch Control

A flush mounted external latch fitting Is located on the left side of the aircraft and permits the canopy to be opened from the outside. The fittingnch square bar extension. Once the canopy is unlocked. It may be raised manually until the prop locks It in the open position.

Conopy External Jettison Handle

The canopy external jettison handle, located beneath an access panel on top of the left chine, permits ground rescue personnel to jettison the canopy. Sufficient cable length Is provided to allow the operator to stand clear of the fuselage during the jettisoning procedure.

Conopy Internal Jettison Handle

A canopyandle Is located on the left console wall adjacent to the pilot's leg. The handle can be used to jettison the canopy without Initiating the seat ejection System. The handle is held in the stowed positionockwlreround safety

pin. Storage for the canopy jettison and seat safety pins la provided at the forward end of the upper right console. Cable travel is approximately six inches.


An Inflatable rubber seal Is Installed in the edge of the canopy frame. The seal seats against the mating surfaces of the canopy sill and windshield to provide sealing for cockpit pressurlsation. The canopy seal preesurlzation lever above the forward right console operates the seal inflation valve. ose hatch seal shutoff lever is also provided to prevent deflation of the canopy seal in the event of nose hatch seal leakage.


The canopy jettison system is designed to unlatch and jettison the canopy from the aircraft by means of explosive Initiators and thrusters. The system consists of two Initiators which are Independently actuated by either the ejectioning or the canopy jettisonanopy unlatchanopy removalanopy seal hose cutter, cable linkage and gas pressure lines. Eithering Initiator or the canopy Initiator or the canopy initiator will fire the unlatch thruster which unlocks the canopy. This thruster then activates the canopy seal hose cutter and fires the canopy removal thruster which jettisons the canopy. Whenever the canopy la jettisoned by use of the canopy jettison handle, the canopy jettison Initiator gas pressureeat Jettison safety valve to prevent initiating the seat ejection sequence untiling ia pulled. fng jettisons the canopy as the initial step in the ejection sequence.


A manually extended rear view periscope is mounted In the top of the canopy to enable the pilot to see the engine nacelles and rear fuselage and rudder area. The periscope.

normally la lockedully retracted It ia moved by using the white nylon pad, mounted on the aft aide of the viewing tube,andle. Pushing the handle to the left unlocks the tube, allowing the periscope to be emended. Then, pushing theprlng-detented position nuke* the rear view available. Cockpit presaurc tenda to assist extension, and resists The diameter of the Instantaneous cone of view is approximatelyhowever, head movement extends the viewing cone lo approximatelytotal angle. Whenthe periscope can be rotatedto move the center of the viewing arc up to 10 from the aft ccnterltne. The de-magnification ratio of the lens system.


The ejection seat system utilizes an upward catapult and rocket thrust to providerisk ejection capability at ground level when airspeed is at leastIAS. The seat incorporates an ejection ring, headrest, knee guards, automatic foot retractors, automatic foot retentionilot-seat separation device, shoulder harness, Inertia reel lock assembly, and anopening seat belt. peed sensor mounted on the fuselage behind the seat automatically selects one of two seatdelays, depending upon airspeed at ejection. (Refer to Ejection Sequence this section.) Quick disconnect fittingson the seat rails and the floor of the aircraft permit disconnection of the oxygen, ventilated suit and electrical lines.

Seot Varllcol Adjustment Switch

The seat may be adjusted vertically by means of an electric actuator mounted on the lower end of the catapult. The three-position switch is located on the right side of the seat bucket. The seat moves In the direction the switch is moved. Power for seat adjustment is furnished by thodc bus.

Shoulder Harness Inertia Reel Lock Lever

A shoulder harness Inertia reel lock lever installed on the left side of the scat bucket is provided for locking and unlocking the shoulder harness. The lever has twoLOCK and UNLOCK. Each position is spring loaded to hold the lever in theposition. An inertia reel located on the back of the seat willonstant tension on the shoulder straps to keep them from becoming slack during backward movement. The rcol alsoocking mechanism which will lock the shoulder harnessorce has been exertedorward direction. When the reel Is locked In this manner. It will remain locked until the lever is moved to the LOCK position and then returned to the UNLOCK position.

Ejection (D) Ring.

An ejection ring, located on the front of the seat bucket. Is the primary control for ejection. An ejection safety pin ts Installed in the ejection ring housing bracket.


The aircraft are equippedackup secondary seat ejection system. andle for this seat ejection system Isand made accessible only by first pulling thelng.


The ejection seat must not be fired by pullingandlo while the canopy Is still in place. The pilot can not eject through tho metal canopy.

When the secondaryandle iseparate initiator fires the seat catapult and seat separation and beltinitiator.


Foot Spun

Foot spurs, attached to the pilot's shoes, are attached to tho ejection neat by cables. Normal foot movement Is In no waysince the cables,light spring tension, rool In and out freely. When the ejection ring is pulled, the knee guards rotate from their stowed position, the cables to the foot spurs are reeled In and the pilot's feet are retracted Into the foot rests. The foot cables areseveredet of cutters as part of the ejection sequence.

Monuol Coble Coffer Ring

The ejection seat Incorporates anmeans for cutting the foot retractor cables. ing, located to the right of the seat headrest, will actuate the cable cutters initiator If the automatic cablesystem falls or rapid abandonment of the aircraft Is required on the ground.


The ejection seat is providedilot-seat separation syatom which operates in conjunction with the automatic seal beltsystem. ind up reel Is mountedthe headrest,ingle nylon web ia routed from the reel to halfway down the forward face of the seat back. From this point two separate nylon straps continue down, pass under the survival kit, and are secured to the forward seat bucket lip. After ejection, as the seat belt is released, an initiator actuates the wlndup reel which winds the webbingross-shaft, pulls the webbing taut, and causes the pilot to be separated from the seatling shot action.


The ejection seat Is equipped with anopening seat belt which facilitates pilot separation from the seat following ejection. Belt opening is accomplished automatically a> part of the ejectionand requires no additional effort on the part of the pilot.


If the pilot Is wearing an automatic opening aneroid type parachute, the parachuteanchor from the parachute aneroid must be attached to the swivel link. As the pilot separates from the seat, th* lanyard, which Is anchored to the belt, servestatic line to arm the parachute aneroid. The parachute aneroid preset altitude ia0 feet.


Pullinging Is normally the onlyrequired to Initiate pilot ejection and results In firing both the canopy Jettison and ejection seat systems. All resultant action: will occur automatically andpecific sequence as explained below.

ing cable fires the ejectionInitiator, actuating the canopysystem and the leg guard thruster. The leg guard thruster rotates the knee guards, retracts tha pilot's feet, activates the cable cutter backup initiator and locks the shoulder harness. Movement of the canopy jettison thruster (final step In canopy jettisonactuates an Initiator whichecond delay catapult initiator and arms the speed sensor. econd delay assures complete canopy separation prior to seat ejection. Gas pressure from the catapult initiator fires the rocket-catapult.

econd neat separation delay Initiator, and enters the speed sensor. If airspeed isIAS, the gas pressure passes through the speed sensor and firesecond delay seat separation initiator. If airspeed IsIAS, the pressure is blocked by the speed sensor.

Initial seat movement upward on the rails disconnects normal oxygen, ventilated suit and electrical lines, and activates theoxygen supply. IAS eitherecond delay may be experienced because of the speed sensor tolerance.

Eitherecond delay initiatorIAS) orecond delay initiatorIAS) actuates the cable cutters, releases the pilot's feet, opens the seat belt and fires the seat separation system.

A static line attached to the seat belt is pulled as the pilot separates from the seat and activates the automatic parachute

If theing ejection sequence was not accomplished, th* canopy must beeither by use of the canopy jettison system or manually. Pullingandle initiates the secondary seat ejection


andle backup ejection sequence does not rotate the knee guards nor retract the foot cables Seat separation delay time willeconds regardless of airspeed.



Refer toor Operatingand Limitations.


Refer toor Takeoff and Landing Information.


PLANNING Refer to Appendix L

toor Weight and Balance Limitations. For detailed loadingrefer to Handbook of Weight and Balance Data. Before each flight, check takeoff and anticipated landing groas weights and weight and balance clearanceF).



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Refer loor engineering,and equipment statu*.


It la not practical for the pilot to perform an exterior inspection while wearing asuit. The exterior inspection should be accomplished by other qualifiedi



A Ladder platform stand which overhangs the chine la used to gain entrance to the cockpit. The canopy is unlatchedby rotating the external canopyclockwlae withnch square bar. The canopy Is manually raised to the full open latched position.


Manual cable cutterecure.

Ejection seat and canopy safety pinsheck.


All circuitn.



Landing gearOWN.

Batteryxt pwr.

Accomplish and check personalhookup. (Hookup will beby personal equipment Refer to.

Suit vent boostetever travel.

Left Comole

N. Set to proper mode and code.

Panel and Instrument lights switches -As desired.

COMM selectorHF.

External light selectorFF.





bypasaoth CLOSED.

Instrument Panel

ay altitude selector lever -CABIN.

Landing and taxi lightFF.


Cockpit temperatureUTO.

temperature switch - ay airN.

Cockpitay auto temperatures desired-


Cockpit airN.

Pressure dumpK If.

Drag chutetowed.

Windshield deicerFF.


Compressor inlet temperature gage -Check needles together and indicating ambient temperature.

Igniter purgeFF (down).

Compressor inlet static pressure gage Check needles together and indicating barometric pressure.

heck for proper Indication.


Periscope MIR SELull- (Projector).

Fuel derichment armingFF.



Inlet air forward bypassUTO.

Emergency fuel shutoff switches -Fuel On (guards down).

Cockpit pressure schedules





Spike and forward bypass position- Check.

Fuel transferFF (guard down).

Fuel dumpFF (guard down).


Air refuelFF.


De structFF (guard down).


Nose hatch seal pressureN.

Pitot pressure selector lever -NORMAL.

Canopy seal pressureFF.

Stability augmentationFF.


Inertlal navigation systems required.





Autopilot and attitude reference selectors desired.

BDHI needle selectorACAN.

TACANnd tuned to desired station.

ADF receiverNT.

Floodlights desired.

Face plate heats desired.

Flight reference system (FRS) compass selectAG.

and SIP power switches -


Lower Instrument Porte 1

Surface limit releaseulled out.

Pitot heatFF.

Hydraulic reserve oilguard down).

Trim powerN-

Nose air conditioningtowed.

Backup pitch damperFF (guard down}.

Pitch logic overrideFF (guard down).

Yaw logic overrideFF (guard down).

Gear releasetowed.

equipment function check


N% and tank lightsest.

quantity indicatorsto zero.

QTY LOW warning lightIlluminate.

Crossfeed and boost pump switches -Press lights on.

Pump releaseUMP REL, then release.

Tank boostheckANK lights on (automatic sequencing).

Crossfeedress (check light off).

quantity indicating system



and warning lights test switch

warning and fire lights

gear unsafe warningsound.


quantity needles willbelow 0.

indicator will decreasezero.

c Spike and forward bypass position indicators increase to maximum forward indication on spike and maximum open on forward bypass.

>- Headset plug and oxygenonnect (if pressure suit is not used).

ndxygenN (if pressure suit la notheck system pressures.

1. Tape and flightN.

starting engines


Before starting an engine,that the wheels are firmly chocked since brakes areuntil hydraulic pressure is available ond no parking brake is installed.

. Determine that intake and exhaust areas are clear of personnel and ground equipment. The ground personnel using Interphoneequipment will be in position to observe the exhaust nozzle and nacelle inspection panels during starting.

. Do not move the control stick until at0 psi hydraulic pressure can be maintained onystem gagesontrol system Inspection will be necessary.

Check with INS crew prior to starting engines.

Fuel low pressureff,


Ground startinanetruct ground crew to rotate engine for start.

DLE when rpm la Indicated.


Engine light up will be Indicatedecondsontinuous rpm Increase andise In EGT.

heck formax during acceleration.

9. Ground startingignal ground crew for starter OFFpm.

10. Idlepm.


Idle rpm increasespm perabove

1. Engine and hydraulic pressure- Check normal,

heckpounds per hour).

- Check



Discontinue start if oil pressure rise is not observed withineconds from start of rotation.

d. Hydraulic system pressures -Check.


Start other engine using above

If engine does not accelerate smoothlypm,throttle to OFF and then quickly advance to IDLE. This "double clutching" momentarily leans the fuel:alr mixture and properly positions the flame front In the burner cans. Count as another TEB shot.

14. TEBheck.


If throttle la inadvertently retarded to OFF do not advance In an attempt to restart engine. In case of false start use engine clearing procedures, this section. Afterburner duct must be visually checked and un-burned fuel removed prior toanother start.


alse start occurs, trapped fuel and fuel vapor may be removed from engine by using tile following procedure:


Ground startingN for1 minute. Then signal ground crc* for ground startingFF,


Do not rotate the engine with fuel shut off (Emergency Fuel Shutoff- UP, Guard up) except In case of emergency( because damage to the engine may result.

before TAXIING

UHF andheck.

s required.

GeneratorESETat idle rpm. Check with INS crew prior to resetting.

BatteryAT (within 3

Generator outheck Off.


If the generator out warning lights fail to extinguish, return the battery switch to the EXT PWR position and repeatbove.

DEST /FIX switch -

INS modeAV. Check with INS crew prior to actuating switch. Press tlie STORE button and check BDHIteering needle forright indication and Distance Toautical mile readout.

INSeport Destination Coordinates, Distance To Go and Groundspeed when slewing is completed.

INS DEST/FIXelectFIX and press STORE button. Check INS FIX REJECT light on.

INS DEST/FIXelectDEST and press STORE button. Check INS FIX REJECT light off.

INS umbilicalheck(confirmed by INS crew).

Externalignal for disconnect.

Inlet air forwardheck open. Ground crew will confirm open.


SAS channelll ON.

SAS recycleress (all lights should go out).

SAS light testress (all lights should illuminate).

Autopilot pitch and roll engage switches -ON.

Autopilot disengage switch (controlress. Check that autopilot disengages.

SAS channelFF. Pitch andnd Roll disengage lights illuminate. Both MON lights must stay out.


2). Surfaceheck for properwith ground crew and eel to zero.

Controlheck for properof movement. Individually check each axle in both directions and have ground personnel verify proper deflection of control aurfacea.

Packages required.

Canopy and seat safetyemove and stow.

lose and lock.

Canopy seal pressureN.


Navigationheck operation of ADF, TACAN, and INS.


All taxiing and turns should beat slow speeds ao as to limit aide loads on the landing gear. Fast taxiing should also be avoided to prevent excessive brake and tire heating and wear.

before takeoff


The canopy should be opened or closed only when the aircraft Is completely stopped, Maximum taxi speed with the canopy open is approximatelynots. Gust or severe wind conditions should be consideredortion of thenot limit taxi speed.

Rear viewheek.

Taxibtain clearance from control tower.

Chocks and downlockignal for removal. Observe ground crew for clearance to taxi,

Nosewheelngage and check operation.

TAXIING 1. heck.

s required.


If engine trim run is required, EOT values appropriate for ambient temperature will be supplied during preparation for flight.

During trim run at Military rpm:

ay auto tempif necessary.


Adjust both controls towardtemperature positions If necessary, to eliminate cockpit fog if fog la encountered at lower temperature settings. 0 o'clock settings are normally aufftcient. Lower temperature settings are desirable when local humidity and ambient temperature conditions permit, in order to assure personal and equipment cooling.

channelll ON.

SAS recycleress, If necessary (lights should go out).

Surface trimheck for aero Betting.

Tanksheck ON.

heck and fix as required. At designated runway position, selectSTORED FIX position and fix. Check INS FIX REJECT Light off. Select STORED MAN. Reset DEST/FIX briefed initial destination position, and store. Check distance to go aftercompleted, then reset DEST FIX to STORED AUTO if desired.

heck. Check andFRS and check INS if Return INS mode selector switch to desired position. Check Standby Compass against runway heading.

Pitot heatN.

Warningll Off.

External lightsCN (If


Flightycle and checkpressures,

Suit vent boostORM.






Engine turbine life can bedecreased by too rapid throttle movement. The time for throttle advancement from IDLE to MILITARY should be no less than one second,

4. elease0 rpm.


The tires may skid if the brakes are held on at high thrust.

heck atthrust.



- Advance to afterburnerposition after enginesrpm.

N and


Fuel derich armingRM.

Elapsed timetart.


To prevent overspeed, afterburner ignition must not be accomplished before the engines reach MILITARY rpm.
















Afterburner Ignition should occureconds.

Abort the takeoff if one or both afterburners do not ignite.

Advancing the power lever to initiate afterburning results in momentary noetic excursion, and engine transient speed oscillation maypm.


9. heck Indicated air-apeed against computed acceleration check speed at selected acceleration check distance. Refer to performance data. Appendix I, ior takeoff

10. egin at computed airspeed approximately five seconds before reaching takeoff speed. Apply smooth, constant back pressure on the stick so that required stick deflection andto takeoff attitude occurs atspeed. Refer toorand takeoff speeds.


The time for throttle advancement should be no leas than one second.


Use indicated airspeed during takeoff and climb until proper climb schedule speed is reached on the triple display indicator.

echeck atTHRUST.


Exact readouts on theseIs time consuming. The readout should be anticipated and needle position checkedlock position. If there Is anyof Improper engineduring powerthe 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. control can be maintained with nosewheel steering up tolift off speed.

During crosewlnd takeoffs the aircraft tonda to weather vane into the wind. This will be noted when tho nosewheel lifts off andsteering Is no longer available. pressure must be held to counteract the crosewlnd effect. efinite 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 noshould be encountered when taking off during reasonable crosewlnd conditions.

rotation technique

During takeoff, the maximum load on the main wheel tires occurs during rotation to takeoff attitude.


Avoid abrupt rotation since thia can Impose an excessive load on the tires and cause blowouts.

In general, the tires are more criticaltakeoff than at landing because of the higher ground speeds and gross weights Wing lift quickly relieves the gear load as the noae is raised. Start rotation approximately five seconds beforethe scheduled takeoff airspeed. nosewheel lift off should be avoided because the unnecessary drag extends the ground run. Delayed rotation also extends the ground run and may result in excessive tire speeds.

after takeoff

When definitely airborne:

1. Landing gearP.


The gear will retracteconds. Observe landing gear limit speed while gear is extended.


Single engine operation Is critical Immediately after takeoff. airspeed and decreasing angle of attack has greaterthan gaining altitudeaximum rate.

After gear retraction is complete: 2. limb power.

Minimum afterburning ie normally set after takeoff. When flight plan deviates from normal climb procedure,maximum afterburning or reduce power in accordance with alternate plan.

Engineheck. At:

Surface limiter releasengagt

Rotate handle counterclockwise and stow to engage limlters. Check SURF LIMIT warning light off to confirm

- Establish climbnormal operation:

KEAS while below.


et" Hg at. Above, with CITto


normal climb

The normal climb procedure optimises power and airspeed schedules for aupereont range and is applicable to climbs afteror air refueling. Use of alternateis permitted, but results insupersonic range capability. The general technique for airspeed and power scheduling Is as follows:

a. After takeoff, accelerateEASlimbing flight path, then climb with minimum afterburningEAS. Intercepteet and readjust climb attitude to5 Mach numbe) The autopilot KEAS Hold and Mach Hold

features may he used for thia climb phase. Adjust throttles to maximum afterburning at0 feet.

refueling, set maximumpower and accelerateach number, then5 Mach number. WhenMach Hold feature isMach Hold at

, level out momentarily,the autopilot, and push's. Establishfpm0 fpm rateEAS. Planso as to avoid


It Is mast Important to exceed5 early In the descent, and5 before atartlng the pull-out with sufficient airspeedso as not toEAS.

t.0 ft. maximumband is optimumide range of ambient temperatures when rales0 fpm0 fpm are0 ft.0 fpm may be favored with tropic hot day temperatures. 0 ft.0 fpm may be used with good results when ambient temperatures are below standard.


When possible, check EGT trimstarting the transonicmaneuver. Abnormally low EGT degrades performance.

5 isEAS, start aso as not topeak load factor of upbe required as level attitude is ClimbEAS,autopilot KEAS Hold feature as

, reduce power to obtainflow reduction0 toper hour per engine. KEAS to.

,0 ft.power to maximumand begin decreasingKEASach increase. Ifthe autopilot KEAS Holdthe speed decrease

cruise Mach number and/oraltitude are approached,so as to end the climb andclimb as briefed.

The following procedure Is recommended after air refueling or when the after takeoff procedures are completed:

After refueling, or atfter takeoff:

aximum afterburning.

Mach Hold may be used if desired.

Check as briefed.

ay auto temp controlsto Individual settings as required.

4. HF radio and


5- EOTheck.



Allowable rpm vs CIT tolerance ii rpm.

ormal EGT is in


above table, or base trim check onsupplied by tanker while refueling.


- StartEAS.

Disengage autopilot and0 fpm0 fpm rate of descent. After5 attained, round-out toclimb speed. Do notEAS.



- Reduce fuel flowper engine.


bypasset bothpen).


At approximately tothere willlight but noticeable yaw as the

compressor bypass bleeds open

if the left and right engines do not

operate on exactly the same


heatFF below.


The PITOT HEAT warning light will illuminate if pitot heat Is left on abovehile climbing.

IFFVSIFs briefed at.

Beacon and fuselageff abovr.


aximum afterburning.


Decrease KEASnotsach number increase in speed above. The KEAS Hold function of the autopilot should maintain this schedule automatically if engaged.

Mach No.


bypasset bothpen).


bypasset bothposition.

Reduce equivalent airspeeds If climb is to be continued after reaching the desired Mach number.

Desired supersonic speeds may beby throttling to partial afterburning settings. Maintain EGT by use of trim switches.

17. OscillographFF or briefed.


Deviations from Normal Climb procedures are permitted when limitations ofre observed. Maximum Thrust may be used continuously, but fuel economy will be less than for normal climb procedures. See. The recommondetl Military Thrust climb speedEAS. EGT can be expected to decrease il CIT decreases. The recommended Maximum Thrust climb speed for subsonic operationEAS to approximately, andbove that altitude. WhenThrust is used continuously afterefinite rotation is required to establish initial climb attitude. Beginsufficiently in advance of reaching the climb speed schedule to avoid overshoot Refer toor climb performance.


Transonic accelerations can be started by usingEAS climb speed schedule, starting at0 feet, or byavel transonic acceleration at an altitude00 feet.

Clifbing Acceleration Procedure

When this procedure is used, accelerate from takeoffEAS and rotate to climb attitude.


Begin the rotation sufficiently in advance of reaching climb speed to avoidEAS. IfIs delayed, it Is possible to overshoot the airspeed by anamount.

EAS at0 feet and climb at this speed using maximum afterburning thrust. Mach number willwith altitude andill be reached0 feet.


Observe limitations of Section V.

Center of gravity control Is important for optimum cruise performance. Fuel load distribution and automatic tank sequencingorward eg for takeoff and Initial climb. During supersonic climb and cruise, automatic sequencing provides an aft eg tolevon deflection and resulting trim drag. Supplemental manual control of fuel usage is also possible, but should only be used In tha event of malfunction of the automatic sequencing system.


Spike and forward bypass knobs must be in AUTO position when cruising0 feet.

For long range operation,hrottle setting for the applicable cruise KEAS/altitude weight schedule; then only make minor adjustments as necessary to maintain the schedule.


Exhaust gas temperature and engine speed limits vary with CIT. Refer to EngineLimits, Section V, for limit



As Mach number is increased, caution ia required in Ihe rate of throttle movement following afterburner Ignition and during afterburner shutdown.


Oil pressure should be monitored closely. Mach number should be reduced If pressure does not remain within the limits listed Inr if the oil temperature warning light illuminates.


Bet rimming of EGT should not be required prior to start of descent unless manual up-trimming has been accomplished during climb or cruise. The amount of downtrlm required will be approximately equal to the total prior uptrlm. Pilot Judgement must govern its use. eneralGT at start of deceleration should prevent overtemperature conditions and provide normal engine operation at lower Mach Rctrim if necessary and accomplish the following before descending in order to obtain scheduled descent distance.

lowly retard to minimum afterburning position.

Spikeheck AUTO.

Inlet air aft bypassheck normal schedule.

Inlet air forward bypassheck AUTO.


Aircraft deceleration rates are limited by maximum tolerable temperature transients within the engines. Engine cooling rates will be satisfactory when deceleration rate Is not greater than prescribed in Section V. Use of Military Thrustpeed scheduleEAS during deceleration toatisfies this requirement when the spikes are set in AUTO. Descents can be made at engine speeds below the Military rpm schedule betweennd. Throttles may be set as desired at subsonic speeds.


Monitor fuel tank pressure during descent, and reduce rate ofif necessary In order to maintain positive fuel tank

A high descent rate belowanthe LN2 system ability to pressurize the fuel tanks. Negative pressure allows atmospheric oxygen to enter the tanks through the vacuum relief valve. If fuel vapor temperature in the tanks la high, above approximately(orand tank internal pressure is equivalent0 feet pressure altitude, orritical percentage of oxygen can result in fuel vapor Ignition.


In the event of inlet roughness set thebypass doors open, then set the spikes forward and increase rpm If necessary. Refer to appendix for normal descentand for performance withbypass open.









- Auto hadura


AH- NotPoLtchedule


oei lo0 fur4,5



Aft bypassosition A.

EGTown trim if required.

- AdjustEAS. cruise altitudeEAS Rate of decelerationexceed allowable Mach rate.

Fuel tankheck. At:

Aft bypassosition B. At:

- Adjust0 rpm. at0 rpm


Set forward bypass open, spikes forward and Increase rpm as required If inlet roughness is encountered.

IFF/SIFs briefed at.

INS modeRS.



The PITOT HEAT warning light will illuminate If pilot heat ia left off belowhile

lightss desired.


bypass switches:

- Check0 or above.

Maintain at0 during remalnce: of descent to subsonic speed.


djust as required. Rate of descent must not result In negative fuel tank pressure. Avoidpm to prevent cycling of engine start bleed valves.

djust aa desired.

Forwardheck closed

DefogN and HOLD if re-required.



Use pitot static system for descent.


Either of two methods of handling power dux ing refueling may be used. Whenever the Initial fuel quantity remaining Isounds it is possible to use minimum afterburning on one engine and less than Military thrust on the other. This allows refueling to be accomplishedonstant altitude of0 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


air refueling director lights



section n

power limited at Military thrust. The conventional procedure of completing refueling without use of an afterburner can also be used;obogganeet will be necessary after the tanks are filled

Prior to air refueling, stabilize and trim at refueling speed for contact. Observe the tanker for director light signals and aas directed by the lights. Aconnection is confirmedild jolt to the aircraft, steady illumination of thelight panel and the extinguishing of the READY light. Slight maneuvering may be necessary at this point to Illuminate the azimuth and elevation neutral lights during fuel transfer. Contact can be maintained between the aircraft and tankerurn orescent. 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.


Accomplish the following prior to refueling:

Radars briefed.

Air refuelEADY.


Amplifier requires up tofive minutes for warmup.

transfer0 lbs).


If lessull fuel load is on-loaded, it is possible for anaft c. g, to develop.

Fuel quantity indicator selector -TOTAL. Monitor total fuel quantity.


When in observation position afterwith tanker.

UHF radio INT-EXT mode switch -INT.

READYush on (green) if

transfer switch - Stabilize in pre-contact position.

Beacon lightUS.

Observe tanker director lightsand boom in ready for contact position.


Normal refueling is accomplished as follows:

Establish contact. After contact Is made:

READYheck out.

Total fuelonitor. When refueling Is complete:

Control stickress.

Air refuelFF. Check ready light off.

Tanksheck ON.

Trim engines to EGT supplied by tanker.

H. RadarFF.

Inydraulic pressure is lost, Rmay be utilized for refueling by moving the brake switch to ALTRAKE position.



Do not leave the brake switch in the ALT STEER A> BRAKE position after refueling.

alternate refueling procedure

The boom may be latched in the refueling receptacle manually as an alternateby using the following procedure:

Air refuel switchMANUAL. Check READY light on.

Control stickresa and hold.

When nozzle has bottomed In the receptacle:



6. Accomplish steps,f Normal Procedure.


alfunction occurs whichdisconnecting the boom, place the Air Refuel switch In the MANUAL position and depress the IFR DISC trigger. If disconnect is notproceed with brute force pull out by retarding throttles.

before landing0 feet:

1. Cockpitay auto temp controls -Adjust to approximately two-o'clock position or as required to avoidfog.


the disconnect trigger isbefore the nozzle is in the bottom of the receptacle. It Is possible for the nozzle to damage nozzle latches, preventing any further refueling.

Fuelonitor TOTAL fuel. When refueling Is complete:

Control stickress.


The automatic limit disconnect system Is inoperative. Allmust be Initiated by the receiver aircraft, since the tanker operator Is unable to release the nozzle latches during manual boom latching.

.ay and cockpitto avoid equipmentIf possible.

. Keep UHF radio transmissionsecond maximum if possible while defogging step is employed.



Whenontains fuel,00 lbs forward tolight nose up pitch trim.

Surface limiterull out and rotateCW at.

Periscope MIR SELull


PATTERN (Typical)


Fuel transferFF.


Shoulderanually locked.



Traffic pattern0 feet above field elevation.

0 feet above field elevation.

Landing gearOWN (check gear down and locked).


Normal gear extension time Is approximatelyeconds. gear limit speed with gear extended.


FinalIAS minimum0 pounds of fueL


Base minimum final approach speed on Intended touchdown speed. Do not use maximum performance final approach speed unless operating conditions require minimum roll or runway is wet or Icy.

Seeypical landing pattern.

16. Landing and taxi lightss required.


Refer to the Appendix for landing ground roll distances. If airspeedigh sink rate will develo resultingard landing. During the flare, throttles are reduced to IDLE and touchdown Is made at approximatelypitch angle (nose approximately on the horizon).

The following procedures should be

1. DLE.


Throttle movement should follow quadrant curvature so that the hidden ledge at the IDLE position can prevent inadvertent engine cutoff.

Touchdowna required.

Hold nosewheel off.


Fuselage angle must not exceed 14 to avoid scraping the tali.

Drag chuteull to deploy. Chute deployment takes approximately three seconds.

Lower nosewheelIAS.

Engage nosewheel steering forcontrol. Steering will notuntil rudder pedals align with nosewheel position (straight ahead) and weight of aircraft is on any one gear.

pply after chute deployment. Moderate braking may be used prior to chute deployment.


If the chute does not deploy observe the brake energy limit speeds in Section V. Brake switch shouldin the ANTI-SKID position If runway ia dry. Refer to Drag Chute Failure, Section UL


Drag chuteurn and push to jettison chute.


The drag chute should be jettisoned while the aircraft still has forward motion to prevent drag chute collapse. The aircraft should not be taxiiedollapsed drag chute.




The traffic patternrosswind landing should be normal, making proper allowance! for velocity and direction of the cross wind. Proper runway alignment on final approach can be maintained by crabbing or dropping one wing;ombination of the

two is recommended just prior to flare. Remove crab before touchdown, using wing low technique to prevent side drift. Reduce sink rateinimum to accomplish smooth touchdown. At Increased cross wind components, sink rate must bedue to Increase of side loadson the landing gear. With more0 knot crosswind component it may be advisable to lower the nose and engage



"leering prior to drag chute With lees thannot crosswind component, rudder control Is sufficient to offset the crosswind effect on the drag chute.


Set brake switch NORMAL and, when field length would be critical in tbe event of drag chute failure, use minimum roll technique. Landing roll will increase due to reduction In available braking force. Use lightest brake pressure consistent with stopavailable.


Tests indicate that the aircraft will plane with heavy wateron the runway. With this condition, directional controlrosswind may be difficult.

maximum brakingchute deployment. may be used prior to chute

to IDLE during flare orafter touchdown.

engine throttle OFF after


Retarding both throttles to OFF further reduces thrust, butnosewheel steering and braking. If the brakes are burned out at the end of the runway, and speed willafe turn off, the nosewheel steering system will "save" the landing.

Throttle technique depends upon the pilot's judgement of the particular field conditions.


Same as wet runway oxcept brakingIs further reduced.


touchdown close to the end ofat minimum airspeed. uccessful short

the drag chute as quicklyafter touchdown. Lowerwhile the chute Is deploying.

Engine shut down will result in loss of hydraulic actuating pressure for the following systems:

engine shutdownbrakes and

englno shutdownand anti-skid brakes.


A go-around may be Initiated anytime during the approach, or during landing roll when sufficient runway remains for takeoff.

Drag chuteurn and push to jctttaon chute. If deployed.

TV. rotILITARY thrust.thrust If required.

Landing gearP after positive climb established.

s necessary.



SAS channel switches- OFF (before taxiing).

i. Lightings required.

vent boostet.

cockpitayfor comfort and



The engine should be operated at IDLEinutes (including taxi time) before engine shutdown to permit uniform turbine cooling and prevent possible rotor seieure.


Canopy seal pressureFF.


s briefed.


The INS should not be operated moreinutes after opening the canopy to avoid tlie possibility of excessive INS component

If taxiing with the canopy open is desired, the canopy should be opened only when the aircraft is completely stopped and canopy seal pressure ia off. It should only be opened If both engines and both air conditioning systems are operating normally and after the normal cockpit post-flight check of INSay and associated equipment has beenand this equipment turned off. The maximum taxi speed with the canopy open and latched tsnots. Gusts or severe wind conditions should be consideredortion of the limit taxi speed.


In the event of engine fire during shutdown, the engine can be motored with fuel OFF to blow out fire If starter unit is Refer to Section TH.

Igniter purgeUMP. Holdeconds.


Externalonnect, If available

BatteryXT PWR or OFF as required.




mergency procedures




Abandoning The




Double Engine


Afterburner Nozxle

Fire Warning-Takeoff






Bailout With Seat

Fire Warning

Smoke Or


Definitions Of Landing

Brake, Steering, Or Tire



Drag Chute


Fuel Pressure

Landing Gear and

Main Or Nose Gear Tire

Emergency Gear


Fuel Dumping

Normal Fuel

Emergency FuelFuel Transfer and


Forced Landing Or


Propulsion Syi'em

Duct Unstort

Inlet Control Malfunction Automatic Spike Control Automatic Forward Bypass Control Operation With Manual Inlet Control

Inlet Unliable

Failure Of Spike To Schedule Or Unliable

Compressor Stalls

Acceleration And/Or Over trim Compressor Stalls In Descent

Engine Flameout

Double Engine Flameout


Single Engine

Simulated Single Engine

Single Engine

Landing Gear System

Gear Unsafe

Gear Emergency Extension Partial Gear Landing Main Gear Flat Tire Landing Nose Gear Flat Tire Landing Heavy Weight Landing




This acction provides recommended procedures for use in the event of emergency or abnormal operating conditions. It does not cover multiple emergencies. Pilots must recognize that single malfunctions will often affect operation of other aircraft systems and require corrective actions in addition to those containedpecific emergency procedure.

of Checkliih

Critical emergency checklist items are those actions which must be performedif an emergency is not to be These steps appear in CAPITAL, letters to permit Immediate identification. They must be committed to memory toaccomplishment without reference to the Abbreviated Checklist.

Definitions ol Landing Situation!

The terms 'landwhen practicable" and "land as soon aa possible" are not used interchangeably. Tbe direction to "land when practicable" means to land at home base or other suitable alternate. Airis allowed when necessary tn order to reach the suitable destination. Alteration of the original flight plan may or may not be required, depending on the flight limits which are imposed because of the emergency or abnormal operating situation.

The direction to "land as soon as possible" means land at the nearest suitable facility.




If there is evidence of fire during ground start, attempt to keep the engine rotating until the fire is out. Apply chemicals from outside the engine onlyast resort.

ire ia evidenttart, or on notification:




Continue motoring the engine when the starter remains engaged and fire Is contained in the tailpipe. If the starter unit has disengaged, it can not be re-engaged until the engine has comeomplete stop.



Abandon aircraft.




Use applicable steps of Engine Fire During Ground Start procedure.


In an emergency requiring groundthe primary concern is to leave the immediate area of tho aircraft ae soon ae possible. The following procedures should be used when fire or explosion are probable. Salvaging emergency and survivalhas not been considered. Theseprovide the fastest means ofthe aircraft and they should beas rapidly as possible after the decision to abandon the aircraft Is made.

This procedure may be Initiated while the aircraft is in motion; however, the lap belt should remain fastened until the aircraft Is stopped.

To accomplish an emergency exit on the ground) proceed as follows:

1. Ejection seat safetyime permits.

Survival kit releaseull.

Seat belt and shoulder harness -


Parachute harness attachments -

Footanually release, (use cable cutter if otherwise unable tospurs).

nlatch or jettison as

Evacuate aircraft.


Without anti-skid operating, extrememust be utilized to prevent wheel skid, as skidding is hard to delect due to aircraft size and weight. Tires may failkid condition can be recognized andain landing gear tire blow-out may be sensed by the pilothump or muffled explosive sound.

If the ANTI-SKID OUT warning lightor anti-skid braking is not effective:


If normal brakes and/or nosewheel steering are not effective, orystem hydraulic pressure is not available:



U both engines are shut down with the aircraft moving, the brake switch should be left in the ANTI-SKID OFF position and steady brake pressure appliedomplete stop. The brakes should not be pumped, as accumulator pressure would be lost.

At landing weights, the aircraft can be laxied safely so long as one tire per main gear remains inflated. At takeoff weights, taxi distance should be minimized if one or two tires per main gear are flat in order to minimize the probability of further tire failures. Taxiing as necessary is permitted tounway with all tires failedain gear, as the massive tire bead tends to protect the wheels for some distance.


In the event that qualified ground personnel are not available, emergency entrance to the aircraft can be accomplished using the procedures illustrated by.


crash rescue procedures

osivfiociwiM io





The components considered as parts of the propulsion system Include the main engines, afterburners, inlets, nozzles, tailpipes, fuel controls, and fuel-hydraulic,and ignition systems. If abnormalof any of these components isprior to reaching the acceleration check distance, the takeoff should be Refer to ABORT procedure, this sec-Lion. The following procedures apply after satisfactory completion of the acceleration check.

thrust failure during takeoff, takeoff refused

If the acceleration check speed is marginal, or If the thrust of either engine decays or falls, and conditions permit:


Refer to abort procedure, this section.

engine failure immediately after takeoff

If an engine falls Immediately after takeoff and the decision is made to continue)Maximum thrust on the operating Lateral and directional control can be maintained when airspeed remains above the minimum single engine control speed. See. However, ability toaltitude and to accelerate or climbon weight, drag, altitude, airspeed, and temperature. Refer to the appendix for takeoff climb capability data. When at heavy weight for the existing airdumping fuel may reduce weight sufficiently to remain airborne.

If able to maintain altitude or accelerate:


Rechcck position of both throttles to assure that maximum power is being obtained,



Fuel dumpUMP (if

Fuel dumping in addition to consumption by operating engine lightens the aircraft at an appreciable rate. If turning at sufficient speed, theengine will also discharge fuel from Its afterburner.

s necessary.

Bank and sideslip toward the operating engine as necessary to maintaincontrol and minimizeegrees of rudder trim with bank and sideslip as needed to maintain course yields minimum drag in the critical speed rangeIAS.



Positively Identify the failed engine before retarding the throttle.

If not mechanical failure:

AIR START (refer toProcedure this section).

For obvious mechanical failure:

fuel shutoff switch -








r sior.su p





ouble engine failure occurs,oil owe:




Decay of engine rpm will result In rapid lossydraulic system pressure and subsequent loss of aircraft control.


If an afterburner falls before leaving the groundecision is made to continue, control failed engine as follows:


AXIMUM THRUST. If unable to light afterburner:


s necessary.

Abort mission.


Nozzle failure may be Indicated by nozzle position, excessive rpm fluctuations, or failure of the engine to control to scheduled speed. This may be accompanied bystall and exhaust gas over Engine shutdown may be necessary.

Nozzle Foiled Open Immediately After Takeoff

In the eventozzle failed open

Affected engine:

fterburner range.

RPM faaintain within limits.


In the event of extreme engine over-speed, if flight condition permits, retard throttle below Military or shut down.

as soon asIs Foiled Closed

In the eventozzle failed closed

Affected engine:

- Military or below,

Do not attempt to rallght theae the engine may flameout (after which it cannot ba restarted due to reduced rpm).

andaintain within limits.

Compressor stall Is likely, and EGT will probably rise.

J, Land as soon as practicable.



If either fire warning light illuminatesleaving the ground and the takeoff is refused:

Accomplish ABORT procedure, this section, as necessary.

-engine only.

Positively identify the affected engine before retarding the throttle.


Shutdown operating engine after stopping.

Scatnsert If time permite.

Abandon aircraft,

Engine Monpgement

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 willcause airspeed to continue increasing until engine rpm begins to decrease. The planned rotation speed may be exceededesult; 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 shutdownailed or flamed out engine.


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

The abort procedure assumes that ato abort will be made before rotation speed Is reached. Aborts from above ro-tatlon speed are not prohibited, but the risks associated with aborting fromigh initial speed at takeoff weight must be balanced against those ofakeoff when making the decision. Inafter rotation speed is reached, the most reasonable course of action is to con tinue rather than abort unless theis such that the aircraft can not fly.

Lower the nose and energize tho brakeswith nosewheel contact. When rotation is well advanced, the aircraft may accelerate beyond takeoff speed and lift off before rotation can be checked. In this case, hold the aircraft off sufficiently to regain control and then touch down without sideslip if possible. Fly the aircraft back to the runway, attempting to regain the center.

Changed IS3

Chuie Deployment


drift chut*econd* for deployment after drag chute actuation. It is permissible to actuate the deploy handle while decelerating in anticipation ofIAS; however, premature deployment can result in destruction of the chute. of the chute system so as toIAS simultaneously with loading of the chute is not recommended unless the risk is justifiedery marginal distancesituation.

Broke Switch

The normal ANTI-SKID ON brake switch setting provides nosewheel steering and braking power fromydraulic system and anti-skid protection. Ii 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 STEERRAKE causes theOUT warning light on tho annunciator panel to illuminate.




Do not unfasten the lap belt or shoulder harness until thehas cometop.

The landing gear should be left In the extended position.

1. DLE.

Retard both throttles to IDLE. Do notto shut down either engineunless failure to do so would vitally endanger the aircraft.

For dry ronuay: use moderate to heavy brake pressure

For wet runway; light to moderate brake pressure.


The limit airspeed for drag chute deploymentIAS.

s required.

Set the brake switch to ALTRAKE whenydraulic system is below normal pressure due to system or left engine failure.


Selection of ALT STEER g< BRAKE changes the source of brakefromoydraulic system and disables the anti-skid system.

6. Shut down one engine (if necessary).

Shutdown of one engine is considered necessary in the event of drag chute failure.

If drag chute falls 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 loft engine if it has failed.



SEC'l ION ill


Dry R

If the drag chule should fall to deploy and stopping distance is critical, proceed as follows:

Wet Or






Icy Runway










This wet or icy runway technique will probably blow the tires early In the landing roll; however,control can still be maintained and the blown tires will remain on the wheels. Additional pedalwill be required aa each tire

blows. Maximum wing aerodynamic braking Is more effective than wheel brakinget or icy runway until the nose is lowered but the nose up attitude must not be heldoint that the nosewheel will slam onto the runway. Use of maximum possible up elevon after the nose is lowered while keeping theon the runway providesdrag and additional down load on the main wheels.



If one or both FUEL PRESS LOW warning lights illuminate during takeoff, abort if airspeed and runway length remaining If airborne or If an abort is not


Tanks withress on.

Analyze difficulty and attempt tonormal sequencing.

Illumination of both fuel pressure low warning lights indicates loss of all boost pumps. This can only result from multiple failures. If this occurs during takeoff, tank pressurlzatior. will supply sufficient fuel to the enginepumps to maintain engine operation.

Fuel can not be dumped withboost pump failure. Use caution and observe operating limits ofeavy weightIs required.

After fuel prossure restored:

4. ress off.

If normal operation can not be restored:

6. Land as soon as possible.

With crossfeed on, more fuel may tend to feed from the forward tanks and cause an. shift. Before. should be checked carefully.



failureain gear tire during takeoff will overload the remaining tlrea on that side when takeoff weightb. This may be precipitate additional tire failures before normal takeoff speed can be reached or before the aircraft can bedepending on speed and the time of failure. As each main gear tire lossthe available brake energyby one-sixth, ability to stop from high speed le Largely dependent on effectiveness of the drag chute.

Failureosewheel tire is not expected toecond tire failure, but it may not be possible toose or main gear tire has failed. In either case, engine or structural damage may be sustained from tire fragments.

Depending on the airspeed attained and whether or not engine damage isakeoff may be preferable to aborting.

The following procedure Is recommendedain or nose gear tire failure la suspected during the takeoff run:



Set the anti-skid switch OFF prior to brake application. Brake with steady application of pressure to avoid aof the blown tire.

If takeoff is continued:

NOT RETRACT GEAR until checked.


Antl-skld off must be selected in order to stop the wheels rapidly after takeoff, as braking is disabled with anti-skid ON when gear down selected and there Is no weight on the gear.


The blown tlre(s) must be stopped in order to minimize the posaiblility of damage to the aircraft.

confirmation of tire andcondition.

The gear should not be retractedisual check has been made by another aircraft or by ground personnel. If loss of one or more tires is verified, the gear should be left extendedanding made as soon as practicable.

when practicable.


If the gear lever cannot be moved to the UP position after takeoff:

Gear overrideress and hold.

Landing gearP.

This overrides the solenoid which isactuated by the landing gear switch.


Improper use of this procedure may cause gear retraction while on the ground.

Once energized, the gear lever must be recycled to the DOWN position In order to bring the ground safety ewitch back into the circuit.



Escape from the aircraft in flight should be made with the ejection seat. The followingummary of ejection expectations:

sea level, wind blast exertsforces on the body op toappreciable forcesKIAS: and excessive forcesKIAS. The aircraft limitbelow these speed*.


flights with oxygen mask andare restricted to belowndEAS 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.

atIAS and aboveor landing run results Inchute deployment.

The zero altitude capability of this aircraft should not be usedasis for delaying ejection if ejection Is necessary. Aircraft accident statistics emphaticallyrogressive decrease In successfulas ejection altitude iseet; therefore, whenever possible, eject0 feet.


Before ejection, when time and conditions permit:

educe so that the pressure suit is not essential to survival.

educe to subsonic and as alow aa conditions permit.

Head aircraft toward unpopulated area,

Transmit location and Intentions to nearest radio facility.

MER position.

Lower helmet visor.

The free fall from high altitude down0 feet with drogue chutewill result In stabilized descent In the quickest manner.

During any low altitude ejection, the chance for success can be greatly increased by zooming the aircraft to exchange excess airspeed for altitude. Ejection should be accomplished while the aircraft isevel or climbing attitude, limbing or level attitude will resultore nearly vertical trajectory for the seat and crew members, thus providing more altitude and time for seat separation and parachute deployment.

Greenull if0 feet.


To accomplish an emergency escape using the ejection seat proceed as follows:


Sit erect with head against headossible, cross arms to pull ejection ring to assist in keeping arms close to body,


If seat falls to eject after normal delay,with the following:

section m


Use canopy jettison handle. If canopy still does not jettison, open canopy and allow It to blow off into the air stream.



Do not pullandle with the canopy still in place.

Keep elbows close tond feet firmly against seat while pulling theandle since the foot retractors and knee guards will not actuate.


After clear of aircraft if not automatically separated from seat;

Manual cable cutterull.


Kick loose from scat.

Parachute armingull.

If at high altitude after drogue chutefree fall:

Extend arms to control spinning. When drogue chute releases:

Feet together.


After drogue chute separation, backward tumbling tendency Feet together prevents pilot chute deployment between legs.

PARACHUTE LANDINGS After main chute opens: Overigh Alt!rude

0 feet,releasekit. Pull handleof the kit.

0 feet.roll -Jet roll bars up.

to land.

landing,release one side ofto prevent being dragged

Overow Altitude

kit Immediately aftershock.

ifjJ> Altitude

face plate and extendto hold open.

emergency oxygen

chest strap.


Failure to loosen chest strapinflating flotation gear may result in inability to breathe.

out life vest oral inflation tubeopen valve.

vent hose.

lifeime permitting,inflation valve.

0 fccbrolease survival kit.

riser release safety bars to tbe(RJ releases).

1. Place left forearm through the "V" formed by the left risers.

j. Place right hand on left riser release, feet together and knees slightly bent.

k. Push up on the left riser release on contact with water, releasing canopy.

1. Release other (right) side of the canopy.

m. Pull raft to you for additional support.

n. Disconnect the survival kit lanyard from the right side of the parachute harness.

o. Remove spurs before boarding raft.


Spurs must be removed to prevent puncturing raft.

p. Remove parachute harness before boarding raft.

q. Board raft.

Overow Altitude

inflate outerparachute opens.

survival kit.

Rocket Jet release roll barscanopy upon contact with water.

standard procedures in water.


if the seat fails to eject, the followingshould be used to leave the aircraft.





Disconnect oxygen supply hoses at the quick disconnect, and suit vent hose at the controller.

Trim full nose down, roll inverted.

Lean forward.

Release seat belt {and control stick, simulataneously) and drop out.

When clear of aircraft:

Pull parachute arming lanyard.

Prepare for landing.

Preparations for landing are the same as for ejection procedure.


warning in flight

IlluminationIRE warning lighta nacelle compartment temperature above0 F. Ancheck should be made for abnormal EGT and, ii possible, lor trailing smoke or any other indication o! fire. In caae of doubt, assumeire does exist and proceed as follows:


If light remains on:


If light still remains on;

- OFF.

If fire warning light extinguishes while shutting down the engine, do notestart.




If it is the left engine which Isand has been shut down, the cockpit air switch should be placed in the EMER position.


At pilot's discretion. If fire confirmed or confirmation not possible and light remains on:

Attempt to descendhigh altitude prior

If there is no fire:

as soon as possible.

smoke or fumes

The pilot cannot delect fumes whenressure suit. The helmet oxygen system is independent of the cockpit and suit air supply. Smoke can be eliminated promptly by dumping cabin pressure unless smoke is entering the cockpit from the airsystem.

off tho fuel If speed above approximately, may cause engine oil to overheat and result in engine failure. Shutting off the fuel may also cause additionaldue to loss of the associated aircraft cooling systems. Reduced Mach number decreases cooling requirements because of lowertemperatures.


Cockpit depressurizatlon will occur at an extremely rapid rate and the pilot will be dependent on his pressure suit for altitude protection.

if the smoke is Introduced by the cockpit air supply system, switch the cockpit system

to EMERG. The defog system should be off

at all times when not required.

If smoke Is entering the cockpit from the air conditioning system:

1. Cockpit airMER.

DefogFF If not required. If smoke or fumes cannot be controlled:

Initiate emergency descent.


The pilot must depend on visual detection of electrical fire whenressuresince he cannot smeU the characteristic odor.

the malfunction.

Turn off electrical systems in order to isolate the Ifdeactivate suspected systems by pulling circuit breakers. The battery and one generator may be turned off without adverse effect on essential systems; however, both generators should not be offunless absolutely necessary as this would shut down all fuel boost pumps.

Leave faUed system off. If required;

Cockpit pressure dumpUMP.

Land as soon as possible.


If extreme conditionsapid descent:




When initial CIT ie high, engine damage can be expected as the deceleration Mach rates specified iniU be exceeded.


Setting this configuration provides the least probability of asymmetric unstart, high drag, and the best means for avoiding inlet roughness during the descent.


Set the aft bypass CLOSED if engine speed is maintained at or near the Military rpm schedule. Engine staUs will occur belowf the forward and aft bypasa are open with rpm at or near MlUtary speed.

4, djustEAS.


Do notEASoad factor.

If necessary, reduce rate of descent to maintain positive fuel tank pressure.

Increase rpm if high suit inflow temperatures are experienced.

5. Forward transferWD TRANS.

For rapid descents during which aircraft control has become or may becomeilot emergency, aft c. g. location with boost pumpsinimum use of flight controls Is recommended. This may Include non-turning flight until lower speeds are attained. If aircraft conti

isriticalow oxygenpiral descent is very effective inore rapid loss of altitude.

In descending through the transonic region, the nose will be betweenndegrees below the horizon.


Turns causing appreciable load factor should be avoided while descending throughoot level athe pitch SAS gain switching willransient "bump" which may increase load factor to near limit value.

When subsonic:

6. Landing gear lever - DOWN. (Below gear limit speed.)


Gear extension at supersonic speed is forbidden.


If the landing gear is extendedEAS or, the landing gear doors will be damaged if sideslip exists.

Extending the landing gear at speeds aboveay cause heat damage to tires and resultazardous landing condition. With geararge nose-up pitch moment occurs in the speed range of. Full nose-down elevon will be insufficient tolight at high KEAS and/or. in this area.

level in0 pounds. When the EMER position is selected, dumping will continue until all fuel excludings expended. To increase the dump rate,select boost pumps for all tanksfuel (except


Accomplish normal fuel dumping as follows:

Fuel dumpUMP.

Fuellternately monitor TOTAL fuel anduel.

Fuel dumpFF0 pounds remain in tank 4.


If the fuel level inas prematurely reached0 pound level and dumping is required (excessive fuel in tanksroceed as follows:

Fuel dump switchEMER.





Normal fuel dumpingeans of reducing gross weight rapidly in the event of an emergency. All tanks containing fuel except forill empty in the normal fuel tank usage sequence. ill not be dumped, as its boost pumps are held off by actuation of the fuel dump switch and manual actuation of theoost pump selector turns dumping off. When the fuel dump switch is in the DUMP position, fuel dumping will continue only until the fuel

lternately1 and 4.

Whenuantity0 pounds:

Forward transferFF, When required amount of fuel remains:

Fuel dumpFF.


boost pumps Inrcuntil the EMER (lump Is turned OFF orumps are selected manually. Thedump must be turned to OFF or DUMP to assure automaticol fuel remaining int termination of fuel dumping.

At least one engine must be operatingorced landing Is to bo attempted. All forced landings should be made with the landing gear extended regardless of terrain. High airspeed or nose high angle of impactlandings with gear retracted causes the aircraft to "slap" the ground on impact, subjecting the pilot to possible spinal injury. It is recommendedear up landing not be attempted with this aircraft; EJECT Instead.


Forward fuel transfer and fuel dumping may be accomplished simultaneously as follows:

Fuel dumpUMP.

Forward transferWD TRANS.

Fuellternately monitornd 4.

Whenuel quantity0 pounds:

Forward transferFF.

Fuel dumpFF0 pounds remain in tank 4.


Ditching, landing with both enginesor other forced landing ahould not be attempted. Ejection is the beat course of action. All emergency survival equipment is carried by the pilot; consequently, there is nothing to be gained by riding thedown.


The following procedures are to bein the event of abnormal operation or failureropulsion systemnlet, engine, afterburner, nozzle, fuel control, or lubrication, fuel-hydraulic, or ignition system.


Inlet duct unstarts can only occur after supersonic speeds are reached and an inlot has beenhat Is, supersonic flow conditions established inside part of the Normally, the supersonic flow region extends from the cowl entranceosition near the inlet throat when inlet floware optimized. hock wave Is formed at the boundary between supersonic and subsonic flow conditions In the inlet. When an inlet unstarts, the internal shock wave is expellednormal" shock wave forms ahead of the cowl. Flow within the Inlet becomes subsonic and pressure In the Inlet decreases. When an Inlet alternately starts and unstarts rapidly, the change in inlet pressure which occurs results in severe airframe roughness.

Shock expulsion, or unstart, may be caused by Inlet airflow becomina ureatcr than

engine requirements and duct bypass spike position too far aft. or abrupt aircraft attitude changes. Improper spike or door positions can result from inlet control error, loss of hydraulic power, or electrical or mechanical failure. Unstarts are usually associated with climb or cruise operations abovehen at normal engine speeds; however, they may beduring reduced rpm descents at speeds above. Between, when near Military rpm, recovery procedures using the restart switch ON position may result in compressor stall.

Unstarts are generally recognizable byroughness, loud "banging" noises, aircraft yawing and rolling, and decrease of compressor inlet pressuresi. Fuel flow decreases quickly and themay blow out. EGT usually rises, with the rate of increase being faster when operating near limit Mach number andaltitudes. istinct increase in drag and loss of thrust occurs because ofair spillage around the inlet andairflow through the engine.

The aircraft yaws toward the unstarted inlet during an unstart. This yawoll in the same direction. Pitch rates are not developed by the inlet unstart. but pitch control problems can occur duringmaneuvering and will be accentuated by altigh Mach numbers and any pitch rate which existed prior to Inlet During the unstart, primarymust beplaced upon maintaining pitch control in order to prevent nose up pitch rates and angles of attack in excess of eight degrees. Thrust asymmetry should be reduced as soon as possible.

Aileron effectiveness is reduced at high altitudes and high angles of attack. Roll control may become critical if the unstart occurs on the inboard inletank. At altitudes0 feet, aileronmay be ineffective in controlling roll during an unstart unless the angle of attack is immediately reduced. Aileronincreases rapidly as the angle of

attack is reduced and only moderate aileron inputs will be required to control the roll. An excessive nose down attitude mayin an over speed in KEAS and Mach if the Inlets are restartedecovery maneuver. Therefore the restart switches should remain on until Bpeed and attitude are fully under control.

The roughness usually clears after thebypass doors open and the spikes are started forward manually or automatically; however, as much as five to eight seconds may be required for the spikes to reach the full forward position. Roughness mayuntil the spikes are fully forwardrestarts at design Mach number.

Inlet pressure should be checked during Moderate CIP increases will occur as the Inlet "clears" or restarts, and when the spike retracts to form the inlet throat farther aft. Return of the forward bypass doors to their normal operating schedule should resulturther CIP increase to normal operating values.

In automatic operation, unstarts which are caused by Improper spike scheduling limit aircraft speed to Mach numbers below that for the unstarted condition. Manualprocedure is necessary if theIs to be accelerated further. If an unstart results from marginal bypasshowever, it may be possible toat speed by adjusting the forward or aft inlet bypass doors to positions which maintain stable flow conditions. In general, if engine speed is maintained, less bypass area Is required as limit Mach number is approached.

hows the operating conditions where airframe roughness will occur due to unstable inlet airflow conditions. Theboundariesunction of Machengine speed, and spike and bypass door positions. The smallest roughness area occurs below the idle rpm range with the forward and aft bypass doors open and spike full forward. ore extensive area occurs with the bypass doors open but with





nd 'Al Fixo. Al 4O0 "IASwl bavridoiiai win intraoi*

ap;. nl.ly ISO

da> A Jp*>di< Moth And




** 0

2 o


f 2)

the spike moving in accordance with the automatic schedule. In both cases, theof inlet airflow instability occurst higher engine speeds, with thedoors closed. At windmllling rpm, heavy roughness will occur in the speed range aboveegardless of spike and door positions.

In the event of an unstart, accomplish only those of the following steps which areto clear the inlet and return to normal operation.


Shut down the engine if an EGT overtcmperature exists for more than five seconds, then restart the inlet and the engine as soon as possible.

In the event an inlet duct unstarts, proceed as follows:





K roughness docs nol clear aftereconds;

AFT BYPASSPEN. When roughness clears:

Aft bypassormal schedule.


FFinlet starts:

derlchment arming switchbelowEGT if

- As required.

If unstarts repeat or inlet roughness does not clear:

and inlet instrumentheck.

unstarts persist:

inlet restart andmanual inlet controls.

not attempt to clear the un-started inlet by placing only one restart switch on.





AutomoHc Spike Controlunction

Manual spike control is necessary if an automatic spike control manfunctions. In this event, the spike and forward and aft bypass must be operated manually as pre-scribed in the schedule table. Use of the AUTO forward bypass setting results in open forward doors when manual spike positions are selected.

Automatic Forward Bypass Control .Malfunction

With the automatic spike control operating normally, there are two options available (or control of tho forward bypass doors In the event their automatic control The manual bypass schedule table may be used, or, if the opposite aide inlet controls are operating normally, thebypass manual setting may be adjusted to provide CIP which is atel below the normal side indication. Note thatautomatic spike and manual bypassbypass position is controlled only by the pilot, and bypass position Is notby spike position. Therefore, it is necessary to anticipate changes in flight speed or attitude which affect matching of the CIP Indications.

than indicated by the TDI. After completion of the turn, the inlet controls may beto the manual schedule. The spike should be reset first, then the forward


The following schedule must be used when automatic scheduling is ineffective.

ag Mach numberach

atch Mach number.

ead Mach number by

ach (e.pike. Mach settingach on TDI).

With Manual Inlet Conttol

Maximum allowable speed ia.

Manual inlet scheduling must not be used0 feet.

To increase longitudinal stability, sufficient fuel should be transferred forward to obtain at least 0 pitch trim. This decreases the possibility of making inadvertent attitude changes which would affect CIP matching. Nose down pitch trim is an indication ofeg for this condition. However, the need for forward transferring should be weighed against the decrease in celling and range capability associated with increased pitch trim requirements at forward c. g.

Maximum bank angles ofegrees are permissible at speeds up to. mall heading changemaller bank angle willthe possibility of an unstart.

Whenegrees bank angle will bethe forward bypass should be adjusted to one position number lower than specified in the manual schedule; then the spike should be adjustedach number position less


The following schedule must be used with manual spike scheduling. It is optional when the spike and opposite Inlet arenormally.

Condition Mach Fwd Dyp. Aft Byp.

Pos. B

et at least


1 psithe


ration) pointer.




Deceler- ALL ation


Unstable inlet conditions which produce inlet and airframe roughness occur at supersonic speeds when an inlet alternately unstarts and restarts rapidly, usually duringat reduced rpm. Inlet unstaitare used first, except that the

throttle normally ia reset to providerpm instead of afterburning thrust after the unstart Is cleared. Subsequent settings may be made as desired. Refer to procedure for compressor stalls in descent.

Failurepike lo Schedulelet^ Sprite Um'i-J'bTe

A combination of unsymmetrlcal thrust and low compressor Inlet pressure on one side when accelerating betweenndndicatespike has failed to move aft on the proper schedule. This may be caused by failure of the spike forward lock to disengage0 feet altitude.

Spike instability is reflected by fluctuations of the respective hydraulic pressure gage. If spike oscillations are of large amplitude the gage fluctuations will be severalpsi and will be indicated on tbe spike position indicators. If an unstable spike or failure to schedule Is suspected, proceed as follows:

Spikeheck whileach number,

Spikeycle FWD then return to AUTO.

If condition continues:

bypass control -

anualhigher Mach number is reached:

and forward bypass controls

If condition recurs or continues:

per spike and bypass


Acceleration ond/or Qvetfrlm

These stalls usually result from EGT up-trim and are most prevalent during throttle application at subsonic speeds and lowinlet temperatures. They may alao occur with constant throttle settings at or below Military at any airspeeds. the throttle should result In thestall recovery. Downtrim andof the throttle should result in proper engine operation.


EGT trimOLDeconds.

s desired.

If stall persists repeat above procedure.

If stall cannot beand as soon as possible.

Compreisor Stalls In Deacenr

The airframe roughness characteristics felt during compressor stalls at supersonic speeds are very similar to those whichduring inlet unstarts. If roughness is encountered, an unstart condition Is more likely while abovehen spike scheduling le at fault. ompressor stall condition Is more likely at lowerspeeds when at or near Military rpm with excessive bypass door opening. The normal descent procedure tends to avoid conditions which may result in compressor stalls, but excessive rpm reduction or spike too far aft precipitates unstarts. (Seet is best to employ theprocedure first in the event of Inlet disturbances until it is apparent that the spike Is scheduling and that spike forward






lo, Sid OarIAS


i Induoa* Ih*llpaitfand U> Caiiftaviailoii


wait PaHnad B>ur.*.


Is ineffective In clearing the roughness, The throttle should then be retarded slowly to the idle stop if necessary, until roughness stops and the compressor stall is cleared. Maintain this configuration for theof the descent until subsonic airspeeds are reached.


educe rpm slowly until stall clears.

When subsonic:


engine flameout

Windmilllng operation at speeds betweenay result In heavy inlet roughness as illustrated by. If an immediate airstart cannot be obtained before the engine stabilizes at wmdmllllng speeds, adjust airspeed toEAS or ateia CD? before making further attempts. Engine flameout with afterburners on or off should be treated identically except for initial throttle positioning after the flame-out occurs. If flameout occurs withON, the throttles should beto minimum afterburner position to reduce thrust asymmetry. If afterburners are OFF at flameout the operating engine should be set to the thrust required by flight conditions. When an engine flameout isby crosschecking EGT, fuel flow, rpm and ENP, proceed as follows:



ACCOMPLISH AIRSTART PROCEDURE. If start Is notailed engine:




Compressor inlet pressure normalhecksia.

6. alf open. (Check TEB


If necessary, continue airstart attempts as long as TEB supply remains unless an obviousfailure has occurred.

After start:

9. Throttles and cockpits

If mechanical failure obvious or unable to start engine:


Failed engine Inlet air forward and aft bypass doorPEN above.

Cockpit airMER if left engine failed.

IS. Establish single engine cruise.

double engine flameout When altitude permits:


When altitude is critical, or engines will not start:




The reason for Initial enginemust be considered prior to Initiation of restart.

The recommended condition for airt any Mach numberEAS withal or greater. Airatarts should not be attempted at lower Inlet pressures, and airspeeds In excessEAS should not be required Insla. Conditions for air start are more favorable when stable Inlet condition exist; however, starts have been obtained while In roughness. Monitor rpm, EGT and fuel flow while making the start attempt. Alloweconds, afterthe throttle, for rpm and EGT rise as an indication of successful start. The recommended procedure for airstarts is as follows:





- HALFstart:

and cockpits

If start unsuccessful (aftereconds):


START attempt (check TEB counter).


. The engine shall not be Intentionally windmtlled at subsonic conditions when CIT is less thanIf It Is necessary to windmill the engine for more than five minutes the engine should not be restarted.

. If the engine must be restartedan In-flight emergency after wlndmllllng In excess of the above Limit maintain as high an airspeed as possible to raise the inlet air temperature prior to starting. The engine should then remain at idle until there Is an indication the oil has warmed up either byof the oil temperature warning light if it haB ILlumlnated orormal response of oil pressure to throttle movement to slightly above IDLE.

. If, following windmill operation in excess of the above limit, the engine must be restarted and operated at high thrust levels while the oillight is illuminated,of such operation shall be as brief as possible.


The wlndmllllng Glide Distance chart,, shows zero-wind distances with both engines wlndmiUlng. The glide speeds are In the same range as for airstart. slower speeds provide greater range, but reduced capability for successful air-starts. There le sufficient engine rpm for adequate hydraulic pressureeet.


Landing with both enginesshould not be attempted.



Engine shutdown should be accomplished In the event of complete engine failure such as seizure or explosion, or in the event of mechanical failure within the engine oraccessories In order to avoid or delay complete engine failure. Mechanical failure situations Include uncontrollable oilEGT, or rpm, and abnormal oil pressure, fuel flow, or vibration. Complete failure probably will not permit normaloperation but, if the engine continues to rotate, cooling fuel will circulate through the engine and aircraft cooling loops with the throttle OFF. An alrstart should not be attempted since doing so can result In fire or explosion. Normal wlndmllllng speeds can be expected after shutdown forfailure and fuel cooling will continue unless the fuel Is shut off. In some cases, alrstart may be attempted after mechanical failure when conditions are favorable for control of oil temperature or pressure or EGT.


Positively identify the failed engine before employing the engineprocedure.

If engine shutdown Is necessary;


For the affected engine, select the OPEN position of the inlet aft bypass switch in order to delay onset of roughness or inlet unstart when the engine is shut down.

N fin roughness).

Sot the Restart switch for the affected engine Inlet ON when roughness

This causes the forward bypass to open and the spike to move forward. will be encounterednd may persist to low supersonic speed. If an engine is shut down at subsonic speed, setting the Restart switch ON only opens the forward bypass as the spike Is already forward.

switch (affected engine)

Setting the generator switch OFFthe automatic cutout featureavoids the possibUty oftransients which might affect the navigation system.

fuel shutoffuel off

Fuel shutoff stops flow through one fuel cooling loop system. Depending oncircumstances, this step may not bo desirable or necessary.


Shutting off fuellndmllllng engine while at high Mach numbers may cause additional emergencies due to loss of cooling fuel for the engine and aircraft systems.

left engine shutdown:

2. FF.

6. Cockpit airMER.

As the throttle is retarded, pause

momentarily at the Military and Hydrauliceview SAS and

services available.

Refer to procedures for SAS, flight control system, and hydraulic system emergencies for operating procedures

c. g.

Refer to use of forward transfer andescribed under fuelemergency operatingto control c, g. during single engine operation.

as soon aa possible.


Tbe aircraft design le such that no flight system Is dependentpecific engine; therefore, the loss of an engine will not result In subsequent loss of all hydraulic or electrical systems. If an engine fails at low speed just after takeoff, tho large amount of asymmetric thrust may require bank toward the good engine and full rudder for directional control. Itefer toor minimum single engine control speeds. After regaining control, however, 7 udder trim with bank and sideslip toward the good engine provide minimum dragacceleration to climb-out speed. Charts showing single engine cllmbout capabilities are Included In the performance data Acceleration to climb speed and climb to landing pattern altitude can bewith Maximum thrust on theenginelimb capability exists for the operating condition. During single engine cruise, or after cllmbout, reduction to zero rudder trim and use of bank and sideslip to maintain course providesdrag. Up to 10 bank toward the good engine may be required.

Pitch trim changes can be expected while dumping fuel due to shifting center of gravity as the tanks empty. Directional trim la quite sensitive to changes In airspeed and power settings during landing pattern

At high speed, engine failure or engine flameout could cause yaw angle to become critical at high rates if an effective damper were not operating. Temporary thruston the good engine helps tothe asymmetric thrust condition. rudder action Is necessary. If large yaw angles develop. Inlet duct airflowmay cause the other engine to stall or flame out.

Roughness, if encountered, is more intense with Increasing KEAS and Mach number. The maximum structural loads Imposed are severe, but are well below design limits.

If alrstart attempts are unsuccessful, or If engine failure hasescent to Intermediate altitudes will be necessary. The spike should be forward and bypass doors open on the wlndmllllng engine toonset of roughness. Note the effect of Mach number and engine rpm on Inletas shown by the Inlet Unstartchart,. Descent range can be extended by decelerating with minimum afterburning or Military thrust on the good engine. Base the choice on the powerto be used for single engine cruise. When no alrstart la to beEAS until subsonic cruise altitude Is reached. ank of upith zero rudder trim should be used to achieve beat cruise performance.

Fuel management during protracted engine out operation should be directed toward maintaining optimum center of gravitymaking all of the fuel available to the operating engine and, when possible, continuing the fuel cooling of necessary systems. Improper c. g. conditions will be indicated by abnormal pitch trim


Single-Engine Ai< Refueling

Single engine air refueling may beualng cither normal or alternate refueling procedures. Approximately the same control trim and forces as for single engine cruise may be used with bank angles up to Afterburning on the operative engine will be necessary when0 feet at normal refueling speed, and amay be necessaryounds of fuel are on board. Refuel hook-up may be accomplished with the operative engine near Military power at low altitudes, although lack of excess thrust will make the hook-up more difficultontlnurd descent may be necessary as fuel la onloaded.


. Trimming EGT up toward limit values improves refuelingcapability.

. if the left engine or the leftsystem Is inoperative, right hydraulic pressure may be used by placing the brake switch in the ALT STEER i. BRAKE position.

. When using minimum afterburner at intermediate altitudes or with small quantities of fuel remaining, it may be necessary to hook-up while climbing in order to avoid overrunning the tanker.


hrust and Military thrust provides the best levels of single engine cruise performance. Military provides the beet range performance, but penalises the aircraft in altitude capability especially at heavy gross weights. good range performance with an ample altitude capability. Theingle engine cruise has poor rangeand should be only used in cases where the required cruise altitude le higher than theruise capability.

Since hot temperatures adversely effect aircraft ceiling, an altitude capability lower than shown on the charts must be expectedot Day.

Refer to Appendix Part IV for single engine cruise performance.


Afterburner flameout Can be expectedesult of engine stall or abnormal Inletor Insufficient airspeed at altitude. Afterburner flameout may be detectedoss of thrust and by comparison of nozzle position Indicators. Tho flamed-outnozzle will be noticeably more closed. Fuel will continue to flow from the spray bars until the throttle is retarded to MILITARY. Correct the Inlet, engine, or airspeed and altitude condition beforeafterburner relight. At high Mach numbers, the minimum airspeed necessary for afterburner operation is lower withscheduling than with spike forward.

In the event of afterburner flameout, attempt to relight as follows:

etard to Military.


Note TEB shot counter number and fuel flow Increase.


Check for more open nozzle position.

If relight not successful:

4. ncrease trim.

For CIT abovetrim towardEGT.

For CIT belowtrim towardtoEGT range.


Uptrlm to the-EGT range carefully due to possibility of engine surge.

If relight by catalytic Igniters not

5. Igniter purgen for two seconds.


The TEB supply will be depleted rapidly If the igniter purge switch is held on for more than two

It relight not successful: 6. ilitary.

afterburner nozzle failure

Nozzle malfunctions may be indicated by the nozzle position indicator, excessive rpm fluctuations, or failure of the engine to control to schedule speed. This may be accompanied by compressor stall andgas overtemperaturc. Precautionary engine shut down may be necessary.

A nozzle failed open condition will be more difficult to recognize at high altitude during afterburner operation near limit KEASopen nozzle position Is normal in these conditions. As altitude increases and KEAS decreases, the nozzle gradually closes topen as limit altitude is approached. ailed open nozzle willin abnormally high engine speeds under these conditions. An increase Inthrottle positioneduction in cruise altitude while maintaining cruise Mach number (increasing KEAS) may permit cruise to be continued until the scheduled descent point Is reached. Nozzle position and rpm of the normally operating engine can be useduide in selection of the lower cruise altitude range where an open nozzle position is normal. Be prepared to use less than Military throttle when the afterburner is shut down.

cutoff failure

If the afterburner does not cut off when the throttle Is retarded to Military, an attempt can be made to vary the thrust by retarding the throttle below Military. The engine should be shut down If thrust cannot be modulated satisfactorily. The fuel may have to be shut off if the flowmeter indicates that tho afterburner is dumping fuel.

At intermediate altitudes, the nozzle failed open condition may be recognized byof thrust and an Increase In rpm. At low altitude and Mach number it will be necessary to rapidly retard the throttleoint midway between IDLE andto keep rpm within limits. The same procedure will apply when altitude and Mach number are decreased and the nozzle failure is detected. If the thrust requirement le critical, such as forit may be practical to retain Maximum thrust, even with engine overspeed, until safe airspeed and altitude are attained.

A nozzle-closed (allure can, in moat cases, bo detected by referring to the nozzleIndicator on the Instrument panel and by analyzing engine symptoms. There are no obvious symptomsozzle failed closed without afterburning because the noaila Is already closed, or nearly so, at Military thrust. EGT and rpm maytogether. Either down trim the englno or retard the throttle slightly and check for rpm suppression or ENP change. Afunctioning noszlo will open slightly to maintain engine rpm. In the case of the nozzle falling closed, do not attempt to light the afterburner because the engine may flame out (after which It cannot be restarted due to reduced rpm). If the nozzle falls closed wilh afterburning, rpm suppression will occur, probably un starting tbe inlet Shockwave. Compressor stall andflame out are extremely likely and EGT will probably rise.

Nozzle Foiled During Crulie

eduction In thrust or rpm isor nozzle failed closed:

ILITARY or below, as

RPM andaintain In limits.


A low oil pressure generally indicates an oil system malfunction. If the malfunction causes oil starvation of the engine bearings, the result willrogressive bearing failure, loss of oil. and subsequent engine seizure. Bearing failure due to oilis generally characterized by rapidly increasing vibration. If this occurs Inwith gage indication of pressure loss, reduce Mach number and altitude and do the following:


Land as soon as practicable.

High Oil Pressure

High oil pressure does not necessarilya hazardous engine operationunless accompanied by high oilhowever, the high pressuremust be reported after flight and the landing should be accomplished as soon as practicable.

Land as soon aa practicable.


Abnormal high or low oil temperature laby Illumination of the OIL TEMP warning light. It le unlikely that low oil temperature will occur in flight with theoperating,igh oil temperature should be assumed. Abnormally high oil temperature could be caused by deficientflow or insufficient fuel/oil cooling. Abnormally low oil temp may be Indicated before start or after extended wlndmllllng operation at subsonic speeds. In the event ofIL TEMP warning lightIn flight, proceed as follows:

Oilheck for normal

Speed andeduce as required If at high Mach number.

Fuelaintain0 pph (if practicable).

If temperature can not be controlled:


Land aa soon as possible.


IL TEMP warning light Illuminates after extendedoperation, refer to AtRSTART procedure, this section.


uel control malfunction Is suspected, minimise throttle movements and monitor rpm and EGT closely.


Fuel hydraulic system failure may be causedailed pumproken line or connector. ailed pump Is indicated by Inoperative exhaust nozzle and start and bypass bleed valves. Line failure isby excessive fuel flow.

If engine fuel-hydraulic system failure Is suspected:

flow -fuel flow Is excessive:

- Military.


Overepeed may occur.


If exhaust nozzle position Indicator does notore closed position:

etween Idle and Military.


A fuel flow of00 pph above normal willroken line.

When below:


Emergency Fuel Shutoffuel off.

fuel system

fuel quantity low warning

U ihe fuel quantity low warning light comes on with appreciably more0 pounds of fuel indicated remaining inotal fuel from the individual tank quantities. Monitoruantity and land aa Boon as practicable. Quantity indications are affected by pitch attitude and longitudinal acceleration. Total quantity indication is also affected by the fuel distribution in the individual tanks.

If the fuel quantity low warning light does not come on with less0 pounds of fuel indicated In tankeat warning light and land as soon as possible.

fuel tank PRESSUR1ZATION failure

Fuel tank pressurlzation failure Isby the tank pressure gage andof the TANK PRESSURE LOWlight. It may be confirmed by liquid nitrogen quantity remaining gage Indications. Impending failure la Indicated byofTY LOW warning light.

No corrective action la possible after both liquid nitrogen Systems are depleted except to limit rates of descent to minimize the difference between fuel tank and ambient pressures. In descent, the fuel tank suction relief valve in the nose wheel well opens when slightly negative tank pressures occur. Rates of descent should be limited ao that tank pressure does not become less2 psi. .

pressure low

If one or both FUEL PRESS LOW warning lights illuminate:


Tanks containingress on.

Analyze difficulty and attempt tonormal sequencing. The difficulty may be due to low rpm while transfer-Ing or dumping.

When fuel pressure is restored:

- Press off.

Limit tank pressure2 psi. This limit is based on structural capabilities of the fuselage tanks.


so that minimum tanklimit la not exceeded.

Adjust power and airspeed as required.

If flight included cruise over:

at subsonic longinutes If possible.

pressure cannot be restored: 5. Land aa soon as possible.

If loiter not possible:

Descend fromos slowly as possible.

Continue descent so that minimum tank pressure limit is not exceeded.


Cooling will be accelerated and pressure may be relieved faster after reaching subsonic speeds If the nose gear is extended.


Loss of all boost pumps can only result from multiple failures. It would beby illumination of both fuel pressure low lights.


Incorrect automatic fuel sequencing Isprimarily by the fuel boost pump lights. ight may illuminate out of norm; sequence, or fall to illuminate onn this event, control the boost pumpsuntil correct automaticanding is made. It le possible that faulty fuel sequencing may manifestby secondary Indications, suchue low level light coming on prematurely, or an abnormal adjustment required in pitch trim due to c. g, change by faulty fuel Note that. require; increased power to maintain speed and If normal sequencing does notnd manual sequencing Is either Inconvenlen or Impossible, turn crossfeed on or transfe fuel to ensure that any available fuel will ge to the engines.


Fuel cannot be dumped withboost pump failure. Use caution If heavyweight landing is required.

Partial boost pump failure may not beby the fuel pressure low lights. fuel sequencing and center-of-gravity shift may be the first indication. Proceed as directed for Fuel Sequencing Incorrect. Crossfeed may be required; however, when crossfeed is on, more fuel will tend to feed from the forward tanks which have boost pumps operating. This could cause an aft c. g. shift which might be hazardous when operatingailed pitch SAS.

Do notanually selected fuel boost pump to continue running In an empty fuel tank. The boost pump will be damaged,


Crossfeed may be required tofuel to both engines during fuel sequence malfunctions.

Fuel System Monqgement With Engine Shutdown

During single engine operation with the left engine failed, the crossfeed valve should be opened afterre emptied by right engine consumption. If the right engine has failed, emptyy successive forward transfer operations. This accomplishes the dual purpose of maintaining c. g. and using all available fuel.


Fuel transfer capability Is lost when operating on battery power and the crossfeed valve position cannot be changed. An. shift can be expected as fuel Is consumed.

Fuel cooling is continued automatically when the inoperative engine is windmllling unless Its emergency fuel shutoff switch is actuated.

Crossfeed should never be used duringtransfer when fuel remains inr 6. If it were, most of the fuelwould come from the operating tank(s) of group.ecause of the aircraft nose up attitude and lower fuel pressure head these pumps would have to overcome.mall. shift would


The design and specification operatingof thengine necessitates operationuel having special During high Mach numberthe fuel serves not only as the source of energy but Is used In the engine hydraulic system and serves alsoeat sink for cooling the various aircraft and engine accessories heated by the highair temperatures. Thisuel having high thermal stability so that it will not break down and deposit coke and varnishes In the fuel system passages. igh luminometer number (brightness of flame) Is required to minimize transfer of heat to the burner parts. Other Items are also significant, such as the amount of sulphur impuritiea tolerated. An advanced fuel.E, was developed to meet the above requirements.

In addition to the fuelpecial lubricity additive is used withE to insure adequate lubrication of fuel and hydrauUc pumps.

Fuels such asnday be used only for emergency requirements such as air refueling when standard fuel is not available and air refueling must beor risk loss of the aircraft. Air refueling procedures with JP fuel are the same as for the approved fuel.

When these JP fuels are used, operation should be restrictedaximum speed of.


Failure of one ac generator will be indicated by illumination of the warning light. One generator in normal operation Is sufficient to support the entire electrical load. In the event of generator failure, proceed as

ESET then

If the generator fault has been corrected, the generator will be reconnected to the system and the warning light will go out.

If the light remains on:

Failed generatorBD?.

Land as soon aa practicable.

If flight Is continued with an Inoperative engine or generator:


Depending on flight conditions and priorossible failure of the otherorward transfer may be advisable.

AffectedRIP. lf^ ^equipment Is operating:



section hi

. Do not manually select additional fuel tank pump*.

. ME radio transmissions are limited to one minute of any ten minute period,


If both ac generators fall the ac buses and the dc monitored bus will be dead. All equipment normally powered from the ac generator buses will be Inoperativethe fuel tank boost pumps which willdumping and forward transfer. Manual pitch yaw and roll trim and EGT trim will also be inoperative. Instrument and panel lights and landing and taxi lights will bebut cockpit lighting will befrom the cockpit spot anday equipment will be inoperative. The INS requires considerable electronic power and should be turned off unless essential for navigation. Tbe autopilot and attitude selector switch should be placed in the FRS position. All engine and airplaneand annunciator panel indicators will be available except the standby attitudeand compressor inlet pressure

All other ac and dc equipment will be powered by the two emergency batteries but for maximum operating time oftwenty five minutes thendnverters must be turned off. This will deactivate the SAS B, monitor channels and INS. Thehannel and most engine and aircraft instruments will be powered by thenverter with thenverter available for backup.

For maximum effective operating time of at leastinutes the following equipment should be turned off by the individualswitch.

UHF, HF, TACAN. and ADF Radio No.ndnverters


Q-bay Equipment

Autopilot andhannels

Anti-collision Light*

and Beacon


Afterinutes various relays may begin to "drop out" and athenverter will deactivate.AS channel and most electrical engine and aircraft instruments will become Inoperative. The ARCnverter willto operateeduced rate anddeplete the battery.

When subsonic the Inlet control and bypass circuit breakers should also be pulled.

Batteryheck BAT.


If only one generatorand as soon as practicable.

If neither generatoronserve batteries.


The minimum windmill speed at which the ac generators will supply powerpm. If the left engine0 rpm the HF radio will be Inoperative. The left generator may be tripped to restore operation if the right engine is0 rpm.


One transformer rectifier will supply the normal electrical demands. Variableac power systems will continue to operate normally. ouble failure of Ihe transformer rectifiers removes power from the dc monitored bus but the INS will be op-


erated from the INS battery and No. 3 The batteries will operate the dc essential bus and should be managed ae for double generator failure.


The Inverter powered systems operate from separate inverters. Refer to Electrical Power Distribution diagram. Section I.nverter is the most important withndollowing in order of lesser importance. Thenverter is Installed to serveackup in the event of failure of any one inverter. It Is placed Inby turning the failed inverter switch to the EMER position. econd Inverter failure should occur thenverter will power the lowest numbered inverter bus whose switch Is in the EMER position.

If an Inverter failure la indicated byof an INVERTER OUT warning light, proceed as follows:


Check that INVERTER OUT light

SAS recycleress.


With both engines out, the hydraulic pumps provide sufficient flow for satisfactory flight control system operation at windmill speeds0 rpm. Reduced control system capability Is available downindmllling speed of0 rpm. With one engine windmllling, alland most utility services arc supplied by the operating engine hydraulic systems. The windmllling engine utility systemand flow may be sufficient to supply service until the engine Is almost stopped.


The loss ofydraulic system will be indicated by the warning light on the pilot's center console, the master caution light, and theydraulicgage. Reduce epeed to lessEAS ifystem falls and turn Reserve Hydraulic Oil System switch on* operative system This willinimum of atours of flight control time remaining at high speed cruise schedules.

Disengagement of the failed hydraulicSAS channels Is necessary to maintain full yaw and roll damping capability. ydraulic system failure Is not sensed by the SAS equipment. It la necessary to double the SAS algnal gain of the operating channel to give the equivalent control response in yaw and roll. Airspeed reductioningle hydraulic systemrecautionary procedure which allows for the reduction in available hinge moment capability. of the failed system SAS pitch channel Is not mandatory, but it may be more desirable to disengage all threethan only the yaw and roll switches. Monitor all system operations closely and attempt to determineomplete failure la Imminent. Be prepared for ejection prior to complete failure.


The lossydraulic system will be indicated by theydraulic pressure gage. If the pressure ofystem fallsel.for gear retraction la automatic. The manual release must be used to lower the gear. Items which are affected by tlie l, hydraulic system are normal brakes,steering, aerial refueling system and the left inlet control actuators. Items which are affected byydraulicare right inlet control actuators,steer and brakes and air refueling

- 42



With both engines out, the ac generators furnish rated electrical power al windmill speeds0 rpm, The emergency batteries provide SAS operation at lower windmill speeds. There Is sufficient hy-drauUc flow to operate the control surface! at satisfactory rates0 rpm and operation at reduced rates Is availableindmllling speed of0 rpm.


During single engineindmllling engine may not develop sufficient system hydraulicto maintain operation of Its aaaoclated 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 channel. The operative engine SAS roll channel must be cycled OFF then ON to maintain damping In the Roll axis.



If eitherydriiullc eyetemhe control forcos will not change. Eitherwill operate the control surfaces, butlower rate and with some reduction in control responsiveness at high KEAS and Mach numbers.

If control difficulties are encountered:

ydraulic system Ifydraulichas failed proceed ae directedydraulic system failure this section.

Disengage autopilot and check control.

Check SAS warning lights. If SAS failure has occurred, proceed asunder SAS Emergency Operation this section.


roll channelN.


When one roll SAS channel laor turned off the simplified logic circuit will disengage the other roll channel. The desired rollswitch must be turned OFF and thon re-engaged to regain single channel roll SAS operation.

hydraulic oil switch -system



If bothydraulichave failed all night controls will be Inoperative.

Reduce speed to lessEAS.


Do notEAS with eitherydraulic system If cither system fails above this speed, reduce speed as soon as possible. Flight control responsiveness will be reduced during single hydraulic system operation at high KEAS and Machnd flight maneuvers under these conditions should be heldinimum.

SAS yaw and pitch switches -




SAS emergency operating procedures and the applicable flight limitations should be used whenever there hashannel disengagementeduction in SAS Disengagement may result from failures of any of the following systems or components: SAS gyro or electronicsflight control servos, or electrical power supply. Disengagement or loss of effectiveness may occuresult of complete or partial lossystem hydraulic power. Disengagement of any channel la 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.




alfunction la Indicated in any SAS axis. Initiate the following preliminary actione:

ydraulic system pressuresnormal. If hydraulicis indicated, follow AB Hydraulic Systemthis section.

OUT Warning Lights

If illuminated, use Inverter Failure procedure, this section.

to appropriate Pitchaxle or Roll Axisthis section.

A single failure or sequence of failures in the pitch and yaw axes which leaveshannel operating in each of these axes does not change the aircraft flight characteristics. However, somecross-coupling in the pitch and yaw axes may rosult from failure of one roll channel. Characteristics which changeesult of failures affecting bothhannel servos in an axle are described as second condition failures with theprocedures. Also refer to the SAS Warning Lights charts.. 'which illustrate the probable causes of failureremaining capabilities,and limits which apply after channel disengagement.

Pitch end Yow Axis Failures

A "first" condition failure exists afterto extinguish one or more recycle lights are ineffective and eitherhannel Is operating (light Off) In each of the pitch and yaw axes. first" condition failure existsingle A, B,hannel light illuminated or in some cases after simultaneous or progressiveof two or more of these lights, as illustrated by the SAS Warning Lights Chart,,nd 2.


Consider that no failure exits when all pitch and yaw recycle lights have been extinguished, regardless of previousof illumination. If normal operation of the recycle lights Is verified by depressing the SAS Lights Test button.

Flight may be continued without restrictionirst condition failure exists except that maximum airspeed is limitedEAS in the case of combined channelduew hydouHc system prsflwuy^c*

A "second" condition failure is defined as existing wheneverecycle lights in one axis remain illuminated after attempti to extinguish them are Ineffective. second" condition failure exists, flight speed is restricted toEAS. Transfer fuel as required to obtain either 2 nose up trim0 pounds in tank 1.


Each instance of recycleew situation and. If the light(s) can not be extinguished, the condition must be determined asflret" or "second" condition of failure in accordance with the definitions provided above.

Logic override procedures are usablesecond" condition failure when theof light illumination indicateshannel with operative servos Is available. Refer to After Second Failures, SASLights Chart. When use of logicis effective, flight characteristics are the same as with SAS fully operational. Howeverrecaution against subsequent hardover failures signals, the autopilot must not be engaged in that channel and second condition failure limits apply.


If logic override la recommended, uae it only in the channelsand only after decelerating to second condition failure limit speeds In order to preventstructural loads which could resultardover failure at higher speeds.

Neither logic override nor BUPD operation should be attempted with either channel known toailed servo.

BUPD plus logic override procedures are availablesecond" condition failure in the pitch axis. The BUPD Is optimised for operation at air refueling speeds, and it should not be operatedEAS5 Mach. It may or may not improve flight characteristics at other flight

Logic override or BUPD plus logicmay not be usable or effectiveecond condition failure In the pitch axle. If neither can be employed, someovercontrol probably will occur when at high Mach numbers. Observance offailure limits is required, and descent to subsonic operating speeds iswhen practicable. Air refueling and landing may present some difficulties In maintaining precise attitude control. With pitch SAS off and neutral c. g. there Is no tendency for the aircraft to returnrimmed attitudeisplacementat landing pattern speeds. However, divergent speed and attitude tendenciesslowly enough to be completely Minimum airspeed limita with pitch SAS inoperative) should be observed.

If logic override procedures are notor possibleecond 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. Automatic scheduling of the inlet components may induce neutrally damped directional oscillations while above. Directional and roll control could become difficult In the event of an unstart or flameout while aboveesult of large bank angles generated by yawing motion. Pilot rudder Inputs usually tend to aggravate this condition. These conditions could also result Inrudder surface loads at airspeedsEAS. Use of both restart switches Is recommended whilein order to avoid asymmetric nacelle drag conditions or unstarts.

recycle llght(s) -release.

If light stays on or reilluminatea:


No further action is requirednd condition failure exists.

If another failure should occur In the same axis:

Illuminated recycleepress and release.

If lights do notomply with limits.

For second condition failure aboveEAS:

Nduring climb.


If climbing, bleed speedEAS.





ot eandjtory

um ilkn of ypt loqM OvwrUfr

*havwu wrrtlt

"otp Chair* iwtlch OH. Io uie 6upo1

AhKiruli Off. H

*wrklt 8nb one channel on.

n II no linpronmtrt.


IM BO WASIt memoryowua trimax fc. Mh owrrlorHo nAioibl (MlUPD t tnd













o o


o o


Itjhb on bul opwxfcn poor









A and B

ChKl Circuit


3SAS pfen-yw run

b Eis DC bus-SAS

Check circuit" SAS yen A

b (ss oc bus-


Check circuit

SASss DC bus-


Check circuita OC

ofl II pressure Is kw

With rwmilycle roll channelress recycle ftjhts off

reis recycleoff

roll channel milchW Ii on

not uct logic override

M Channels will be Inoperative

Any COmttnatMn, end/or roll lightsoccur

Lights my Illuminate stsulbrveousfy or progrtMhtly



KEAS aolmum

Figuref 2)

epood is stabilized belowEAS:


Aft bypassormal schedule.

s required.

transferransferto maintain at least 2 noseup0 lb.

If SAS lightsood servo is

ogicngage asby servo availability.

Pitch or yaw logic overrideosition depending on failure analysis.


Refer to SAS Warning Light Chart.

b. hannel switch-Beep ON.

Recycle light should extinguish.

C If control does not improve -ChannelFF.

d. Logic overrideFF.

For pitch axis second condition /allure; when speed isEAS5 Mach;

12. ngage as required.

nd B- OFF.


logic overrideposition as Indicated by

Aeeplight should extinguish.

c. If control does not improve -Channel switch OFF.



on failure analysismay be repeatedSAS channel if indicated.

Roll Axlt Follures

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 systems or electrical power supply which would

cause disengagement, checkransient disengagement as follows:

1- hannolFF, then ON.

A transient or Intermittent fault existed if the light then remains off. If the light does not extinguish, or rellluminates whilea first condition failure exists In the roll mode.

irst failure:

Unless the failure can be associatedpecific hydraulic or electrical powerregain the use of one channel by the following arbitrary step sequence:

3, hannelN.


Be prepared to move the switch to OFF Immediatelyardover signal results, indicating that the failed channel was Inadvertently selected.

Operation with only one roll channel engaged results Inof logic circuitry. There is no automatic protection against inadvertent selectionailed channel, or against subsequent failureroperly operating channel which has been engaged.

ard-over signal Is obtained onor during subsequent operation, or If no improvement is noted in flight

A channelFF.

B channelN.


Be propared to disengage theImmediatelyardoverresults.

ard-over signal is obtained onor during subsequent operation, or if no improvement Is noted Inual or second condition failure exists.

econd condition failure:

channel switchesBoth OFF.


Some undesirable cross-coupling may occur during single roll SAS channel operation. This appears aa 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 rollIs provided by pitch and yaw axisoperation.

Scheduled activity may be continued for the remainder of the flightingle roll SAS channel operating. The roll autopilot may be engaged and the automatic navigation feature of the INS used as desired.


. Operation with both roll channels disengaged Is permitted if croes-coupllng about the pitch and/or yaw axes prevents precise aircraft control with one roll channel

. In the event of single engine failure at low speed, or during single engine landing, failure of one roll SAS channel and simultaneousdisengagement of the other roll channel may occur due to loss of hydraulic power from the wind-milling engine.

To avoid changes In control characteristic!ritical time during single engine land' Inge, either make the approach with both roll SAS channels disengaged or with the roll channel which Is powered by theengine disengaged.

A second roll SAS channel failure while at high speed will probably be indicated bypitch transients and small roll transients without illumination of cither 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,both roll 5AS channels and check for control Improvement. If nola noted, the single roll channel may be reengaged If doelred.

Failure or Intentional disengagement of both roll SAS channels is expected topilot fatigue, reduce missionand will disable the roll autopilot; however, no hazard to safety should result and there are no flight restrictions onoperation.


Pitch, yaw or roll trim malfunctions may be of the inoperative type or the runaway type. Runaway trim failures In pitch may occur at slowsec change in olevon deflection) If due to automatic trim motor operation or at fast speed if due to manual trim motor5 /sec change In elevon deflection). ow speed runaway type of malfunction will be apparent by the need for constant manual pitch trimming. The runaway yaw trim rate is per second trim change. The roll trim rate Issec. Runaway yaw trim will be accompanied by rudder pedal deflections as the surfaces move. Runaway pitch or roll trim will not beby stick movement due tomovement.

the event trim runaway failure Isproceed as follows:


II circumstances permit:

speed toEASMach number.

With runaway noae up pitch trim:

luel forward to reduceforce requirement.


Do not transfer fuel If nose down pitch trim has occurred.

When time and conditions permit:

N. Check for control

Affected trim circuitull.


hase circuit breakers must be pulled on the suspected circuit.

Trim Malfunctions:

runaway slow speed pitch trimauto pitch trim

runaway high speed pitch trimmanual pitch trim

runaway roll or yaw trim -or yaw circuit breakers,


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 piloteed for maintaining stickto maintain attitude. The pitch autopilot can also be used when the auto trim motor Is Inoperative, but automatic pitch trim synchronication will not be


Disengagement of the pitchwhen not In trim mayonsiderable transient.

If the trim malfunctionunaway In the roll axis, right or loft stick deflection will be required for the rest of th* flight but stick force will not be more than normally required for the same amount of deflection. If the malfunctionunaway in the yaw axis, rudder pedal force will beto maintain neutral rudder pedal position.


If malfunction or failure of the air data computer (ADC) Is suspected, proceed as follows:

check TDI Instrumentoperated airspeed

If cross check shows TDI to be Inaccurate:

Revert to use of pltot-statlc operated Instruments for aircraft control.



Under some conditions both of the pltot-statlc operated systems may becomeor inoperativeommon malfunction. Failure of the pitot heater may slmultaneoualy affect both normal sys-

section m

in Icing conditions. The pitotbe pluggedoreign body ofboth systems fall,follows:

Attempt to restore operation byalternate static source.

Maintain aircraft control by use ofand power indicating Instruments.

Request escort aircraft for letdown and landing.

air conditioning and PRESSUR1ZATION

If cockpit altitude does not decrease:

Nose radio equipment UHF. HF andFF.

Nose hatch sealFF.


If cockpit repressurizes the pres-surlaatlon loss Is due to failure of the nose hatch seal and periodically the nose air and desired radio may be turned on for possible short time usage.

engine of. cockpit system inoperative

If the left engine Is shut down:

1. Cockpit systemMER.

cockpit depressurization

Cockpit dapressurlsationeet will be indicated bysuit Inflation. If suit Inflates, proceed as follows:


If Increasing or at actual aircraft altitude:


During thia time the pilot will be depending on the pressure suit only for altitude protection.

If cockpit still does not repressurlze:

air handle and nose hatch seal -


Suit ventilation boostMERG.

Reduce altitude.if possible.

RadioN only asafter altitude la reduced.

cockpit and ventilated suit abnormal temperature


When the nose air valve Is OFF It will shut off pressurlzation and cooling air to tha nose compartment and possibly result In loss of UHF, HF and IFF equipment. TACAN and normal ADF equipment located Inay will still be available.


Cockpit temperatureheck. If temperature Indication is abnormal:

auto temperature rheostatas desired.


The temperature control bypass valves are motor operated and travel from full hot to lull cold or vice versa inoeconds.

If auto temperature control is not effective and cockpit temperature remains too high or


temperature control switchin COLD or WARM as desired.


In this position the temperature control bypass valvesoeconds to travel from full hot to cold or the reverse.

If no correction In temperature occurs Ineconds:


If cockpit temperature is etlll abnormal:

or emer cockpit auto- Rotate toward COLDas required.

If suit temperature cannot be controlled by the preceding steps:


altitude and speed as required.

q-8AY abnormal temperature

IfAY temperature indication isproceed as follows:

auto tempotateCOLD or WARM as necessary,


The above step should beIn Increments as there willag in the temperature

If auto temp control is not effectiveay temperature remains abnormal:

and cockpit EMERGENCY- HOLD in COLD oras necessary.


The temperature control valve will takeoeconds to travel from full hot to full cold or the

ay temperature control system should fall In the cold position and heavy cockpit fog occurs:


Normal cockpit airotateWARM as necessary to dissipate cockpit fog.



Rite ofsi ronoe

System Indicating normal pressureFF.

Visor openingeconds. This allows Increasedflow between visor and seal.

System indicating normal pressureN.

above steps ifno correction noted:

is safe to continue flight Ifnotsi.

If pressure risessi:


Land when practicable.

Drop of Pressure Belowsi

Accomplishbove. If no correction noted:

Land when practicable.

No Pressure on System


Accomplishbove. If no correction:


Land when practicable.


No suit pressure when TEST IND button pressed

Descend below suit inflation altitude.

Land when practicable.

Reduced oxygen flow inBoth systems)


Descend to safe altitude.

Land when practicable.

Constont oxygen flow in helmet

BothFF then ON.




Repeat on other system If necessary. If no correction:



Descend to safe altitude.

Land when practicable.

Green apple loose from snap (possible active lystem)

pressure gage on greenindicator shows full:

green apple and continuepilot's discretion.

Poor or no communlcotlora

I. Check communications lead fordisconnect.




A single engine landing it basicallyormal landing, except that tho pattern may be entered at any point and la expanded to avoid ateep turns. Airspeed Is maintained above the normal value on final approach. 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 may be anticipated and rudder trim set to neutral on the trim Indicator after final approach Is established. heading la maintained by rudderuntil thrust Is smoothly reduced during the flare. The landing gear may be lowered after lining up on final approach with the left hydraulic system operating; however, at leasteconds must be allowed for emergency gear extension If the left system Is inoperative,


Hydrauliceview services available.

If left engine is inoperative, brakeLT STEERBRAKE.

Inoperative engine SAS pitch and yawFF.

SAS rolloth OFF.

Operative engine SAS rollN.

Landing gearOWN.

Establish steeper than normal final approach.

9. IAS minimum untilla assured.


If It la necessary to land with more0 pounds of fuelincrease minimum approachnot for each0 pounds.

10. Ruddereutral.

U. When landingetard throttle.

12. Make normal landing.


Directional trim changes will be moreduring an actual single enginewith one engine windmllling,

Retard one throttle to IDLE.

Follow Single Engine Landing procedure.


Make decision to go around as soon asand definitely prior to flare.

s required.

Continue approach until go-around la assured.

Landing gearP, aa

Delay gear retraction until there Is no possibility of contacting tho runway.

s necessary.



An unsafe Indication could be caused byydraulic ayatem pressure or malfunction within the landing gear extension orsystem. Upon detecting an unsafe gear Indication, proceed as follows;

Land goar control and Indicator circuitheck IN.

L hydraulicheck.

Recycle landing gear lever to down position, repeat as desired and pull emergency gear release handle If

If landing gear still indicates unsafe:

Landing gearequest visual confirmation.

If all landing gear appear fully extended,ormal landing on side ofaway from suspected unsafe gear. Observe the following precautions:

anually lock.

weight off unsafe gear aspossible then allow gear torunway as smoothly as If nose gear is heldnose at approximately

aircraft to roll to aahead, haveprior to furtherengine shutdown.

If any gear remains fully retracted, use Emergency Extension procedure.

If all gear are not fully extended, ref to Partial Gear Landing procedure, this section.


. Increasing airspeed may assist Inartially extended nose gear.

, Yawing aircraft may assist inartially extended main landing gear.


The emergency landing gear extension sya tern unlocks the landing gear uplocks and allows the landing gear to free fall to the down and locked position. ydraulic system pressure Is available, the landing gear handle must be placed in the DOWN position or the landing gear control circuit breaker must be pulled to permitextension. time requlrod for emergency gear extension lao 90 The emergency landing gear handle must be pullednches for full actuation. If It Is not fully actuated, one or more gear may fall to extend.

Ifydraulic system has decreasedel or normal gearis unsuccessful, proceed as follow:

Landing gearOWN.

Emergency landing gear release hand! PULL.

Verify gear down and locked.

If Li hydraulic system pressure low:



Alternate nosewheel steering Is availableystemdecreases0 psl.

If landing gear remains retracted or landing gear handle sticks In the UP position:

Landing gear control circuit breaker -PULL.

Repeatf this procedure.


When the landing gear controlbreaker Is pulled nosewheel steering will be inoperative.


The landing gear must not beif the manual release handle Is being held in the free fall (full out) position as damage to the system can result. The GEAR RELEASE handle should beto return to the stowed position before attempting tothe gear with the landing gear lever.


A landing with the nose gear retracted or with all gear up should not be attempted. Under idealanding with the nose gear extended and both main wheels retracted may be possible. If this configuration can be accomplished, base a

decision to land or eject on whether other factors are favorable or not. Windand direction are Important In selection of the landing heading.

ecision Is made to land, conventional final approach and landing speeds andare recommended. This will result In the tall touching while the nose Is at less than normal height. An attempt to hold the aircraft off byigher pitch angle Is not recommended because of the greater possibility of high impact loads as the nose gear slaps down. An emptyondition Is desired.

nose gear onlynecessary as follows:

gear CONT circuitin.


gear CONT circuit

landing gear releaseto release nose gearlock releases nosenose gear downN.

Do not transfer fuel forward.

Fuel dumpUMP, Ifto reduce weight.

Igniter purgeump during approach.


Inertia reel lockOCK.

Canopy jettisonull, if desired.


If the canopy lo not jettisoned prior to landing. It should not be unlocked until the aircraft has stopped.

Make normal approach and landing.

Drag chuteull.

Use rudders for directional control.

ThrottlesOFF. when directional control is no longer possible.

Abandon aircraft as quickly aa possible.

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 isthat only one or two of the three tires has failed. If only one tire has failed, little danger exists when landing at low weighttwo tires have sufficient strength to support the aircraft.

Touch down on good tires.

Drag chuteull, as soon as possible.



Hold weight off bad side as long asusing full aileron.


Maintain IDLE rpm until fire-fighting equipment arrives. Engine shutdown allows fuel to vent in the vicinity of the wheel brake area, thusire hazard.


If it is necessary to landlattire or tires,. if possible and proceed as follows afterormal touchdown.

Drag chuteull.

Noseold off.

Hold the nosewheel off as long aa practicableIAS) and then lower gently to runway.

nosewheel steering andto maintain diroctlonal control.

After stop, before shutdown:



Use normal procedure, observing operating limits of Section V.


The emergency abbreviated checklist la furnished separately.



uxiliary equipment


Communication* and Associated Electronic Equipment


Lighting Equipment Flight Recorder Dictet Tape Recorder


Autopilot System Navigation Equipment lnertlal Navigation System Periscope Destruct System



The communication, electronic navigation and instrument approach equipment Includes the following:

HF Communication, direction finding and ranging

DF ReceiverACAN Equipment ILS Equipment

) Radio EquipmentEquipment

IFF Equipment

nterphone Equipment


quipment Is capable ofand receiving on any0 channels in the frequency range05 MC. The equipment can be operated In either of two basic modes; an internalband) mode in which Its operation is compatible with any conventional UHF radio communication set, and an external {wide band) mode In which It has high resistance to jamming and low detectabillty. In this mode it incorporates message privacy and range measurement functions. When used In conjunction with the UHF DF system andet it provides direction finding capability in either mode.

In the internal mode power outputatts minimum while In the enternal mode the power output is approximatelyatts. The power output in either mode may be reduced Intepsb incrementsractionatt.













Moit of Umquipment le mounted in the pressurized and cooled nose compartment and Includes the blower cooled translator group, the receiver-transmitter groupeparate inverter. irection finding equipment and the flush antenna are also mounted in the nose

Theontrol panels are mounted on the pilot's left console. ange indicator is mounted on the instrument panel. Thefinding equipment is also connected to theeedle of the BDHI when the equipment is operating. Theantenna is mounted in the lower right chine. Power for the equipment Isby the essential dc bus. The leftbus supplies blower and heater power.


The translator control panel labeled UHF COMM is located on the left console. Itprovisions for control ofand power output, mode of operation and receiver volume. It provides the pilot withreset frequency channels,for manually selecting anynd controls for operation of the separate fixed tuned guard channel receiver.

Function Selector Switch

This four position rotary switch Is labeled OFF, MAIN, BOTH and ADF. In the OFF position the equipment le not energized. In the MAIN position the translator group equipment is energized with only theand main receiver operative. In the BOTH position the equipment is energized with the transmitter and both main and guard receivers are operative. In the ADF position the0 equipment Isand the main receiver and the transmitter are operative. Theeedle of the BDHI is also disconnected from theDF receiver or TACAN receiver and receives directional signals from thequipment.

Manual-Preiet-Guord Selector Lever and Indicator

This selector lever controls the manner of frequency selection. In the MANUAL (left) position the manual frequency selector switches are activated and the frequency selected is visible in the cutouts above each switch. In the PBE5ET (center) position the preset channel selector switch isand the channel selected Is displayed In the window below the Preset Indicator. In the GUARD (right) position the guard channel frequency la set on the mainand transmitter, and GUARD isIn the window below the selector lever knob.

Preset Channel Selector Switch

This switch, located in the center of the panel, selects one of twenty presetwhen tlie manual preset guardknob Is positioned to PRESET andthe frequency channel selected in the window beneath the Preset Indicator. The channel numbers are blanked out when MANUAL or GUARD Is selected.

Frequency Selector Switches ond lodicoton

Five rotary switches across the top of the panel permit manual selection of any one0 frequencies in0.c range. These switches are activated when the manual-preset-guard selector knob la positioned to MANUAL. Each switch Is used to select the digit displayed In theabove.

'NT-EXT Mode Switch

This two position switch la labeled INT-EXT. In tha INT position tlie translator transmits and receives narrow band AM signals independent of theequipment. This position Is used for conventional UHF transmitting and In tlie EXT position the signal translator nnd receiver-transmitter are used together to receive and transmit the wind band pseudo-noise encoded signals. The EXT position is also used for direction finding and/or ranging functions. The power selector switch may be used tothe transmission power In either position.

Power Selector Switch ond Indlcolor

The rotary switch Is the larger of the two concentric knobs and controls the output of the transmissions from tbe translator and receiver-transmitter combination. It has ten positionsoo permitpower output from the maximum of)ow ofb increments In the INT or narrow band mode. In the EXT or wide band mode power is increasedaximum ofatts but also may be reduced inb units toicrowatt. The digit In the cutout above the knob Indicates the power output selected. osition,for an additional amplifier, labut power output Is the same aa theosition.

Volume Control

This is the smaller center knob concentric with the power selector switch and adjusts the audio level of the receivers. rotation will Increase volume.

Tone Bolton

This button Is located on the right side of the panel. When0 cycle tone la produced for audio checking or transmission on either INT or EXT mode.

receiver-Transmitter control panel

The receiver-transmitter control paneljust aft of the translator control on the left console Is used when operating in tho oxtornal or wide band mods only, In this mode of operation thehe following functions:


voice communications.

Automatic Direction Finding.

Automatic Range Measurement.


Ranging and Direction

contains rotary selector and

address switches with positionpushbutton Interrogation andranging switches and indicating lightseparate response Indicator light. These switches and lights are only operative when the INT-EXT mode switchn the EXT position.

Code Selector Switches

Five rotary type digital Indicating selector switches labeled Sl-il, are provided on the panel. The digital window type Indicators have positions labeledo 7. stations roust have identical code selecting settings In order to establish wide band communication.


In this Installation only the let and 5th selector switches are operative givingossible code selections. 2nd and 3rd selector switches must bend the 4th select switch must be on 6.

Konge Addreii Switch

The inboard or 6th rotary selector switch labeled ADBS-BGE is used for selective ranging. It has eight positions which show position indicationsnclusiveossiblo range addresses. The other three positions are labeled A, O, and T.osition allowsangeon any terminal regardless of Its ad-drees (ADBS-HGE) code. This isan emergency code. osition Is an off position which prevonts another terminal from ranging although voicecapability is retained. ositionest position for checkinglights on the translator andtransmitter panels.

Ronge Interrogate Switch ond Indicator Light

This pushbutton switch containing an Integral light Is used to make interrogation ofand range In the external mode. When the translator function selector switch is In the ADF position the one time bearingwill be Indicated on theeedle of the BDHI. If the translator function switch is in the MAIN or BOTH position pressing the JJNT button willne time range measurement tn nautical miles and tenths. Normal time for one time directional or range indicationeconds. The button Is also used to establish automatic ranging and automatic ranging and direction finding in combinations with the CONT button. The

button will be momentarily depressed and the light illuminated while the ranging or direction range is being obtained. The light will be extinguished aftereconds.

Continuous Ronge Switch

This pushbutton switch and Integral light labeled CONT, when pressed, setsontinuousranging or combination ranging and DF conditions. This condition is activated by the Interrogate button provided that the other station has previously activated their continuous range operation. The lightilluminated at both stations whileranging or ranging and DF operation is in effect. The range indicator andeedle of the BDHI will be updatedeconds. Either station pressing its CONT button or MIC button will terminate the automatic cycle. When the cycle isbroken the range Indicatorsiles.

Response Light

The response light, labeled BESP, will be illuminated when thesa range measurement Interrogation from another aircraft or station.


The distance indicator is mounted on the upper left side of the instrument panel and displays the distance between two rangingets. Negative contacteading.


The UHF communication antenna is located on the lower right chine and remains ex-

tended. Provisions are available to hy-draulically retract the antenna flush. The UHF-ADF antenna ia mounted on the bottom of the nose compartment and is theantenna at any time the function switch on the translator is in the ADF position. It is optimized for the DF function andand ranging will be inferior when using this antenna for other than the direction finding function.


Internal Mode UHf Cormminicotiont

Microphone selectorHF.

FunctionAIN or BOTH.


Power selectoret.



If GUARD is selected frequency selection will be automatic.


Frequency selectoret. If PRESET:

Channel selectoret.


9. Volumes desired. External Mode Ul IF Encoded Communication noti

External mode Is applicable only when communication Is with anothertation.

Microphone selectorHF.

FunctionAIN or BOTH.


Power selectoret.



If GUARD Is selectedwill be In the narrow band conventional mode even though INT-EXT switch ts In EXT.


Frequency selectoret. If PRESET:

Channel selectoret.

Code selectoret.

Range addressr as desired.

Mlcress (tone will be heard forecond).

Volumes desired.

section rv

Automatic Bonging

A one time range interrogation le made as follows:

Select proper frequency and power.

FunctionAIN or BOTH.


Code Selectoret.

ADBS-BGE selectoret.

TUTress. Light willforeconds.

When the light extinguishes range Indication may be read.

To update range reading:


To communicate with range partner:

OLD. Walt for tone

Automatic Ranging

Automatic continuous ranging with bothreceiving continuously updated range informationeconds Is accomplished as follows:

Frequency andet.

FunctionAIN or BOTH.


Code selector and range address- Set.

Bequest selected range partner to press CONT button.



CONTheck on.

INT and BESPheck alternate Illumination. Both stations willupdated range reading"econds.


ontinuous rangeanging interrogation cycle Is not completed, thawill automaticallyonce. The digital rangewill be held during this period for approximatelyeconds and, If ranging Is not ro-establlahed, then reset to zero.

To resume communication:

ress. Tone0 seconds dependingpart of the ranging cycle Is


First transmission after muting will be to ranging partner only. Subsequent transmissions will be heard by all stations havingcode selections.

control -Opetoflon

During ADF operation thequipment and directional antenna are used for receiving, and the direction of signals from the responding UHF station will beby theeedle of the BDKL

section rv


Inferno! Mode Direction Finding

For DF operation in conventional narrow band mode proceed aa follows:

Select proper frequency and power.



Fequest communicating station fortransmission or tone.

Bearing to transmitting station will be indicated by the BDHIeedle.

External Mode Direct'.on rinding

For semi automatic or one time ADFproceed as follows:

Select proper frequency and power.



Code selectoret.

Range addresset.

INTress momentarily. Light will Illuminate foreconds. When light la extinguished, bearing will be indicated by tho BDHIeedle.

To update bearing;

TNTress. To resume communication:


For continuous updated ranging anddirection bearing proceed as above except:


Bequest ranging partner to press CONT button.




Holding the Mlc button until tone8 seconds one way ADF2 seconds two way ADF) terminates the automatic ranging and ADF functions. The CONT and INT buttons will re-establish the continuous ranging and ADF cycles.



The bearing, distance, heading Indicator located on the left side of the Instrument panelotating compaeaange shutter labeled OFF covering thedistance readout, andndirectional Indicating needles. Tho card displays true or magnetic headingon the position of the INS mode switch. In the NAV position,'true heading using the INSeference will bo Indicated at tho lubber line. In the FRS position, magnetic heading from the FRS will be Indicated. Theeedle will read an ADF bearing from thenless the AN/ARCsIn the ADF mode; or TACANdepending on the position of theeedle selector switch. Theeedle will indicate the steering direction from the INS. When reliable TACAN Information Is being received the range shutter will be up and the range readout will represent slant range to the TACAN station being

BDHIeedie Selector Switch

This two position switch Is located on the right consols. It selects either the TACAN orDF inputs to theDHI needle, provided thes notIn the ADF mode. In lhe TACAN {forward} position, theeedle ol* the BDHI Is connected to the TACAN receiver and the needle Indicates the bearing to the selected TACAN station. In tbe ADF (aft) position, theeedle Indicates the bearing to the selected station. When thes operating in the ADF function, the switch Is inoperative and theeedle Indicates the bearing to the selectedtation.


Theourse indicator Is Installed on the center Instrument panel. It Is used In conjunction with the BDHI to Indicate course deviation when operating the TACAN system. It is also used to Indicate course and glide slope deviation and marker beacon passage when operating the IXS receivers. ourse setourse selector window to show courseertical CDI course deviation needle and dot deviationorizontal glide slope Indicator needle and dot deviation scale,O-FPOM Indicator window,DI warning flagseading pointer with right and left pointer ncaleearker beacon light. The indicator is powered by signals from the respective receivers.


Tong range airborne single side band (SSB) voice communications transceiver which transmits and receives inoegacycle range. Thocan be tuned in one kilocycle steps. Tbe primary operating mode is SSB, using either the upper or lower side of thesignal, which allows all the power to amplify the side band selected. Thecan also transmit and receive AM signals.

The equipment consists of the transceiver with an anteruia tuner which is mounted in the pressurised nose compartment. The antenna is the pitot boom and Insulatedportion of the aircraft nose. This equipment has been modified to use fixed frequency ac power from thenverter for those circuits which are frequency critical and variable frequency ac power normally furnished from the left generator for non-critical main ac power. Asensing relay is provided to transfer this main ac power source if the left orgenerator bus dropsngine rpm) to thef the CO MM selector switch Is in the HF position. Control circuit power is supplied by the essential dc bus.

F Control)

The control panel for the HF equipment Is located on the left console and contains the following:

Service Selector Switch

This switch turns the equlpmont on or off and selects the desired operating mode. In the USB (upper side band) position, only the upper side band signal is transmitted or received. This is the sum of the voice signal and the radio frequency (rf) signal. In the LSB (lower side band) position, only the lower side band signal is transmitted or received. This signal is the difference of the voice signal and the rf signal. In the AM position the signal Is amplitudeand both side bands and the original rf signal are transmitted and received.


section iv

Selector SwItches

The first switch selects the properpoint as Indicated by the digits In the first two windows. It will indicate from The frequency will Increase as the knob ia rotated clockwise and decrease as the knob Is rotated counterclockwise.

c knob selects the proper one hundred kilocycle point and Indicatesn the third window. Thec knob selects the desired ten kilocycle point and Indicatesn the fourth window. The one kc knob selects the desired one kilocycle point and indicatesn the right window.

Volume Knob

This knob Is used to adjust the audio level In the headphones.



When background sound is again heard In the headphone:

- Press. Wait forto0 cps toneheard until tuning is complete.

When the equipment is tuned0 cps tone):

djust so thatIn headphones is barely audible.


hort circuit exists In the output of the powerrotective circuit turns off the equipment. Restore to operation as follows:

1. Service selectorFF, then back to desired operating mode.

selectoret to desired mode. This will turn the equipment on. For normal voice communication this may be USB, LSB or AM.

Frequency selectoret to desired operating frequency. The muting of sound in the headphones will Indicate the equipment is setting to the new frequency.


The service selector switch may have been moved from the OFF position to an operating mode with the desired operatingalready set up. In this case, rotate the ten kc select knob one digit off frequency and then back to the operating This will allow the equipment to return to thefrequency.


When the antenna coupler Isto complete severaltuninghermal relay will de-energize the Restore to operation as follows:

selector switch -two minutes thewill cool.

selectorooperating mode.

If HF and/or BW operation is required with Inoperative engines or generators:

W powern.

COMM selectorF.


The COMM selector switch must be In the HF position to provideHF or BW communication with wind mil ling engines and/orgenerators. In thisautomatic transfer of main ac power from the left generator to thenverter is accomplished If the frequency of the left generator bus dropsycles0 engine rpm. At0 engine rpmycles the automatic bus transfer occurs and if the right engine or generator bus is above I ngine rpm) the trequency relay will reconnect the main ac power to the rightpower source.


The Sel Call Decoderonvenient method for the selective reception of HF transmissions. It willallelected channel of the HF receiver and unmute the receiver when the proper call signal Is received. The decoder operatesreset Sel Call coder frequency and will recognize only this channel. omentary contact switch and Indicating light is on the left console to MUTE or UNMUTE the HF audio circuit. The Indicator light iswhen the decoder is In the muted mode. The Sel Call Decoder Is alsounmuted when the transmitter key is pressed which provides audio sidetone during transmission to the pilot's headset. Power for the Sel Call Decoder is furnished by the dc essential bus.


compartment with Lhe flush antenna mounted on the lower fuselage just aft of the nose boom. The EGG beacon transponder and antenna are mounted on theay hatch. Both beacons are controlledosition toggle switch located on the left console. The switch is labeled EGC-OFF-TNKR. Power for the transponder isby the dc essential bus.


The TACAN system provides continuousof bearing and slant distanceelected surface beacon or to anothercontaining the necessary transponder equipment. Tho system transmitspulses which trigger responding 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 visuallyon the bearing, distance headingon the instrument panel. Theis capable of operation on any one ofhannels andange ofautical miles. The transmittingrange50 megacycles. Frequency ranges for reception are; low band4 megacycles, air toegacycles; highegacycles, air toegacycles. Power for the equipment Is furnished by the left acand essential dc buses.

ontrol Panel

A control panel Is Installed on the right The panelhannelswitch, mode selector switcholume control.

aircraft ia equipped withand beacon and an EGG beacon. and beacon transponder is located In the nose





Selector Switch

A channel selector Is used to select any onethe available channels. Selection Isby setting the desired number In the channel window using the concentric knobs. The outer knob selects the first two digits and the inner knob selects the third digitesired channel.

Volume Control Knob

Audio level of the TACAN stationsignals is increased by rotating the volume (VOL) control clockwise.

Mode Selector Switch

he set is de-energized.

he set Is energized and presents bearing and course Information on the BDHI and course Indicator.

ame as the BEC position and also

presents range In nautical milesACAN station on the BDHI.

ame as the REC position and also presents range In nautical miles and bearing to another properly equipped aircraft.


L, INS modeRS if operative.

2. TACAN mode selectorEC.

mode Selector Switch has four positions.







Alloweconds lor warmup;

Channel selectoresired channel.

Adjust VOL as desired snd verifyIdentification.

eedle selectorACAN.

Observe bearing pointer on BDHI; To-Frotn indication on course Indicator.

Mode selector,.

Observe range to station or aircraft on BDHI,

Bearing, Distance, Heading Indicator (BDHI)eedle

The BDHIeedle may be connected to the TACAN receiver by the BDHIeedle selector switch. If the receiver is tunedACAN station, theeedlt will Indicate the bearing to the station. Refer to BDHI this section.

nd all tenthscs.VOR signals will not beon theourse Indicator. The associated glide slope frequenciesegacycles will botuned when the receiver Is tuned to the desired localizer frequency. TheON-OFF-VOL control also activates the fixed tunedegacycle marker beacon receiver and marker beacon signals areby coded audio tones In the headset and coded flashes of the single marker beacon light on theourse Indicator. As the marker beacon antenna la located on the Inside of the nosewheel door the marker beacon will only be usable with the landing gear down. The XLS ON light Is provided to li-.dlcate that the localizer is furnishing signals to theourse Indicator and that TACAN signal Inputs are disconnected from that Instrument. TACAN bearing and range to selected stations will still beon the BDHL All receivers are solid state and operate with power furnished by the essential dc bus.

ILS Control Panel


ILS equipment consisting of localizer, glide slope and marker beacon receivers are provided for ILS approaches. In addition the equipmentontrol panel and Indicating light, the ILS converter andantennas. Localizer, glide slope and marker beacon signals are reflected on theourse Indicator. Localizer signals are not reflected by theeedle of the BDHI which continues to show TACAN or HF/UIIF ADF bearings as selected.

The localizer receiver tunes odd tenthlocalizer frequencies00 mc. It will also tune VOB voice or tone signals between oven tenths from

The ILS control panel is located on tho lower right side of the Instrument panel. The panel controls consistN-OFF-VOL control concentricarger(megacycle) selector on the left side of the panel. Tho small knob turns the ILS equipment from off to on and furtherrotation will Increase the volume, of voice reception or tone identification. The larger knob selectsigit megacycle frequency of the desired localizer station which Is Indicated In the window In the center of the panel. The right hand side control knobs are also concentric with the small center knob to eliminate tonefrom the headset. The larger concentric knob controlsigit tenths and hundreth mc frequency selector which ia Indicated in tho frequency window. The ILS ON light Is located just above the panel

adf control panel

and Indicates that the TACAN receiver la disconnected from theoureeend course and glide slope deviation Indications arc from the LLS equipment.

OporoHon of IIS


Desired localizerelect.



Front courseelect.

heck ON.

Glide slope and localizer warning flags-Check not visible.

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 OFF position the equipment is de-energized. In the ADF position the equipment functions as andirection finderontinuousof the bearing to the radio station shown on the BDHI If thes not operating In the DF mode. In thisboth the sense and loop antennas are connected to the receiver. In the ANTreceived signals are obtained only from the sense antenna and ths equipment functionsonventional aural radio In the LOOP position receivedare obtained only from the loopand the equipment functionsanual direction finder to enable the pilot tothe bearing to the radio station by aural null procedures.


ThoDF radio receiver is anor manual direction finderow and broadcast range aural receiver. The equipment consistsadioontrollush senselush fixed loopDHI and thecabling, antenna coupleruad rental error corrector. The receiverrequency range5 megacycles in three bands. Power for the equipment is furnished by the essential dc bus andvolt Instrument transformer.

Bond Selector Switch

The band selector switch is the larger of the concentric knobs located In the outboard side of the control panel and Is used tothe desired frequency band. Thefrequency scale will also appear In the frequency indicator window for the band selected as follows:

Band Frequency


Low Frequency Band

Control Ponel

The ADF control panel is Installed on the right console of the cockpit. The controls are described below.


5 MC

CDistress Frequency and Lower Broadcast Band

0 KC Upper Broadcast Band

Tuning Control

The tuning control Is theof theconcentric knobs and tunes thewithin tha frequency band selected. The tuned frequency is indicated on the scale of the frequency Indicator. Thela alao rotated slightly for maximum reading on the tuning meter.

Loop Control

The control labeled LOOP Is used totha electrical equivalent ofthe loop (gonlo) antenna. The control is labeled L, and It and the left or righteffect will be apparent in the headset and the tuning meter. The speed of theeffect may be slowed by turning the loop control approximately half way toabeled position.

Goin Control

The gain control la the smaller of theconcentric knobs and Is provided to adjust the audio level of the receiver.


Opetotlon at the ADF Receiveronventional Rod io Receiver


Band selectorelect desired band.

Tuningotate to desired frequency and adjust for maximum reading on the tuning meter.

djust as desired.

The BFO switch can be used to tune In continuous wave signals or to aero beat modulated signals.

Operation o( the ADF Receiver at an Autorrotk Direction Finds*

Tune receiver as above and positively Identify the station.



The BFO switch when in the BFO positioneat frequency oscillator to aid In tuning the receiver or to receive coded transmissions.

Bearing, Dlitonce, heading indico'oreedle

ADF bearing Indication Is provided by theeedle of the BDHI when the BDHIeedle selector switch Is In the HF/ UHF (aft) position and the AN/ARC-SO ADF is not operating. Refer to BDHI, this

Tuningune for maximum signal reading on tuning meter.

BDHIeedle selector switch -ADF (aft).

Read bearing to station on BDHIointer.

Operation of the ADF Receiverongol Direction Finder (Aurol Noll)

Tune receiver as above and poatively Identify th* station.

Tuningune for maximum signal reading on tuning meter.

BDHIeedle selector-ADF (eft).


Loopr L, as necessary, to acquire null.


If the AN/ARC-SO ADF function is operating, theDHI needle will remain connected to the UHF equipment.


Theransponder providesdetection, decoding, encoding and transmission of signals in the IFFSIF) system andocally installediscrete operating function. The transponder will alsonterrogation: however, the set will notoreply without accessory equipment. Any one of numerous codedavailable for Modesan be selected by rotating the appropriate selector switches on the panel. The set is capable of transmitting an emergency reply regardloss of tbe Interrogation mode.rovision ia also incorporated to Identify position of the aircraft. Power for the set Is furnished by the essential dc bus. of theapability deletes (heunction from the transponder.


The transponder control panel la Installed on the upper left console. The paneltwo code selectors forndodes.ndoggle switches,witch, IFF powerswitch and an emergency switch bar.

Power Switch

The IFr power switch has three positions: Off, LO, and ON. When the switch is placed in the LO position only local (strong) interrogations are recognised and answered. With the switch in the ON position, there is full sensitivity for recognition and reply. The IFF power switch activateshen in the ON or LO position. Response tondnterrogations depends on the position of theoggle switches. When the Emergency switch bar Is up, the power switch Is forced to the ON position. 0 second time delay Isin the power switching before the equipment la operative.

Mode Switches

Two two-position mode switches, one fornd one for Modeontrolofndeplies. Correctly coded interrogations willode has been made active by selecting the Di position. rwitch la in the OUT position, that mode le not active and does notupon interrogation except In Emergency.s active at all times when the power switch is In the ON or LO position and is not affected by theroggle switch positions.

Code Selectors

Two rotating type code selectors are The code selector forf two rotary digital indicating switches. The first digit window,,,r 7. The second digit window will Indicate,rheode selector will indicate,,,or each digital Theode selection alsotheode transmission.


Emergency Switch Boi

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 emergencypulse groupa transmitted onach time an interrogation is made on Mode X. re also turned on by the emergency bar irrespective of the position of then-Out switches. In the EMERGENCY positionill respond on the code selected butill respond onf code selected,


The ground radar scope indication from this transponder Is codedifferent manner than the normalransponder.

Identifjcotion of Poiltion Switch

The Identification-of-position) switch is used to control transmissionulse groups. The switch has three positions; MIC, OUTosition. When the switch ia momentarily Inosition,imer is energised foreconds. If an Interrogation is recognized on any active mode within thisecondeplies will be mado. When the switch Is In the OUT position,ofulse groups Is withheld. The MIC position is inoperative.


PowerN or LO.


1 .andN-OUTrequired.


peration Is continuous when the power switch Is in the LO or ON position. For secure IFF operation, both thendoggle switches must be In the OUT position.

s required.

Codes required.

To make an emergency response to Modendnterrogations:

ush up.

Call Knob

The call knob is inoperative.

Normal-Aux-Llsten Switch

The Normal-Aux-Listen switch has two positions; NORMAL, and AUX LISTEN. The NORMAL position allows all audio signals to pass through themplifier. The volume control knob on theanel Is used to adjust the audio signal The AUX LISTEN positionthe amplifier and audio intensity must be adjusted with the individualvolume control. The switch Is safety-wired to the NORMAL position.


Annterphone control panel Is located behind the pilot's seat. The panelall knob, Normal-Aux-Listen switcholume control knob. Due to the location of the panel, the volume control must be preeet prior to flight. No On-Off switch Is provided and the equipment Iswhenever the essential dc bus le energized.

Microphone Switch

A transmitter-interphone mlrcophone switch is Installed on the control stick. TheTRANS position (up) Is used for UHF, or HF depending on the position of the microphone selector switch on the left con-sole. The INPH position (down) provides interphone operation for communication with the ground crewnterphone during refueling contact. This position Is also used to activate'the dlctet recorder for pilotif recorder switch Is in RECORD position.

Throttle Microphone Sutton

A microphone button is provided on the right throttle for use during taxi, takeoff and landing when the nose steering must be held engaged. Thisushbutton switch which must be held down for radio

Communication Selector Switch

A three poeltlon roUrylabeled COMM located on the left console selects the radio or Interphone to which theoutput will be connected. In the HF position the microphone output will betoT HF radio. This poeltlon also automatically provides ARCc inverter power to the HF when the leftIs below correct frequency. In the UHF (center) position the microphone output la connected to tha ARCHF radio. The right position labeled SXL disconnects the microphone from ail transmitters to prevent inadvertent transmissions. The microphone connection to the interphone system and to the tanker is through the refueling probe and is accomplished by using the normal INTFH position.

IFR Volume Cont-ol

The XFR volume control is Located on the upper left console and when turnedIncreases the Interphone audio volume.


Two retractable lights are located near the midpoint of the fuselage. One Is on the top of the fuselage and the other on the bottom. When the lights are retracted they are flush with the fuselage contour and when turned on willhite light from above and below. The lights will extend approximately two inches.and, when In this position and turned on, the red lights and reflectorsatpm, giving the effect oflashes per minute. The lights are powered by the essential dc bus and the rotating andmechanism Is powered by thenverter.

Beocon and Fuielage Light Switch

This three position switch is located at the forward end of the upper left console. In the center OFF position the lights areand turned off. In the BCN LTS (forward) position the lights extend,and rotate. Extension and retraction time Is approximatelyeconds. In the FUS LTS (aft) position the white lightsin the retracted position.

landing and tcxj lighti

0 watt landing lightatt taxi Light are mounted on either side of the nose gear strut. Power for the lights isby the left generator bus.

Landing and Taxi Light Switch

A luminousot) switch located on the left side of the instrument panel operates the landing and taxi lights. The switch has three positions; LANDAXI LT (down) and OFF (center).

INTERIOR LIGHTING Cockpll Lighting System

The Instruments and consoles arewith edge and post lighting. Intwo flood lights are provided on each side of the cockpittility spotlight is mounted above each console. Theare detachable and may be moved about the cockpit. Rheostats on the aft end of the spotlights are used to vary their Each spotlight Is providedushbutton switch which enables the pilot to obtain maximum brill ance without use of the rheostat. Red or white light may beby rotating the lens color selectors on the front of the lights. Power for the instrument and console lights is furnished by the left generator bus. Power for the floodlights and utility spotlights Is furnished by the essential dc bus.

Cockpit Light Switches

Rheostat type instrument and panel light switches are located on the cockpit left console. Ten rotary positions areto vary light intensities from OFF to BRT. The floodlight switch located on the outboard side of the right console varies the intensity of both lights from OFF to BRT.


An automatic, continuously operating flight recorder Is normally mounted In the right chine of the aircraft to record airspeed, altitude, vertical acceleration, heading and elapsed time on an aluminum foil tape. The recorder has Its own pitot static system which may also be used ae an alternate for the normal pltot-statlc system. Heading information for the recorder is furnished by the FRS compass system. Ac electrical power from thenverter is used topring motor wound so that all in-

formation except heading will be recorded for approximatelyinutes afterpower is interrupted. The recorder pltot static system remains available an an alternate airspeed System when theis not installed.

Flight Recorder Switch

This toggle switch is located outboardthe right console and has labeled positions ON and OFF.

Pitof Pressure Selector Lever

This lever Is located on the forward right side of the cockpit wall. It Is normally safety-wired In the NORMAL, position. In the eventalfunction of the normal pltot static position system, the lever may be moved to the ALT position. Thispltot static pressure from the flight recorder system to the aircraft flightand ejection seat speed sensor.


The Dictet Tape Recorder Is located on the left side of the canopy. as two levers; one labeled REWIND, RECORD andand one labeled ON and OFF. It ts preset prior to flight and is activated by the interphone switch. The tape is In motion only when the Interphone switch Is used and provides up to two hours of recording time,


The autopilot portion of the AFCS relieves the pilot from manual aircraft control andeans for automatic navigation when coupled to the output of the INS. The autopilot functions are:

Pre-engage synchronisation.

Attitude hold In roll and pitch.

Pitch and turn wheel inputs.

Automatic pitch trim.

Heading hold.

Mach or KEAS hold.

Auto navigation.

The autopilot Is optimised for basic missior cruise speed and altitude but may be used at other flight conditions.

There are no restrictions on use of the roll autopilot. The autopilot authority Is limitec to prevent severe maneuvers due to anmalfunction. Tbe maximurnoltch authority0 feet up and down eleven. 0 feet theauthorityp and down eleven. The maximum roll authoritylevon. The autopilot signals arewith SAS signals and produce control surface motion through the SAS electronics and servos.


Do not use the autopilot when using BUPD.

Autopilot control movement of the elevons is not reflected in control stick motion. Automatic pitch trim Is operative when the autopilot pitch channel le engaged. The slow speed pitch trim motor operates to correct for long period pitch trim changes and there should be no pitch transient at disengagement. Preengage synchronization of autopilot pitch and roll trim operates when the pitch or roll channels are

Aulopilol and Altitude Reference Selector Switch

This selector switch Is Located on the right console outboard of the INS control panel. The switch has three positions: FRSOFF (center) and INS (aft). In the FRS position directional signals from the FRS compass and attitude signals from the FRS pitch and roll gyros are supplied to bat' the autopilot and the attitude Indicator. In the OFF poaltlon the autopilot can not be engaged but pitch and roll signals from the FRS are furnished the attitude indicator. In the INS position the INS stable platformpitch and roll signals to both the autopilot and the attitude Indicator and true heading directional signals aro furnished to the autopilot. In the OFF and FRS positions Inverter power for autopilot, air dataand TDI Indicator is furnished by thenverter bus. In the INS positionpower for these Items Is transferred to thenverter bus. This switching provides the same phase of power for the autopilot and the air data computer as that provided for the FRS or INS.


Avoid excessive switchingFRS and INS positions as the resulting powertend to degrade INS


The autopilot controls and Indicators are or the SAS panel located on the right console. The control stick Is equipped with controlcommand and emergency disengage The circuit breakers are on the right and center console circuit breaker panels. Power Is from the essential dc bus and thernverter.

Autopilot Pitch Engage Switch

A two-posltlon pitch engage switch laon the Inboard side of the autopilot control panel. In the ON (fwd) position, the pitch autopilot is engaged in the attitude hold mode.


At least one active SAS pitch channel must be engaged and bank angle must be less than 30 before the pitch autopilot can be engaged.

The switcheld In the ON positionolenoid. The pitch channel may beby placing the switch to the OFFby using tlie disengage switch on the control stick, or by turning the autopilot selector switch OFF.

Autopilot Pitch Trim Synchronization Indicator

The pitch trim synchronization Indicator shows the amount of pitch signal existing prior to engagement. An up or downof the needle indicates theof the transient which will occur when the pitch channel Is engaged.


The pitch trim synchronization needle will normally be centered within one needle width. of the autopilot pitch channel with more than one needle width of misalignment Is not

Autopilot Pitch Control Wheel

A. serrated pitch control wheel le located just forward of the pitch engage switch. The wheel Is used to make pitch attitude corrections when engaged in the attitude hold mode. Forward rotation of the wheel commands nose down and aft rotationnose up. Pitch attitudeor 20 of wheel rotation.


AutopUor Rol! engage Switch

turn Control Whcr

two-position roll engage switch la located on tlie autopilot panel. In the ON (fwd) position, the roll autopilot in engaged In the attitude hold mode.


At least one SAS roll channel and one active SAS yaw channel must be engaged before Lhe roll autopilot can be engaged. Bank angle must be less than

A serrated turn control wheel Is located on the autopilot panel. It allows the pilot to make roll attitude corrections when engaged in the attitude hold mode. Right rotation of the wheel commands right roll and leftcommands left roll. Roll attitude changes 1 for 10 of wheel rotation. The pilot can command up to 50 of bank angle

In the attitude hold mode. Above 50 of bank the roll autopilot automaticallyto prevent the steady pitch rate from bottoming the pitch servos, as this would

eliminate pitch damping capability.

switch Is held In the ON positionolenoid. Autopilot signals are supplied by either the FRS or the INS, depending on the position of the autopilot selector switch. The roll channel may be disengaged by placing the switch to the OFF position, by using the disengage switch on tho control stick or by turning the autopilot selector switch OFF.

Autopilot Roll Trim Synchronlaction Indicator

The roll trim synchronization indicator shows whether oroll signal exists prior to engagement. The needle always deflects to the right and does not Indicate the direction of the transient which will oc* cur at engagement.


Roll engagement Is notif the needle Is deflected to the side of the dial,ardover signal.

Moch/KEAS Hold Switch

A Mach/KEAS hold switch is located on the inboard side of the autopilot panel. The Mach or KEAS hold mode is engaged when the switch is In the respective position, provided the pitch autopilot is engaged. Tbe switch is held In by solenoid action. The autopilot then controls the pitch attitude to maintain the same Mach number or KEAS that existed at the time of engagement. When the Mach or KEAS hold is engaged, the pitch attitude hold is discontinued and the pitch control wheel setting should not be changed. Mach hold reference signals are supplied to the autopilot from the air data computer.


Do not use the Mach/KEAS hold mode when the TDI indication is known or suspected to be

Auto Nov Switch

AUTO NAV switch i* located between the Mach/KEAS hold and heading hold switches. The auto nav mode is engaged when the switch Is in the ON position provided the

roll autopilot Is engaged. The switch in held on by solenoid action. Steering signal are furnished by the INS and the autopilot controls the aircraft to follow the selected great circle course. If the heading hold mode was previously engaged. It will be disengaged when auto nav is selected. The bank angle ia limited to 30 in the auto nav mode.

and heading to be changed without opposition from the autopilot. When the switch Isboth the roll and pitch axes arein the attitude hold mode, regardless of the mode that was engaged prior tothe CSC switch.

Autopilot Emergency Disengage Switch

Hold Switch

A heading hold switch is located on theside of the autopilot panel. Thehold mode Is engaged when the switch Is in the ON position provided the rollis engaged. The switch is held on by solenoid action. Heading signals from cither the FRS compass or INS control the roll axis of the aircraft to maintain the heading existing at the time of engagement. Heading hold may be engaged whileank. The autopilot will roll the aircraftings level attitude and lock on the heading at time of engagement. Thehold and auto nav switches areto permit only one to be engagedime. The auto nav switch will bewhen the heading hold switch Is on.


When In heading hold mode the drift rate Is eirnllarree gyro rate and will be approximately 8 per hour increasing to 15 per hour In polar areas.

Control Stick Command Switch (CSC)

A control stick command switch is located on the right side of the control stick. While the switch Is depressed, both the roll and pitch autopilots revert to the preengage synchronization mode. This allows attitude

A trigger-type switch located on the forward side of the control stick will disengage the autopilot completely. The autopilot is not reengaged when the switch is released.


The autopilot Is placed in normal operation as follows:

Check SAS engaged, recycle lights out.

Check pitch and roll trim preengage synchronization Indicators aligned.

Pitch and roll engage switches ON. These switches may be engaged together or separately as operation of the two Is completely Independent.


Bank angle must be less than


Headings desired. If Auto Nav le required:

Autopilot selectorNS.

Auto NavN.

Do not operate manual roll or pitch trim when Lhe autopilot is engaged.

s required.


Maintain stabilized KEAS or Mach conditions foreconds.

To change attitude or heading:


After attitude and/or heading change;


Mach/KEAS holdFF,

Autopilot selectors desired.

Heading hold or autos desired. To disengage autopilot:

1, Autopilot disengageress.


Do not engage Mach/KEAS hold during turns or other maneuver" as undesirable transient will be produced. Mach/KEAS hold may however be loft engaged during turns If already on.

6. Mach/KEAS holdN as desire*


The pitch control wheel must not be used during Mach/KEAS hold operation to prevent rapid pitch motion or disengagement.

2. Pitch and roll engageFF.


3. Autopilot selectorFF.

Moch/KEAS Hold Engagement

Prior to engagement of Mach/KEAS hold the pilot will accomplish the following:

1. Attain desired KEAS, altitude and Mach number.

To minimize altitude excursions during turns;

radually advance during roll in.

radually retard during roll out.

If changing flight conditions, retrim when power settings arc changed more


2. Throttle s required.


The Flight Reference System andavigation system which supplies information for indication and control of aircraft heading and attitude. It can be used independently of the lnertlal Navigation System. The FRS consistslightplatform, turn rate servo,compass transmitter, heading andcouplers for the autopilot, control panel, and the rotating compass cards of the BDHI. irectional gyro (DC) or magnetic slaved (MAG) mode can beto provide directional reference to all latitudes. In either mode:

Heading information is furnished -

To the autopilot when the autopilotswitch Is in the FRS position.

To the BDHI compass card when the autopilot selector switch is In the FRS position.

Attitude Information Is furnished -

To the autopilot when the autopilotswitch Is in the FRS position.

To the attitude indicator when theselector switch is in the FRS

useful when the magnetic field Is weak or distorted or when navigating in the polar regions. It is more reliable than themode at latitudes near the magnetic poles. When In the DC mode, with proper hemisphere and latitude selection made on the control panel, the gyro is made toto compensate for apparent gyro drift due to earth rale at the selected latitude.

Mognetlc Sloved Operoting Mode

When operating In the magnetic slaved mode, the FRS isyro stabilisedslaved to the Induction compass This mode provides headingnortherly turning error or oscillations. It Is less reliable than the DC mode atnear the magnetic poles as the MAG mode Is subject to severe magneticnear those poles,


The COMPASS controls are located on the right console, immediately forward of the circuit breaker panel. The panelunction selector switch, set heading knob, latitude selector knob and indicator window, synchronization indicator, malfunctionhemisphere selector switch,ake command button.

Take Command Button

Directional Gyro Operating Mode

When In the directional gyro mode' ofthe FRS is free of magneticand operatesirectional gyro, indicating heading relative to an arbitrary reference heading selected by the pilot. It may be used at all latitudes, but is most

A combination button and light on thepanel provides for transfer of control of the FRS by depressing the button andthe green light. It is not operative on this installation.


F'-ne'io" jcicclor Sv/ifcn

The two position function selector switch allows selection ofree gyro reference. The DG (right) position xelects directional gyro mode; the MAG (left) position selects the magnetic slaved mode.

Hemisphere Selector Switch

The Hemisphere Selector switch must be set to correspond to the hemisphere In which the aircraft is located. The left (S) position is used when in southern latitudes. The right (N) position ia selected for northern latitudes.

Lctirudi- Selector Knob ond Indicator

The latitude selector knob may be rotated to select and display latitude in degrees and tenths of degrees in the indicator window. The knob Is used only In the DG mode. The latitude setting Is used In the DG mode to correct the directional gyro for the apparent drift due to the earth's rotation. Foroperation of the FBS in the DG mode, the latitude indicator must be act towith the actual latitude of the aircraft at all times.


The proper corrections will not be made if the hemisphere selector switch setting does not correspond to the hemisphere In which the aircraft Is located.

Malfunction Indicator


monitored functions from normal operation will cause the Indicator to display three white triangles,

Hecdlriii Se: Ki-ob anc Sy-icliron motion Hdiculor

The heading set knobeans to fast slave or eynchroniae the rotatingcard of the BDHI to the correctheading or desired gyro heading,on the position of the functionswitch. When in the MAG mode, initial synchronization with the compass transmitter heading le obtained by pushing and holding the heading set knob until the synchronization indicator becomes centered. In the DG mode, the heading is set to the desired Initial indication by pushing and turning the heading set knob. Turning the heading set knob clockwise produces anheading, with the rate of change being indicated by the deflection angle of the synchronization Indicator.


Function selectorAG or DG, as desired.

Hemisphere selectoret to correspond with aircraft location in Northern (N) or Southern (S)

Latitude selectoret towith existing Latitude when DG mode selected.

Heading setynchronize or slave to heading desired.

Autopilot selectorBS.

malfunction indicator is provided which monitors the power supply and other prime system functions. Any deviation of the



The normal ul living rate of the FltS Is per minute. The gyro may be as much from the proper heading when the compass system la energlxed before takeoff, and as muchours would be required to slave to the correct heading at normal slaving rates. Manual fastis provided hy pushing and holding the heading set knob depressed. This Increases the slaving rate per minute and willSO error in IS seconds.

If the compass is properly slaved before takeoff, no In-flight manual fast slaving is required unless free directional gyroi> selected. When operating in the free gyro mode, the desired heading can be established by using the heading set switch.


The roll autopilot must be disengaged before attempting manual slaving when the FflS Is being usedeading reference.


Thelight reference platform consistsingle axis directional gyro which is attitude stabilizedwo axis vertical gyro. ompass transmitter Is provided which establishes the directional reference while In level flight by detecting aircraft heading with respect to the horizontalof the earth's magnetic field. When the system is operated in the magnetic mode, the directional gyro is slaved to the compass transmitterateegrees/minute. When operating in thegyroode, the compass transmitter signal is disabled and thereference is established by thegyro operatingree gyro (except

for earth's rate latitude correction). gravity sensors are used Inwith pitch and roll torquer motors to eroct the attitude gyro to the local During periods of acceleration or deceleration along th* flight path, heading and pitch attitude errors can be Introduced due to the following affects:

a. The pendulously supported compass

transmitter la displaced from theplane and becomes sensitive to the vertical component of the earth's magnetic field. This results in anheading reference. Theof this errorunction of aircraft heading, transmitter tilt angle and the relative magnitude of thefield component. This error Is introduced Into the system at theslaving rateegrees/ minute.

h. The pitch erection aensor, which Is acceleration sensitive, provides an output signal to th* pitch torquer causing it to process th* attitude gyroalse verticalormal rateegrees/minute.

In order to minimize th* above deficiencies, an electrolytic fore or oft accelerationsensor (similar to the pitch erection sensor) Is provided on the pitch gimbal of the attitude gyro. This sensor disables the pitch erection and slaving circuitshreshold settinglong itsaxis is exceeded. ree vertical gyro. It Is subject to an apparent drift from the vertical due to the effect of the earth'a profile and earth's rotation. These effects, coupled with the gyro free drift rat* ofotal drift from the vertical ofegree/mln. This displacement of the attitude gyroravityto appear along the sensitive axis which actsias to the) which initiates slaving and pitch erection cutout. When ue

bias signal acta in opposition Lo the sensed acceleration signal thevalue may drop belowhreshold, thereby restoring pitch erection and slaving while the aircraft is still accelerating. Thegyro will then erectalse vertical at the normal erection rate and the compass transmitter will precess the directional gyroalse heading as determined by the transmitter tilt angle. When the attitude gyro driftegrees, and the bias signal acts to aid the sensedsignal; the system will maintain the cutout condition for an indefinite period after aircraft acceleration has ceased. In order to prevent this condition, the system is designed to limit pitch erection cutoutaximum periodinutes Independent of acceleration.

In operational use theystem performs In the manner described above duringof prolonged acceleration such asacceleration-climb to supersonic cruise speed after takeoff and after refueling.

During cllmbout, pitch attitude and heading errors increase toegreesegrees respectively. Those errors are eliminated at the normnl rateB whenacceleration ceases. The heading error can be washed out very rapidly by pushing and holding the heading set knob on the FBS control panel until theindicator becomes centered. During aircraft turns in excessegrees/min the system operates as designed to cutout roll erection and slaving. urn Is Initiated immediately following cllmbout, the accumulated cllmbout heading error will be increased and be maintained throughout the turn.


The inertial navigation system is self-contained and Operates in all modesthe use of electromagnetic radiation or external references. The system consistsyro-stabilized platform, platform electronics, coupler and power Supply, and converter assembly, digital computer and computer power supply,panels, and dlstance-to-go, ground-speed,irection indicator.

In operation the system displays present position, ground speed and the direction and distance to go to any ofreselected positions as continuous readouts. When operated in autopilot AUTO NAV, and INS STOBED AUTO mode, the aircraft will be steered automatically to each point in the night plan sequentially, with no pilot action required. If the flight plan Is being flown in sequence in the STOBED AUTO mode, the destination select light will illuminate if the destination displayed on theselect panel does not agree with the destination towards which the aircraft is flying. This light is extinguished when the pilot sets the selector panel to the number of the stored destination being approached.

The destination select panel providesof destination The firstreselected 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 greatcourses. Turns from one course to another will be made with bankaximum bankor the groundspeed and heading change required. Theointer of the BDHI will point toward the optimum path to follow to place the aircraft on the next course. If the pilot switches to adestination in STOBED MANUAL

before completing the route segment he is on, the turn will be made in accordance with computer program directions.

Positionsorovide ADF typofor courses to these points and 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 pathosition Included in theufficientof alternate destinations is available to provide adequate coverage throughout the mission. Duplication of any of the firstosition* in this groupteering indication on the BDHIointerthat of ADF navigation,he pointer points directly to the5 degrees needle deflection).

The basic reference of the inertialsystem Is provided by three single-axis accelerometers mounted at right angles to each otheryro-stabilized platform. The platform employs three floatedgyros, also mounted at right angles. The platform Is initially aligned with areference frame, representedlane tangent to the surface of the earth and oriented to any convenient azimuth at the point origin. The platform stableis isolated from the airframeystem of three gimbals whichegrees freedom of rotation In yaw and roll, and pitch angles0 degrees. All platform outputs are changed to digital form before entering the computer. In normal operation the platform also providesoutputs in analog form through resol-vers and synchros to the autopilot, and the attitude indicator. Conversion of present position to latitude and longitude readout Is accomplished continuously by the digital computer when in operational mode. air, necessary to the system. Isby the aircraft airconditlonlng and pressurization system. elf-contained heating system is Incorporated in theto ensure that gyros and precision sensing components are maintained atwithin an optimum operating range. The system Is powered by thenverter, the LH generator, and the monitored dc bus.


Accuracy of INS Information will be slightly degraded If pressurealtitude data supplied by the air data computer Is lost or is

The INS is controlled from two control panels, the navigation panel and theselect panel. (See.


The navigation control panel, located on the right console, consistsEST/FDC selector switch, STORE pushbutton, MODE selector switch, FIX ADJ knob, two sets of geographic coordinate digital readoutlabeled PRESENT POSITION and DESTINATION/FIX POSITION, aINPUT Indicator labeled LAT and LONG, with thumbwheels for manualof geographic coordinateswitch for selectionatitude. Theand indicators are as follows:

Mode Selector Switch

The mode switchotary selector switch with five positions, labeled as follows: OFF, RST, ALGN, NAV, and FRS.


During flight the MODE switch must not be switched to any position other than NAV or FRS, otherwise the INS will be deactivated and will notuntil the switch is moved through OFF, RST, and ALGN positions In conjunction with the ground operating equipment and normal TN5 preflight procedure.

ins panel and indicators


Do not move the MODE selector switch from the OFF position in flight if the INS hss not bcon cycled from OFF to the NAV mode prior to flight. The INS system will be damaged.

RST Mods

The RST (reset) mod* is used only on the ground during TNS preflight when thehas reached operating temperature, tt permits the ground operating equipment (COE) operator to check correct power switchover from ground to aircraft povter, start the gyro spin motors, and make the computer ready for us*.


The INS must ba completely warmed up, stabilized, and alignedoordinateframe before it can be operated. This Is necessary lo minimize the drift of the stable reference platform once it is aligned to the coordinate reference frame. The complete warmup and alignmentat normal ambient conditions takesours. During this period the destination loading operation ianormally by us*unched tape. However, the coordinates of the present location andestinations or targets may be set in manually by the VARIABLE INPUT thumbwheelselector and entered Into the computer memory by pushing the STORE pushbutton for each position.eriod of gyro stabilisation, the platform is torqued to the coordinate reference frame and the gyroa are drift-trimmed. The two transverse horizontal acceler-ometers 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 accclerometerthat the platform is not level in that axis. While level aignment of the platform is being accomplished automatically,azimuth Is alignedelectedwhich Is transferred to theby the ground operator. The platform is drift-trimmed at the reference points thus established, and the drift will beto certain preeetablished ratesthe system can be operated. Thereetent between NAV and ALGN positions and the MODE switch cannot be moved either way between these two positionsit is depressed.

NAV Mode

Switching to the NAV mode permits the COE to be disconnected, and places the platform In the operational mode. The gyros arc essentially memory devices that memorize the coordinate frame established. The system operates using thesecoordinates to perform the navigation problem, and the accelerometers measure translations of the platform caused by movement of the aircraft. Theoutputs are integrated once tovelocity on each axis,econd time to establish their displacement from the point of origin. These displacements (distances flown) arc translated intoposition coordinates by the In addition to indicating position coordinates to the pilot, this position Is also used to torque the platform to thevertical and azimuth as the aircraft changes position. The coordinate frame thus rotates about the earth to retain its orientationlane tangent to theof the earth at the position of the

FRS Mode

The flight reference system la the primary backup for the INS. Normally, the INS is operated with the switch In the NAVbut the pilot may switch to the FBS position at any time to check FBS operation. When the switch Is In the NAV mode, the BDHI rotating compass card indicates INS true heading; when in the FBS mode, the card indicates magnetic heading. When the switch is moved from the NAV to the FRS mode, the INS system continues to operate normally.


If the INS should fall, the MODE switch should be moved to the FRS mode without delay in order toeading Indication on the BDHI display.


The DEST/FIX (destination or fix) switchive-position rotary selector switch with positions as follows:





STORED AUTO. The INS will automatically sequence consecutively through there-storcd destinations as each Is reached when the switch Is In the STORED AUTO position.

STORED FIX. Toix point, the switch is set to the STORED FIX position, tbe destinationpanel is set to the desired destination

number, and the STORE pushbutton iawhen the fix point crosses theline on the periscope screen.

STORED MAN. To select any ofoordinate positionsestination, out of the automatic consecutive sequence, the switch is set to the STOBED MAN (manual) position, the destinationpanel is set to the desired destination number, and the STORE pushbutton is

VARIABLE FDC. Toariable (un-stored) fix pointoint of reference, the switch is set to the VARIABLE FDC position, the VARIABLE INPUTare set to the fixpolnt coordinates, and the STORE pushbutton is depressed when the fix point crosses the horizontal line on the periscope screen.

VARIABLE DEST. Toariable (unstored) destination, the switch is set to the VARIABLE DEST (destination) position, the VARIABLE INPUT thumbwheelselector are set to the desiredand the STORE pushbutton Is


The fix-adjust knob, labeled FDClight cursor on the periscope and is used to update the INS by means of visual 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 FDC ADJ knob to coincide with the fix point as it crosses the horizontalline on the display. (Refer toof fix-taking for further Information.)

STORE Poihbutton

The STORE pushbutton is used to store in the computer memory cither selectedinformation or position information which ha* been selected by the VARIABLE INPUT thumbwheelselector. It also initiates the computations required to navigate to the coordinates selected.


Do not push this buttonourse change or fix is desired.


The DEST/FDt pushbutton on the destination select panel isin function to the STORE button on the navigation panel. They may be used Interchangeably.

the desired destination coordinates with the VARIABLE INPUT thumbwheels; selecthemisphere withelector and depress the STORE pushbutton. The DES-INTATION/FLX POSITION indicator will read out the new coordinates immediately after the STORE button Is depressed, and the INS will navigate the aircraft to the new destination using ADF type steering. update fix coordinates are Inserted In tlie computer in the same way as aexcept that VARIABLE FDC ison the DEST/FDC switch.

present position indicator

The PRESENT POSITION indicator Is set at the geographical coordinates ol the flight origin site prior to takeoff. In flight It continuously indicates the coordinates of the aircraft position as computed by the INS.

Hemiaphere Selector Switch

elector switch may be placed inr S, depending in whichthe desired destination or fix is This selector is only used inwith the variable Input thumb wheels to manuallyestination or fix point in flight.


The VARIABLE INPUT indicator hasthat are used to manually Insert any desired reference coordinates into thethus giving the pilot added flexibility of operation In flight. {It is good practice to put the DEST/FIX switch in theDEST or VARIABLE FIX position prior to setting the coordinates in the) To Insert variable destinationinto the system, select VARIABLE DEST on the DEST/FDC switch, then insert


The DESTINATION/FIX POSITION indicator normally displays the Latitude and longitude coordinates of the destination to which the INS is navigating. This display may be the coordinates of any selected destination from therestored positions, or theof any selected variable destination. This coordinate display normally changes at such timeB as the computerew courseewly selected For STORED MANUAL orDEST modes, this change will occur upon depressing the DEST FIX or the STORE pushbutton. For sequential or out of sequence destination selections in STORED AUTO mode, the destinationdisplay will change coincident with roll out to the new destination course. The minutes counter portion of the latitudemay also changeix is taken. WhenTORED FIX orfix is taken; the calculated

recti on (in nautical miles) is displayed on ths latitude: minutes display, withoutlongitude, or the degrees portion ofon the DESTINATION/FIX POSITION Indicator. The portion of the latitudeused for the fix distance indication is blocked off In white on the indicator (see. The calculated fix correction Is displayed upaximum value ofautical miles whether position la updated or whether the fix is rejected. Thefix correction will continue to beuntil another fix is taken, orew destination is selected and displayed.ew destination is selected, theminutes counters will revert to aof destination latitude until such time as another fix is taken.


The destination select panel, labeled NAV, is located on the instrument panel. The panelwo-place digital counter,by thumbwheels, and apushbutton switch which read out DEST FDC when lighted. The numbertored destination or fixay be set on the counter manually and inserted into the IKS computer by depressing either the DEST FDC or the STORE pushbutton.


Positionshroughan bebut are inoperative.

The DEST FIX pushbutton illuminates when the destination number on the panel and the destination approached by the aircraft are not the same. When they are again the same (thumbwheels must behe light will go out. In all modes the light will come on when pilot action is required. When tho DEST/FIX switch is placed in either STORED or VARIABLE FDC, the light

will come on. When the STORE pushbutton is deprossed the light will go out. In any mode inew destination is selected by depressing the STORE pushbutton, tho light will go out when the system accepts the new destination selection. When aInside tlie aircraft's minimum turn radius is selected In the STORED MAN or VARIABLE DEST mode, the DEST FDC light will blink on and off. When the aircraft's location falls outside the minimum radius path, the blinking DEST FDC light willand the destination will be accepted. In the STORED MAN mode, the light will also come onestination is passed over byiles withouta new (DTGM or greater and


A distance-to-go and groundspeed Indicator Is installed on the Instrument panel. Digital indicators display the distance between the aircraft position and the destination, and the groundspeed, in unitsautical mile and knots, respectively. ew destination is selected either automatically or manually the Indicator will change to show the new distance-to-go. The dlBtance-to-gowill decrease toward aero whilethe destination, then increase after passing the destination if flight ison the same course. Distance-to-go will not read aero at destination if thecross-course distance is greaterautical mile, since readout resolution is to the nearest nautical mile.


The INS computes true heading and steering information and this Information can be die-played by the BDHI installed on thepanel. The rotating compass card of




4 -4

' K'.'l IO;, IV

BDHI receive* Iho true heading signals as long as the MODE switch on the INS NAVIGATION control panel is In the NAV position. When the MODE switch is in the FRS position the compass card ia driven by the FRS signals, although the INS system stiLl generates true heading. f the BDHI is driven by the ADF or TACAN as selected by theeedle selector switch. s driven by the steering signal of the INS when the MODE switch on the NAVIGATION control panel Is In the NAV or FRS position. oints to the direction of the great circle course or in ADF steering mod* will point towhich iire withinegrees of the aircraft heading (or indicate direction to turn if angular difference Is greater thanegrees).


. The aircraft will automatically fly the course computed by the INS and selected by the pilot only if the autopiloti the AUTO NAV mode.

degrc* turn Indication on the BDHI0 degree bank angle to be made by the autopilot. The bank angle command is proportionately smaller when smaller turn angles areon the BDHI.

course selection

In the STORED AUTO mode, the INS is capable of providing steering information to any selected destination when the path from source to destination is greater thanautical miles but las*ilesegreeegrees of great circle arc). In the STORED MAN mode, the above restriction* exist only for destinations numbered The sequence in which courses are providedupon the position of the DEST/FIX

switch on the navigation control panel. In STORED AUTO position, course direction* will be provided to stored destinationsin their numerical sequence; however, an out ofdeviation can be made In STORED AUTO by selecting the desired out ofdestination number on the destination select panel and depressing either the DEST FIX or STORE pushbutton. After tho out of sequenceother destinations will then continue to be automatically selected In numerical sequence. In the STORED MAN orDEST positions, steering directions to Individual destinations are supplied after each destination la selected by depressing either the DEST FDCar STORE pushbutton. For STORED AUTO or STORED MAN modes, the steering information provided by the computerreat circle flight path only if the destination selected is one of the firstets of stored. ADF type steering will be commanded for STORED destination selections numberedr greater and for all VARIABLE DEST mode selections. In STORED MAN mode, the computed course starting point laas follows;

position of the currentia selected by thethe starting point for theIf the aircraftIsilespoint when the STOREdepressed.

computed position of theis selected by thethe starting point for theif the distance toiles fromdestination.

ourse has been selected andand either great circle or ADF type steering provided to navigate toward the course destination point, the INS wiU contlnue to navigate to that point regard-





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1 of any change of position of tha DEST/ FIX switchew destination Isby either automatic sequencing in the STOKED AUTO position or by depressing the STORE pushbutton in the STORED MAN or VARIABLE DEST positions. If aselection is made in which the new destination Is aft of tlie present courseby an angle greaterhe initial steering direction is Indetermlnant and the aircraft may roll out cither right or left in turning around to the new course.

Fixed-Poth Flight Plon

A pre selected-path flight plan will bo flown In AUTO mode. Consecutive destinationshroughill be selected automatically. The point-to-point paths will be segments of great-circle arcs for destinationsnd direct for destinationshe use of STORED AUTO mode results in smooth entry turns at required bank angle upaximum ofegrees to the next course. Turns will be initiated before reaching the destination and the turn point will depend on aircraft groundspeed and the degree of coursa change required.

Deviation from Fixed-Path Flight Plan Uilng Stored Dci'friatloni

Stored destinations may be selectedIn any arbitrary sequence, and acan be selected any number of timesission. Any partial sequence of the stored destination fixed-path plan can be used by manually selecting the firstof the sequence, then switching to STORED AUTO mode until the desiredla accomplished. Then,ew stored destinationew course to be computed as described above.


In the STORED MANUAL mode, if the aircraft flys over theIn great circleew destination, the DEST/FDC light comes on and the vehicle will alternate between right and left steering signals. The DEST/FIX light operates similarly in ADF steering;the aircraft will fly in circles, always coming back over the selected destination.

- Ihjn-Stored Auto

One or more destinations can be skipped by selecting the destinations desired on thecounters of tho DESTINATION SELECT PANEL and depressing the STOREwith tha DEST/FDC switch in the STORED AUTO position. The INS willthe track In progress when the STORE pushbutton ia depressed but die next automatic sequence will select the course to thedestination.

In the STORED AUTO mode, the destination select lightxtinguished when the number on tlie destination select panel agrees with the stored destination which Is presently selected. The stored destination whichthe light will be the same as the stored destination toward which the aircraft is flying except whenestination out of sequence in the STORED AUTO mode.

Example: The aircraft is flying towards destinationn the STORED AUTO mode ands selected on the destination select panel. The destination select light Is pilot decides tond fly from destinatione selectsn the panel and depresses

the store button. The light will now beonly on destinationven though he is still flying This Indicates to the pilot thatas beenas the next destination.

Ute of the VARIABLE INPUT Indicator For Unstored Destination!

Use of destination coordinate* set on the VARIABLE INPUT indicatorrequires that the DEST/FIX switch be set to the VARIABLE DEST position. ADF-type steering to the point selected is provided when the STORE pushbutton Is The initial ADF-type steering heading is based on computed present Coordinates of stored destinations can be duplicated,


The maximum great-circle arc between source and destinationegrees todefinition of direction. Thisa distance ofautical miles from source to destination.

Mini tip Path LeigUi

In the STORED AUTOourse cannot be selected when the distance from the start pointtored destination or thepresent position) to the nextIs less thanautical miles. The computer will ignore any attempt to selectestination. In the STORED MAN mode theile restriction exists only for destinations numbered all destination selections are restricted by comparing the desired destination'slocation with the aircraft's minimum turn radius capability. (The minimum turn radius is computedunction of ground speed.) The destination is accepted if it is

outside the minimum turn radius path. If the desired destination is inside theturn radius path, the DEST FIX light on the DESTINATION SELECT PANEL will blink on and off, indicating that the computer has acknowledged the destination. Thewill continue on its same course until its location falls outride the minimum turn radius path. At such time, the DEST FIX light will extinguish and the destination will be accepted. (See)

Minimum Distoncc Between Petti not! org

In the STORED AUTOoursebe selected when the distance from the start pointtored destination or the aircraft's present position) to the next destination is less thanautical miles. In the STORED MAN mode theileexists only for


Since all rotating gyros are subject to some drift, alignment of the coordinate reference frame established by the gyro platform tends to depart from the true coordinate frameeriod of time. Thiserrors in position and azimuth which increase with time. (See) The indicated position can be updated by taking visual fixes when the coordinates are known. These fixes are taken by use of the perl-scope and are inserted Into the INS as

select the desired prestoredon the destination selectset the coordinates of the fix point

in the VARIABLE INPUT indicatorelector.

the DEST/FDC switch to theSTORED or VARIABLE (Use STORED FIXthe fix to be made is at apoint.)

Pull the MIRROR SELECT handle to the aft position for surface viewing andthe narrow view magnification with the PERISCOPE control. When the fix point is identified visually, position the periscope cursor with the FIX ADJ knob so that it will intersect and track the fix point. Continue tracking until the fix point crosses the periscopereference line.

Depress the STORE pushbutton at the instant the fix point crosses theof cursor and horizontallines. (At highec-ond delay in depressing the STORE pushbutton will resultosition fix error ofauticalhe computer will make the Axas follows;

the inserted fixrepresent the position of abelow the aircraftinstant of STORE

the fix position withcomputed position atof STORE pushbuttonand displaying theposition on the PRESENTIndication.

the difference is greaternautical miles, thewill not makeix was notIndicated by illumination ofFIX REJECT light on thepanel.

difference willaximum valuehe latitude minutes counterDESTINATION/FIX(section of thethe white outlinefix difference will be displayed

whether the fix updates or is The display will remain until either another fix is taken or another destination is selected.


The fix correction only updates the coordinates displayed in the PRESENT POSITION windows and does not realign the platform. The rate of error buildup accrues from the time the DNS system was switched to the NAV mode.

A stored destination may be usedix by selecting the destination number on the destination select panel, moving the DEST/ FDC switch to STORED FDC position, and depressing the STORE pushbutton.

Fix Sequence

No position fix should be taken before atours have elapsed (Including ground operating time) in the NAV mode ofix should be taken as soonas practicable. The optimum time to take the first fix Isours after selecting NAV operating mode. Subsequent fixes should be taken at intervals notours.

Fix Limit

For all fixes except those taken on stored positionshe maximum position fix corrections that will be accepted areautical miles of latitude correction and/ orautical miles of longitude correction. An attempt toosition fix thatthese values will cause the master caution light and the INS FDC REJECT light on the annunciator panel to illuminate. The INS FDC REJECT light will remain onubsequent acceptable fix is taken, orthe DEST/FIX switch is moved to a

fix position. Inability to obtain anfix correction can bo due ta Incorrectly stored fix point coordinates, incorrect fix point Identification, or degraded INS More than one attempt to achieve fix corrections ahould be made beforethat tho INS ll not reliable. STOBED FDCESndo not use thoautical mile update limit for fix reject, butariable limit which Is loaded to the desired value during prefligM preparation. This variable limit capability can be used for INS performance prediction priorpecific mission. For example, if It Is known that INS had to be accurateertain limit topecifica STOBED FDC on citherrust previous to entering the mission area wouldriteria for mission abort. Positionrixes are the same as normal stored fixes except for the variable fix rejection criteria.


Maximum position error will accrue at an average rateautical milos per hour during theours, and at an average rateautical miles per hour thereafter.

STOKE or DEST FDC pushbutton -Press.


This procedure updates the INSautical mile or less from the starting point coordinates.

Acceptance of the positionix Indicates that the INS errorautical miles or less in error in either latitude or longitude and that computed north and east velocities are each lesseet per second. These INS performance criteria are based on an anticipated time duration ofinutes from NAV entry until the positionrix.

If INS FDC BEJECT light comesMS accuracy may be marginal.


Destination selects briefed.

STORE or DEST FDC pushbutton -Press.



Reliability Check

STORED FDCESndrc designed to check INS performance before takeoff to attempt to predict INS accuracy during the flight. The aircraft is accurately positionednown spotTORED FDC is take as follows:

Stop aircraft at designated runway position.

Destination select


Effects ofnverter Foilure on INS

Thenverter supplies power to the INS. nverter failure may have catastrophic results on the INS. The system performance may be degraded after switching to the emergency Inverter. The degradation of system performance will directly depend upon the elapsed time between thenverter failure and switch over to the emergency Inverter. If thenverter falls and the INS outputs are no longer meaningful, the pilot should turn the INS MODE switch on the Nav Panel to OFF. This will lessen the possibility of damage to the system.


The periscope viewing system provides the piloteans lor observing or making visual lixes on terrestial objects which cannot be seen directly from the cockpit. It can also functionky compass, andisplay unit which projects maps and selected data on the presentation screen in the cockpit. Ther periscope windows, viewing optics, and projection equipment are located forward of the cockpit pressure bulkhead.


The basic downward looking function of the periscope systeminified imageround object,ixed lens two-field system. The wide or narrow angle field of view la controlled by thecontrol. The modified wide-angle lens systemoverage of85 forward of nadir and is Intended to be used for observations of largeground objects. INS update fixing is possible when in tho wide angle field of view. The modified narrow angleoverage of approximately 47 forward of nadir and is intended to be used for update fixing of the INS system by fixing on pro-selected ground objects. The forward look distance possible with cither field of view is dnpondont upon the altitude and attitude of the aircraft at the time of the observation. 5 gives thelook rangeunction of the aircraft attitude and altitude. Tha resolution of the optical system in all modes is better than that of the unaided eye. Due to the minifi-catlon imposedixed lens system,the pilot is only expected to Identify prominent ground objects suchoast line, lake or town. Also, due to theslant viewing angle, there will be some apparent distortion when an object appears near the top of the reticle plate; especially when using tho wide angle field of view. It la equippedrl-prlem plastic grooved dlffuaer which provides for two eye viewing of the periscope Image.


The periscope mirrors can be shifted somm strip film projector displays maps or other selected data on thescreen. The pilot may regulate the projector light Intensity and advance or reverse the film as noccssary in order to refer to the desired Information. ilm deBtruct capability Is available and Ismanually or automatically In case of ejection.


A six Inch presentation screen Is installed at the top center of the cockpit instrument panel. The ground area displayed depends on the lens system selected! wide angle or narrow angle. When using the wide angle lens, the circle drawn near the center of the reticle plate Indicates what will be visible with the narrow angle lens. The nadir point of the wide angle view is Indicated by the intersection of the chord drawn across the lower half of the narrow angle view circle and the center vertical line on the reticle plate. The nadir point of the narrow angle view is Lndicatod by th* intersection of the horizontal line and center vertical line on the reticle plate. The nadir point, ason the rotlclo plate, has beento correct for the angle that theaxis has been shifted forward due to the normal level flight attitude of tlie aircraft and the physical placement of the periscope system In the aircraft.

A compass rose is Incorporated In theplate; tho numbers appear around the edge of the plate and are back lighted when the system Is In the sun compass mode.

A pair of vertical dashed lines arcto show the pathround object as It moves from top to bottom on thescreen. Tho dashed lines are tangent to the narrow angle circle and show the pilot what will be visible In the narrow angle view as soon as the object appears in the wide angle view.


A movable cur nor, remotely controlled from the INS navigation control panel, is used with the narrow angle "nadir" line tothe INS for position error. The reticle plate or cursor is not illuminated. Athumb knob In the lower right side of the presentation screen is used to rotate tiie reticle plate.

Periicope Lem Control Hondle

A handle, labeled PERISCOPE and located on the lower left side of the cockpitpanel. Is used to select the desired lens system wide or narrow angle view. The wide angle lens is selected by pulling the handle to the out (aft) position. The narrow angle lens Is selected by placing lhe periscope handle In the in (forward) position.

Mirror Selector Handle

andle, labeled MIRROR SELECT,the mirrors within the basic periscope system. The handle can be positioned to one of three positions as follows;

(forward) position selects film

detented position selects thefor overheadistance to OUT).

(aft) position selects theperiscope.

Projector Light Switch

A rheostat switch, labeled PROJ, controls the film projector light intensity. Thelight is switched OFF at the full counterclockwise position. The projector light is turned on by rotating the rheostat

toward the DRT position. Intensity of the image on the presentation screen isby further clockwise rotation. Power for the light Is furnished by the essential dc bus.

Projector Film Switch

A momentary three-position toggle switch, labeled PROJ. controls movement of map and data film. When the switch is held In the down position tho film will advance. When the switch is held in the up position, the film strip will rewind. The centeris OFF. Power for the switch tsby the essential dc bus.


The sun or sky can be observed with the periscope by shifting the mirrors forviewing. idpoint detent is provided in the mirror selector control forthe periscope optics In this position. nurled thumb knob located at the lower right of the presentation screen Is used to rotate the reticle plate,oggle switch labeled sun compassositions located on the periscope control panelan electric motor which rotates the polarized disk. The sun compass Is usedackup to other heading devices when in locations where other devices might not function accurately. It is used to r emergency heading determinations. It is also used to make periodic cross checks of other headingat any point along the flight path.

The sun compass utilizes the precomputed azimuth (Zn) of sun and the sun's image on the presentation screen. Accuracy Is2 when the sun's elevation ts *6 tot elevation values abovehe heading error becomes greater. It also utilizes the precomputed azimuth of sun and the sky polarization phenomenon. Accuracy is within 2 when the sun's elevation is -8 to


altitude above terrain

angle field of view

angle field of view









th* mmlmateor an object to wear at tlwiep of toa rrfitfe platstwitojr llx una.mi anof*s asiumrt.

FORWAR0 rake in nautical miixs5

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Th* following definitions arc applicable loBun compass procedures;

he red pointer on the electrically driven polarizer disk. It takes the place of the sun when the sun is too low in the sky.

Lubberhe fixed pointer at the top of the presentation screen.

Compasset of numbers around the periphery of the manually rotatedplate.

zimuth ofrue bearing of th* sun relativearticular position on the groundpecific time.

elative Bearing ofh* horizontal angle between the true heading of the aircraft and the true bearing of the sun.

TH- True Heading ofheof aircraft relative to the north pole.



Heading Method

Determine the precompuled value for azimuth of sun (Zn).

Manually rotate compass rose until Zn value is over the sun's Lmag*.

Read True Heading (TH) of aircraft on compass rose, indicated by the lubber line.

Bearing Mctliud

Manually rotate compass rose so the lubber line indicates zero.

Determine the precompuledBearing (RB).

Read the Relative Bearing ofon compass rose at pointby sun's image.

eflection of direct sun image as it appears a: the edge of the reticle plate.

Polarized Skyhe characteristic of the atmosphere to polarize sun light by scattering. The maximumirection 90 from the sun.

Sun Composi Swi'ch

A three- position momentary toggle switch, labeled SUN COMP, on the periscope panel controls the rotation of the sun compaas polarizer disk. The switch positions arend R. Inosition th*disk will rotate in adirection. Inosition thedisk will be rotatedlockwls* Power is provided by th* essential dc bus.

POLARIZED SKY LIGHTtoA. True Heading Method

Determine the precompuledof sun (Zn).

Electrically turn polarizer pointer toward the visible sunlight.

Adjust the central disk as dark as possible when the concentric rings are of equal brightness.

Manually rotate compass rose until Zn value Is In line with thepointer.

}. Read the True Heading (TH) ofon compass rose, as indicate, by Lubber line.

B. Relative Bearing Method

Manually rotate compass rose so that lubber line indicates zero.

Determine the precomputedBearing (RB).

Electrically turn polarizer pointer toward the visible sunlight.

After landing nitrogen pressure will not be available to assist in raising the canopy if the destruct system has been actuated. The canopy may be jettisoned ifis depleted and help is not available.


If pointer is not positioned toward ambiguity will be

Adjust the central disk to be as dark as possible when thering* are of equal brightness.

Read the Relative Bearing ofon compass rose, asby the polarized pointer.


A destruct system Is Incorporated in the airplane to destroy the projector film, and the maps. The film strip in the projector are destroyed by electrically igniting small thermite assemblies which burn inimum ofeconds. The water soluble maps are destroyed by forcing watermall reservoir into the map case using nitrogen gas pressure from the canopy accumulator. The destruct system Is actuated manuallyuarded switch labeled DESTROY located on the right forward panel. The system Iswhen the switch is placed In the up position. oller type micro switch on theejection rails will automaticallythe destruct system if tho pilot ejects. Power for the system is from the essential dc bus.



perating limitations


Introduction Instrument Markings Engine Operating Limit* Moximum Weight Limit! Maximum Altitude Limit Airspeed*

Load Focror Limits Prohibited Maneuvers Center Of Gravity Limits fuel Loading Limitations Aircraft Systems Limitations


This section provides general aircraftand engine operating limitsbe observed in normalspecific limits may be changedtoflight tests

are presently extending operationalmaking continued review of thenecessary. When necessary, to avoid delay in providing current limits to operating personnel, these specific limits will be supplied by the manufacturer's flight test organization at the operating site.


The Instrument markings shown Inre self evident and are not necessarily repeated elsewhere in this section.


Pilot preflight briefing must includeand limitations information pertinent to individual engines Installed. General engine operating limits are summarized In. Thrust rating definitions are provided in Section I.




perating limitations


Introduction Instrument Markings Engine Operating Limits Moximum Weight Limits Maximum Altitude Limit Airspeeds

Load Foe to* Limits Prohibited Maneuvers Center Of Gravity Limits Fuel Looding Limitations Aircraft Systems Limitations

This section provides general aircraftand engine operating limitsbe observed In normalspecific limits may be changedtoflight tests

are presently extending operationalmaking continued review of thenecessary. When necessary, to avoid delay in providing current limits to operating personnel, these specific limits will be supplied by the manufacturer's flight test organization at the operating site.


The Instrument markings shown inre self evident and are not necessarily repeated elsewhere In this section.


Pilot prefllght briefing must includeand Limitations information pertinent to individual engines Installed. General engine operating limits are summarized in. Thrust rating definitions are provided in Section I.








-limil CGI for









r Ma-hum

Ir 2)







ih;iyhies may bp operated continuous-ly at all ratings when within the normal exhaust gas temperatureowvvcr, no mure Elian one hour may be accumulatedn excesb of the normal liii.itndmidi bf reduced immediatelyn emergency limit lom-oeriture exceeded. (Sec Ijf'T Limits and


ccumulated operating time In thene for more ttianinutes may ro'iuire engine removal.


The nominal operating band, normal Limits and emergency exhaust gas temperature operating HihoAiles arc prescribedunction of compressor Inlet temperature ani: in. T's forhen compressor inlet temperature ia above,heti CIT is bclmv. The netting at which the red warning light on Ihe EOT sagend He fuel dcrichment system operates, if armed,l jiach is above the normal operating temperature limit schedule.


At compressor inlet temperature? belowthe possibility of engine hi.ill exists at EClT's between the maxim-.im permissible valac and the nominal operating banc-.

In the event tbat emergency engine operation is required, EGT may-V'C when aboveGIT, or towhen belowCIT; however, an accurate accounting of operating time in tie emergency op-crating zone must be maintained.

N ij u-

. Any operation in or above the emergency operating zone requires special maintenance action.

. The permissible emergency EflT level at low CIT's is .lbovc the derich system actuation point; therefore, the derich system must be disarmed if this level is to be atl.une-!.

DC No. 8


Angines may belud continuously at all ratings when within iho normal exhaust gas temperature limits; however, no more than one hour may bewith EGT in excess oi thr norma) limit schedule, and EOT must be reduced immediately if an emergency limitis exceeded. fSeesniLa and

1 1

Continuous or accumulatedtime in the emergency EGT operating aone for more thaninutes may require engine


The nominal operating band, normal limits and emergency exhaust gas temperature operating schedules arc prescribedundi in of compressor inlet temperature as shown in. Limit EGT's foroperation areheninlet temperature is above.hen CIT is btiow. The setting at which the red warning light on the EGT gage illuminates and the fuel dcrichment system operates, if armed,alue which is above ths normal operating temperature limit schedule.


. Any operation in or above the emergency piratin; xoiespecial maintenance action.

. The pormissible emergency EGT lavsl at low CIT's Is above the derich system|nnt; therefore, the derich system must beif this level is to be attained.


The maximum allowable compressor inlet temperature. In addition,must be monitored so that engine cooling rates will not be excessive. While above an airspeed of, the aircraft maximum rate of descent should be such lhat rate of deceleration does notach in throe minutes. There ia noon rata of deceleration while below.


The minimum pressure recommended for airstarts from stabilized windmllling speedsel. This pressure is markedreen radial line.

At compressor inlet temperatures, the possibility ofstall exists at EGT's between the maximum psrmissible value and the nominal operating band.

In the event that emergency engine operation ia required, EGT may be increased towhen iiboveClT, orhen ba-InuIT: however, an accurateoi operating time in the emergencyzone must be maintained.

engine speed

Military and afterburning engine speeds arc the same and are automatically scheduled by the fuel controlunction of Compressor Inlet Temperature. The normal schedule is shown by. Engine overspeed0 rpmisual inspection of the turbine. Notify the0 rpm Is ever exceeded. Each instance of overspeeding should be reported as an engine discrepancy and should include the maximum rpm attained.



j4 -JM



approved fuel isTC. pproved source of lubricity additive,, must be mixed with the fuel in the ratio9 gallonsO gallons of fuel. Fuels such asnday be used only for emergency requirements such as air refueling when standard fuel Is not available and air refueling must beor risk loss of the aircraft. with emergency fuels should beto speeds below.


The approved oil isB. Ifbecause of low ambient temperatures, it may be diluted with Trlchloroethylene, Federal. Typen accordance with Maintenance Manual

Oil Prenure

Oil pressures belowsi are unsafe and requireanding be made aa soon as possible, using minimum thrust required to sustain flightanding can be Normal oil pressure is fromosi. Except at idle throttleoil pressures betweensi andsi are undesirable and should be reported after flight. radually Increasing oil pressure up tosi is acceptable at high Mach numbers provided the Indicationto normal values after aircraftto subsonic speed.

Oil Tempe'oture

Oil temperature must be at) prior to starting unless previously diluted with Trlchloroethylene. Engine oil temperaturesre unsafeanding should be made as soon as passible if the temperature cannot be maintained bslow tbla value. An engine should not be restarted after windmilling at subsonic speed when CIT is less than) for moreinutes. Ifoparation above IDLE with Oil. TEMP warning light illuminated shall be as brief as possible.

Maximum gross weight is not limitedby takeoff performance capabilities. Base maximum takeoff weights onprovided in Part II of the Appendix.


Maximum altitude with derichment installed and operational0 feet; maximum altitude without derichment0 feet.


(Refer toor the limit flight speed and altitude envelope.)


The stall warning light on the annunciator panel and the master caution light illuminate when angle of attack reaches 14 Inone is also produced in tha pilot's head-aet. WhenIAS, the speed at which stall warning occurs is ths minimum airspeed restriction for ths existing vehicle weight, c.nd load factor unlessis governedigher valua ofKEAS as displayed by ths TripleIndicator. Minimum airspeedEAS0 feet.


The Mach-airspeed indicator limit hand ia set to indicate airspeedEAS. However,EAS limit applies only at altitudes0 feet, and at airspeed below,0 feet, limit airspeed decreases linearly with altitudeEAS0 feetEAS at sea level. Above, limit airspeed decreases linearlyEAS atFAS at. Seeor variation of KIAS with altitude for KEAS.


Maximum recommended operating speeds are at leastEAS less than limit airspseds. EAS) is not recommended0 feet.





A red radial lineIAS represents the minimum subsonic speed restriction0 feet when the stall warning light Is off.

To avoidafe angle of attack positive g's are limited'sach. (This is equivalent to5 bank level turn.)



triple display Indicator Is not marked however, the limit equivalent speeds are as follows unless:

Mach-alrspeed Instrumentequals either the limitIndication or the minimumrestriction.

stall warning light illuminatesstall warning tone Is heard.

The aircraft shall be operatedanner to avoid full stalls, spins, and inverted flight. Normal bank angle when operating abovesegrees.


Bates of descent must be limited so as to maintain positive fuel tank pressure when sustained cruise speeds have exceeded.

TDI Airs

limit airspeedEAS at sea level, increasing linearly with altitudeEAS0 feet pressure altitude;EAS0 feet and the altitude for. Limit airspeed then decreases linearly with Mach numberEAS at. Normal operation cruise speedach.

Minimum TDI Alrtpeed

The minimum airspeed restriction varies linearly with Mach numberEAS0 feetEAS0 feet, and isEAS0 feet


The aircraft shall be operated. g. rangeAC while subsonic.. must be forwardAC for takeoff and ehould be as nearAC as possible with existing fuel for Landing.

The aft c. g. limitAC while This limit results from stability considerations at high Mach number. stability exists at farther aft centers of gravity betweenndut for simplicity the aft limit la not changed. The purpose of elevon trim limits imposed In this Mach region la to alert the pilotajor malfunction in the fuel


The maximum allowable positive load factor's in symmetrical maneuvers's in roll maneuvers asby. The maximumload factor0 whenEAS varying0's at higheras shown by.

On those aircraft, if an. emergency exists and EMERtransfer is operated to place more0 lbs innd total fuel Is less0 lbs, the aircraft should be limited tc maneuvers causing not more.

As elevon trim can be used as an indication of abnormal c. g. condition, ths following pitch trim limits apply:

Whileo more thannose do*-.


While nose down fromo nose down above.

At Initial cruise the trim limit isnose down. As altitude increases and KEAS decreases,. trim limit becomes approximatelymore nose up perEAS decreaseEAS.

(In addition, expect approximatelymore nui" up trim (or each percent. is forwardAC).


These limits to be supplied at the operating site.



The SAS shall be on for all takcoffs and



The spike and forward bypass controls must he operated in the AUTO mode at all times when0 feet. When inlet controls must be operated manuallyallowable speed is.


The canopy shall be opened or closed only when the aircraft is completely stopped. Maximum taxi speed with the canopy open isnots. Gusts or strong winds should be consideredortion of thenot speed limit.


do notEAS oraximum of b sideslip with gear When sideslip angle exceedsoperation with gear extendedlimited LoEAS. Operation al super-some speed with gear extended is prohibited. The landing grur in designed for landing sink speeds at touchdown which decreasePS0 poundsPSounds. Side loads during takeoff, landing, and taxiing must be kept to aas landing gear side load Strength is critical during ground maneuvering.


The maximum taxi speed recommended isnots forires. The rated ground speed limitnots. 0 feetnots correspondsIAS withambient temperaturealm day,IAS atambient temperature. Limit indicated airspeed on the groundby the amount of tailwindalong the runway and increases by th; headwind component. Refer toor ratcd'speeds at other altitudes and temperatures.

Taxiing Restrictions

A heat check is required for tires, wheels, and brake8:

a. Prior to takeoff when taxiing has exceeded one statute mile.

When continuous taxi distance hastatule miles.

When clear of th> runway after an aborted tak-inffeavy weight landing.


initial braking speed for stop using raied brake capacity







Section VI



Configuration Effect!

Stability Charocterlitlei

High Angle Of Attack Conditions


Control Effectiveness Single Engine

Normal Operating Character it tics Takeoff Climb Crate

High Altitude Turn Technique


Ah Refueling





and Landing

This aircraft operates within anlarge Mach number and altitude envelope but the equivalent airspeed, angle of attack, and load factor envelopo Is relatively Typical takeoff and landing speedsnots respectively, and the Cruise spend Is0 knots atuch number. Sustained cruise altitudes at high Mach number range0 feet to0 feet.

The aircraft is designed to obtain maximum cruise performanceach number. The external configuration, air inlet system power-plant and fuel system sequencing are optimised for this flight condition. hree axes stability augmentation system is an integral part of the aircraft control system design and is normally used for all flight conditions. The normal flight characteristic discussed in this section assume proper SAS operation, unless stated otherwise, andof tha limits specified in Section V.

Changed is8


External configuration feature! which affecc flight characteristics include the delta wing, fuselage chines and the engine nacelle

The normal delta wing characteristics are present In these aircraft. There Is no stall at normal operating speeds and flightinstead therearge increase in drag as the limit angle of attack Is This characteristicelta wing can cause very high rates of sink to develop if the aircraft is flown at toopeed. The stall warning light Is to limit the angle of attackafe value so that stall la not encountered.

The dihedral effect Is positive, butat the higher Mach numbers. Poll damptng Is relatively low over tho entirerange of these aircraft and the lateral-directional qualities are relatively poor with SAS off.

The chines extand from the fuselage nose to the wing leading edge. At subsonic speeds they have the beneficial effect of Increasing directional stability with increasing angle of attack. At supersonic speeds, theyLift andeed for canard surfaces or special noae up trimming The automatic fuel tank sequencing shifts tho center of gravity aft duringto correspond with the aft shift of center of lift with increasing Mach. Then It maintains c. g.elatively constant optimum location during cruise. Thisof the center of gravity don to the center of lift decreases pitch trimand minimizes the thrust and fuel flow required for cruise. This also reduces the static longitudinal stability margin, but the SAS compensates for the reduction and provides satisfactory handling qualities.

The mid span location of the enginesdrag and Interference effects of the fuselage. The Inboard cant and droop of the nacelles gives maximum pressure recovery at the engine Inlets at the angles of attack normal for high altitude super tonic cruise. However, the location results In sensitivity of the aircraft to asymmetric thrust During afterburner cruise, throttle and EGT trim adjustment to equalise fuel flows minimises thrust differences. Engine EGT and fuel flow values should be matched by throttle adjustments during subsonic cruise. Indicated flows duringoperation may include heat sink aystom requirements after hot flight with low fuel remaining, so that flowmeter values may not be representative of engineand thrust.


The augmented dynamic stability is positive and flight tests have demonstrated that the dynamic damping characteristics aredeadbeat. No unusual staticcharacteristics have been disclosed when operating within the c. g. and angle of attack limits. Positive static stabilityto exist. is somewhat aft of the limit while at intermediate supersonic speeds (fromo at leasthe. limit Is violated while near the design cruise Machtatic Instability In pitch may be If pitch rates are then generated and not arreated within the angle of attackitch up can develop and result In structural failure of the aircraft* I

The aircraft is controllable withoutaugmeautlon to. Without SAS it ia also controllable during climb and descent, during Inlet unstart up toEAS and during unstart and engine flame out up tond during twenty degree bank turns in heavy turbulence at

-# i



low supersonic and transonic Mach numbers. However, control with SAS off Is sensitive and control movements should be kept to the necessary minimum. Thrust asymmetry should be minimized, particularly at the higher Mach numbers. Sustained cruise or maneuvering without pitch and yaw axisaugmentation is not recommended near design speed.

At cruise Mach number, the pitch stability is only slightly positive and disturbances are only lightly damped. Sudden loss of all pitch SAS while In maneuvering flight willitch transient which willincrease the load factor for the same stick position.

Without SAS the yaw stability may vary from positive to very slowly divergent. Response of the automatic air inlet system toronounced effect onmotion of the aircraft. Unlessby the pilot, phasing of the spikes and forward bypass can tend to either drive or damp the yaw oscillations.

Emergency operating procedures for use in the event of SAS failure are given In Section UI.

high angle of attack conditions

Minimum airspeed restrictions and aangle of attack warning light areto prevent approach to pitch-upand to maintain adequate ground clearance at takeoff and landing consistent with performance objectives. There is no stall In the classic sense where an abrupt lose In lift would occurritical angle of attack. (See, Lift vs Angle ofose up pitching moment develops instead, as angle of attackwhich becomes uncontrollable with lull nose down elevon as the critical angle of attack boundary is reached. (Sec Figure

ubsonic Critical Angle ai Attackn uncontrollable pitch-up will not occur until after limit angle of attack as given ins reached. The SAS will tend to maintain apparent stability about all three axes until loss of control occurs, then the aircraft will pitch-up with little or no warning. Note that there Is an airspeed margin of fromoIAS when subsonic and at the aft c. g. limitAC. The margin Is less at supersonic speeds and varies with Mach number, . 's aft of normal limits will materially reduce the margin. When near limit angle ofilot induced rapid nose up pitch rate may require more margin for recovery than is available.


An uncontrollable pitch-up maneuver will result when the critical angle of attack boundary is reached. from this condition Isunlikely. Attemptedmust not be continued to the point where insufficient altitude for recovery or ejection exists.

Pitch rates which accompany Increasing angles of attack must be checked and load factor relievedufficient rate toairspeed when the critical angle of attack boundary is approached. Whenand terrain clearance permits,should be increasedIAS before resuming level flight. Care must be exercised to Insure that recovery load factors will neither cause Limit angle of attack to recur or Impose load factorsallowable values. When supersonic and near limit Mach number, it may be necessary to reduce power or increase drag (or both) while recovering ao that limit Mach number will not be exceeded while airspeed is Increasing.


Intentional spina arc prohibited. Thetechnique ia auggeated in the event of an inadvertent spin; however, ejection may be the beat course of action afterexisting altitude, airspeed, spin rate, attitude, and fuel loading conditions, as spin recovery has not be demonstrated and is considered extremely unlikely. At the pilot's discretion:

Center the controls, disengage surface llmlters, and determine the direction of rotation from the turn indicator.

Apply forward stick and full roll control into the direction of spin as the nose drops.

Apply opposite rudder to stop rotation.

Center the rudder and roll control as rotation stops.

Start pull-outIAS.

If possible, avoidIAS and limit load factor during recovery.


With uncontrollable conditions, eject at0 feet above the terrain whenever possible.


Generally control effectiveness is good. At high altitude and angle of attack roll control effectiveness Is reduced. This Isroblem if an unstart occurs in the down Lnleturn. Refer to Inlet Duct Unstarts, Section Ul.


The yawing moment resulting fromthrust Is large if an engine falls just after takeoffingle engine go around Is necessary. o full rudder deflection andegrees or more bank Into the good engine will be necessary to maintain control Immediately after loss of power. Drag can then be minimised by reducing pedal force and trimming to 7 to 9 rudder positionsimultaneously using bank andtoward the operating engine asto maintain the desired flight path. The SAS automatically responds withcontrol at the time of engine failure or go around power application and Itsrate is faster than pilot reaction time. However, rudder control follow up by the pilot Is necessary ae the yaw SAS authority Is limitedegrees rudder The SAS continues to apply rudder deflection as longideslip Is maintained, but this deflection is not indicated by pedal position or by the rudder trim indicator.

The amount of rudder deflection required during single engine operation decreases as airspeed Is increased. During single onglna cruise5 Mach number, the aircraft can maintain course withlimlters engaged. Optimum rudder deflections are maintained by the SASusing rudder trim when bank andtoward the operating engine are used to maintain course. The bank anglesapproach

Above, engine failure, flameout or inlet unstart may require yaw axisaugmentation to avoid excessiveand bank angles which could cause the operative engine to stall or flameout. Inlet unstarts whileEAS and maximum power are quite severe. In these cases, unassisted pilot reaction is too slow toall the control Immediately required. Pilot follow up is necessary after the initial SAS corrections.


Before retarding the throttle to shutdown an engine, care must be exercised to properly Identify the side on which the malfunction There have been cases where the operative engine was improperly Identified as the aourc< of the problem.


Refer toor specificInformation.


The aircraft accelerates rapidly to rotation speed once maximum thrust Is set during takeoff. The nosewheel can be liftedonots below takeoff speed, but this la not advised because the drag that is createdthe acceleration and extends tht takeoff run. With aero degrees pitchtick force of approximatelyounds Is normally required to lift the nosewhoel at rotation speed. Stick force must be relaxed during the rotation In order to check the nose up pitch rate. During maximumtake off speed and attitude must be monitored carefully to avoid over-rotating and dragging the tail.


Normal climbs to supersonic cruise speeds involve three phases of operation. These consistubsonicransonicto the supersonic climb schedule,upersonic climbing acceleration.

There are no unusual characteristics during the subsonic phase exceptightbuffet may be feltach number as airflow conditions near the tertiary doors and ejector flap areas change

A Mach jump on the TDI instrument will be observed83 Machduring transition to the supersonic climb speed schedule. No uniqueoccur In this area; however, there Is an area of decreased excess thrust from5 toive technique is uaad to Improve acceleration through this speed range. The transition should be made without other maneuvering if possible, aa even shallow turns Increase drag sufficiently to decrease acceleration and Increase fuel consumptionoticeable Increase in acceleration can be expected after passing The pull up to establish climb attitude should benots before the supersonic climb speed schedule Is attained. This will reduce the possibility of overshooting the desired speed.

The supersonic climb Is initiated whon climb airspeed Is establishedeet. It is essential totne schedule accurately to achieve beet climb performance. Speeds which are higher than normal should be avoidedlimit airspeed can behort period of time.

The aircraft does not respond Immediately to small pitch commands. Thismakes precise airspeed control difficult until experience Is gained In the aircraft. If significant ovorspeed occurs, the recommended action Is to reduce power until climb speed can be reestablished rathor than pull up sharply and Impose load factors.




A continual variation In trim la required during tlie acceleration to cruise apeed, withEAS schedule requiring more nose down trim thanEAS schedule. The variation of trim at the aft limit isby. This figure alao shows the variation of trim required with airspeed when operating near the. limits.

Occasional periods of Inlet roughness may be encountered In the area between. It may also be encountered at climb speeds In the region above; however, the roughness diminishes as cruising altitudes are reached and theairspeed is reduced from the climb airspeed schedule.

The transition to cruise altitude and speed is accomplished with power being reduced slowly as the initial cmiae altitude la

The following definitions have been adopted in order to categorize supersonic cruise operation.

rangehis type ofmaximum range for thespecified. Powerare in the lower portion ofrange (near the 82 mark).

altitude cruisehisoperation yields altitudeare above the maximumbalow the maximum cellingspecific range which resultsthan for maximum range,efficient cruisingmaintained.

ceilinghisoperation requires continuousat near maximumsotting for the Mach

These types of operationsruise climb thatradual but continuous Increase In altitude as fuel is consumed. The flight parameters are: Mach number, equivalent airspeed (KEAS) and altitude. These three variables are dependent upon one another. Gross weight,. also have primary effects on performance capability.

Mach, KEAS, Alti'ude Relationship

The selection of the values for any two of the Mach, KEAS, or altitude variablesdefines the value of the third. For Instance, if cruise is scheduled fornd the desired cruise altitude is0 feet, the KEAS mustnots.

Effect Of Changing Air Temperate

Ambient air temperature may appear to change abruptly as different air maeaea arc encountered because of the high trueat cruise. Initially, If constantis maintained, flightarmer air mass willecrease In Mach number and KEAS, and the true airspeed (TAS) and compressor inlet temperature (CIT) wiU remain constanthortigher TAS and CIT will result aa the desired Mach number is re-established. The opposite would occuresultolder air mass. New cruise altitudes are usually required to compensate for effects of variations in ambient air

Effect Of Mach Number

Another characteristic of supersonic crulsa is that any given gross weight and CIT, the altitudes for maximum range or maximum celling profiles increase with Mach number.ule of thumb, this increaseeet5 Mach number. elated characteristic is that if the Mach number la allowed to increase slightly above that desired, and if the throttle ia notthe aircraft has an increasingof excess thrust. It is easy to exceed target Mach number inadvertently.


Maximum Rongs (Optimum) Cruiie Profile

At high Mach numbers, the maximum range (optimum) profilea continuous cruise climb with the throttles in the afterburner range near thePLA throttle mark. When at heavy weight, It may be necessary to initiate this type of profile by flyingonstant altitudehort period, slightly higher than the altitude for best specific range, in order to maintain KEAS at ormaximum operating limits. In this case, the initial cruise altitude schedule remains above the optimum until gross weight Is reduced sufficiently toruise climb. Cruise climb should not be continued0 feet (because of present operating restrictions).


Constant altitude turns withbank angle can normally be made by increasing thrust.ule of thumb, fuel flow and angle of attack Increase in proportion to load factor. It ia more economical to allow altitude to decrease while turning, maintainingpower setting and Mach number during the turn, and regaining the altitude lost upon rolling out. KEAS should not beto increase above the maximumlimits during descending turns,

Moximum Ceiling Profile

The maximum calling profile00 feet above the altitude schedule for maximum range. Sec. Stabilisingruise climb profile wherein the aircraft ia constantly flying at its absolute ceiling at maximum afterburner is very The only control the pilot has to maintain constant Mach number is to climb or descend. Therefore the maximumaltitude can ba obtained by usingless than full power.

High Altitude Quite Profle

High altitude cruise profiles schedule ths cruise climb altitude below the maximum afterburning celling. Continuous use of maximum afterburner is not required.

Effect Of Mach Decrease

The Mach number must not be allowed to decroasc more5 Mach number bo-lo* the desired cruise speed. mallIn Mach number and KEAS ataltitude may cause the aircraft to Intercept the celling for that spend andthrust limited. escent of several thousand feet may be required tothe desired cruise Mach number.


Refer toummary of maximum range and cellingfor various Mach numbers, weights and ambient temperatures.

High Altitude Turn Technique

Tumi Leu Than

The techniques described below minimise altitude variations while turning andaltitude losses which can be encountered If turns are Initiated near the maximum afterburner celling for the existing Mach number, gross weight and ambient


a the target speed recommended for turning when maximum altitude Is the primary consideration.


e. When turn entry la scheduledigh altitude cruise profile when at light weight and using partialfor power settings:

Prior to turn, cruise at power required to maintain target Mach number at desired altitude.s recommended when minimum altitude loss is th; primary consideration,

Turn Mach Hold OFF and leave the autopilot Attitude Hold mode engaged.

Enter the turn with throttles set below maximum afterburning. While turning, adjust the throttles as required to control Mach number and adjust the autopilot pitch trim wheel to control Cruise altitude can b= maintained in most cases.


Do not make abrupt pitch attitude changes.

Afterof turn, allow altitude to0 feet if sufficient excess thrust is available.

After completing the turn, engage Mach Hold to maintain the desired cruise climb schedule, and use power as required.


Descent characteristics are not unusualfor the variation In flight path angle encountered during the supersonic Normal deceleration techniques include maintaining an optimum KEAS schedule to obtain maximum range andexceedingollig limitations. When cruise KEAS Is higher than, altitude should be maintained after power reduction until KEAS decreases. When

cruise KEAS Is lower than optimum,should be started immediately after power reduction, maintaining cruise Mach number until desired KEAS is intercepted. The angle of descent varies from1 initially to approximately 7 ass reached.

Air Refueling

Air refueling of these aircraft with the flying boom system of theankers poses no problem of compatability and is normally accomplished00 feet. The aircraft provides anstable platform with the SAS on. The only characteristic that causes some problem is that, without afterburning, the aircraft may become power limited at the higher refueling altitudes before aonload can be completed. Thisusingoboggan techniqueechnique of completing the refueling with one afterburner on.

Forward visibility In the observation and precontact positions is excellent, butdownward, and aft visibility is Rendezvous is easiestlightly low position with the tanker within 60 either side of the nose. The pilot'svisibility Is optimized by lowering his seat prior to contact. Depth perception through the vee windshield is slightlyand some pilots may prefer to use one side of the windshield during contact.

A slight buffet will be felt as the contact position la reached. This ia tanker down-wash and has no effect on thelight decelerating effect. response of the engines isand aircraft drag at refueling speeds produces correspondingly goodresponse.

Overcontrol of the engines should be avoided while gaining and holding position due to non-linearity of throttle position vs engineiven throttle angle change near military power yields more thrust changeimilar change in the throttle mid range. The aircraft may become power limited if the afterburner-on technique is not used, and tobogganing descents of up0 feet


per minute may be requiredhe military power throttle position is approached. Asymmetric thrust Is easily controlled when the afterburner-on technique is used.

Light turbulence encountered while inposes no particular problem with SAS operating normally, and shallow turns of up to 20 bank angle can be made without However, If all pitch SAS including the back-up pitch damper are Inoperative, It is recommended that refueling not beexcopt in an emergency. Thetends to be unstable without any pitch SAS, but control can be maintained under favorable conditions with fuel transferred to. location.

All disconnects should be made with aand slightly downward relative motion with wings level. This will Insureof the boom from the receptacletraight line force. Side or rolling loads or excessive deviations from the desired elevation increase tha possibility of boom and/or receptacle damage during disconnect.

Nigbt refueling is essentially the same as for daytime operations except that added caution and effort la required to avoidand the tendency toward throttle over-control while in contact Is increased.

Normally the aircraft is flown directly to touchdown rather than attempting lo float just off tho runway with subsequentat too high an attitude. Prompt chutewill result in momentaryloads of up to one g. The chute should not bo deployed in the air because of the rapid deceleration and rate of sink that could develop, but it can be actuated before nosewheel contact without any unusual pitching tendencies.

Practice landings with SAS off are not Approach control during emergency landings with all pitch SAS off Is increasingly more difficultr exceeds the aft limit.

Approach and Lending

Handling characteristics during approach and landing with SAS operative are good. Short period disturbances are well damped, and rates of roll available for maneuvering are adequate. The aircraft can be held off the runway to speeds that are much lower than are recommended for landing. The touchdown attitude normally Is from 10 to 12 angle of attack. Thereisk of damage to the aft fuselage if the touchdown attitude exceeds







Normal Trimwilh Surge Semitive Engines

Inlet System

Spike and Bypass


Fuel System

Normal Fuel Tank

Jet Pump


Fuel Tank Empty Light

In-FlightManagement Prior to

Management During

Subsonic Cruise Fuel

Flight Control

Brake SystemI


On many engines, engine trim adjustment can be accomplished onlyest stand or while the aircraft is on the ground. One of the features affecting operation of these aircraft is an engine trim device which is operated from the cockpit. The trimmer is normally used to maintain EGT within aoperating range in flight when at or near Military thrust or when theis operating. It modifies the turbine inlet temperature vs compressor inletscheduling characteristics of the main fuel control. Changes in engine trim are indicated by the EGT gage. Trimming has little direct effect on afterburnerbut the trimmer is the only main engine control available to the pilothrottle is act in the afterburner range.


normal trim operation

heck chocked.

following describes the normal use of trim capability at the present time.

ILITARY momentarily then IDLE.

to Tokooff

A trim run Is usually made on the end of the runway prior to takeoff using tho following procedure;

This serves to unload the trimmer and reduce hysteresis.

3. ILITARY until EGT

EGT trims required per.

dle rapidly.

This throttle chop serves to check the proper sequence of bleed operation by the absence of compressor stall.

Climb and Cruise

Trim as necessary after takeoff and while accelerating toIT to maintain EGT less than EGT must be maintained belowwhen aboveCIT. It is recommended that the EGT be maintained betweenandduring cruise.

Subsonic Operation

The engine fuel control is scheduled toturbine temperature rather rapidly as compressor inlet temperature fallso preclude the possibility of engine stall. The EGT trim may be used in this operating regime to up trim the engine if required. It may be possible to increase tho EGT more thanut in most cases the increase will be less since the uptrim rangeunction of the original trim setting.


Uptrtmming in the low temperature area can cause over temperature during subsequent aircraftorIT unless the EGT trim Is reset to nominal schedule prior to acceleration.

0 feet. iven level of thrust, higher throttle settings andfuel flow are required as EGT is decreased. Full throttle ceilings in cruise and while turning are reduced; this occurs because combined burning efficiency of the engine and AB decreases with lowered EGT. The degradation in thrust for all throttle settings, at0 feet, isercent perf EGT decrease. Although only one flight condition is illustrated, the trend is the same for other flight conditions.

Effect of RPM Suppression on

As EGT decreases, the engine nozzle opens to maintain scheduled rpm. At high Mach number and maximum power, low EGT may cause the nozzle to open fully and any further EGT decrease will result in rpmbelow schedule. When this condition occurs the engine speed willpm for eachf EGT decrease. The airflow through the engine decreases due to tho suppressedigher Inlet duct bypassand opening of the forward bypass doors. Athis resultshrustnd drag increaseercent perf EGTfor each affected engine. If Mach number decreasesesult of the change in thrust and drag, the spikes schedule more forward and the forward bypass doors open further. Performance willrapidly under these cumulative effects and it is recommended that cruise EGT be maintainedndto avoid the possibility of this situation

oi Engine Thrus* Vcviolior> with EGT

llustrates the typical variation of engine thrust with EGT at scheduled rpm.

After Air Refueling.

Therew can advise the pilot of the proper engine trim to be set when the air-


thrust variation with egt




At *PM



partO'CotECt nj threat* wttlog



craft departs the tanker. engine exhaust gas temperature vsndicated free air temperature.


If an EGT increase betweenand maximum afterburning power exists the pilot should also be briefed on the amount of EGT increase.


Some engines have surged during ground operation while at turbine discharge(EGT's) specified for takeoff by the respective EGT trim curve*. Thisto surge, when it appears, is greater at lower ambient temperatures. Surgeengines have not exhibited thisin flight. Most of these engines can be identified prior to flight; however, engines with no previous surge history have developed into surgers after exposure to descents from high Mach numbers.

It an engine surges during pre takcoil trim, down trim to eliminate surge but do not trim lower thanelow the desired trim point for the ambient temperature.


. Surginground run problem only.

. Engine thrust is reducedounds at sea level static for eachoi EGT down trim from the normal trim curve. After takeoff engines down trimmed for surge protection should be up trimmedGT when CIT reaches



When an inlet unstarts aboven.

During normal scheduling a*n, or with variations in angle of attack or yaw angle.

When descending past0 feet.

Spike is moved forward If the spike knob Is not in AUTO:

If FWD position la selected.

osition is selected.

Spike Is moved forward under the following conditions:

The restart switch is actuated to ON.

If L-ydraulic pressure to the spike actuator fails belownd al higher Mach number unless the inlet is unstarted.

(On the left side) thenverter falls, the L.P ICS circuit breaker opens, or if the L. SPIKE t, DR LVUT circuit breaker opens.

(On the right side) the number 3fails,R ICS circuit broaker opens, orR LVDT circuit breaker opens.

The emergency spike forward switch la actuated to the forward positionconfirmed loss of hydraulic

Forwsrd bypass opens or moves toward open when the control knob is in AUTO under the following conditions:


an inlet unstartsn.

Per the automatic schedule.

During periods of rapid RPM decrease.

When manual spike is selected.

The forward bypass opens or moves toward open when the forward bypass control is not in AUTO under the following conditions:

ower number position or OPEN is selected.

unction of the manual spike knob position selected, up to approximately one inch of door opening variation.

Restart switch Is actuated to ON or forward bypass open.

ydraulic pressure to the door actuator fails.

(On the left side) the number 2falls, the Li SPIKE It DR ICS circuit breaker opens.

(On the right side) the number 3fails,R ICS circuit breaker opens, orR LVD: circuit breaker opens.

7. main landing gear doors are open.

The forward bypass is closed when the speed is below, except when the main landing gear doors are open.


The position of the third striped pointer on the CIP gage Is controlled by an output signal from the air data computer from pressures sensed by the pitot static system. The Indication Is scheduled in accordance with automatically computed values of Mach number and KEAS so that the striped pointernormal"sia) (see) for the flight condition ifEAS and. Values indicated while at lower speed conditions are notto be representative of normal inlet pressures. Above the minimum speedubstantial difference between the "normal" and actual CIP pointer indicates improper inlet operation. Higher actual pressures than "normal" indicate possible unstart conditions. Lower than normal actual pressures indicates poor pressure recovery due to improper spike and/orsettings except when at abnormal angles of attack or in yaw conditions where inlet operation is automatically biased to produce less than normal recovery, The normal spread between CIPointers) should notel. The difference betweenointer and the striped pointer should notsi. The striped pointer may be useduide for bypass door settings duringoperation of one or both inlets and It is preferable to keepointer slightly below the "normal" indication toargin below unstart pressures. Continued automatic or manual Inlet oper-cation at pressures substantially below the "normal" indication can result in loss of aircraft range.


As the total tolerance of the striped pointer can be as much asLl psi at maximum Mach number, it is possibleroperly operating inlet to be above the "normal"



The normal sequence of fuel usage with the aircraft fully fueled is completely Afteruel is used, thissequencing maintains c. g, in the rangeAC forhigh speed cruise at altitude. Starting, taxi and takeoff are normally accomplished with the pumps in tanks,eeding the engines. ormally emptiesclimbout or shortly after supersonic speeds are reached) thenontinues feeding the left engine andhe right engine. Aspproachesloat switch starts the pumps in tank 3. Whenecomesecond float turns off theumps. Any residual fuel in the tank is pumped intoy the Jet pump system, The float switch onurns on the pump innd the second (empty) float switch turns off theumps. Aspproaches empty the two pumps ineeding the leftare started and aspproaches empty the other two pumps ineeding the right manifold are started. Aft transfer of fuel too control c. g. isautomatically through the left Aft transfer occurs at any time whenumps are on, space is available,uel Is aboveounds level as shown onnd both throttles are in the afterburner position. When aft transfer is in operation,s feeding ands nearly full the maxi-


















mum aft transfer rate isbs per minute. If Space is available inuch asround takeoffon-afterburning period, the transfer rate isb per minute. These rates are based on the9 size or

preventsrom operating whens running. The operation of thellage system Is automatic and the cockpit fuel pump lights will not indicate when tankrc operating. However, the fuel quantity indicator mayrop inuel.


A system of six jet pumps is installed. These transfer residual fuel from tanks which are almost empty to tanks whose boost pumps are oporating. Jet pumps scavenge tanks I,nd 6. Fuel froms scavenged by the jet pump in tank 2. Boostperates the jet pump which scavenge tank 2. Boostperates the jet pump to scavenge tank 3. Boostperates the jet pump to scavengend tanket pumps to scavenge tanket pump usage is entirely automatic and requires no attention by the pilot. ube from the boost pump outlet to the jet pumpellows whichalvea suction tube from the previous sequenced tank to the low pressure sectionenturl. oost pump is not running, the bellows In ita associated jet pump contracts and closes the suction tube between the tanks.


Through action of bypass and relief valves, excess cooling loop and engine fuelsystem fuel not burned by the engine, or hot fuel not accepted by the fuel control, (smart valve] may be returned to tankfs close to beingual float switch activate"o furnishuel to the fuel manifold and create an ullage space in tank 4. This same float switch also turns off. Whens runningan nol run. econd back up float switch controlsnd


The float switches are wired in series so that If one switch should stick closedill not keep running and soremature depletion ofuel.


When tanksre empty the yellow EMPTY lights Illuminate in the pushbutton switches. The empty lights for each tank will also Illuminate for tankf normal sequence has been used. However, tanks.o not always show an empty Light if fuel has been used out of the normal sequence.

ust be nearly empty beforempty light will illuminate. ust be nearly.empty before theight will illuminate, andust be nearly empty before thempty light will illuminate.ow warning light whichindependently of the tank empty light sequencing at0 Ib level. Thetank quantities can also be checked if out of sequence fuel usage Is suspected.


In flight, all tanks can refuel simultaneously at invidivudal rates which varypm (tankopm (afthe initial transfer rate if all tanks have space to accept fuel is0 ppm with normal tanker nozzle pressure. The rate decreases aa individual tanks are

filled,0pm as the last tanks are topped off. Engine fuelduring refueling increase a" the tanks fill, ranging frompm total.

fuel management prior to refueling

Up0 pounds of fuel should betorior toefueling contact. Ihe transfer Improves the un-augmcntcd pitch stability of the aircraft by moving It* center of gravity forward. With normal SAS operation, there is no marked change in handling characteristics.

In someanker rendezvous may be madeigh speed run withull or almost full, and fuel remaining In tankormal forward transfer would emptyirst. nly forward transfer is more desirable in this case to make the maximum amount of space available in that tank for cool fuel from the tanker. This improves the fuel heat sink capability. This also speeds up the refueling operation. Hot fuel transferred froms consumed fromhen its pumps start. With both generators functioning normally, the air refueling procedure may be accomplished without any fuel management action except that failure toFR Onlyan air refueling will affect theof the fuel infter the air refueling is completed,

fuel management during refueling

During refueling, the engines are supplied by normal pump sequencing. umps, or any other pumps, will remain on if they have been turned on by manual sequencing. With the forward transfer switch off, tank

1 pumps will continue to supply both mani -folds as long aa fuel remains in that tank. With the forward transfer switch on,umps are made Inoperative unlesss selected manually orb*. If not selected manually,ill be shut off wheneceive fuel and the startloat switches in the bottom of those tank* are opened.

subsonic cruise fuel management

The aircraft fuel system sequencing was designed to optimise the. at normal cruise. The recommended c. g. for subsonic flight isa atated in Section V.

For subsonic cruise the following procedure will keep tho CG within the recommended limits for subsonic flight.

0 lbs fuel remaining or afterif initial load is less0 lbs, turn forward transfer switch on, Leave forward transfer switch on until before landing check or when starting acceleration if final portion of flight is supersonic.

fught control system

Do not move the control stick during engine start. The Inboard and outboard elevons may not deflect simultaneously or equally when powered by one hydraulic system If pressure is less0 psi. This is due to unequal friction in the inboard andsets of elevon actuators. reloaded spring in the control pushrod mechanism can move from its detent position as aof unequal elevon movement, of normal hydraulic pressure should cause the elevons to resume normaland restore the spring to Its detented position. However, an inspection lato ascertain that the spring has reset properly.



PitchChcocieriitici Dv to Logoed Pitch Koht pitching

SAS Lagged Pitch Rate switching may cause transient load factors to develop when climbing or descending0 feet. Thisormal SAS characteristic which results from the design of the SAS pitch rate damping circuits.

Signals from the SAS pitch gyros always gotraight-through circuit that varies the pitch rate gain (damping response rate) with pitot differential pressure. In addition to this path, the signals goagged pitch rate circuit that changes tho pitch rate gain when0 ft. altitude.escent, the lagged pitch rate term is switched out0ut this does not Instantly remove the existing command. The signal that existed prior to switching drops Instantlyevel equalf its value for the higher altitude. The remainder ol the signal decays exponetially to zeroime constant ofeconds.

These two pitch rate gains are summed prior to Introduction into the sorvo Therefore, as the aircraft descends0 ft.urn, the SAS is supplying an input to the pitch transfer valvesunction of aircraft bank angle, Mach number, altitude, and pitch rate, causing the elevon surfaces to be deflected from the position commanded by stick position and trim setting as long as there is an aircraft pitch rate. The nose up pitch rate causes the SAS to oppose pilot control and/or trim action byown elevon deflection increment.

When the aircraft passes0 ft. altitude while descending, the lagged pitch rate term is switched out of the circuit, causing the pitch rate gain to reduce to that of the straight through circuit. Response to control stick positioning becomes more positive, SAS opposition to the pilot Induced

pitch rate Is reduced, and the elevons are automatically repositionedew angle which is governed by the lesser SAS gain. The aircraft responds to this surface change byump in the nose up pitch direction. Repositioning of the control stick and/or rotrlmmlng iswith the amount of changeunction of pitch rate desired before and after the transition. The reverse action is prevalent if the aircraft is in an ascendinf turn.

During transitionst. with no pitchtraight Line climb or descent, the gain will switch0 feet with no resulting aircraft movement.


On Drynti-Skid On

stop an airplane, the kinetic energy must be absorbed by aerodynamic drag, braking action and rolling friction. The variable relationship of these factors is modified by the landing or aborted takeoff roll distance, speed at the beginning of deceleration and weight. Rolling friction, although ofeffect, is neglected in thediscussion. Aerodynamic braking is composed of drag effects of the airplanes surfaces and the drag chute. The most effective airplane surfaces are the wing and elevons.

Energise the brakes smoothly withto heavy pressure with the nosewheel on the runway for dry runway conditions. It is unlikely that the anti-skid system will actuate on dry runways at weightsounds. At lighter weights, it is possible to cause momentary wheel spindown and, if this occurs, the center tires are more susceptible to blowout.

action is limited by three factors; braking friction available between the tires and the runway, methods of brakeand the kinetic energy limits of the brakes. Rated brake capability Is shown in, Braking friction availableon runway conditions, weight, and positive or negative aerodynamic lift Other considerations affectingcapability arc tire hydroplaning effects, efficiency of anti-skid devices and pilot technique. Techniques must be variedon conditions existing at the time. Speed at start of deceleration will beon touchdown speed during landings and time required to deploy the chute and apply brakes. Aborted takeoff speed will depend on the abort decision point and speed increase during the period for pilot reaction, Weight factors will differ depending onor emergency landings or aborted take-offs, Stopping techniques and/or procedures must also be varied depending on deployment or non-deployment of drag chute, runway length, weather conditions and tire

Drag chute failure procedures In Section III are mainly intended for landing weightalthough the procedures areapplicable to aborted takeoff weights also. The following braking techniqueis mainly applicable to heavy weight aborted takeoffs.

Pilot judgement regarding braking technique must be used after drag chute deployment. Brake pressure may be relaxed If alow speed can be attainedhort run with ample distance remaining. However, If distance is criticalong run would result, maximum use of brakes must be continued until the stop is assured. This requires hard and continuous brake pressure.


Hard braking may result in brake seizure after stopping, increasing time to clear the runway. Ifkeep the aircraft moving at slow speed until clear of the Taxiing at low speed tounway is permitted with all tires failedain gear. The massive tire bead tends lo protect the wheelshort distance at heavy weight.

Long rune at heavy weights may result in blown tires due to sidewall failures. of one tire will usually overload thetires on that side and probably cause successive failures. Therefore, sufficient brake pressure should beduring and after chute deployment to minimize stop distance. This reduces heat build-up in the tires which is characteristic of extended roll-outs at heavy weight. If

tire failure is known or Suspected, maintain enough brake pressure to prevent wheel spin-up and, possibly, wheel and/or tire disintegration at high rotational speeds,


.ry runway, do not use up elevonethod of increasing braking force because of therisk of tire fatigueaster stop is possible with the rolling stock intact,

. Ii tires blow with either wet or dry conditions, increased brake pressure on that side le required to maintain braking force with the remaining tires,

. Rated brake energy capacities and associated maximum braking speeds may be disregarded during aborted takeoffs. It is considered better to use the brakes at high speed, as tire failure may occur if the roll is extended by delayed braking.

Broking on We* Runwqyi

Energize the brakes smoothly with light to moderate pressure and the nosewheel on the runway for wet or slippery runway If the drag chute does not deploy, select NORMAL {or ALTRAKE if the left engine hasnd then shutailed engine, or shutdown the right engine if there has been no engine failure. In order to reduce thrust and increaseeffectiveness. Also use moderate up elevons so as to provide as much drag as posssible without lifting the nosewheel. The increased gear load may cause tire failure at heavy weight; however, tire failure may be acceptable since the tires will notdisintegrate. Braking deceleration

available is nearly the same for braked tire rolling and blown tire locked conditionset surface. Locked wheel skids of up0 feet have left the wheels undamaged during wet runway testing.

Unless hydroplaning is encountered, good nosewheel and rudder steeringcan be expected and have beenduring stops on wet runways with and without the drag chute, with all main gear tires blown and wheels locked, and with one engine shut down,

Hydroplaning In various formeimiting factor with wet runway conditions and,nosewheel and rudder steeringeffective, wheel braking force is nill until the tires can make contact with the runway. The aircraft tends to follow aand will driftross wind. for the extended stop distance involved, skids across or into dry runway areas are the chief hazard of wet runway stops. The wheels tend to lock-up and cause blown tires while slidinget surface. Dry areas tend to destroy the tires due tofriction or wheel spin-up, This allows the wheels to make runway contact and may ultimately destroy the wheels and then the brake assemblies. Even so, the aircraft can probably survive on the landing gear struts so long as It remains on the main runway, orard surfacewhere theremooth transition from runway to overrun.

ill weather operation




Before Instrument



Instrument CruisingApproach &






Except for repetition necessary for emphasie or continuity of thought, this section contains only those procedures which differ from or are in addition to the normal procedures supplied in Section II.


aircraft handle well during all phases of instrument flight when operated inwith procedures specified in theparagraphs. As with all highjet aircraft, constant attention to flight instruments is required. Normal jet instrument techniques are satisfactory for operation during Instrument conditions. aids include an Inortial Navigation

Syetem. TACAN. ILS and ADF. IFF ia als installed, aa are the directional and ranginf features of theadio.

The ships pitot static system Is the primary speed and altitude reference during takeoff, penetration, approach and landing. Speeds given hore are knots indicated airspeed (KIAS). Equivalent airspeeds (KEAS) and


altitude information from the air datasystem (TDI instrument) can be used; however, TDI response may not be as rapid as the ship system indications duringairspeed situations.


Maximum thrust will be used for instrument takeoffs. using procedures identical to those contained in Section II. The followingsupplement those given in Section II:

stall warning indication is referenced to pitot total pressure and to the attitude probe In the Rosemount pltot-statlc boom. It is independent of pltot-static pressures sensed by the ship and air data computer systems to that extent. Pitot heat should be sufficient to keep both the pitot head and the attitude probe operating during icing


Keep pitot heat on during allinstrument flight operations.


- Begin at computed Apply smooth, constantto establish an2 degrees on theIn about five seconds. will fly off the runway atairspeeds.

oegree pitchwhile accelerating tospeed. The altimeter andvelocity indicator should showclimb Indication beforethe landing gear. Care muatto insureositiveclimb ia maintained duringto climb speed in order tothe aircraft from settling backrunway surface.

aligning the aircraft with the center-line of the runway, check synchronisation of the FRS compass, check the INS mode selector in FRS position, and sat theindicator so that the miniature airplane is level with the horizon line. The BDHIeedle selector switch is placed in the TACAN position to display TACANInformation on theeedle of the BDHI.


The FRS compass will be used for heading reference during all take-offs and Instrument departures.


Initial Indications of the altimeter and vertical speed indicator may be thatlight doscent.

gearP whenairborne.

- Minimum afterburningup Is indicated. Militarybe setIAS isa Military power climb iainstrument departure.


Use Indicated Airspeed duringand climb until proper climb speed schedule is reached on the TDI.



Instrument climb* mng tbe normaland afterburner schedule* can be made safely, it may be desirable to main-lain maximum afterburning alter Ukroff at heavy weights, but allowances must be made for the more rapid acceleration and steeper than normal climb attitude. It is highly recommended that the normalschedule be used alter takeoff. This minimizes the possibility of exceeding the desired climb speed schedule. It alsomore time for EGT control if thi* is required. Maximum thrust may be resumed If desired after stabilizing at the proper climb speed.


. Reduce climb speed if rough air is encountered as described inIn Turbulence, this section.

. The TDI and ship system pi tot-static flight instruments should be cross checked periodically during instrument flight toproper operation.

Restrict all turning maneuvers tobank angle during low altitudeflight.


The optimum VFR Military thrust schedule is suitable for instrument climbs toaltitudes. As soon as the climb schedule is intercepted, the TDI become* the primary pitch control instrument for the remainder of the climb.


Establish cruising airspeed at the desired altitude and retrim the aircraft. After

Instrument Departure Instructions have been accomplished, the BDHI may be switched to display INS navigationas required for mission completion. Readjust the horizon bar on the attitudeto indicate level flight attitude when the aircraft is in level flight at cruising airspeed. These aircraft have excellent handling characteristics throughout their normal flight speed range if properly trimmed and flown by reference to the flight instruments.


Below Flight, themeat be set to station pressure and used to mainUin assigned altitude.


A constantangle of bank may be used for all turn* except rate turns whenor desired.


Any angle of bank exceedingiaa stoop turn. The aircraft Is easily controlled on instruments in banks up toowever, due to structural loadbank angles In excess ofshould be avoided.


Holding patterns and descent* between holding levels should be flownIAS at altitudes fromO0 ft0 feet. (KEAS rangesnots at theeo altitudes.) 0 rpm will be required. At normal weights, average fuel flow while turning varies from0 pph per engine at the higher altitudes to0 pph




0 lent. Tlie rate decreasespli per engine during straightomewhat lower airspeeds can be used, it* desired, il there ia little or no turbulence. To descend between holding levels, reduce power0 feet above thealtitude. Refer to Appendix for loiter performance.


The IKS mode selector must be placed In the FRS position prior to Initial station passage and entry into the holding pattern.

. Check the FRS compass forand that the BDHIeedle selector switch is in the desired TACAN or ADF position.

. When the BDHIeedleswitch is placed in the TACAN or ADF position and the INS mode selector switch is placed in the FRS position, theeedle of the BDHI will display magnetic bearing to the selected ADF or TACAN station provided theunction switch is not in the ADF position. For the same conditions except with DNS mode selected, true heading will be displayed.


The ships pitot-static instruments are the primary flight instruments during aprocedure descent. These penetrations are flownIAS with power set0 rpm. Initial rate of descent will00 fpm. 0 pph per engine fuel flow can be expected for normal weights when starting0 feet. The initial rpm should be maintained and fuel flow allowed to increase as altitude is lost,


Engine speeds0 rpm should be avoided to preventof the engine start bleed valves. TACAN will beif left engine rpm Is below

The landing gear may be used for additional drag during the penetration If desired, but should be extended no earlier than middle station passage and no later than the turn to final approach whenrocedure turn to final approach heading. ormal teardrop penetration or straight in approach tho landing goar should be extended prior to the final approach gate. At normalgross weights,IAS after level off through the turn to final approach. Total fuel flow increases to0 lb/hr with the gear down. Final approach speeds are identical to those for normal traffic patterns and landings and will be adjusted for existing gross weight. ingle engine approach the gear should not be extended until final approach Is initiated. Minimum approach speedIASingle engine


Fuel requiredypicalpenetration is00 pounds.


These aircraft are equipped to make either TACAN, ILS or ADF approaches. Approach Radar (PAH) approaches may also be made. When flown asaircraft control response Is good at all times. The downwind or outbound portions of all approaches are flownIAS with the landing gear down. The base



incrrai final1 knot br tacitw 5co0iw. olluil remain lr<i


leg or procedure turn portions areIAS. The minimum final approach speedIAS and should be Increased by one knot for0 pounds of0 pounds under normalconditions. With one enginehold gear extension until the final turn is completed andinimum final approach speedIAS.


. When the left engine has failed, the landing gear must be extended using the Emergency Landing Gear Extension procedure. The pilot should be aware of the time required and of the other aircraft systems which are affected by loss of the left engine.

. Altimeter position errorare small at instrument approach speeds and may be neglected,

. Use the rain remover anddefog and deice systems as needed.


Apply Military thrust as soon as Ito-around is necessary, Use afterburning or Maximum thrust tf necessary. Raise the landing gear onlylimb has been established, and climb to the missed approach altitudeIAS. When positive rate of climb has been established adjust power as necessary toIAS and00 foot per minute climb. In theingle engine missed approach is necessary, follow the single engine go-around procedures in Section III andthe single engine minimum control Speed.


Fuel requiredissedand GCA is0 pounds. FR closed pattern go-aroundounds.


Information on flight through icing conditions is not conclusive at this time. Flight to and from terminal areas where heavy icing conditions and/or heavy rain are present Is undesirable. Extended Sight in any known icing conditions is prohibited. If icing conditions or heavy rain at near freezing conditions is encountered In flight, the engines must be examined for damage during post flight inspection.


Without hoi air delclng, forward visibility through the windshield ia unsatisfactory under all icing conditions at penetration and approach speeds. Ice buildup occurs very rapidly and dissipates very slowly,with heavy build-up. even afterto lower, warmer altitudes. Ice will build up on the spikes at penetration andspeeds and enter the engine as It breaks off upon descent to warmer altitudes. Engine damage due to ice ingestion is not normally severe enough to causa engine shutdown and can be minimised by reducing

rpm. Hot air (low on the windshield is satis factory for de-icing and inhibiting iceif used prior to the time that icingare encountered. If windshield icing is anticipated or encountered:


In rain, forward visibility is obscuredater film which extends over almost all of the windshield area. Use of the rainliquid during light and moderate rain conditions improves visibilitysable condition at approach speeds. Visibility is momentarily obscured as the liquid Is ap-aplied. then the windshield clears and beads of water form which stream across the glass. Pain remover application is needed at ten to fifteen second intervals for best effectiveness. The hot air deicer should not be used in light to moderate rain, as the hot air by itself does not clear theand the rain remover liquid isblown away before it can become effective. The rain remover system is not effective with very heavy rain conditions, and, although hot air deicing provides very slight improvement, visibility remains


Reduce speedIAS before applying rain remover fluid.

[ CatlTION

Do not apply rain repellentry windshield. Prolongedmay result.

Rain removalUSH. NOie

Momentary cloudiness will occur.

2. Repeat as required when visibility deteriorates.

high humidity conditions

If fog emanates from cockpit overheadducts:


If condensation forms on inner or outer glass:

Windshield defogNCREASE as required.

Windshield deicerN R, or& R, as required.

turbulence and thunderstorms

should not be scheduled through areas where extreme or severe turbulence is forecast. In the event lhat suchare encountered however, airspeed should be maintainedEASeneral rule. Refer to the Structural Capability In Gusts chart,.

operation in turbulence

Gust conditions are defined in terms of "extreme", "severe", and "moderate to mild" conditions. (Refer USAF5. Aircraft structuralin turbulence air do not penalize

mal operating procedures except when0 leet. In the event ol0 feet, airspeed should be maintainedEAS.hows that the aircraft can be operated safely in severe turbulence at the design speed unless in the altitude range00 feet. peed reductionEAS should beby reducing power if severeis encountered while operating at high speed in this area. (The normallimb speed in this area isormal descentEAS,the optimum path for rough air penetration.

Transonic Acceleration in Turbulence

The probability of encountering unforecasted severe or extreme turbulence in clear air is relatively small. However, if there is a

reasonable possibility that this may occur during the transonic acceleration phase, modify the normal climb procedure. After reaching supersonic conditions, climbEAS insteadEAS while0 feet. Increase the climb speed above this altitudes to reach normal climb Speed00 feet. Be prepared toEAS if severe turbulence is encountered0 feet.

Jet Penelrofion ond Landing Approach

Normal penetration and approach speeds are compatible with rough air penetration schedules. However, the normal turn to final approach speed may be increasedIASIAS in order to avoid the possibility of maneuveringduring this phase. Standard rough air penetration techniques apply to this aircraft.


Detailed cold or hot weather procedures are not available. The pilot should always be aware of the effects of non-standard temperatures on takeoff and landing distances and minimum single engine control speeds. The pilot should also be aware of the effects of wet, Icy, and slush covered runways on takeoff and landing distances and on groundcharacteristics. Refer toor cold weather Oil Temperature operating limits.


Detailed specific night flying procedures are not required; however, the normal precaution of memorizing the positions of switches located in dim or unlighted locations should be Lower fire warning light covers to reduce glare In event of illumination.

rformance data


I -1

n Field Length

HI Climb And Descent

Subsonic Cruise

Supersonic Cruise



List of Illustrations


Position Error Corrections vs Mach

Position Error Correction* vs

Approximate Difference* Between IAS and

Position ErrorlternateAl-4

Airspeed Compressibility Correction

True Mach Number vs Equivalent Al-6

emperature Conversion

Standard Atmosphere

Standard Units Conversion


number indicators. The TDI providea

digital values for equivalent airspeedperformance charts are baaed onpressure altitude, and true Mach

data with VJ and/or YJ-l The differences between indicated

airspeed (KIAS) and KEASunctionaltitude, and ship systemand Fuelomparison of KEAS from the

triple display indicator and KIAS fromdata are applicable to aircraftnormal ayatem Instruments la shown

withE withdditive. Figure Al-3. For example: EAS,

viatlona from the nominal fuel densitynormal ship system will

ounds per gallon have negligible effect KIAS0 feet indicated pressureperformance as long asIAS0 feet indicated

gross weight and fuel load are known. How- altitude. Other combinations ofwith all tanks filled to capacity,and KIAS can be determined from

maximum fuel load0 pounds for Figureor use in the eventound per gallon change In

density. This effect on operationalmust be considered,

POSITION ERRORDisplay Indlcoior

Airspeed, altitude, and Mach numberTriple Display Indicator of the Air

available from the ship syatemcomputer system Is the primary in-

and from the triple display Indicatorfor climb and for all operations

The ship normal and alternate ayeteme sup- above. Its digital Indicationsconventional altimeter andcompletely compensated for position


error and compressibility effect a; however, the TDI airspeed lags during takeoff. IAS from thr normal ship system ahould be uaed until climb speed Is attained. The shipahould also be used for patternand landing, althoughlhe TDI follows these speed and altitude changes without excessive lag.

Normal Ship Syitefn

Figuresndhow position error corrections for the altimeter and alrspeed-Maeh number indicators when the normal (Rosemount pitot static) system Is selected. Corrections for operation at subsonic speeds were obtained by calibrations in flight and during ground runs. Corrections provided for supersonic speeds were obtained from wind tunnel results. Standard corrections for compressibility effects must be applied (subtracted) after the position errorara made in order to obtainairspeeds. The compressibilityIs supplied on Figure A The position error and compressiblltyhave been combined on3 In order toirect comparisonKIAS and KEAS instruments.


The alternate system senses pitot static pressure by meanslush static port and total head tube located under the right hand chine. When the alternate system Is selected, these pressures operate the ship system altimeter and alrspeed-MachIndicator. The position errorare provided by


KIAS after the airspeed position errorare made In order to obtain KEAS.


Figure Alhows the relationship between true Mach number, pressure altitude, and equivalent airspeed, based, the standard atmospheric parameter.


Ambient air temperature and true airspeed can be obtained from the TDI Mach and CIT gage as shown on the Mach-Airepeed-Tem-perature Chart, figure Al-7- For example,DI Mach5 and CIT, the ambient air temperature Isnd the true airspeed0he affect of adiabatlc compression and temperature rise on atmospherichas been Included by using aa* parameter.


6 ARDC standard atmospheric table. Figurerovides referencepressure, air density, and sonic speed Information which may be ofIn over-all flight planning.


The standard units conversion chart,,eans for direct conversion of temperature, distanco, and speed between English and metric units.

corrections for compressibilityon KIAS are provided by Figurehese correction should be subtracted from














apfucabuhip normal or alternate airspeeo svs1ems after position error correction

cfrom kcas to obtain keas

pffw, 'ii4-










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29 64

























9 OOQ704






a xero


8 *F

Standard Values at Seain. He



IbiecViV Siandardat Altitude 2 It jMithrrmjt temp F


TjRT or P^ mRT ConstantPfT. Constant Pressure ra/Kai/Tt Constant Temperaturer/V>

Reversible Adiabat*^

l 2)






Normal Performance

Maximum Performance


Refusal Speedr.

With Drag Chute. Dry

With Drag Chute. Wet

Without Drag Chute, Dry

Without Drag Chute, WetAZ-7

Maximum Weight For Single Engine Flight

Gearn Ground

Geari Ground

Gearut of Ground

Landing Speed

Landtngormal Performance

With Drag Chute, Dry

With Drag Chute.

Without Drag Chute, Dry

Without Drag Chute. Wet

Landingaximum Performance

With Drag Chute, Dry

With Drag Chute, Wet

Without Drag Chute, Dry

Without Drag Chute, Wet

Takeoff performance data are supplied for two type* of takeoff operations. One applies to normal operation when field length is not critical with respect to takeoff distance. This type of takeoff is basedotation speedIAS for airplane gross weightsounds or moreIAS for all' lower gross weights. This permits a

variable liftoff speed and allows'afor takeoff with tatlwindrestricted by tire limit speedknots

The other type of takeoff appries to maximum performance operation where takeoff and rotation speeds are varied with gross weight so that takeoff Speed correspondsift coefficientirspeeds at takeoff and the resultant ground run distance with this schedule



always less than (or the normalschedule. Pitch angle atis approximately 11 . g. This schedule results in minimal tail clearance at liftoff. Therefore, this schedule Li recommended only when the normal schedule or an intermediate schedule is inadequate for takeoff orrequirements and/or tailwlnd*.

Normal Takeoff Performance and Rotation Speed Schedule

Thr normal performance takeoff speed is basedonstant nosewheel liftoff speednots for gross weightsounds or morenots for gross weights lessounds. Time from nosewheel off to main gear off Is assumed tobased on average rotation time from test data. (Start of stick input to initiate rotaticn should anticipateoff by aboutnots The normal performance takeoff speed schedule based on constant rotation speed is incorporated in the takeoff ground run distance chart. When the takeoff is to be madeallwind component, the nosewheel liftoff speed may be decreased slightly, ifto avoid exceeding tire limit speed. Takeoff speed is also reduced by one knot per knot decrease in rotation speed. Reducing the rotation and takeoff speeds resultsne percent decrease in ground run distance per knot decrease In airspeed.

The normal performance takeoff speedargin of control during cllmbout in the event of engine failure Just after liftoff if immediate corrective

action lb taken. Takeoff speeds are higher than the minimum single engine control speeds for steady flight (speeds at which 5 to 10 of bank with upideslip can be used to maintain flight path, utilizing up to full rudderwith maximumn theengine. See Aftercidealip and bank toward the operating engine are recommended with up to 9 rudder trim to minimise drag. The trim required decreases withairspeed.

The takeoff speed schedule for normal performance is Incorporated in theground run distance chart,

Maximum Performance Rotationkcuff .Speed Schrdule

The maximum performance rotation and takeoff speed schedules are shown in2 as part of the takeoffchart. Maximum performance takeoff speeds are based on the takeoff attitude whichift coefficient0 In ground effect. g. Therefore, takeoff speedirect function of airplane gross weight and takeoff speeds are listed on the gross weight lines of the takeoff distance chart. Rotation or nOBewheel liftoff speed Is that speed where the airplane rotation is initiated and is scheduled toeconds prior to liftoff. speeds are plotted on the chartunction of ground run distance and as affected by wind and slope.


Figures A22 present normal and maximum performance takeoff ground run distance*unction of ambient temperature, pressure altitude andeight. Correction grids for wind and slope effects are included. Limit tailwinri arMl are incorporated for reference and may be used aa anof takeoff overspeed margin (the difference between takeoff speed and tire limit speed). The takeoff distances are based on test data and have noallowance included.

Normal Perf<irm.ir.cf Takeoff Ground Run Distance

1 shows normal performance takeoff distances obtained by use of the normal performance rotation speed and the takeoff speed schedules incorporated in this chart.


For an ambient temperature ofpressure altitude0 feet andgroHs weightounds, determine the ground run distance and takeoff speech

Enter1 at the ambientand pressure altitude, proceed horizontally to the gross weight and read down toero wind zero slope distance0 feet. Continue down tonot rotation speed line (applicable to weightsnd read to the right horizontally tothe aero wind and aero slope

takeoff speednots. For aofnotsownslopeontinue to read downistanceeet. The wind does not affect the takeoff speed but theownhill slope decreases the takeoff apeednots.

Maximum Pf rformar.ee Takeoff Ground Run Distance

2 shows maximumtakeoff distance! consistent with theaximum performance takeoff speeds and the rotation speed schedules which are also shown in this chart.


Using the same conditions as in theexample, determine the maximum performance ground run distance,speed and takeoff speed. Enter2 at the temperature and pressure altitude condition, proceed horizontally to tbe grops weight and read down to the zero wind zero slope distance0 feet. The takeoff speed listed on the distance lineoundsnots. Continuing to read down, intersect the rotation speed curve, read to the right for the zero wind and zero slope rotation speednots. Continuing vertically, determine the distance with wind and slope0 feet. For the wind and slope case, the wind doen not affect the rotation speed but the slope decreases thespeednots.

Intermediate Takeoff .Speed Schedule

When the normal performance schedule is inadequate because of runway length and takeoff weight and temperature, or other conditions, an intermediatefcpeed should bo considered before reverting fully to the maximumbchedule. When rated tire .speed is the limiting factor becauseail-wind condition, base theced on the rated tire vpeed. Rated tire speed can be determined from. ominal "pad"nots is desired, the intermediate takeoff speed will be the rated lire Hpeed minusnotJ. The corresponding takeoffcan be obtained by using themethod: At the samepressure altitude, gross weight, wind and slope conditions, find theperformance takeoff ground run Subtract the intermediate takeoff speed from the normal takeoff speed. This number is equal to the percentof tlie normal performance Do not schedule takeoffs at speeds lower than the schedule forperformance.

cratlon check speed below theof this guide line and the checkvalue selected. The break in the acceleration line* atnot*the change in acceleration at. maximum thrust is obtained.


For example, inoundperformance takeoff example where the aero wind distance0 feetIAS. the guide line illustrated shuws that an acceleration check speedIAS would be reachedfoot check distance. In the case where the takeoff run0 feet0 knot headwind anduide line drawnoint0 feet KIAS intersectsfoot check distanceIAS. foot check distance point is recommended for operation at weightsb, since airspeeds reached at this distance allow reasonable accuracy in making the speed check. foot checkhowever, is recommended at light weight or low temperature

Check Speed

3eans fora speed check at any requiredduring the takeoff run. Use tlie takeoff speed and distance values found from12 to locate the positionuide line on the acccl-

Fuel Allowance

A- 12


Planning data (or determination ofcheck and refusal speeds may be obtained from tiie refusal speed4 Refusal speed without chute may be used asacceleration check speed instead of the normal line speed check, if desired, provided the corresponding checkis also computed. Either the scheduled rotation speed or therefusal speed with chute,is lower, is the maximum speed alecision toakeoff is recommended.

The charts show actual maximumrefused takeoff capabilityconservatism or service allowances. However, various factors mayto performance less than optimum.lown tires, delay in drag chute deployment, etc. Brake energy capacity is assumed tof full rated one-stop capability, thus allowing someservice use prior to the refused takeoff. This factor is presented as-maximum refusal speed. For abort conditions where brake burn-out might occurtop can be made the aircraft is assumed to free roll atlow speed and not stop.

Refusal Speed With Drag Chute

The refusal speed with drag chute charts are shown in45 for dry and wet runways, respectively. The abort speeds are given as function of temperature, pressure altitude and gross weight for available runway length. Assumptions made in the refusal speed

calculations are as follows:

Normal rate of acceleration is maintained to the refusal speed at whichoniplctc andloss of one engine occurs.

Ma:-irnurn afterburning thrust is maintained on the operaling engineoi obtained from the failed engineecond: before the throttle on the operating engine is retarded to idle-


No rotation is attemptedakeoff emergency occursreaching rotation speed, even though airspeed may exceed rotation speed during thisand action period.

If the takeoff is aborted afterhas been initiated, aerodynamic braking is utilized until the aircraft has decelerated to chute deployment speednots. (Refer to ABORT procedure, Section III. )

The nose is lowered before brake application if aerodynamic braking has been required. bove.)

Braking torque is obtained oneafter retarding the throttles unless rotation has been initiated. The allowance for chute5 seconds after drag chute switch actuation.

Optimum wheel braking is continued until the aircraft is stopped.

7. Dragn jettisoned atnotsry runway and Is retainud to lull stopet runway. (The chute should always be retained until stopping is assured. )

3. Zero wind and zero sope.

ard surface runway. (The effect' o! water,oronhave not been considered. )

10. The takeoff is continued if rotation has been initiated prior to an engine failureositive climbout capability exists.


Using the same conditions as tn theexample,0 feet pressure altitude,ounds gross weight, find the refusal speed with drag chute0 foot dry runway.

Enter4 at the temperature and pressure altitude conditions,horizontally to the gross weight, then downward to the available runway length (accelerate and stop distance available) andefusal speednots. (This refusal speed is usable onlyf the brake energy isn't) Prm rrd downwardb dashed line andaximumspeednot" and interpolate Che distance to accelerate and stop0 ft.

Using the same procedure for the Wei runway case, enter5 andefusal speed

Befu.al Speed Without Drag Chute

speeds without drag chute are presented in67 for dry nnd wetespectively. Theeed* are givenunction of temperature, pressure altitude and gross weightie available runway length. Assumptions made in the no chute refusal speedre the same as those made for calculatingspeeds with chute with theexceptioni:

Drag chute deployment is attempted a* the nose Lf loweredIAS (if rotation has been accomplished) or at start of braking if abortingower speed. Drag chute failure recognition is normal chute deploy time5 secondseconds recognition time.

etay, one engino is s'cut dovni after recognition of chute fail ere,rag is usedhe nosettitudenti-skid is turned ', off andtop is with all tires lown.

Using the same conditions as in theexamples, find the refusal speed for dry and wet runway*.

ry runway, enter6 at the temperature and altitudeproceed horizontally to the gross weight, then proceed downward to the

maximum refusal speed linebs andaximum refusal speednots. Using the camefor the wet runway conditions.

enter7 andefusal



Three curves of single engine climb capability are supplied to show, the effect of speed and temperature as an aid in judging performance at takeoff weight with maximum thrust. 9. show the effect ofon maximum weight for gear down in ground effect, gear up in ground effect, and gear up out of ground effect, respectively. The illustratedrepresents wind tunnel tests egradation factor0 Ib should be applied before use and ascale incorporating this factor is provided. The values shown are forAC. For deviationg, decrease gross weight0 lb for each one-percent forward shiftg, or increase gross weight0 lb for each one-percent aft shiftg. If operating at lest than theweight for climbout, the excess thrust Indicated by the maximum weights shown can be used for acceleration and delayed climb. Instead of immediate climb at low airspeeds. Thisla usually permitted by the takeoff situation at the operating base. est climb speed with gear down close to the ground isIAS. The gear would ordinarily be retracting during an acceleration to this Speed, so the valuearget speed for transitionhallow climb attitude. The best single engine climb speed with gear up away from ground effect isIAS provides an adequate angle for single engine climb to pattern altitudes. In normal operation,IASarget speed forpultup to normal climb speed,


Using the same conditions as on the takeoff case for gear down in ground effect find the single engine capability speed. EnterS0 ft altitude and using the temporary decrement weight scale gives the speed asEAS. Use of the gear up out of ground effect curve,. gives the speedEAS.


Landing speed schedules and landing rollout distance information are provided for dry and wet runways, with anddrag chute, and for normal andperformance techniques.

normal and maximum performance landing speed schedules

hows the normal and maximum (heavy weight) performance landing speed schedulesunction of gross weight down0 pound*. At gross weights less0 pounds the landing speedIAS for normal performanceIAS for maximum performance operation. The maximum performance landing speeds arenots less than normal landing speeds for all weight conditions.

In addition to thr landing *peedwhich tire limit speed

will be exceeded may be obtained from this chart.


a normal landing atgroi' weight,approach and landinglimit tailwlnd for0 feet andof.

Enter the chart0 pounds, proceed up to intercept the normal approach *perd lineIAS and the normal landingIAS. Enter thealtitude chart0 feet, read up to intersectnot landing speed line. Thin determines the limit tailwlndto avoid exceeding tire limit speed as overnots.

final approach andspeeds and limit tailwindwiounds,pressure altitude andtemperature. Enterounds grossup to intercept the finalspeed schedulethe heavy weight landinglineIAS. Enterchart athorizontally to theline, read up tonot landing speedreadnots at, the limitcomponent.


Normal Performance Landing Ground Roll Distance Willi Drag Chute

The normal performance landing ground roll distance with drag chute for dry and wet runways is shown inespectively. is givenunction oftemperature, pressure altitude and gross weight. Wind and slope effect! are included. Standard landing technique assumes chute deploy switch actuationecond after touchdown, chute fully deployedeconds after touchdown, and the nose downIAS. Full braking prenun requiresecond after initial pedal depression. The chute Is assumed to be Jettisoned atnots for the dry runway case and Is retained to full stop for wet runway conditions.


For conditions ofair0 feet pressure altitude0 pounds gross weight, find the normal performance landing ground rollwith drag chute.-

1. Enter2 for the dry

runway at the temperature andcondition, proceed horizontally to the gross weight and readto determine zero wind, zero slope ground roll distance0 feet. eadwind ofnot*ownhill slope ofhe ground roll distance would0 feet, as shown in the chart.


Applying the VMM procedure in thanw.iv chart,.rro wind, zero0 (cel. For the wind and -dope case the distance would0 feet.

Normal Perform ante Landingil Roll Distance Without Drag Cl.ute

Normalanding ground roll distance for Un.lnig without chute isnI4or dry and wet run<vays. Normalwithout chute assumes the same sequence of events aa for landing with chute. The nose of the airplane is lowered al or beforeIAS and brakes applied. In the no chute wet runway case it is assumed that one engine is shut down and anti-skid is turned offart of braking.

Using the same conditions aa in the landing with chute example, find the normal performance landing distance without drag chute.

For the dry runway case, entert the ambientand pressure altitudeproceed horizontally to the gross weight, then read downward toero wind, zero slope ground roll distance0 feet. With *ind and slope,0 feet as shown in the chart.

etondition, the zero wind and slope distance0 feet, as shown in. The wind and0 feet.

Maximum Performance Landing Ground Roll Pittance Willi Drag Chute

When u'ing the minimum roll landing technique,ith drag chute may beior dry and wet runways, Performance is given as aof temperature, pressure altitude, ami gross weight. Wind and slope effect' are incljded. Minimum roll technique assume* touchdown atnots slower speed than normal. If touchdown is at speeds lessnots, chuteW assumed to be initiated and the no.ic lowered as soon as the main gear touches. If at speedsnots, chute deployment and lowering the nose is delayednotsreached. Brake: are applied ashe nosen the ground. The chute is assumed jettisoned atnots for the dry runway and is retained to full slop for the wet runway.


Using the same conditions an In thelanding examples0 lb find the maximum performance landingwith chute.

For the dry runway case, enterI6 at the ambientand pressure altitude, proceed horizontally to the gross weight. Read downward toero wind, zero slope ground0 feet. With wind and slope the distance0 feet.

In the caseet runway, the zero wind, zero slope distance0 feet as uhown in. The windope effectistance0 feet.

a.' '

MaximumLanding Ground Roll Distant Without Drag Chute

Maximum performance landing distances without drag chute are presented inft and9 for dry and wet runways, respectively. Theicnce-of event' is assumed as forwith the drag chute. hreerecognition time, of chute failure ia incorporated before shutdown of the right engine and turning anil-skid off for the wet runway landing.


Using the same conditionsusedfind the maximum performance landing distance without chute.

1. For the dry runway case entert the ambient temperature and pressure altitude, proceedto the gross weight. Read downward toeroro slope distance0 feet. Continuing through the wind and slope correctionsistance0 feet.

2, In the wet runway case, the zero wind and zero slope distance0 feet as shown in. With wind and slope, the distance0 feet, as shown in the chart.




Prm olt 0

Runway temp 84 Wind^-T) rode 1 Up^Dn) Runway et

% (Up-Dn) Runway Dry -

Runwayt Pren0 Ft Runwayind0 Kn Grade 0

check tpeed


Predicted ground



Min ilngle engine (georind)

Runwoy length available

Max refusal with chute (0

Max braking wlrh

Aoproacl. Land Chute No 0 I 0 Ft

Chute No chute No I

limit IAS

Landing Immediately After Lb 0 0tAKINC- SPEEDS: Fuel remaining

No Limit


Londing Immediately After Lb

re final no chuteind) Max broking no chute


123 KIAS

per knot head or tail wind component.

? >

h r. z





_ l


1 mi



a ^


i IT




Normal ClimbRDC Atmosphere)

Phase I, Subsonic climb from brake1

Phase IA. Subsonic climb from off2

Phase JJ, Transonic

Phase DJ. StdT, Supersonic4

Std Day, Supersonic

Std Day,flT, SupersonicA3-6

Normal ClimbEAN TROPIC" Atmosphere)

Phase I, Subsonic climb from brakeA3-7

Phase IA, Subsonic climb from offA3-6

Phase U, Transonic Maneuver 9

Phase UJ, Moan TropicT, Supersonic

Mean Tropic Day, Supersonic

Mean Tropic AT, Supersonic

Military Thrust Climb Performance, Std


Std Day

Std Day

Normal Descent


EAS Descent

Stogie Engine DescentRDC Atmosphera)




Stogie Engine DescentEAN TROPIC" Atmosphere)




Stogls Engine Turning


1resent normal performance to cruise altitudes foroperation6 ARDC Atmosphere and "MEAN TROPIC" Atmosphericrespectively. The dataomputed from results of Flight Test and Operational Testing withngines. The climb ie segmented In three phases and Includes the effects of varying gross weights and air temperatures on fuel used, time, and s the subsonic portion of the climb from brake release at sea level0 feet0 Mach. Corrections for time, fuel, and distance are listed In the chart for takeoffs from other field Phase IA is the subsonic portion of the climb from various refuel altitudes0 feet0 Mach. Phase II is the transonic acceleration portion of the climb0 feet0 Mach0 feot5 Mach utilizing tho "dive through" technique. Phase UI Is the supersonic portion of the climb0 feet5 Mach to the altitude at which cruise Mach number is first attained. Phase UIA la the constant Mach portion of the climb from the end of Phase UI to the altitude for start of cruise. The followingabulation of the average results of flight test* for Phase UIA.



Power Lb/Mui.

Long Range 0


High Altitude 0 Max



Obtain the time, distance, and fuelfrom brake release for takeoffield elevation0 feet0 Mach0 feettandard day. Fuel load at brake release0 lb afterground fuel allowances for normal ground operation. (See Appendix, Part U, for ground allowance computation) Find the initial gross weight at brake release by adding the zero fuel weight and the fuel load remaining. If the zero fuel weight0 pounds, the Initial gross weightounds.

Enter1 at the initial gross weight at brake release and read fuel used, time and distance for07 minutes,6 nautical miles, respectively.

From the table In table In1 for0 foot field elevation, reduce time, fuel, and distance0ounds,autical miles. Therefore, fuel used, time, and distance for07,0 the gros* weight at the end oflimb. Enter3 at the recomputed gross weight and read fuel used, time andfor00 minutes2 nautical mUes. ofnd Phase II result* In the fuel used, time, and distance to the start of Phase UI07.2 nautical. The recomputed gross weight for entering Phase HI wlU. Enter5 with the recomputed gross weight and at 0 feet andead fuel used, time, and distance for Phase UI0inutes,autical mUes, respectively. Add aU three phases and obtain fuel used, time, and distance




Normal Climb Performance From Brake

RDC Atmosphere)

Phase I. Subsonic1

phase II. Transonic A-

Phase III. Std Day. AT, Supersonic

Std Day. Supersonic

Std Day, +C AT, Supersonic

Normal Climb Performance After Air Refueling

RDC Atmosphere)

Normal Climb Performance From BrakeEAN TROPIC" Atmosphere)

Phase I, Subsonic Climb

Phase II, Transonic

oan Tropic Day AT,9

Mean Tropic Day, Supersonic

Mean Tropic0 C AT, Supersonic

Normal Climb Performance After Air Refueling


Military Thrust Climb Performance, Std

Std Day




Single Engine Descentax

Single Engine

Single Engine Turning


1S"resent normal climb performance from brake release to cruise altitudes for supersonic operation6 ARDCand "MEAN TROPIC" Atmospheric

conditions, respectively. Tho data isfrom results of Flight Test andTesting withngines. The climb is segmented in three phase* andthe effects of varying gross weights and air temperatures on fuel used, time, and distance. a the subsonic portion of the climb from brake release at sea level0 feet0 Mach. Cor-



Profile Scheduled


rcctlons for time, fuel, and distance are lintod in the chart for takeoffs from other field elevations. Phase II Is the transonic acceleration portion of the climb0 fent0 Mach0 feet5 Mach utilizing the climb and dive Phase III i9 the supersonic portion of the climb0 feet5 Mach to the altitude at which cruise Mach number is first attained. Phase UIA is the constant Mach portion of the climb from the end of Phase HI to the altitude for start of cruise. The followingabulation of the average results of flight tests for Phase 'ITA.


Avg. Total Fuel Flow Lb/Min.

Long Range 0


High Altitude 0 Max



Obtain the time, distance, and fuol required from brake release for takeoffield elevation0 feet0 Mach0 feettandard day. Fuel load at brake releaae0 lb after ground fuel allowances for normal ground operation. (See Appendix, Part U, for ground allowance computation procedure. Find the initial gross weight at brakeby adding the sero fuel weight and the fuel load remaining. If the aero fuel weight0 pounds, the Initial gross weightounds. Enter1 at the initial gross weight at brake release and read fuel used, time and distance for0inutes,iles, respectively.

From the table tnor0 foot field elevation, reduce time, fuel, and distance0b,mi, respectively. Therefore, fuel used.

time, and distance for04, respectively. Recompute the gross weight at end of climb. Enter2 aa the recomputed gross weight and read fuel used, time, and distance for Phase II0inutes, andautical miles, respectively. Recompute the gross weight at end of Phase. Enter4 with the recomputed gross weight and0 feet andead fuel used, time, and distance for0inutes,autical miles, respectively. Add all three phases and obtain fuel used, time, and distance03 minutes,autical miles, Fuel remaining0 feet0.

Service allowances and/or allowances for deviations from the normal climb schedule can be applied to an affected phaae when (Forubsonic cruise operation prior to reaccderating might be scheduled in the flight plan.) The effect of such an allowance must be accounted for when computing the initial weight to be used for the next phaae of the climb.

after air refueling

6resent normal climb performance from the end0 foot refuel (refueling with one AB on) to the altitudes at which cruise Mach number is reached6 ARDC Atmosphere and "MEAN TROPIC" Atmospheric conditions, respectively. Adjustments which should be uaed for otherltitudes are listed In the charts. The data la computed from Flight Teat and Operational Testing resulta withngines. The assumed fuel load at the end0 lb. A resulta are identical to the tabulation in the previous discussion.


Turns during climb are not recommended, however, if mission requirementsurn, compensation for range lost due to the turn must be included in the flight plan. For example,5 heading change0 bank at an initial altitude0 feet.

To minimize the rate of climb loss due to turning, the recommended procedure is to advance power touring the turn and maintain the speed scheduleEAS. Resume the normal climbon completion of the turn.

Comparison of straightaway climb andclimb on time, fuel and distance results In an overall range loss ofiles0 profile. On completion of theturn,0ime, fuel andto that altitude will be5b,i greater thanormal climb with no turn.


I6 present Military climb performancechedulenots equivalent airspeed (KEAS)eet0 Mach number when higher altitudes are attained. This power and speed schedule provides the most climb distance for the fuel consumed whencruising flight plana, such as for ferry or buddy missions, are used.


Find the time, distance and fuel required to climb0 feet from S. L.ay with an Initial gross weightb. Enter0 ft, andb initial gross weight6 miles0 lbs. Adding takeoff allowances results In time, distance and fuel values5 +

20or climb from sea level0 feet.


Find the time, distance and fuel required to climb0 feet0td day with an Initial gross weightounds. Enter0 feet andound initial gross weight;Uesounds. Reenter0 feet and an adjusted Initial gross weightounds;0 miles0 pounds. Adding takeoff allowance and subtracting values for climb from sea level0 feet results in time, distance and fuel values8)0) for climb from takeoff00 feet.


On course descent performance is ahown oa,. resents descent performanceEAS schedule. EAS descentwith forward bypass doors in the automatic and open positions respectively.


resent singl* engine descent performance0 feet and7 KEAS). The data Is based on flight test with the Inletas listed in the charts. Time, dietance, and fuel required are plotted versusushover at constant Mach Is required to Increase airspeedEAS toEASEAS schedule. Better range is obtained whenEAS schedule is used, reducing airspeedEAS while maintaining constant altitude. These effects are in-


in the performance data. Specific range begins to decrease rapidly0 feet; therefore, the charts arcto an altitude0 feet soower reduction technique can be used and the resultant change In performance can be determined. The effect of changingi the indexedeet has not beenby flight tealingot included in_ the data.

ummarizes the effect ofon Maximum AB descent performance for constant values.EAS. the effects of decreasing power at the index altitude0 feet for constant airspeeds,EAS, Z4 gresents the effects turn at 35 bank angleEAS descent. 0 feet of altitude Is required to complete turn. For convenience in missionround track profile is also


Find the track time, distance, and fuelto descend0 feet and0 usingEAS descent0 turn is to be completed0 feet, and Minimum AB is to be used0 feet. Enternd note on ths penetration distance curve thatf turn is completed0 feet altitude. Bead time at that altitudeinutes and fuel used0 pounds. On tho ground track profile note that the distance traveled isautical miles. Enter0 feet (end of turn altitude) and read time, distance, and fuel required0 feetinutes,autical miles,0 pounds. Reenter at the final altitude0 feet on the Minimum AB line and road time, distance, and fuel requiredinutes,autical miles0 lb of fuel. Add the incremental readings and5miles,0 pounds of fuel.

Find the time, distance, and fuel required to descend on course0 feet and0 usingEAS descent schedule and Minimum AB0 feet. Enter0 feet and read the time, distance, and fuel required0 feetautical miles,0 pounds of fuel. Reenter at the final altitude0 feet on theAB line and read time, distance, and fuel requiredinutes,autical miles0 lb of fuel. Add the result* and6autical miles,0 pounds of fuel.

03 minutes,autical miles, respectively. 0 feet0

Service allowances and/or aUowances for deviations from the normal climb schedule can be applied to an affected phase when (Forubsonic cruise operation prior to reaccelerating might be schedule in the flight plan.) The effect of such an aUowance must be accounted for when computing the initial weight to be used for the next phase of the climb.


Obtain the MEAN TROPIC day time,and fuel required from refuel0 feet0 Mach0 feet {start of Phasenter80 feet and read fuel used, time, and distance for Phase IA00 minutes,6 nautical miles, respectively. The recomputed gross weight for enteringill.


Turns during climb are not recommended, however, if mission requirementsurn, compensation for range lost due to the turn must be Included In the flight plan. For example,5 heading change0 bank at an initial altitude0 feet.

To minimize the rate of climb loss due to turning, the recommended procedure Is to advance power touring the turn and maintain the speed scheduleEAS. Resume the normal climbon completion of the turn.

Comparison of straightaway climb andclimb on time, fuel and distance results in an overall range loss ofiles0 profile. On completion of the 45 turn,0 feet; time, fuel andto that altitude will be5b,l greater thanormal climb with no turn.


resent Military climb performancechedulenots equivalent airspeed (KEAS)eet0 Mach number when higher altitudes are attained. This power and speed schedule provides the most climb distance for the fuel consumed whencruising flight plans. Such as for ferry or buddy missions, are used.


Find the time, distance and fuel required to climb0 feet fromay with an initial gross weightb. Enter0 ft, andb Initial gross weight6 miles0 lbs. Adding takeoff allowances results in time, distance and fuel values2)0) for climb from sea level0 feet.


Find the time, distance and fuel required to cUmb0 feet0 foot takeofftd day with an Initial gross weightounds. Enter0 feet andound initial gross weight;tle>ounds. Reenter0 feet and an adjusted Initial gross weightounds;0 milce0 pounds. Adding takeoff allowances and subtracting values for climb from sea level0 feet results in time, distance and fuel values8)0) for climb from takeoff00 feet.


On course descent performance Is shown on, resents descent performance for theEAS schedule. EAS descentwith forward bypass doors in the automatic and open positions respectively.


Single Engine Descent data Is presented for Military. Minimum afterburning andafterburning power. EAS6 ARDC and Mean Tropic Atmosphere conditions. Refer toB.

Allowances For Deceleration To Deicent Speed:

When cruisingigher KEAS than the desired descent schedule, the constantdeceleration is made at the same power setting as the constant KEAS descent. The constant Mach lines show the beginning point of the deceleration for each Mach number. In the situation where the cruise KEAS is less wan the desired descent KEAS, the constant Mach descent is made with Maximum afterburning power. The constant Mach lines show the descent for different Mach numbers.

tude0 feet noechnique of power or airspeed change can be used and the resultant effect after power change in performance can be determined. The effect of changing KEAS at the0 feet has nol been defined by flight testing and is not Included in the data.


Wheningle enginewith the operating engine in Military power, the Mach rate limitach In three minutes wtU be exceeded.

Single Engine Turning DeicenT

resents the effects of aturn at 35 bank angleEAS 0 feet of altitude is required to complete turn. For convenience in missionround track profile is also provided.

Comparison Of Dsicent Power and Speed Scheduled

The Maximum afterburning descent, as compared to the Minimum afterburning and Military power descents, resultsongeionger elapsed time and more fuel used. EAS descent astoEASlightly longer distance, less elapsed time and more fuel used. overall range resultsescent speedEAS Is used and if Military power Is used in the descent and for cruise. There wiU be little overaU range lose If either Minimum afterburning or Maximum afterburning deacent power la used as long aa the cruise Is accompUshsd In MlUtsry power. The charts are Indexed to an altl-

Somple Uie Of Chorti:

Find distance, time and fuel to deecend0 feet0 feet, usingafterburning powerEAS. Initial spaed la7 KEAS). (ARDC Standard) atmoaphere conditions are expected. Refer to. Enter tha chart0 feet and located the Minimum afterburning line for7 KEAS) condition, and readtime and fuel0 feet.

ilesinutes0 pounds






DAIA lASIli lism Flight







or A


Li i










t. Oalq i

3 Add0 lb..







u oooijaniuiv ianssisid












i All' ClOHO Al



Changed ISB



Lilt of Illustrations



Subsonic Long1

Subsonic Maximum Range Cruiseach

Maximum Subsonic Specific Range

Subsonic Range Factor

Buddy Mission70

Subsonic Specificach
















Long Rangefterburner

Long Rangeilitary


Single Engine Cruisefterburner fr

This part of the appendix supplies two engine cruise and loiter performance data and single engine cruise performance data. The material for two engine operationong range cruise chart, maximumrange summaries for long range cruise-climb anduddy missions, loiter performance, and specific range charts for altitudes0 feet0 feet. The single engine data show cruise climb range capability with andafterburner,pecific range chart for operation at Military thrust.



The two engine performance data applies to operation with YJ orngines when. IsAC. Operation at more forward c. g. conditions reducesrangeor each one percent shift In c.s noted on the specific range charts.


1 presents the constant altitude, maximum range cruise climb, and Military thrust cruise climb capability of thein terms of distance to go0 lbs gross weight0 lbs fuel remaining). The additional distanceto lower gross weights Is also provided. Cruise speeds for constant altitude cruise are tabulated on the chart. The chart can be used on an incremental basis for anystart and end cruise condition.


Determine the range available00 feet, and by cruise climbing with an initial gross weightb If cruise is to be terminated0 lbs fuel remaining0 lbs gross weight). 1 shows that by cruising0 feet the range will0 nmi. This range Increasesmi0 feet. Maximum range isby cruise climbing8 Mach number. Under this condition cruise would be initiated0 feet and ended0 feet0 lbs fuel remaining. Distancewould0 nmi.


2 presents the distance available0 lbs gross weight0 lb fuel remaining) for maximum range cruise climb8 Mach numberi . The chart can be used on an incremental basis for any desired start and end cruise condition.


3 presents the maximum specific range summary for cruise climb at various Mach numbers. Note that the optimum cruise climb occurs at This summary Is obtained from the subsonic range factor chart,y the equations Range Factorpecific Rangeand Its corresponding pressure altitude) =. (Refer to section on equations).


4 presents the subsonic range factor for long range cruise climb at any Mach number. The chart shows thereange factor and corresponding cruise climb schedule)iven cruise Mach number. Thisuick means for calculating best range available for any given cruise Mach. The chart also shows that the optimum rangeb-nmi/lb) occurs at8 and the corresponding cruise climb scheduleb.

Definition of Terms ,

KTAS nmi

Specific Range (instantaneous) -,

F lb

Rangepecific Rangeuel Used, nmi

Range Factorpecific RangeSixlb

Range Factorpecific Rangeinstantaneous).

ii ti Range Factor (avg) x

Range Available,

W (avg)

Fuel Used, nmi

final W

Rangeange Factornitial W


or Range Factor (avg)


x nm;



Distance flownnitial W.


Initial pressure

The final pressure ratio,

the standardith the Initial andratloa, and determine theInitial and final0 ft0 feet,


Determine the cruise fuel required and cruise climb schedule for cruise5 Mach. The planned cruise distancemi. tandard day with sero wind. Planned Initial cruise gross weight


tclimb schedule7 )lb and the rangeml/lb.


section on equations. In'w ('final

Determine the range available and the cruise climb schedule for cruise0 Mach. (Note that this la not the optimum cruise speed.) The Initial cruising weightb,0 lb of fuel are to be used. tandard day with sero wind.

gross weight0 lb.

tclimb schedule* )lb and the range factor la- nmi/lb.


Distance Range Factor

W (final) 0 '


0 lb.

Therefore, cruise fuel0 lb.

c. Using the same method aa In theexample, the approximate initial and final cruise altitude*0 feet0 feet, respectively.


5 presents the distance available0 lbs groas weight0 lb fuel remaining) for Buddy Mission cruise at70 feet. The speed and altitude schedule Is compatible withanker performance The chart can be used on anbasis for any desired start and end cruise condition.


6 presents specific range data at The Buddy Mission altitude la listed on the chart. If desired, greater range la obtained by cruise climbing.


7 presents loiter performance as minutes0 lb of fuel used. Thespeed schedule is listed in the chart.


Determine the loiter time available0 feet for an Initial gross weight0 lb. 0 lb of fuel is to be consumed. Enter70 lbs gross weight0 feet and9 minutes0 lb of fuel. Reenter0 lbs0 feet and2 minutes0 lb of fuel. The average value5 minutes0 lb of fuel. This5 minutes for the0 iba of fuel consumption.


The specific range charts8) present cruiee data for varlouaaltitudes0 ft0 ft) throughout the speed range from rnaacimum

endurance to Military thrust. Each chart presents nautical miles0 lb of fuel (nml/Klb)unction of Mach number and gross weight with aubscales of KEAS and KTAS for standard day. Also included are an overlay grid of fuel flow per engine, the maximum range speed schedule, and the recommended loiter speed schedule.


The single engine performance data applies to operation with YJ engines. iveservice allowance Is Included. Refer to text for other Items affecting theresults. The long range cruise data for both Military and Afterburnercan be used in conjunction with the single engine descent information in Part ill. Transition from end of descent (asin the single engine descent curves) to start of single engine cruise Isby drift down. Duration of drift down la Indeterminate and is largelyon piloting technique. Drift down consistslow sink period during which fuel economy Is above the corresponding cruise values for the same weight aa long as the actual altitude is above the scheduled cruise altitude. The difference In miles per pound can be neglected In planning and provides an operational contingency pad. Refer to Section IU for fuel management during single engine cruise.


resents single engine long range cruise performance for afterburner operation in terms of distance to go0 lbs gross weight0 lbs fuel remaining). The chart Is based on zero wind distance without turns at teat day conditions. Test EGT was trimmedor CIT range ofto The long range


Figure No. . 1



Turning Performance


Specific Range, Ambient Temp. 2

0 AS-



Long Bange6 ARDC Atmosphere

Fuel and Time16

ruise Intercept2


High Altitude6 ARDC Atmosphere

Fuel and Time17

ruise Intercept2

elling Cruise18

(With STD

Long BangeEAN TBOPIC Atmosphere

Fuel and Time19

ruise Intercept2

High Altitude CruiseEAN TROPIC Atmosphere

Fuel and Time1

ruise Intercept2


eiling Cruisef 2



Long RangeRDC Atmosphere

Fuel and Time1

ruise Intercept2


High Altitude6 ARDC Atmosphere

Fuel and Time1

ruise Intercept2


ellini Cruise1

Specific Range. Ambient Temp.


(With STD DAYZf>sz

Long RangeEAN TROPIC Atmosphere- -

ciJimhanVTlmC1 .



LiU of.lKuttraHortt (Con'f)

High AltitudeEAN TROPIC Atmosphere

Fuel and Time

ruise Intercept Points

Cruise Performance

eiling Cruise Profile

(With MEAN TROPIC climb)


Specific Range, Ambient Temp.


Long Range6 ARDC Fuel and Timeruise Intercept Points


Long RangeEAN TROPIC Atmosphere

Fuel and Time

ruise Intercept Points


Performance Mission Planning Factors for Supersonic Cruise

Rapid Deployment to6 ARDC Atmosphere

Profile of Rapid Deployment tu ARCP.


Rapid Deployment toEAN TROPIC

Profile of Rapid Deployment to ARCP

(MEAN TROPIC Atmosphere)



f 3

f 3

f 2

f 2


1 presents generalised turning performance at constant Mach numbers for various ambient temperatures and bank angles. Turn radius, distance, and time are plottedelected range of Mach numbers, ambient temperatures, bank angles, and degrees of turn.


0 turnorecast ambient temperature. 30 bank angle, of turn, find the turn radius, distance, and time. As shown In the chart,Qentsr1 at0mbient temperature and note that true0 knots. Proceed horizontally to 30 bank angle and read turn radius5 nautical miles. Proceed downward of turn and readautical miles flown. Proceed horizontally0 KTAS and read the turn timeinutes.


Specific range charts are presented for speeds of0 and for four ambient temperature conditions at each speed as shown by the llstrof illustrations. The data Is computed from Flight Test and Operational Testing results withngines. Correctionsange of bank angles are included on each chart to show the effect bank angle has on specific range and altitude capability while turning. scales provide KEAS-altitudeand fuel flow conversions.


Refer to, Specific Range data for0 cruisembient temperature. Locate the Max Range cruise schedule line. At long range cruise power0 pounds gross weight the cruise climb altitude0 feet and the aero bank angle specific range0b of fuel. urn at the same power setting,0 degree'bank angle, the specific range0b of fuel and the altitude0 feet. The fuel flow per engine0 lb/hr at sero bank0 lb/hr ategree bankmbient temperature day. At this temperature,orresponds7 KTAS as listed in the chart.


Long range cruise summaries are presented fornd High altitude cruise summaries are presented for0 The high altitude profiles are based onines shown on the Specific Range charts, except that theshown conforms with the0 ft altitude restriction. These data are presented for both6 ARDCand the "MEAN TKOpig"as shown in the list! of illustrations. The climb and cruise data are computed from Flight Test and Operational Testing results withngines.Descent data is based on Flight Test and Operational testing

at near standard temperatures. There are three sheets for each figure. The first sheet provides cruise summaries showing distance and time from end AR0 feet through the climb, cruise, and descent0 feet with0 lbs0 lbs of fuel reserve. The second sheetclimb-cruise intercepts which are to be used in conjunction with sheet 3. The third sheet presents performance and flight planning data. The initial conditlone shown are end AR0 feet, and brakewith0 lbs0 lbs fuel remaining UBing the normal climb schedule. The effect of variousIs shown for climb and cruise The descent performance shown is based on operational testing and does not Include the effect of temperature. DescentTropic" atmosphere may beby Increasing the presenteddata by the following increments:

0 miles



Use of the chart is illustrated by theexample:


Refer tondf 3.

Find the total distance capability and time requiredigh altitude cruiseorecast ambientt cruise. rofile is planned consistingeavyweight takeoff at sea level with standard day climb, cruise without turn, normal descent,0 lb fuel reserve0 feet. Planned fuel load at brake release0 lb.

Enterfb gross weight, sea level altitude^ standarc day climb temperature,ruise temperature and read the0 feet. Read climb distanceiles, climb time1 minutes

and fuel remaining0 lb. Referring tofhe intercept of the standard day climb line andruise line is shown. The lower portion ofhows cruise distance and cruise time to zero fuel remainingunction of fuel remaining and cruisetemperature. Entering the portion of the curve at the fuel remaining value0 lbruise, read the cruise distance5 miles and cruise time8 minutes. Then read on the cruise line (fromof0 lb descent line) the0 lb. Reading the distance and time to zero fuel remaining,iles and the time is 24 This gives the incremental cruise distance5 miles and the cruise time8 min-

itcs. The descent0 ftiles8 minutes as shown by the vertical scales at the right side of the profile portion of the chart.

Distance and time from brake release at sea level0 lb fuel0 feet0 lb fuel remaining is:

7 miles

7 minutes

J0fhow standard and tropic day missionfor five representative Mach numbers, and portray the climb, cruise andsegments of the missions. fhow the corresponding time and fuel remaining for the presented profiles.

ive thedetail information forlight of specific length. These curves present the overall mission time from brake release to ARCP, cruise Mach number, altitude to Initiate constant Mach climb, cruise altitude and the DTG to start deceleration to arrive0 feelointiles from the ARCP. 5 is the minimumcruise Mach recommended, as this speed Is the "break point" for minimum time between subsonic and supersonic flight plans. ission distance of lessiles, the flight should be made1 Mach. Missions longeriles would be Mown at the Mach number given.


To select flight plan for minimum time to ARCP, with Mean Tropic day temperatures, andiles from takeoff point.

Refer to, "Rapid Deployment to ARCP".

Mission time from brake release to ARCP5 minutes.


Start constant Macheet. Cruiseeet. DTG at startiles.


eiling Cruise summaries arc presented for00 aa shown in the list of illustrations. The data were calculated from Flight Teat andTeating results withngines. There are two sheets for each figure. The first sheet presents cruise summaries showing distance and time from end AR0 feet through the climb, cruise, and descenteet with0 lb*0 lbs fuel reserve. The second sheet presents cruise summaries which0 lb fuel remaining at altitudero distance and time). The initialshown are end AR0 feet and brake release0 lbs fuelusing the normal climb schedule. and time allowances for reserves0 lbs0 feet are shown in the charts. To obtain the total distance and time, add the twoand times for the desired profile.


Refer to,fnd tho example figure on the following page.

Find the total distance and time0 Machelling cruiseorecast ambient temperature ofat cruise.rofile is planned consistingeavyweight takeoff at sea level with standard day climb, cruise without turns,0 lb reserve0 feet. Planned fuel load at brake releaso0 lb. Enter,ft the climb line for the sea0 lb fuel remaining case and read distance and time9 nmi.5 min. Reenter at0 lb reserve descent line0 feet and read distance and timemi7 min. Add the distances and times and4 nmi2 min.

If forecast temperatures indicate standard day climb and cold day cruise,the distance will be increased by two small Increments. The cruise distance will be longer due to the colder temperature, and

the climb distance will bo longer due to the climb to higher altitude. Referring to the text illustration below, which isb gross weight0 lb fuel remaining at brake release, the shaded triangles show where the standard day climb Intercepts the four cruise lines. The cold day interceptistance5 nmi. Extend the climb curve to the altitude where the cold day cruise begins andistance5 nmi. The difference between these) la the increase in range due to cold day cruia* conditiona. The corresponding time incrementin. for themi of cruise. This resultsotal range and time9 nmi5 min.

mission planning factors table

A Mission Planning Factors Table Isonor quick reference in mission planning.

rapid deployment to arcp

resent the datainimum time profile from braketo ARCP.

The profile Is defined as:

ounds fuel remaining at brake release.

Normal climb schedule to cruise Mach number.

Climb to cruise altitude at constant Mach number.

Cruise for two minutes atPLA.

Normal decelerationEAS.

EAS descent to0 ftointiles from ARCP.

The data are presented for both6 ARDC and Mean Tropic atmospheres.



ndngines may becontinuously at all ratings when within the normalrmp?rature limits; however, nothan one hour may hewith EOT inthr normal limit achedule, and EGT muat be reduced immediately if an iiner<cncy limitia exceeded. (See EGT Limits and

| mow j

Continuous or accumulated opcr-ating time in the emergency EGT operating /one for more thaninutes may require engine


The nominal operating band, normal limits and emergency exhaust gas temperature operating schedules are prescribedunction of compressor Inlet temperature as shown in. Limit EGT'e foroperationhen comprea-?or Inlet temperature is above, andwhen CIT ia bslo* The aetting at which the red warning light on the EGT gage Illuminates and the fuel derichment system operates, if armed,alue which ia above ths normal operating temperature limit schedule.


At compressor inlet temp?ratures, the possibility ofstall exists at EGT'a between the maximum permissible value and the nominal operating band.

In the event that emergency engine ops rat ton is required, EGT may be increasedhen aboveIT. orIT: however, an accurateo( op'ratlnR time in the emergencyenne must be maintained.


. Any operation in or above the emergency operating soiespscial maintenance action.

. The permissible emergency EGT level at low CIT's is above the dench systempoint: therefore, the dericb system must beif this level is to be attained.


The maximum allowable compressor Inlet temperature. In addition,must be monitored so that engine cooling rates will not be excessive. While above an airspeed of, the aircraft maximum rate of descent should be such that rate of deceleration does notach In three minutes. There is noon rate of deceleration while below.


The minimum preaaure recommended for alratarts from stabilized wlndmllllng speec*si. This pressure la markedreen radial line.


Military and afterburning engine speeds are the same and are automatically scheduled by the fuel controlunction of Compreaaor Inlet Temperature. The normal scheduleahown by. Engine overapaed0 rpmisual inspection of the turbine. Notify the0 rpm is ever exceeded. Each instance of overspeedtng should be reportec as an engine discrepancy and should include the maximum rpm attained.



approved fuel IsE. pproved source of lubricity additive,, must be mixed with the fuel in the ratio9 gallons0 gallons of fuel. Fuelanday be used only for emergency requirements such as air refueling when atandard fuel ia not available and air refueling must beor risk loas of the aircraft. with emergency fuela should beto apeeds below.


The approved oil ia Ifbecause of low ambient temperatures, it may be diluted with Trlchloroethylene. Federal, Typen accordance with Maintenance Manual

Oil Prewjrs

Oil pressures belowal are unsafe and requireanding be made as soon aa possible, using minimum thrust required to sustain flightanding can be Normal oil pressurs Is fromosi. Except at IDLE throttleoil pressures betweenal andsi are undesirable and should be reported after flight. radually increasing oil pressure up toal is acceptable at high Mach numbers provided the indicationto normal values after aircraftto aubeonic apaad.

Oil Temperotmo

OH temperature must be at) prior to starting unless previously diluted with Trlchloroethylene. Engine oil temperaturesre unsafeanding should be made as soon as possible if the temperature cannot be maintained below this value. An engine should not be restarted after windmllling^ at subsonic speed when CIT is leas than) for moreinutes. Ifoperation above IDLE with OIL TEMP warning light illuminated shall be as brief as possible.

Maximum gross weight ia not limitedby takeoff performance capabilities. Baae maximum takeoff weights onprovided in Partf the Appmdix.


Maximum altitude with derichment inatalled and operational0 feet; maximum altitude without derichment0 feet.


(Refer toor the limit flight speed and altitude envelope.)


The stall warning light on the annunciator panel and the master caution light Illuminate when angle of attack reaches 14 Inone la also produced in the pilot's WhenIAS, the speed at which stall warning occurs is tbe minimum airspeed restriction for the existing vehicle weight,nd load factor unlessIs governedigher value ofKEAS as displayed by the TripleIndicator. Minimum airspeedEAS0 feet.


The Mach-airspeed Indicator limit hand is aet to Indicate airapeed (KIAS) correapond-ingEAS. However,EAS limit appliea only at altitudes0 feet, and at airspeed below.0 feet, limit airapeed decreaaea linearly with altitudeEAS0 feetEAS at sea level. Above, limit airapeed decreases linearlyEAS atEAS at. Seeor variation of KIAS with attitude for KEAS.


Maximum recommended operating apeeda are at leaatEAS laaa than limit airapteda. EAS) ia not recommended3 feel.







A red radial lineIAS represents the minimum subsonic 'peed restriction0 feet when the atall warning light is off.

To avoidafe angle of attack positive g's are limited'sach. (This la equivalent to5 bank level turn.)


The triple diaplay Indicator Is not marked however, the limit equivalent speeds are as follows unless:

Mach-alrapsed instrumentequals either the limitIndication or the minimumrsatrictlon.

atall warning light Illuminatesatall warning tone la heard.

Maximum TDI Airspeed

The limit airapeedEAS at eea level. Increasing linearly with altitudeEAS0 feet pressure altitude;EAS0 feet and the altitude for. Limit airapeed then decreases linearly with Mach numberEAS at. Normal operation cruise speedach.

Minimum TDI Airspeed

The minimum airapeed restriction varies linearly with Mach numberEAS0 feetEAS0 feet, and IsEAS0 feet


The maximum allowable positive load factor's In symmetrical maneuvers's In roll maneuvers aaby. The maximumload factor0 whenEAS varying0'a at higheras ahown by.


The aircraft shall be operatedanner to avoid full stalls, spina, and Inverted flight. Normal bank angle when operating aboveaegrees.


Ratea of descent muat be limited so as to maintain positive fuel tank pressure when austalned cruise speeds have exceeded,


The aircraft shall be operated. rangeAC while subsonic.. must be forwardAC for takeoff and should be aa nearAC aa possible with existing fuel for landing;.

The. limitAC while This limit results from stability consideration* at high Mach number. stability exlata at farther aft centers of gravity betweenndut for simplicity the aft limit is not changed. The purpoae of elevon trim limits Impoaed In this Mach region la to alert the pilotajor malfunction In the fuel aye-tem.

On those aircraft, if an. emergency exiata and EMERtransfer ia operated to place more0 Iba innd total fuel ia leaa0 Ibe. tha aircraft should be limited to maneuvera causing not more.

Aa elevon trim can be used as an indication of abnormal c. g. condition, the following pitch trim limits apply:

Whileo more thannoae down.



m m

iii iii

iii iii


m m




m fiSE

m in


ib sb


. I

* ?

* e

c 5


While nose down fromo nose down above.

Ai initial cruise the trim limit" nose down. As altitude increases and KEAS decreases,. trim limit becomes approximatelymore nose up perEAS decreaseEAS.

(In -idriition, expect approximatelymore nose up trim for each percent. Is forwardAC).


These limits to be supplied at the operating site.


Do notEAS oraximum of 5 sideslip with gear When sideslip angle exceedsoperation with gear extended is limited toEAS. Operation atspeed with gear extended is prohibited. The landing gear is designed for landing sink speeds at touchdown which decreasePS0 poundsPSounds. Side loads during takeoff, landing, and taxiing must be kept to aas landing gear side load strength is critical during ground maneuvering.




The SAS shall be on for all takeoffs and



The .tpike and forward bypass controls must be operated in the AUTO mode at all times when0 feet. When inlet controls must be operated manually,allowable speed Is.


The canopy shall be opened or closed only when the aircraft is completely stopped. Maximum taxi speed with the canopy open isnots. Gusts or strong winds should be consideredortion of thenot speed limit.

The maximum taxi speed recommended isnots forires. The rated ground speed limitnots. 0 feetnots correspondsIAS withambient temperaturealm day,mbient temperature. Limit indicated airspeed on the groundby the amount of tailwindalong the runway and increases by th? headwind component. Refer toor rated speeds at other altitudes and temperatures.

Taxiing Restrictions

A heat check is required for tires, wheels, and brakes:

Prior to takeoff when taxiing has exceeded one statute mile.

When continuous taxi distance hastatute miles.

When clear of th; runway after an aborted takeoffeavy weight landing.


stop capability

ound capacity dry ano hard runway

iotorwind, zero slope

roscmoun1 photdown


If required, cooling should be accomplished until ground inspection reveals that the tires and wheels are sufficiently cooled foroperations (temperatures relatively tolerable to the touchj.


Cooling may be accelerated by use of fane.

Abort during takeoff rollire change.


The one-stop energy rating of the brakesoot-pounds. Speeds corre-sondlng to these energy ratings from which stops can be made vary with gross weight, ambient temperature, altitude, wind, and whether or not the drag chute is deployed. Corresponding indicated airspeed, altitude, temperature, and weight conditions are shown byor the above rating for stops on dry runways with zero wind component. Headwind components may be added to values shown, and tailwlndmust be subtracted from values There is no limiation onfor brake application at normalweights when the drag chute is used. Refer to Part II of the appendix for detailed information related to maximum refusal speeds and heavyweight landings with various operating conditions.

Brakesew conditionapacity for one hard stop from rated speed. Ifsooner, they will burn out prior to stop. In normal operations, delaying brake application untilf limit speeds reduces wear and conserves brake capacity for high energy stops.

If brakeB chatter at slow speeds during taxi, turns, and at the end of the landing roll, light braking only must be used to avoid "walking the gear" and cyclic loads on the airframe structure.


The maximum speed for dragIAS.

LN2 QuonHly Indicator

The LN2 quantity indicator is markedellow caution arcnditers remaining. Approximatelyiters ofnitrogen are requiredormalfrom cruise altitude to landing. Initiated with less thaniters per system oriters total may deplete the

syetem and require use of emergency procedures for fuel tank pressurisatlon failure as lower altitudes are reached.


Updated autopilots now have no limitations when operating within the normal flightenvelope.


An autopilot hardover failure at speeds in excessEAS0 feet or speeds in excessEAS00 feet will cause excessive load factor if Immediate pilot corrective action Is not accomplished.


Plightsressure suifcusing the oxygen mask and regulator,are restricted to altitudes belownd speedsEAS.






Hi hi

iii Hi

= 1








in :n

ill in


IS! Ill


Sii iii

iii ii'

e 1





m m




m m

1 i

S3 :k


> :



^ ja-


B Si!



s; =

i i



r i




Normal ClimbRDC Atmosphere)

Phase I. Subsonic climb from brake1

Phase IA, Subsonic climb from off2

Phase II, Transonic

Phase m, Std Day AT,4

Std Day, Supersonic

StdT, SupersonicA3-6

Normal ClimbEAN TROPIC" Atmosphere)

Phase I, Subsonic climb from brakeA3-7

Phase IA, Subsonic climb from offA3-8

Phase II, Transonic9

Phase HI, Mean TropicT, Supersonic

Military Thrust Climb Performance, Std



Mean Tropic Day, Supersonic



Normal Descent Performance

EAS Descent Performance

EAS Descent

Single Engine DescentRDC Atmosphere)




Single Engine DescentEAN TROPIC" Atmosphere)




Single Engine Turning Descent



ProfUe Scheduled


1resent normal performance to cruise altitudes foroperation6 ARDC Atmosphere and "MEAN TROPIC" Atmosphericrespectively. The data Is computed from results of Flight Test and Operational Testing withngines. The climb is segmented In three phases and includes the effects of varying gross weights and air temperatures on fuel used, time, and s the subsonic portion of the climb from brake release at sea level0 feet0 Msch, Corrections for time, fuel, and distance are listed in the chart for takeoff* from other field ele-vaUone, Phase IA la tho subsonic portion of the cUmb from various refuel altitudes0 feet0 Mach. a the transonic acceleration oortlon of the climb0 feet0 Mach0 feet5 Mach utilizing the "dive through" technique. Phase IU la the supersonic portion of the climb0 feet5 Mach to the altitude at which cruise Mach number la first attained. Phase UIA la the constant Mach portion of the climb from the end of Phaae Ul to the altitude for start of cruise. The followingabulation of the average results of flight tests for Phase UIA.


Avg. Total Fuel Flow Lb/Min.

Long Range 0 Cruise

High Altitude 0

/ PPH/eng)



Obtain the time, distance, and fuelfrom brake release for takeoffield elevation0 feet0 Mach0 faettandard day. Fuel load at brake release0 lb afterground fuel allowances for normal ground operation. (See Appendix, Pari U, for ground allowance computation) Find the initial grosa weight at brake release by adding the zero fuel weight and the fual load remaining. If the zero fuel weight0 pounds, the Initial gross weightounds.

Enter1 at the Initial grosa weight at brake release and read fuel used, time and distance for07 minutes,6 nautical miles, respectively.

From the table In table in1 for0 foot field elevation, reduce time, fuel, and distance0ounds,autical miles. Therefore, fuel used, time, and distanca for07,0 the gross weight at the end oflimb. Enter3 at the recomputed gross weight and read fuel used, time andfor00 minutes2 nautical mUea. ofndesults in the fuel used, time, and distance to the start of Phaae Ul072 nautical. The recomputed gross weight for entering Phase UI wlil. Enter5 with tho recomputed gross weight and0 faet andead fuel used, time, and distance for Phase UI0inutes,autical mUes, respectively. Add aU three phases and obtain fuel used, time, and distance


part rrz

part m


mi,Climb Performance From Brake

RDC Atmosphere)

Phase I, Subsonic AJ-1

Phase II, TransonicA3-2

phase III, Std AT. SupersonicA3-3

Std Day, Supersonic

Std0 C AT, Supersonic

Normal Climb Performance After Air Refueling


Normal Climb Performance From BrakeEAN TROPIC" Atmosphere)

Phase I,7

Phase II, Transonic

Mean Tropic AT, Supsrsonlc ClimbA3-9

Mean Tropic Day, Supersonic

Mean Tropic DayC AT, Superaontc

Normal Climb Performance After Air Refueling


Military Thrust Climb Performance, Std


Std Day

Std Day




Single Engine Deacentax

Single Engine

Single Engine Turning


15hroughresent normal climb performance from brake release to cruise altitudes for supersonic operation6 ARDCand "MEAN TROPIC" Atmospheric

conditions, respectively. The data isfrom results of Flight Test andTesting withngines. The climb is segmented in three phases andthe effects of varying gross welgois and air temperatures on fuel used, time, and distance. s the subsonic portion of the climb from brake release at sea level0 feet0 Mach. Cor-





Avg. Total Fuel Flow Lb/Min.

rectlons lor time, fuel, and distance are listed in the chart for takeoffs from other field elevations. Phase Ii is the transonic acceleration portion of the climb0 feet0 Mach0 feet5 Mach utilising the climb and dive Phase III is the supersonic portion of the climb0 feet5 Mach to the altitude at which cruise Mach number is first attained. Phase IIXA Is the constant Mach portion of the climb from the end of Phase III to the altitude for start of cruise. The followingabulation of the average results of flight tests for Phase IHA.

FPM Power

Long Range 0


High Altitude 0 Max



Obtain the time, distance, and fuel required from brake release for takeoffield elevation0 feet0 Mach0 feettandard day. Fuel load at brake release0 lb after ground fuel allowances for normal ground operation. (See Appendix, Part II. for ground allowance computationind the initial gross weight at brakeby adding the aero fuel weight and the fuel load remaining. If the xero fuel weight0 pounds, the initial gross weightounds. Enter1 at the Initial gross weight at brake release and read fuel used, time and distance for0inutes,iles, respectively.

From the table In figure. for0 foot field elevation, reduce time, fuel, and distance0b,mi, respectively. Therefore, fuel used.

time, and distance for04, respectively. Recompute the gross weight at end of climb. Enter2 as the recomputed gross weight and road fuel used, time, and distance for Phase II0inutes, andautical miles, respectively. Recompute the gross weight at end of Phase II. Enter4 with the recomputed gross weight and0 feet andead fuel used, time, and distance for Phase IU0inutes,autical miles, respectively. Add all three phases and obtain fuel used, time, and distance03 minutes,autical miles, Fuel remaining0 feet0,

Service allowances and/or allowances for deviations from the normal climb schedule can be applied to an affected phasewhenre-qulred. (Forubsonic cruise operation prior to reaccelerating might be scheduled In the flighthe effect of such an allowance must be accounted for when computing the initial weight to be used for the next phase of the climb.


6resent normal climb performance from the end0 foot refuel (refueling with one AB on) to the altitudes at which cruise Mach number is reached6 ARDC Atmosphere and "MEAN TROPIC" Atmospheric conditions, respectively. Adjustments which should be used for otherltitudes are listed in the charts. The data Is computed from Flight Test and Operational Testing results withngines. The assumed fuel load at the end0 lb. Phase IHA results arc Identical to the tabulation In the previous discussion.




Turns during climb arc nol recommended, however, if mission requirementsurn, compensation for range lost due to the turn must be Included In the flight plan. For example,5 heading change0 bank at an initial altitude0 feet.

To minimise the rate of climb loas due to turning, the recommended procedure Is to advance power touring the turn and maintain the speed scheduleEAS. Resume tha normal climbon completion of tho turn.

Comparison of straightaway climb andclimb on time, fuel and distance results In an overall range losa ofiles0 profile. On completion of the 45 turn,0 feet; time, fuel andto that altitude will be5b,l greater thanormal climb with no turn.


resent Military climb performancechedulenots equivalent airspeed (KEAS)eet0 Mach number when higher altitudes are attained. This power and speed schedule provides the most climb distance for the fuel consumed whencruising flight plans, such as for ferry or buddy missions, are uaed.


rind the time, distance and fuel required to climb0 feet from S. L.ay with an Initial gross weightb. Enter0 ft, andb Initial groin weight6 miles0 lbs. Adding takeoff allowances reaults in time, distance and fuel values5 +

20or climb from sea level0 feet.


Find the time, distance and fuel required to climb0 feet0td day with an Initial gross weightounds. Enter0 feet andound Initial gross weight;ilesounds. Reenter0 feel and an adjusted Inllial gross weightounds;0 miles0 pounds. Adding takeoff allowance* and subtracting values for climb from sea level0 feet results in time, distance and fuel values8)0) for climb from takeoff00 feet.


On coursr descent performance is shown or.,. resents descent performance for theEAS schedule. I9EAS descentwith forward bypass doors In the automatic and open positions respectively.


resent single engine descent performance0 feet and7 KEAS). The data Is based on flight test with tha Inletas listed in the charts. Time, distance, and fuel required are plotted versusushover at constant Mach is required to increase airspeedEAS toEASEAS schedule. Better range Is obtained whenEAS schedule la used, reducing airspeedEAS while maintaining constant altitude. These effects are In-


eluded kit lhe performance data. begins to decrease rapidlyfeet; therefore, the charts areto an altitude0 feet ao thatreduction technique can be usedresultant change In performance can The effect of'ri'l har. li.i1 Im.-h -

find by flight testing and is not included in the dala.

ummarUes the effect ofon Maximum AB descent performance for constant values,EAS. the effects of decreasing power at the index altitude0 feet for constant airspeeds,EAS. resents the effects turn at 35 bank angleEAS descent. 0 feet of altitude Is required to complete turn. For convenience in missionround track profile is also

Find the time, distance, and fuel required to descend on course0 feet and0 usingEAS descent schedule and Minimum AB0 feet. Enter0 feet and read the time, distance, and fuel required0 feetautical miles,0 pounds of fuel. Reenter at the final altitude0 feet on theAB line and read time, distance, and fuel requiredinutes,autical miles0 lb of fuel. Add the results and6autical miles,0 pounds of fuel.


Find the track time, distance, and fuelto descend0 feet and0 usingEAS descent0 turn Is to be completed0 feet, and Minimum AB is to be used0 feet. Enternd note on the penetration distance curve thatf turn Is completed0 feet altitude. Read time at that altitudeinutes and fuel used0 pounds. On the ground track profile note that the distance traveled Isautical miles. Enter0 feet (end of turn altitude) and read time, distance, and fuel required0 feetinutes,autical miles,0 pounds, Reenter at the final altitude0 feet on the Minimum AB line and read time, distance, and fuel requiredinutes,autical miles0 lb of fuel. Add the incremental readings and5miles,0 pounds of fuel.


3 minutes,autical miles, respectively. 0 feet0

Service allowances and/or allowances for deviations from the normal climb schedule can be applied to an affected phase when (Forubsonic cruise operation prior to reaccelerating might be schedule in the flighthe effect of such an allowance must be accounted for when computing the initial weight to be used for the next phase of the climb.


Obtain the MEAN TFOPIC day time,and fuel required from refuel0 feet0 Mach0 feet (start of. Enter8eet and read fuel used, time, and distance for Phase IA00 minutes,6 nautical miles, respectively. The recomputed gross weight for entering Phase II will.


Turns during climb are not recommended, however, if mission requirementsurn, compensation for range lost due to the turn must be included in the flight plan. For example,5 heading change0 bank at an Initial altitude0 feet.

To minimize the rate of climb loss due to turning, the recommended procedure la to advance power touring the turn and maintain the speed scheduleEAS. Resume the normal climbon completion of the turn.

Comparison of straightaway cUmb andclimb on time, fuel and distance results In an overall range loss ofiles0 profile. On completion of the 45 turn,0 feet; time, fuel andto that altitude will be5b,l greater thanormal climb with no turn.


resent Military climb performancechedulenots equivalent airspeed (KEAS)eet0 Mach number when higher altitudes are attained. This power and speed schedule provides the moat climb distance for the fuel consumed whencruising flight plans, such as for ferry or buddy missions, are used.


Find the time, distance and fuel required to climb0 feet from S. L.ay with an Initial gross weightb. Enter0 ft, andb initial gross weight6 miles0 lbs. Adding takeoff allowances results in time, distance and fuel values2)0) for climb from sea level0 feet.


Find the time, distance and fuel required to climb0 feet0 foot takeofftd day with an initial gross weightounds. Enter0 feet andound initial gross weight;ilesounds. Reenter0 feet and an adjusted initial gross weightounds;0 miles0 pounds. Adding takeoff allowances and subtracting values for climb from sea level0 feet results In time, distance and fuel values8)0) for climb from takeoff00 feet.


On course descent performance Is shown onI9. resents descent performance for theEAS schedule. EAS descentwith forward bypass doors in the automatic and open positions respectively.

AST ni


Single Engine Descent data Is presented for Military, Minimum afterburning andafterburning powerEAS6 ARDC and Mean Tropic Atmosphere conditions. Refer toB.

Allowance* For Deceleration To Descent Speodi

When cruisingigher KEAS than the desired descent schedule, the constantdeceleration le made at the same power setting as the constant KEAS descent. The constant Mach lines show the beginning point of the deceleration for each Mach number. In the situation where the cruise KEAS is less than the desired descent KEAS, the constant Mach descent is made with Maximum afterburning power. The constant Mach lines show the descent for different Mach numbers.

Comparison Of Descent Power and Speed Schedules:

The Maximum afterburning descent, as compared to the Minimum afterburning and Military power descents, resultsongeronger elapsed time and more fuel used. EAS descent astoEASlightly longer distance, less elapsed time and more fuel used. overall range resultsescent speedEAS Is used and If Military power le used in the descent and for cruise. There will be little overall range Iosb If either Minimum afterburning or Maximum afterburning descent power Is used as long as the cruise Is accomplished in Military power. The charts are indexed to0 feet soechnique of power or airspeed change can be used and the resultant effect after power change in performance can be determined. The effect of changing KEAS at the0 fee; has not been defined by flight testing and is not Iniluded inthe dala.


Wheningle enginewith the operating engine In Militsry power, the Mach rate limitach In three minutes will be exceeded.

Single Engine Turning Descent

resents the effects of aturn atbank angleEAS 0 feet of altitude is required to complete turn. For convenience in missionround track profile is also provided.

Sample Use Of Charts:

Find distance, time and fuel to descend0 feet0 feet, usingafterburning powerEAS. Initial speed is7 KEAS). (ARDC Standard) atmosphere conditions are expected. Refer to. Enter the chart0 feet and located the Minimum afterburning line for7 KEAS) condition, and readtime and fuel0 feet.

ilesinutes0 pounds


Enter the same chart0 ieet and read distance, time and fuel0 Ieet0 feet.

5 milesinutesounds

Add the above values to obtain distance time and fuel0 feetach0 feet in Minimum afterburningEAS.

iles3 minutes0 pounds


Find the track time, distance, and fuelto descend0 feet and0 usingEAS descent0 turn is to be completed0 feet, and Minimum AB la to be used0 feet. Enternd note^ on the penetration distance curve thatf turn ia completed0 feet altitude. Bead time at that altitudeinutes and fuel used0 pounds. On the ground track profile note that the dlatance traveled isautical miles. Enter0 feet (end of turn altitude) and read time, distance, and fuel required0 feetinutes,autical miles,0 pounds. Reenter at tha final altitude0 feet on the Minimum AB line and read time, distance, and fuel requiredinutes,autical mllaa0 lb of fuel. Add the Incremental readings and5autical mllea.0 pounds of fuel.



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Original document.

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