The E single-place and the F two-place high-performance, multipurpose tactical fighters are produced by the Northrop Corporation, Aircraft Division. In addition to twin-engine reliability, the aircraft are capable of supersonic flight. Similarity of operating procedures and flight characteristics will allow a pilot qualified in either aircraft to fly the other with minimum of training. The F rear cockpit is equipped with dual controls and instrumentation to allow the aircraft to be used as a pilot trainer or dual-piloted tactical fighter; however, minimum crew requirement is one pilot. Thrust is provided by two turbojet engines equipped with afterburners. An automatic auxiliary intake door on each side of the fuselage above the wing trailing edge provided additional air to the engines during takeoff and low-speed flight. The fuselage is an area-rule (coke-bottle) shape. The wing, horizontal tail, and vertical stabilizer are moderately sweptback. The F wing is “wing fences” to improve boundary layer control. Each wing is equipped with leading and trailing edge flaps used for takeoff, landing, cruise, and maneuvering flight. The maneuvering flap position provides automatic control of flap position by central air data computer (CADC). Deceleration equipment includes a speed brake under the central fuselage, a drag chute to decrease landing roll, and an arresting hook under the aft fuselage for runway arrestment. The tricycle landing gear has a steerable nosewheel and a two-position extendable nose gear strut used for takeoff. Flight controls are hydraulically actuated by two independent hydraulic systems equipped with artificial feel devices to stimulate feel to the pilot. The cockpit(s) are enclosed by manually-operated clamshell canopy(ies). Fuel cells are in the fuselage, with additional fuel carried in external tanks. The fire control system includes a fire control radar with search and range tracking capability, a lead computer optical sight, and a sight camera. Basic armament includes two 20mm guns in the nose (F left gun only), and air-to-air missile on each wingtip. Additional weapons consisting of various bombs, rockets, and flares are carried on five jettisonable pylons
|Aircraft Code||Radar||Optical Sight||Recon Camera||Windshield Rain Removal|
The aircraft is powered by two J85-GE-21 turbojet engines equipped with afterburners. Sea level, standard day, static thrust at military (MIL) power is 3200 pounds and at maximum afterburner (MAX) power, 4450 pounds. Air to each engine enters thru an air inlet duct on the side of the fuselage and is directed into the engine compressor section by a variable geometry system consisting of inlet guide vanes (IGV) and variable stator vanes. The variable geometry system reduces the possibility of the compressor stall. Compressor bleed air is used to provide anti-icing to the inlet guide vanes, bullet nose, and T2 sensor of the engine and pressurization to the radar waveguide, windshield and canopy seals, anti-G suit, and external fuel tanks from transferring fuel. Compressor bleed air also provides windshield and canopy defog, cockpit pressurization, and pressurization and cooling of the aft electrical bay and the forward avionics bay. The nine-stage axial flow compressor is coupled directly to a two-stage turbine. Exhaust gases from the combustor section pass thru the two-stage turbine section and are discharged thru a variable exhaust nozzle (VEN). An exhaust gas temperature (EGT) control system electohydraulically varies the opening of the VEN to provide overtemperature protection and maintain EGT within allowable limits in MIL and afterburner (AB) power ranges.
Normal airstarts can be made with doors failed in the open or intermediate position.
The throttle for each engine provides main engine control from OFF to IDLE, IDLE to MIL and afterburner (MAX) control from minimum to maximum afterburner operation. Each throttle controls respective engine fuel supply, fuel shutoff valve, main fuel control throttle angle and stopcock valve, main and afterburner ignition circuitry, engine speed, and afterburner control operation. The left throttle also controls crossbleed operation. The left throttle also controls crossbleed start valve circuitry. Fingerlifts on the forward side of each throttle (F front cockpit) provide a stop detent at IDLE. Raising the fingerlift permits retarding the throttle from IDLE, to OFF and the IDLE to MIL range, throttle friction is constant. A spring detent between MIL, and MIN afterburner must be passed over for afterburner or nonafterburner operation. Afterburner thrust modulation is provided throughout the afterburner range. Throttle friction in the afterburner range is slightly greater than that provided from IDLE to MIL position. Throttle friction is preset and not adjustable by the pilot.
The ignition system provides electrical ac power for starting either engine on the ground or during flight. The ignition system for each engine consists of an engine start button, arming circuits, 40-second ignition timer, and main and afterburner second igniters. AC power can be provided by an external electrical power unit, aircraft generator power, or aircraft battery powered static-inverter. Engine start buttons are provided in both F cockpits. With the battery switch OFF, the engine start button ignition circuits are inoperative. With the battery switch at BATT and the throttle at OFF, pushing the engine start button arms the ignition circuit and starts the ignition timer. The ignition circuit is completed to the main and afterburner igniters when the throttle is positions at IDLE. When the throttle is advanced from MIL into AB range, (with or without external power) the ignition circuit is completed to the main and afterburner igniters, starting the ignition timer for approximately 40 seconds. Afterburner ignition and timer operation may be discontinued at a time by regarding the throttle out of AB range. For ground starts only, the ignition duty cycle is: 3 attempted starts, 3 minutes off, an addition 3 attempted starts, and 23 minutes off.
The engine fuel control system meters the proper amount of fuel to the engine for optimum performance throughout the engine operating range.
The engine-driven main fuel pump is a combination boost and high-pressure pump mounted on the engine accessory gearbox. The main fuel pump also provides servo fuel pressure to the afterburner control servos and the afterburner shutoff valve.
The hydromechanical overspeed governor is provided to limit engines speed to a maximum steady state of about 106 % rpm if the main fuel control should fail.
Variable exhaust nozzle (VEN) operation is controlled by throttle positions and egt. When the throttle is advanced slowly to MIL, nozzle opening decreases toward 0% until approximately 85% rpm. At this point, the nozzle remains constant at a fixed cruised flat position (16% to 22%) until the throttle is advanced to where the nozzle starts to further close toward 0%. The engine delivers best cruise power performance with minimum fuel consumption when on the cruise flat. When the throttle is advanced beyond the cruise flat toward MIL rpm, the nozzle will continue to close until anT5 amplifier control to maintain egt within limits. This is called T5 modulation. Just prior to T5 modulation, the nozzle is still mechanically controlled by the throttle. A throttle setting just prior to T5 modulation will improve fuel consumption rates. When T5 modulation occurs, the nozzle will open slightly. During a rapid throttle burst from IDLE to MIL or MAX, the nozzle will close to 43% to 53%, and stay at that opening momentarily. Nozzle hesitation at this point during acceleration minimizes exhaust back pressure to provide rapid acceleration and to preclude compressor stall. The nozzle will then close to 0% to 3% until T5 modulation occurs. At high altitude, low airspeed, when a throttle burst from IDLE or cruise to MIL or MAX is made, the nozzle will open toward the 43% to 53% area, then close to approximately 7% to 12% to minimize rpm rollback and compressor stall prior to T5 modulation. During a throttle burst to AB range at low altitude, the main afterburner fuel flow is delayed by a sequence valve, momentarily causing the nozzle to pause (approximately 6% to 14% above MIL steady-state nozzle position) to allow afterburner pilot fuel to light off first; permitting a softer afterburner lightoff, thus reducing rpm rollback and compressor stall. In the event of engine overtemperature during nozzle modulation, the nozzle will open to approximately 28% to 38% to maintain safe egt operation. This nozzle position is known as the T5 lockout area. At high altitudes and low airspeeds, MIL nozzle opening may be larger and egt lower than observed at low altitudes and high airspeeds. During ground operation at MIL power, nozzle opening should be approximately 10%. As the throttle advances into the AB range, opening should approximate 25% to 50% in minimum afterburner, increasing to approximately 80% at maximum afterburner. Nozzle indication of 75% or higher indicates a full-open nozzle (nozzle-limited) condition. Under this condition, fuel flow to the affected engine will be reduced to maintain EGT within limits. If the T5 amplifier fails during MIL or AB power, retard the throttle to maintain egt within limits if flight conditions permit.
Engine Controls/Indicators omitted by author
The T5 amplifier system maintains a preset turbine discharge egt within allowable limits during MIL and AB power operation by varying the exhaust nozzle opening. Operation is automatic with ac power supplied by the engine tachometer generator. If EGT is higher than the reference temperature, the amplifier causes the VEN to open; if loser, the VEN will close. The system operates primarily in MIL and AB power ranges.
The T2 sensor and the T2 resistance-temperature-detector (RTD) are two engine components that indirectly control MIL/AB rpm and egt. The T2 sensor in the main fuel control repositions the three-dimensional ?? to schedule rpm, variable geometry system, and set the proper acceleration fuel flow schedule during throttle transients through the operational envelope. T2 temperature controls MIL/AB rpm. For example, as airspeed increases, T2 temperature increases and MIL/AB rpm will increase. When T2 temperatures decreases, as in a sustained climb, MIL/AB rpm will also decrease. With T2 temperature of -43° C and below, MIL/AB rpm may be as low as 90%. The T2 RTD biases the T5 amplifier at cold engine inlet temperatures to cut back fuel flow and corresponding egt to prevent compressor and turbine stresses.
Afterburner operation is initiated by advancing the throttle from MIL to AB range. Afterburner lightoff on ground should occur within approximately 5 seconds.
The engine-driven afterburner fuel pump is a single-stage centrifugal pump. The pump supplies fuel to the afterburner fuel control during afterburner operation. The afterburner shutoff valve, actuated by fuel pressure from the main fuel control, prevents fuel supply to the afterburner fuel pump inlet until the throttle is positioned in the afterburner range and the engine is operating at nearly military rpm.
The hydromechanical afterburner fuel control contains a fuel metering section, a computing section, and a VEN control section. Fuel is scheduled to the afterburner main manifold spraybars as a function of throttle position, compressor discharge pressure, and nozzle position, and to the pilot manifold spraybars as a function of compressor discharge pressure only.
Each engine has an independent, self-contained oil supply and lubrication system with a serviceable capacity of 4 quarts. The system consists of an oil reservoir, a lubricating and scavenging six-element pump, oil filter and bypass, and an oil cooler (oil-to-fuel heat exchanger) with a pressure-controlled bypass valve. Oil is pumped from the reservoir and delivered under pressure thru the oil cooler and the oil filter to the engine accessory drive gearbox, engine accessory drive gearbox, main bearings, and other internal moving parts. Oil is returned to the reservoir thru the scavenging system. A sump vent system maintains a positive pressure, making the lubrication system insensitive to altitude. Large oil pressure fluctuations and zero oil pressure may occur during maneuvering flight. (See section V, Operating Limitations.)
The fire warning and detection system provides a visual indication of a fire or an overheat condition in the engine compartment. When the system detects a fire or overheat condition, the fire warning light for the respective engine will come on.
An airframe-mounted gearbox is located forward and below each engine. Each engine-driven gearbox operates a hydraulic pump and an ac generator. Automatic gearbox shift occurs in the 68% to 72% engine rpm range.
Starting the left engine requires an external low-pressure air source for initial motoring of the engine. After starting the first engine, the other engine is started by using the same external air source directed by a manually operated diverter valve. With external ac power applied, batter switch in BATT position, and the engine motoring at 10% rpm or above, momentarily pushing the start button arms the ac-powered ignition circuit and permits the ignition timer to run for approximately 40 seconds. The ignition circuit to the main and afterburner igniters is competed and fuel flow starts to the engine when the throttle is advanced to IDLE. Without external ac power and the battery switch at BATT, a batter-powered static inverter will activate to provide ac power for engine start when the start button is pushed. After one engine has been started and the generator is on the line, the static inverter is automatically disconnected.
A crossbleed start capability without external air is provided for starting the right engine after the left engine has been started. Compressed air from the ninth stage of the left engine compressor section is used for initial motoring of the right engine. A crossbleed control valve installed as part of the left engine compressor ducting system is alerted for activation when the left engine throttle is advanced above 70% rpm. Actuation of the right engine start button will then open the crossbleed control valve, permitting air to flow from the left to the right engine. The right engine ignition circuit is then completed by movie the right throttle from OFF to IDLE position. In order to ensure an adequate flow of air for starting, the left engine should be operating at approximately 95% rpm. The crossbleed control valve closes and power is removed form the valve-open circuit any time the left throttle is below approximately 40 seconds after the right engine start button has been actuated.
If the throttle is at OFF, the airstart is accomplished by pushing the engine start button and advancing the throttle to IDLE, the same as ground starts. If the throttle is in the IDLE, to MIL range, alternate airstart is accomplished by advancing the throttle into AB range, which activates the engine ignition circuits to the main and afterburner igniters and allows the timer to run for approximately 40 seconds. If the throttle is in AB range, the throttle must be cycled to MIL and returned to AB range to activate the ignition circuits and timer; or a start may be obtained with throttle in AB range to activate the ignition circuits and timer; or a start may be obtained with throttle in AB range by pushing and holding engine start button until lightoff occurs. The battery switch must be at BATT to activate the static inverter when the engine start button is pushed to complete engine start. With no ac power, the batter switch must be at BATT to provide ignition when throttle is moved into AB range.
A compressor stall is an aerodynamic interruption of airflow thru the compressor. The stall sensitivity of an engine is increased by foreign object damage, high angles of attack at lock airspeeds and high altitudes, abrupt yaw impulses at low airspeeds (below approximately 150 KLAS), temperature distortion, engine anti-ice system in operation, and ice formation on the engine inlet ducts or inlet guide vanes. (See discussion in section VII, Adverse Weather Procedures.) Compressor stalls can also be caused by component malfunctions; engine rigged out of limits; throttle bursts to MIL or MAX power at high altitude and low airspeed; hot gas ingestion from other aircraft or during gun firing at high altitudes and negative g conditions; and maneuvering flight with landing gear down at altitudes above 30,000 feet. Variable inlet guide varies and variable stators have been installed in the engine to reduce the possibility of compressor stall. Operation is automatic as a function of engine rpm and inlet temperature. A P3 compressor dump system [T.O. IF-5-740] activates for approximately 16 seconds to reduce the possibility of compressor stall when a throttle is burst to AB range at intermediate or high altitudes; however, installed engines must also be modified with a connecting P3 dump system. During sustained maneuvering without throttle movement, increased stall margin can be obtained by positioning the throttle at 85% to 95% rpm.
Engines without p3 dump system may be installed in aircraft modified by T.O. IF 5-740, or engines with p3 dump system may be installed in unmodified aircraft. In either case, the P3 dump system will be inoperable.
Flameout may be caused by component malfunctions compressor stall and starvation, fuel contamination (water), fuel icing, engine inlet guide vane icing and by throttle transients outside the normal flight envelope. (See section VII) Improper recovery from an unusual altitude such as a high pitch altitude stall (see section VI) can produce an abrupt yaw at low airspeed, causing compressor stall and flameout.
The fuel system consists of three bladder-type fuel cells in the fuselage divided into two independent systems. Each cell contains a network of explosion and fire suppressant foam material. The forward cell supplies fuel to the left engine: the center and aft cells supply the right engine. Either system can supply fuel to both engines. Additional fuel may be carried in jettisonable external tanks. Fuel is transferred from external tanks to the internal systems thru the single-point manifold by air pressure supplied by the compressor ninth stage of each engine. Each internal system contains and individually controlled fuel boost pump, a fuel shutoff valve controlled by either the throttle or a shutoff switch, a fuel flow indicator, and fuel low and pressure caution lights. A dual-pointer fuel quantity indicator serves both internal systems. Fuel quantity and fuel flow indications are provided in both F cockpits. The internal system contains a 2-way semiautomatic fuel crossfeed balancing system controlled by the autobalance switch. A crossfeed switch and left and right fuel boast pump switches (F front cockpit) are provided to manually control crossfeet operation. The internal system contains a common vent to vent fuel vapors overboard at the vertical stabilizer trailing edge just above the rudder. Control of external fuel transfer to internal system is provided by external fuel transfer switches F front cockpit). An external tanks empty caution light (F both cockpits) indicates that selected external tanks are empty.
Rear cockpit fuel indicator lights provide indication of external fuel, boost pump, and crossfeed switch positioning. With fuel system in autobalance operation, the indicator lights will indicate left or right boost pump off and crossfeed on.
Two ac-powered dual-inlet fuel boost pumps provide fuel under pressure to the engine-driven main fuel pump, and during afterburner operation, to the engine-driven afterburner fuel pump. The left system boost pump is in the inverted flight compartment of the forward fuel cell, the right system boost pump is in the inverted flight compartment: of the aft fuel cell. Either boost pump is capable of supplying sufficient fuel to both engines throughout the IDLE to MAX power range with the fuel system in crossfeed operation. If both boost pumps are inoperative, sufficient fuel with flow by gravity to maintain maximum afterburner power from sea level to 5000 feet. Sufficient fuel may flow by gravity to maintain maximum power to 25,000 feet. Reduced power and flight at the lowest practical altitude for terrain clearance and emergency requirements will further assure continued ?? engine operation with boost pumps inoperative. With both boost pumps off or inoperative, crossfeed is not available and less usable fuel is available due to location of the gravity feed fuel inlet in each system.
The low-level volume-sensing float switch in each internal fuel system closes when the fuel level drops to approximately 350 to 400 pounds, depending upon switch positioning, indicating system tolerances and fuel density. A 10-second time delay relay is energized when the float switch closes. If fuel quantity level does not increase and open the float switch, the respective fuel low caution light comes on the autobalance holding solenoid for the opposite system is deactivated. For example, when the autobalance switch is at the left low position and the float switch in the right system closes, the autobalance switch will return to the center position.
JP-4 (NATO F-40) is the normal primary fuel; however, in areas where JP-4 is not available, JET A-1 with FSII (Icing Inhibitor) or JP-8 fuels (both designated F-34) can be used as the primary fuel provided the main and afterburner fuel controls on each engine are adjusted to provide proper minimum fuel flows and acceleration schedules. If JP-4 is used in engines adjusted for JET A-1 with FSII or JP-8 as primary fuel, then JP-4 is an alternate fuel. JET A-1 with FSII or JP-8 is an emergency fuel if used in aircraft with engines adjusted for JP-4 as primary fuel. JP-5 (NATO F-44) and JET A-1 without FSII (NATO F-35) are emergency fuels. See section III for airstart envelopes to be used with primary and alternate fuels and section V for limitations where using alternate or emergency fuels.
Fuel quantity data for the internal and external fuel systems is shown in figure 1-35. Quantities listed for 275-gallon external tanks are shown as minimum capacities for tank manufacture variances or restricted capacity refueling procedure. Actual total weight of fuel depends on the specific gravity of the fuel.
Internal differences in 275-gallon type external fuel tanks can cause fuel capacity to vary. Tank differences are caused by procurement from different manufacturing sources, internal modifications, or procedural restrictions for refueling certain tank configurations. Mixed tank configurations are possible throughout the inventory; therefore, actual total usable fuel quantity for each aircraft should be verified before flight.
Fuel balancing is required on each flight because the right (AFT) system has a greater fuel capacity than the left (FWD) system and because the engines may use fuel at different rates causing unequal fuel quantities During flight, check indicated fuel quantities against known or expected quantities at preplanned flight stages and check fuel quantity gages for proper operation with the FUEL & OXY switch (see figure 1-36 for location and function of controls and indicators). If a malfunctioning indicator is suspected or discovered, fuel quantity can be estimated by using available information such as opposite system quantity or by using fuel consumption vs time. Do no select autobalance or manual crossfeed operation to avoid possible dual engine flameout caused by fuel starvation.
Autobalance operation is initiated by pulling the autobalance switch out of detent and positioning it to the left or right low position corresponding to the internal system with the lower fuel quantity. The switch is held at the selected position by a holding solenoid. Selecting the left low position opens the crossfeed valve and reverse rotation of the left boost pump in permit fuel feeding from the right system to both engines. Autobalance operation ceases when: (1) fuel quantity indicator pointers are within 50 to 125 pounds; (2) the low level float switch in the system supplying fuel closes for longer than 10 seconds, or; (3) the crossfeed switch is positioned to CROSSFEED. When autobalance operation ceases, the holding solenoid is deenergized, allowing the autobalance switch to return to center, the low system boost pump will resume normal operation, and the crossfeed valve will automatically close (unless and crossfeed switch has been positioned to CROSSFEED). Maneuvering flight may produce fuel sloshing sufficient to affect fuel quantity indicator pointers and low level float switches and could cease autobalance operation prematurely. Autobalance operation will function normally with only one engine running (ac power available and both boost pumps operating).
Manual crossfeed is accomplished by turning the crossfeed switch on to open the crossfeed valve and turning off the boost pump switch of the system with the lower fuel quantity. When the fuel quantities of both systems indicated within 100 pounds of each other, the boost pump switch that is off should be turned on. After the pump has operated for a minimum of 2 minutes, turn the crossfeed switch OFF. If the switches are not repositioned after the systems indicate balanced, the systems will become unbalanced in the opposite direction.
If an internal fuel system has less than 650 pounds of fuel, the quantity of fuel falls below the fuel boost pump upper-inlet and the boost pump output is reduced approximately 40 %. During crossfeed operation, if the engines are operated at power settings requiring a fuel flow of 6000 pounds per engine per hour or greater, the low pressure light may come on and engine rpm fluctuations may occur because of insufficient fuel pressure. With a low fuel state (approximately 400 pounds), do not attempt to ensure fuel flow to both engines by selecting crossfeed operation with both fuel boost pumps operating. If the fuel supply in one system is depleted, or is pulled away from the boost pump by g-forces and the boost pump in the other system fails, air may be supplied to engines causing dual engine flameout. There is no cockpit indication of boost pump failure. With both fuel systems below approximately 400 pounds, autobalance operation is not available.
Autobalance operation should be used to maintain fuel balance until approximately 400 pounds remain in each system. Then, with both boost pumps operating, place the crossfeed switch to CROSSFEED and allow the engine to be fed from both systems simultaneously.
When external tanks are carried, use inboard tanks first, centering tank next, and internal fuel last. During ground operation, delay or stop transfer of external fuel when either the left system indicates 1700 pounds or more, or right system indicates 2300 pounds or more. When inboard tanks are empty (indicated when EXT TANKS EMPTY caution light comes on), check fuel quantity indicator for a decrease in quantity to assure that inboard tanks are empty. To transfer centerline tank fuel, turn off PYLONS fuel transfer switch and turn on CL fuel transfer switch. Failure to turn off the fuel transfer switch when inboard tanks are empty prevents EXT TANKS EMPTY light from indicating when the centerline tank is empty. The light will remain on until the switch is turned off.
Fuel balancing should be delayed until external fuel transfer is complete
Fuel may vent overboard if fuel level shutoff valves in the internal system fail while transferring fuel from external tanks.
Fuel venting during ground operation is a fire hazard.
If fuel venting occurs on ground or in flight, discontinue fuel transfer from external tank until fuel quantity indicator indicates less than total capacity. If in a climb, level aircraft and do not climb to a higher altitude until internal fuel quantities have been reduced.
The external stores may be salvo or selectively jettisoned from the pylons. The system consists of an aircraft battery-powered emergency all jettison and select jettison system, and (EXCEPT E-1 F-1) a one-shot thermal battery-powered emergency jettison system. Controls consist of an emergency all jettison button, a select jettison switch and button, and (EXCEPT E-1 F-1) an external stores jettison T-handle. Stores and pylons may be jettisoned on the ground or in flight with the gear up or down. See section V for jettison limits.
All jettison systems will operate regardless of the battery switch position, with or without ac power.
When the emergency all jettison button or the external stores jettison T-handle (if installed) is actuated, the system will jettison the outboard stores first, the centerline store 200 milliseconds later, and the inboard stores or empty fuel tanks 300 milliseconds later (or 800 milliseconds later for tanks containing fuel).
The centerline store, any wing store, or paired wing store (both outboard or both inboard) may be jettisoned individually. Only one release or paired release will occur for each actuation of the select jettison button. The released station or stations must be deselected before another store can be jettisoned. Sequencing logic provides priority to release centerline, inboard, outboard, and wingtips in that order. For example, with inboard and outboard selected the inboard will jettison and must be selected OFF before outboard can be jettisoned.
A single actuation of the select jettison button jettisons wing and centerline stores and also actuates the pylon jettison circuits. If pylons are jettisoned with stores, the stores will jettison from the pylons first followed by the pylons 1 second later.
Following an attempted release or jettison, any munition that does not separate from the aircraft should be considered armed and susceptible to inadvertent release during landing.
Pylons will jettison only if equipped with necessary hardware and explosive bolts.
Electrical power is supplied by two ac systems and one dc system. An external receptacle is provided for ac power input to the aircraft when the engines are not in operation. DC power is supplied by a battery and two 33-ampere transformer-rectifiers.
AC power is supplied by two 13/15 kva 320 to 480 Hz generators, one operating from each engine. Each generator functions independently and supplies 115/200 -volt three-phase power to the ac buses. Normally, power distribution is divided between the right and left systems. One generator will automatically assume the full load, except the corresponding aux intake door, without disruption if the other generator is off or inoperative. Generators cut in individually when each engine reaches approximately 48% rpm and should be on the line at engine idle. Generator dropout occurs at approximately 43% rpm.
Two switches placarded L GEN and R GEN are on the right vertical panel (F front cockpit) Generator caution lights, placarded L GENERATOR and R GENERATOR, on the caution light panel (F both cockpits) will come on any time the respective generator fails or is turned off. Each generator switch has a reset position, permitting the pilot to reset the generators if necessary.
DC power is obtained from each ac system thru a transformer-rectifier which converts ac to dc. A 24-volt, 11-ampere-hour (E-1 F-1 [T.O. 1F-5E-611] 13-ampere-hour) nickel-cadmium battery serves as a standby source of power for all dc circuits and is charged by the transformer-rectifiers. If one transformer-rectifier fails, the other continues to supply all dc power. A caution light placarded DC OVERLOAD (F E-1) on the caution light panel (F both cockpits) warns of a dc overload. See section III for dc overload emergency procedures.
The battery switch on the right vertical panel (F front cockpit) is a two-position switch placarded BATT and OFF. During normal flight conditions, the switch should remain in BATT position.
If the battery relay does not close when batter switch is placed at BATT, a normal start cannot be accomplished.
A static inverter, powered by the dc bus, converts 24-volt dc from the battery to 115-volt ac. The inverter, when activated, provides an alternate source of ac power for the following:
a. Engine ignition on the ground or in flight.
b. Operation of left engine instruments during start of left engine.
c. Fuel and oxygen quantity indicators.On the ground, with dc power only (battery switch at BATT), the inverter is activated when either engine start button is pushed or when the fuel and oxygen check (F front cockpit) switch is held at GAGE TEST or QTY CHECK position. In flight, with dual engine flameout (battery switch at BATT), the inverter is activated when either engine start button is pushed or either throttle is moved into AB range for engine restarts, or when the fuel and oxygen check switch is held at GAGE TEST or QTY CHECK position. In flight, with normal ac-dc power, operation of the static inverter can be checked by positioning the fuel and oxygen check switch to GAGE TEST and observing counterclockwise movements of fuel and oxygen quantity indicator pointers.
Hydraulic power is supplied by two independent systems, the flight control hydraulic system and the utility hydraulic system. Each system is powered by a positive displacement piston-type pump. The right airframe-mounted gearbox, drives the flight control hydraulic system pump, and the left airframe-mounted gearbox drives the utility hydraulic system pump. Both systems operate at 3000 psi. The flight control and utility hydraulic systems both provide the hydraulic power for the flight controls. In addition, the utility hydraulic system provides the hydraulic power to operate the landing gear, gear doors, speed brake, wheel brakes, stability augmenter, nosewheel steering, two-position nose gear strut, gun gas purge doors, and gun gas deflector doors.
The hydraulic pressure indicators on the instrument panel (F rear cockpit vertical panel) provide visual indication of hydraulic pressure in each system. See section V for indicator markings and pressure limits.
The landing gear system provides normal extension and retraction of gear, alternate extension of gear, nose gear strut hike-dehike, and nosewheel steering. The landing gear is extended and retracted by utility hydraulic system pressure electrically controlled by the landing gear lever (F both cockpits). Retraction time is 9 seconds with nose gear strut hiked and 6 seconds with nose gear strut dehiked. Gear extension time is 6 seconds. The main gear is held in the retracted position by individual uplocks hydraulically actuated. The nose gear uplock is contained within the gear dragbrace mechanism. All gears are held down by hydraulic pressure on the gear actuators and locked in the down position by spring-loaded overcenter downlocks. Three green lights, a red warning light, and an audible warning signal (beeper) heard thru the headset are provided to indicate when the landing gear is in a safe or unsafe position. A landing gear alternate release is provided in case of utility hydraulic system or electrical malfunction.
The nose gear strut can be extended (hiked) 13 inches of retracted (dehiked) on the ground by the nose strut switch outboard of the throttle quadrant (F front cockpit). Full hiking of the strut will add approximately 3 degrees to the pitch attitude, which shortens takeoff ground runs. The nosewheel is steerable in the hiked and dehiked positions; however, steering response may be slower during transit. Automatic gear strut dehike occurs anytime aircraft weight is off the main gear, regardless of the position of the gear lever. The strut fully dehikes before it enters the wheel well.
A landing gear alternate release D-handle (F front cockpit) permits gear extension with the landing gear lever up or down should the normal extension system fail. Pulling the handle deenergizes the landing gear hydraulic and electrical systems and releases the main gear uplocks, main gear inboard door locks, nose gear, and nose gear forward door to allow the landing gear to extend, assisted by gravity and airloads.
The landing gear downlock override button to the right of the landing gear lever enables the landing gear lever to be raised to the LG UP position while the aircraft is on the ground with the struts compressed. If the locking solenoid fails to release the landing gear lever from the LG DOWN position when the struts are extended, as after takeoff, the button can be pressed and held to allow the lever to be placed at LG UP.
The nosewheel steering system provides directional control and shimmy damping during ground rotation. With the nosewheel steering button pressed and held, nosewheel steering is controlled by movement of the rudder pedals. Nosewheel steering is available when the aircraft weight is on the right main gear. When the nosewheel steering button is released, the system provides viscous shimmy damping capability. Damping is effected by use of hydraulic fluid trapped within the nosewheel steering actuator and is not dependent upon utility hydraulic system pressure.
Each main wheel is equipped with a hydraulically operated multiple-disk power brake assembly. Brakes are operated by conventional toe-type brake pedals (rudder pedals) and use utility hydraulic system pressure to operate brake control valves. Proper brake disc operating clearances are automatically provided when the brake pedals are momentarily pressed hard while engines are running. Should the utility system fail, the brake valve acts as a brake transfer cylinder, and brake pressure is proportional to the amount of foot pressure applied to the brake pedal. After utility system failure, unlimited brake applications are still attainable.
The drag chute system consists of a 15-foot ring-slot deceleration parachute, packed in a development bag and stowed in an air-cooled compartment at the base of the rudder, and a T-handle (F both cockpits) to deploy the chute.
The drag chute T-handle on the instrument panel is mechanically connected to the drag chute release mechanism. To deploy the chute, the handle is pulled straight out (without turning) to the first stop (approximately 3-1/4 inches). Initial movement of the handle latches the drag chute to the aircraft. Further movement of the handle unlocks the compartment door latch, allowing the spring-loaded pilot chute to deploy and withdraw the drag chute into the airstream. The handle will lock in the deployed position. The drag chute can be jettisoned by turning the T-handle 90 degrees clockwise and pulling it out to the next stop (approximately an additional 3-1/4 inches). The handle is under spring tension during the final pull to jettison chute. When released, the handle will retract to the first stop. To stow, rotate the handle counterclockwise and push it in.
To avoid inadvertent jettisoning of the drag chute, ensure that handle is pulled to first stop and locked without rotation.
The arresting hook system is an emergency system consisting of a retracted hook under the fuselage aft section and a button (F both cockpits) to electrically release and extend the hook for runway arrestment. The hook is held in the up position by a lock assembly. A ground safety pin is provided to prevent inadvertent actuation on the ground and must be removed before flight. The gear lever must be down for hook to extend. For extension, the uplock is released electrically by pushing the arresting hook button. The hook then extends by torsion bar spring force, which maintains a positive downward force on the hook while a self-contained hydraulic dumping unit acts as a snubber to minimize hook bounce. Activation of the arresting hook button will illuminate the light in the button to indicate hook release, and will automatically dehike the nose gear strut, if biked.
An electrically-controlled hydraulically-actuated speed brake is located under the fuselage center section. The speed brake is powered by the utility hydraulic system and controlled by a three-position speed brake switch on the right throttle. The variable speed brake has a full extension of 45° without a centerline (CL) store and 30° with a CL store. After release or jettison of CL store, full speed brake extension is obtained by cycling the speed brake switch. High airspeeds may prevent full extension. The speed brake and horizontal tail are mechanically interconnected to minimize trim change during speed brake operation.
Positioning the switch aft opens speed brake (out); forward position closes speed brake (in). The center (off) position neutralizes hydraulic pressure. Intermediate speed brake positions can be obtained by short intermittent actuation of the switch. For the open and intermediate speed brake positions, the switch (F front cockpit) should be returned to center position after positioning speed brake. The speed brake switch in the F rear cockpit is springloaded to the center position.
The wing flap system consists of leading edge and trailing edge flaps used for takeoff and landing cruise, and maneuvering flight. The flap system provides an automatic flap setting (maneuver) and manual flap setting (up, emergency up, full and cruise). Automatic flap positioning is provided by determined by flap control selection. Each flap powered by an ac motor. The left and right leading edge flap actuators and the left and right trailing edge flap actuators are mechanically interconnected. Both the leading and trailing edge flaps are electrically interconnected and, in turn, are mechanically interconnected to the horizontal tail operating mechanism to minimize trim change automatically when the flaps are operated.
The flaps are controlled by either the flap lever on the throttle quadrant or by the thumb switch on t he right throttle. The flap lever has three placarded positions FULL, EMER (emergency) UP, and THUMB SW. The THUMB SW position transfers control of flaps to the thumb switch which has three placarded position UP, CR (cruise), and M (maneuver). The flap lever overrides all thumb switch positions when positioned at EMER UP or FULL. A flap position indicator on the instrument panel provides visual indication of flap setting.
The front cockpit thumb switch provides M, CR, and UP placarded positions. The rear cockpit has M and UP placarded positions and an unmarked spring-loaded center position, which allows thumb switch control of flaps from the front cockpit. Momentarily positioning the rear cockpit thumb switch control of flaps. The flap system remains in the setting selected by the rear cockpit thumb switch until either another setting is selected by the rear thumb switch, or another desired setting. However, holding the rear cockpit thumb switch in M or UP overrides any cycling or repositioning of the front cockpit thumb switch. Flap lever selected of EMER UP or FULL in either cockpit overrides any thumbs switch setting.
In the maneuver setting, the flaps are automatically positioned by signals from the central air data computer (CADC). Above 550 KIAS or .95 mach, the CADC prevents extension of the flaps by the thumb switch or, if the flaps are extended, initiates a steady audible warning signal. The audible warning signal may be silenced by retracting the flaps or pushing the warning silenced by retracting the flaps or pushing the warning silence button next to the gear lever. The flap position indicator and warning signal operation for maneuvering flap conditions are as follows.
F The audible warning signal may be masked by cockpit noise during low altitude, high speed flight.
Maneuvering flaps are used for takeoffs and landings and may be used for in-flight maneuvering.
The cruise (CR) setting, the flaps are at 0°/8°, which provides improved fuel consumption and buffet control when the aircraft is flow at reduced speed for maximum endurance with stores. Maximum range for all configurations and maximum endurance for a clean aircraft is obtained with flaps up.
The flight control system consists of an all-movable horizontal tail, ailerons, rudder, and a stability augmenter system. All control surfaces are actuated by dual hydraulic actuators, one powered by the utility hydraulic system and the other by the flight control hydraulic system. If either hydraulic system malfunctions, hydraulic power to the flight control system will continue to be available. Artificial “feel” is buil into the system, and electrical trim actuators change the relationship of the “feel” spring to the control stick.
The control stick incorporates a pitch and aileron trim button, weapon release button, gun trigger (F cooperative in rear cockpit), dogfight button (dogfight/resume search switch), nosewheel steering button, and a pitch damper cutoff switch. The nosewheel steering button may be used as an alternate microphone button during flight, with landing gear up or down.
The stability augmenter system (SAS) automatically positions the horizontal tail and rudder to damp out pitch and yaw oscillations and also provides manual rudder trim. With yaw damper off, rudder trim is inoperative and returns to neutral. The system is controlled by pitch and yaw damper switches and a pitch damper cutoff switch. The damper switches are electromagnetically held in the engaged positions and are spring-loaded to the off positions; and will disengage automatically in case of certain system malfunctions or loss of ac power. The CADC senses airspeed and determines the amount of control surface movement required. The aircraft can be safely flown without: augmentation throughout the entire flight envelope. However, augmentation improves handling characteristics and may be desirable for particular missions. The system can be disengaged at any time during flight and may be reengaged during flight provided the SAS limitations in section V are observed.
An aileron limiter, which is mechanically positioned by retraction of the landing gear, provides a spring stop which limits the aileron to one-half travel. To obtain full aileron travel of 35 degrees up and 25 degrees down, additional stick force must be applied to override the aileron spring stop. The aileron limiter is disengaged when the landing gear is in the extended position, allowing full aileron travel.
Maximum rudder deflection is 30 degrees either side of neutral with the landing gear extended or retracted; however, the amount of deflection during flight is a function of dynamic pressure force on the rudder surface and varies with airspeed and altitude.
Maximum E horizontal tail travel is 17 degrees up and 5 degrees down. Maximum E horizontal tail travel is 20 degrees up and 5 degrees down.
The pitot-static system supplies both impact and static air pressure to the CADC and the airspeed/mach indicator. The ultimeter and vertical velocity indicator receiver only static pressure from the system.
Aircraft are equipped with one of the following altimeters: AAU-7A/A pressure altimeter, AAU-19/A altimeter, or the AAU-34/A altimeter.
The AAU-7A/A is an aneroid altimeter which measures and indicates uncorrected pressure altitude based on static pressure inputs from the pilot-static system. A setting knob at the lower left of the instrument face adjusts the altimeter setting and altitude indications. The AAU-7A/a does not receiver corrected altitude inputs from the CADC.
The AAU-19/A altimeter indicates up to 80,000 feet, and is settable to sea level pressure from 28.10 to 31.00 inches of mercury. The drums indicate altitude in 10,000, 1000, and 100-foot increments in a three-digit display, with the last two zeros detected. A single multi-turn pointer rotates around the dial, which is graduated from 0 to 1000 feet in 50 and 100-foot increments. The altimeter has a primary and a standby operating mode. In its primary (servoed) operating mode, the altimeter displays corrected pressure altitude computed by the CADC. In the standby (STBY) mode, indicated by the STBY mode control level position and flag, the altimeter displays uncorrected pressure altitude. The standby mode automatically takes over in the event of CADC failure. Once the standby mode has taken over, the mode control lever must be moved to the RESET position to return the altimeter to the servoed mode.
The AAU-34/A altimeter indicates up to 80,000 feet, and is settable to sea level pressures from 28.10 to 31.00 inches of mercury. Three drums indicate altitude in 10,000, 1000, and 100-foot increments in a five-digit display, with the last two zeros permanently displayed. A single multi-turn pointer rotates around the dial, which is graduated from 0 to 1000 feet in 20 and 100-foot increments. The altimeter has a primary and a standby operating mode. In its primary electrical (ELECT) operating mode, indicated by the ELECT mode control lever position and flag, the altimeter displays corrected pressure altitude computed by the CADC. In the standby pneumatic (PNEU) operating mode, indicated by the PNEU mode control lever position and flag, the altimeter displays uncorrected pressure altitude. The standby mode takes over automatically in the event of CADC failure. Once the standby mode has taken over, the mode control lever must be moved to the ELECT position to return the altimeter to the electrical mode.
The AVU-8 airspeed/mach indicator indicates airspeed in knots from 80 to 850 and in mach number from .5 to 2.2 and is driven by the pilot-static system. The indicator include a maximum allowable airspeed pointer (red) and an index setting pointer. The setting pointer is controlled by a knob in the lower right corner of the instrument.
The CADC converts raw air data inputs into computed outputs. The CADC is equipped with a monitoring system which continually monitors the computing functions. Should a malfunction of failure occur, the AIR DATA COMPUTER, light on the caution light panel will come on. However, failures within the pitot-static system may cause erroneous inputs to the CADC that are not indicated by caution light illumination.
The angle-of-attack (AOA) system consists of a vane transmitter mounted on the fuselage and AOA indicator and indexer in the cockpit (F both cockpits). With landing gear down, the system automatically provides angle-of-attack information thru displays on the AOA indicator and indexer. With landing gear up, angle-of-attack information is displayed only on the indicator. AOA transmitter information is also provided to the CADC for use by the optical sight system.
The AOA indicator is calibrated in units from 0 to 30 and operates in all phases of flight. The on-speed index on the face of the indicator is set at approximately 3-o’clock position (15.8 units), which is the optimum angle-of-attack for normal landing approaches with gear and flaps down. Each F indicator has a maximum rate-of-turn index set at 21 units. When electrical power is removed from the AOA system, an OFF flag will appear on the face of the AOA indicator.
Each cockpit has an AOA maneuver mode switch on the left trim panel, placarded ON and OFF and springloaded to the center (neutral) position. Momentarily placing either switch to the ON position, with the landing gear up, will activate both front and rear indexer lights. The indexer lights can be turned off by moving either switch from the center position to the OFF position while the gear is up. Placing the gear down overrides the maneuver mode switches to provide continuous landing approach AOA indexer indications.
The attitude and heading reference system (AHRS) includes the attitude sensing and indicating subsystem, the heading and navigation subsystem, and the standby instruments. The attitude sensing and indicating subsystem consists of an attitude indicator (AI) or attitude director indicator (ADI) [T.O 1F-5E-611], and rate switching gyro to control and coordinate functioning of the subsystem. The heading and navigation subsystem consists of a horizontal situation indicator (HSI), a compass switch, and magnetic azimuth detector and compensator. Standby instruments consist of the standby attitude indicator and the magnetic compass. A power cutoff switch behind the headrest (F rear seat headrest), used for ground maintenance system check, controls aircraft electrical power to the AHRS and must be positioned at ON for flight.
The ARU-20/A attitude indicator (AI) (F both cockpits is gyro-stabilized to show aircraft pitch and roll attitude. The attitude sphere is stabilized by the displacement gyro (two-gyro platform) powered by the left ac bus and the dc bus. The AHRS rate gyro balances electrical inputs to the displacement gyros so that the attitude sphere maintains position thru all aircraft maneuvering. The AI can be tumbled by power interruptions which cause an OFF flag to appear in the lower left of the indicator face. If power failure occurs in any flight condition other than straight and level, the AI may erect to a false vertical when power is returned. The FAST-ERECT switch on the instrument panel next to the AI (F front cockpit) is provided to expedite gyro erection. When the switch is pressed and held the attitude sphere and the horizon bar on the radar indicator, when turned on, will erect. Inflight erection should be accomplished in straight and level flight. The attitude sensing subsystem provides pitch and roll signals to the fire control radar and roll signals to the lead computing optical sight.
The attitude director indicator (ADI) is gyro-stabilized to show aircraft pitch and roll attitude. The attitude sphere is continuously stabilized to maintain position thru all aircraft maneuvering. The rate gyro balances electrical inputs from the displacement gyros to provide sphere stabilization. The ADI can be tumbled by power interruptions which cause an OFF flag to appear at the lower left of the indicator face. If failure occurs in any condition other than straight and level flight, the ADI may erect to a false vertical when power is returned. The ADI also includes pitch and bank steering bars and localizer (LOC) and glideslope (GS) warning flags. The pitch and bank steering bars function to provide deviation indications when making instrument landing system (ILS) approaches. The LOC and GS warning flags provide warning of failures of the localizer or glideslope functions of the ILS. The ADI is powered by the left ac bus.
The HIS (F both cockpits) indicates heading, course, range to destination, bearing to selected ground navigation aids, course deviation, and system status. The instrument consists of an aircraft symbol, a compass card graduated in 5-degree increments, a bearing pointer, lubber lines (upper and lower), a course arrow, course selector window, course deviation indicator (CDI), TO/FROM indicator, course (CRS) set knob, heading (HDG) set knob, an OFF flag, and a Deviation/DF window. The Deviation DF window and OFF flag provide HSI status indications. The HSI is powered by the left ac bus.
With the compass switch at MAG, the magnetic heading of the aircraft is displayed under the upper lubber line, and the reciprocal heading is displayed under the lower lubber line. When the compass switch is in the DIRECT GYRO position, the heading displayed becomes a random heading.
If DIRECT GYRO is selected with the correct magnetic heading displayed at the time of selection, the heading will probably remain close to the correct magnetic heading, as the gyro has a very slow random drift rate. If DIRECT GYRO is selected when the compass card is not properly slaved to magnetic north, the compass card will be stabilized but will not indicate proper magnetic heading. In this case, the magnetic compass must be used for correct magnetic heading. When the course arrow is set, it will remain aligned (parallel) with the radial or localizer course selected, providing the compass card is slaved to magnetic north.
The bearing pointer indicates correct magnetic bearing to a selected TACAN station when the compass card is functioning in the MAG mode. If the compass card is not aligned with magnetic north, which is possible when in the DIRECT GYRO mode, the bearing pointer will still indicate magnetic bearing to a selected TACAN station.
The bearing pointer will not indicate proper relative bearing if the compass card is not slaved to magnetic north. With bearing pointer of compass malfunctions, the CDI may be used to find magnetic headings to a TACAN station; for this use, center the CDI with the TO indication and fly the course in the course selector window, using the standby compass.
With bearing pointer or compass malfunction, using the CDI to determine the magnetic course to a TACAN station should be attempted only as a last resort if unable to confirm position by radar.
The aircraft symbol is presented at the center of the HSI and is fixed relative to the instrument. Comparison of the aircraft symbol with the compass card, course arrow, course deviation indicator, and heading marker will give a pictorial view of the angular relationship between the aircraft and the displayed navigational information.
Then a TACAN channel is selected and with the compass in MAG mode, the head of the bearing pointer indicates magnetic bearing and the range indicator displays slant range to the station. When the course to the station is selected with the course set knob, a white triangle appears on the same side as head of course arrow (indicating TO), the CDI displays aircraft position relative to the selected course, and the Deviation/DF window will be blank. When ADF is selected, the bearing pointer indicates relative bearing to selected ground or airborne station. In this mode, the Deviation/DF window displays DF, the CDI centers, and the range indicator warning flag appears.
When the NAV MODE selector is positions at VOR/ILS and a VOR frequency selected, the navigation mode indicator displays VOR. With compass in MAG mode, the head of the bearing pointer indicates magnetic bearing to the station. When the course to the station is selected with the course set knob, a white triangle appears on the same side as head of course arrow (indicating TO) and CDI displays aircraft position relative to the selected course. When an ILS frequency is selected, the navigation mode indicator displays ILS and the bearing pointer on the HSI stows at approximately 4 o’clock. The CDI displays aircraft position relative to the localizer course.
The standby attitude indicator (ARU-32/A or ARU-42/A-1) is a self-contained indicator that provides a visual indication of the bank and pitch of the aircraft and should be used when the attitude indicator or AHRS fails. The pitch limits are 92 degrees in climb, 78 degrees in dive, and the roll capability is a full 360 degrees. Approximately 3 minutes are required to erect to true vertical after power is applied to the system. The indicator should be caged and set, if required, following engine start, and left uncaged for the remainder of the flight. The standby attitude indicator is powered by the 28-volt dc bus, or by the right ac bus (ARU-32/A before T.O. 1F-5E-586). When power is interrupted or the indicator is caged, the OFF-warning flag appears on the face of the indicator. Approximately 9 minutes of useful attitude information is provided after power failure.
The indicator may precess following sustained acceleration or deceleration periods and may tumble during maneuvering flight near the vertical.
Avoid snap-releasing the cage and trim knob after setting to prevent damage to the indicator.
A magnetic (standby) compass on the upper right windshield frame (F front cockpit) is provided for use if the primary navigation systems fail. Illumination of the compass is controlled by a switch on the compass mount when the flight instrument light control panel is turned on. Compass correction cards are in the holders on the right interior trim panel of the cockpit and rear cockpit pedestal (F EXCEPT F-1).
The intercom system provides headset amplification for the UHF radio, the radio-navigation systems, maneuvering flaps and landing gear audio warning signals, the AIM-9 missile tones, cockpit-to-ground crew, and cockpit-to-cockpit communications.
The COMM/NAV control transfer system allows transfer of cockpit operating control of either of both the UHF and navigation radio sets. The navigation transfer switch in the front cockpit and a UHF and navigation override switch in the rear cockpit. See figure 1-58 sheets 1 and 2 for location and function of controls.
Aircraft are equipped with either the AN/ARC-150 UHF radio or the AN/ARC-164 UHF radio. The UFG radio provides two-way voice communication at line-of-sight range. An interface with an AN/ARA-50 UHF/ADF provides direction-finding capability. Twenty UHF frequencies may be preset and selected by the preset channel selector control. The system includes a transceiver, a control panel (F both cockpits), an antenna selector switch (F front cockpit), and upper and lower antennas. Frequency range is 225.00 to 399.975 megahertz. A total of 7000 frequencies, spaced 25 kilohertz (.025 megahertz) apart, may be dialed by using the manual frequency selector knobs and windows. The right window contains a basic digit 0 thru 9, and at the 0,2,5 and 7 digits an additional 0 or 5 digit will appear to the right of the basic digit. The ARC-150 and ARC-164 radios operate in the same manner and are unchangeable.
Do not key ARC-150 or ARC-164 transmitter while changing frequencies; damage to the transmitter will result.
On aircraft equipped with antenna selector switch incorporating AUTO mode position, replacement of ARC-164 with ARC-150 radio will cause automatic antenna selection to operate improperly. Manual selection of UPPER and LOWER position is required. AUTO position may be placarded INOP (inoperative) when ARC-150 is installed.
The ARA-50 ADF operates in conjunction with the radio to provide bearing indication to any ground or airborne UHF station to which the radio is turned. Any frequency in the standard UHF communications band may be used. Relative bearing information is displayed on the HSI when the ADF position is selected ?? the radio control panel. For E-1 F-1 and T.O 1F-5E-611, ADF information is displayed on the HSI when the NAV MODE selector is at DF.
E UHF/ADF homing signals may be unreliable with landing gear in down position. F For UHF/ADF operation, the COMM and NAV control transfer switches must be selected to the same position (either FWD or AFT)
Aircraft are equipped with one of the following TACAN systems: AN/ARM-65, AN/ARN-84, or AN/ARN-118. The system provides bearing, range (DME), and course information to a TACAN ground (or airborne) station. TACAN information is displayed on the HSI. For E-1 F-1 and T.O. 1F-5E-611, the NAV MODE selector must be at TACAN on display information on the HSI. The system operates in the UHF navigation band and provides 126 channels. In addition, the ARN-84 and ARN-118 provide the capability of selecting an additional 126 channels and also have an air-to-air mode and self-test function.
The air-to-air mode (A/A or A/R TR position on the function selector switch) provides range to similarly-equipped cooperating aircraft out to 250 mm. Cooperating aircraft must select TACAN channels spaced 63 channels apart. Bearing information for the ARN-84 is not provided and the bearing pointer rotates continuously. The ARN-118 provides range to cooperating aircraft, and bearing and range to specially-equipped cooperating aircraft. In the A/A REC mode, bearing information is provided to specially-equipped cooperating aircraft. To obtain bearing to a cooperating aircraft, UHF/ADF can be used. Both the ARN-84 and ARN-118 provided a self-test capability however, the ARN-118 has an automatic self-test. When the TACAN signal becomes unreliable or is lost, the ARN-118 switches to an automatic self-test. Indications of the automatic self-test air:
a. TEST light will blink.
b. Range warning flag and OFF flag appear on HSI.
c. Bearing pointer slews to 270 degrees for 7 seconds.
d. Bearing pointer slews to 180 degrees, CDI centers, and TO indication appears for 15 seconds.
If the TEST light remains on after completion of the test cycle, the TACAN has malfunctioned. See figure 1-59 for location and function of controls.
The radar beacon encoder-transponder system (skyspot) provides increased tracking capabilities for the X-band ground-based radar. A three-position switch placarded SST-181 on the right vertical panel (F left console) provides selection of OFF, DOUBLE, and SINGLE pulse reply. A 10-position code selector installed in the encoder-transponder is preset by the ground crew before flight for code pulse spacing. If code position 1 has been preselected, the transponder will provide only single pulse coded replies regardless of the position of the switch.
The ARN-127 navigation system consists of a receiver, a control panel, VOR-localizer and glide-slope antenna in the upper vertical tail, and a marker beacon antenna in the lower center fuselage. The system provides VOR navigation, localizer, and glide-slope information to the ADI and HSI. The system operates on odd decimal frequencies from 108.10 to 111.95 MHz for ILS localizer and glide-slope information. Frequency range for VOR navigation information (displayed on the HSI) is the even decimal frequencies from 108.00 to 111.85 MHz and all frequencies from 112.00 to 117.95 MHz. See figure 1-60 for location and function of controls and indicators.
The ILS provides visual indications of glide-slop and localizer course. Paired localizer and glide-slope frequencies are automatically selected when the localizer frequency is selected. The ARN-127 navigation system operates in the ILS mode whenever the navigation mode selector is at VOR/ILS and ILS frequency is selected. The pitch and bank steering bars on the ADI indicate amount and direction of deviation from the localizer course and glide slope. Marker beacon passage is indicated by flashing of the green marker beacon light on the instrument panel.
When making an ILS approach with the antenna selector switch in the UPPER or AUTO position, the ADI pitch steering bar may fluctuate during UHF transmission. Selection of LOWER antenna position will eliminate this operational characteristic.
The IFF/SIF system is an airborne pulse transponder which receives coded interrogations from surface or airborne radar (IFF) and automatically transmits coded selective identification (SIF). It is not capable of interrogating other stations. The system operates in five modes and is capable of I/P (Identification of Position) and emergency identification. The modes are 1-Security Identity; 2-Self Identify; 3-Air Traffic Identify; 4-(Classified)-Security Identify (when installed); and C-Altitude Reporting. The equipment consists of a control panel (F front cockpit), a transponder (transmitter-receiver), an airborne test set/in-flight monitor, and an antenna switching unit (lobing switch) in the nose section. The receiver will respond only to interrogations in the selected mode and code. Mode 2 is preset into the transponder. An altitude encoder in the CADC provides an interrogating ground station with the aircraft altitude. Automatic altitude reporting is corrected pressure altitude computed by the CADC. The system does not automatically trigger an emergency code upon ejection. The emergency code can be manually selected.
Warning, caution, and indicator lights warn of failures critical to flight, hazardous or potentially hazardous conditions, or of a change in system status requiring awareness and possible action. The lights consist of two red FIRE warning lights, a red “gear unsafe” warning light in the landing gear lever, a yellow MASTER CAUTION light, a yellow ARREST HOOK down light, three green landing gear position indicator lights, AOA indexer lights and a caution light panel with 21 individual word capsules (yellow) for individual aircraft systems. A full set of warning, caution, and indicator lights is provided in both cockpits F. A WARNING test switch on the lighting control panel (right console) permits testing and lights and FIRE WARNING circuits. A three-position BRT/DIM switch, spring-loaded to the neutral position, allows a selection of bring or dim operating modes. Warning, caution, and indicator lights are powered by the dc bus in the bright mode and by the right ac bus in the dim mode.
The caution light panel contains 21 individual system word capsules, including seven (F E-1 six) spare capsules. Spare capsules illuminate only when the WARNING test switch is positioned to TEST. Each light when illuminated, except ENGINE ANTI-ICE ON, will remain on as long as the malfunction exists or the status is unchanged. The individual system caution light will not go out when the MASTER CAUTION light is reset to rearm the circuit. The ENGINE ANTI-ICE ON light will go on when the engine anti-ice switch is in the ON position. For functions of other individual caution lights, see the appropriate system description.
The master caution light must be reset after each activation to provide warning of subsequent activation of caution lights.
The warning test switch on the right console lighting control panel (F both cockpits) tests all warning, caution, and indicator lights in the cockpit as well as the landing gear audible warning signal, fire detection sensing loops, and angle-of-attack indexer.
When the test switches in both cockpits are actuated simultaneously, the fire warning lights and the landing gear audible warning signal will not come on in either cockpit. When the warning test switch is release in a cockpit, the fire warning lights in the other cockpit may illuminate momentarily.
F [T.O. 1F-5-826] When the test switches in both cockpits are actuated simultaneously, the fire warning lights will illuminate in each cockpit.
The aircraft is equipped with exterior and interior lighting. Exterior lights controlled from the cockpit lighting control panel consist of dual retractable landing-taxi lights, position (navigation) and fuselage lights, formation lights, and an anti-collision beacon. Interior lights consist of flight and engine instrument lights, console and panel lights, cockpit floodlights, thunderstorm lights (E only), and a utility light.
Two white landing-taxi lights, one under each engine nacelle, are electrically controlled, two-position, retractable lights. In flight with the gear down and NAV rheostat out of OFF detent, the landing-taxi lights will extend automatically to the landing (fully extended) position. The lights will retract and go out automatically when the gear is raised or the position lights are turned off. With the lights extended for landing, the LDG and TAXI LIGHT switch ON position illuminates the high intensity beam in each light. With the aircraft light on the main gear, the landing-taxi lights retract to the intermediate or taxi-light position and each beam is automatically reduced to low intensity for taxing.
Primary position lights are engine nacelle mounted (left-red, right-green). Auxiliary position lights are in outer wing panels (left-red, right-green). The tail position light is in the vertical stabilizer (white). Each auxiliary position light has an inboard white segment which illuminates the aft fuselage and vertical stabilizer for night formation flying. Position lights are controlled by the NAV rheostat. Two white fuselage centerline forward of the landing-taxi lights, come on steady-bright when the NAV control is placed in the FLASH position.
Formation lights, controlled by the FORMATION rheostat, consist of paired white dorsal lights aft of the cockpit and aft end mounted missile launcher lights (left-red, right-green) [T.O. 1F-5-736]. A rotating anti-collision beacon (red) in the vertical stabilizer is controlled by the BEACON switch. On unmodified aircraft, the missile launcher lights are not installed.
Left launcher trail formation lights will be removed when TDU-11/B target rocket is carried on left launcher rail.
Flight and engine instrument indicators on the instrument panel, right vertical panel, and right console are white-lighted by internal lamps. These lights operate off the right ac bus and are controlled by FLT INSTR and ENG INSTR rheostats on the lighting control panel.
Armament panel lights on the left vertical panel provide backlighting for the armament panel and optical sight. Controlled by the ARMT LIGHT CONTROL rheostat switch on left vertical panel, the lights operate off the left ac bus.
Console lighting includes integral backlighting in the consoles, pedestals, vertical, optical sight (unmodified aircraft) and radar panels, controls, illuminated legends and markings on controls, control panels, and instruction plates. Console lighting is powered by the left ac bus and is controlled by the CONSOLE rheostat on lighting control panel.
Floodlights in the cockpit provide while illumination of the instrument panel and each console panel. The floodlights are powered by the left ac bus and controlled by the FLOOD rheostat, from off to bright, on the right console, bypassing the Flood rheostat. The engine instrument rheostat must be out of the off position for the floodlights to be available.
The thunderstorm lights on each side of the cockpit rear bulkhead provide white illumination of the cockpit. The thunderstorm lights are powered by the left ac bus and controlled by the FLOOD rheostat on the right console. The FLOOD rheostat must be rotated past the 3 o’clock position for the thunderstorm lights to come on.
The utility light is on the right interior trim panel of the cockpit (F both cockpits). The light is controlled by a self-contained rheostat switch which can be rotated to turn the lamp on and vary the lamp intensity. A lens cap provides selection of red or which spot or floodlighting. In an emergency, pressing the pushbutton switch on the light assembly provides full intensity of the lamp and permits use as a signaling light when the pushbutton is intermittently pressed. The light, equipped with an extension cord, is hand portable and can be detached from its support to allow use anywhere in the cockpit. Auxiliary mounting supports are provided for relocation of the light, if desired (lower right corner of the cockpit windshield frame.)
Stow utility light after use to prevent interference with the ejection seat and possible inadvertent initiation of the manseat separation system.
A 5-liter liquid oxygen system supplies breathing oxygen. An oxygen regulator on the right console controls the flow and pressure of the oxygen and distributes it in the proper proportions to the mask. The oxygen regulator contains a gauge, a blinker type flow indicator, emergency flow lever, oxygen diluter level, and supply lever. Controls and indicators are provided in both F cockpits.
A combination pressure breathing diluter demand, oxygen regulator is used in conjunction with the oxygen mask. The oxygen system is controlled by the supply, diluter, and emergency levers. And interlock between the supply lever and diluter lever causes the diluter lever to trip to 100% position when supply lever is at OFF, preventing any flow of air thru system. Gaseous oxygen is supplied to the regulator in the range of 65 to 110 psi. The regulator reduces the oxygen pressure, mixes oxygen with air in varying amounts, depending on altitude and demand, and delivers it thru a flexible hose to the oxygen mask. At high altitude, the regulator supplies positive pressure breathing. System operation is indicated by the flow indicator and oxygen pressure gage on the oxygen regulator panel. The emergency lever should remain at NORMAL unless an unscheduled pressure increase is required.
When placing the emergency lever at EMERGENCY or TESY MASK, it is mandatory that the oxygen mask be fitted to the face and not removed. Continuous use of positive pressure with a leaking oxygen mask or the mask removed for extended periods will deplete the oxygen supply rapidly.
The cockpit (F both cockpits) is enclosed by a manually controlled one-piece clamshell type canopy. The canopy is counter-balanced throughout its travel limits. The canopy drive mechanism is protected against excessive loads by a hydraulic damper, which also restricts canopy opening and closing speeds. An inflatable seal in the canopy will inflate only when the canopy is locked and an engine is operating. Exterior and interior normal and jettison controls consist of locking handles and jettison handles and a canopy caution light. The exterior and interior locking handles must be used only to lock and unlock the canopy. Raising and lowering the canopy must be done by hand pressure applied to the canopy frame.
Damage to canopy drive mechanism may result if the locking handles are used to raise and lower the canopy.
The canopy jettison handle in the cockpit (F both cockpits) is safetied by a removable safety pin. After the pin is removed, a spring clip which safeties the handle must be overridden when the handle is pulled. See figure 1-66 for location of controls and caution light.
The F-5E is equipped with either the Standard or improved rocket catapult ejection seat. The F-5F (each cockpit), is equipped with the improved rocket catapult ejection seat. Both type seats include: a seat adjusting unit and control switch, an automatic-opening safety belt, shoulder harness, inertia reel locking lever, headrest, canopy piercer, calf guard, two leg braces, two catapult firing triggers, a jettison initiator, a survival kit container, a man-seat separator system, and a sequenced seat ejection system (F only). The Improved seat additionally includes a drogue chute, which stabilizes the seat (and pilot) during ejection. Either seat will eject thru the canopy if canopy jettison fails. See section III for ejection envelopes and escape parameters.
Leg braces with handgrips incorporating firing triggers are interconnected and attached to the seat. Raising the leg braces to the fully up and locked position with the handgrips kicks the shoulder harness (E only) and exposes the firing triggers. After the leg braces have been raised to the locked position, they cannot be lowered to the stowed position.
With the seat fully down and the leg braces raised, space between the firing triggers and consoles is severely reduced.
An inertia reel lock consisting of a reel (F gas-driven power reel) and cable attachment provides mechanical locking and unlocking of the shoulder harness controlled by an inertia reel lock lever. With the harness locked, (LOCK position) any slack remaining in the harness can be reduced by sitting back in the seat. The slack will then be reeled in to assume a new locked position. When unlocked, (AUTO position) the harness is free to reel in and out. A rapid acceleration of 3g or more will automatically lock the reel and keep it locked until the lock lever is cycled. In the E, when the handgrips are raised, the shoulder harness is locked. In the F, when the firing triggers are squeezed, the power-reel is actuated causing the shoulder harness to be forcibly retracted and restrained by gas pressure, regardless of the position of the lock lever.
The ejection seat is equipped with an HBU safety belt. The belt incorporates a 1-second (.65 second in the Improved seat) delay initiator to provide automatic opening of the best during ejection. Use of the automatic-opening feature of the belt decreases seat separation and parachute deployment time, which reduces the altitude required for safe ejection. The buckle on the left half of the belt incorporates a rotary latch mechanism consisting of a belt latch, lanyard latch, interlock device, and a serrated manual release handle spring-loaded to the locked position. The interlock device prevents fastening the belt without first attaching the automatic parachute arming lanyard into the lanyard latch. Actuation of the handle is not necessary when manually attaching the lanyard anchor and connecting the right half of the belt. Full counterclockwise rotation of the manual release handle releases the lanyard anchor and the belt link. See figure 1-69 for proper connection and operation.
The man-seat separator is an inverted Y-shaped Web strap assembly routed along the back of the ejection seat. The upper end of the strap is attached to a gas-operated ballistic reel behind the headrest and the lower end of the straps are routed under the survival kit and attached to the forward edge of the seat bucket. During ejection, high pressure gas from the safety belt initiator activates the ballistic reel, which draws the Web straps taut, forcing the survival kit and pilot to separate from the seat.
The anti-G hose on the left side of the seat next to the headrest (figures 1-67 and 1-68) is held in the stowed position by a flexible spring. A spring-loaded dust cover on the end of the hose must be opened to insert the anti-G suit hose connector.
The ejection seat may be equipped with either the BA-22 or BA-25 personnel parachute. The ejection seat is compatible with either parachute; however, the BA-22 parachute equipped with a zero-delay lanyard must have the lanyard attached to provide a similar minimum altitude ejection (below 2000 feet AGL) capability (see section III).
The BA-22 automatic-opening parachute can be equipped with either an aneroid device incorporating a 1-second delay timer or a .25-second delay timer connected to the parachute arming lanyard. The BA-22 parachute with 1-second delay timer is also equipped with a zero-delay lanyard to the automatic-opening safety belt connects the parachute arming lanyard and timer. The zero-delay lanyard, connects the safety belt and the parachute ripcord to bypass timer operation. Major differences of the BA-22 parachute which affect ejection performance are:
The BA-25 automatic-opening parachute is equipped with an aneroid device incorporating a .25-second delay timer connected to a parachute arming lanyard. Connecting the parachute arming lanyard to the automatic-opening safety belt connects the parachute arming lanyard and timer.
Above a preset altitude, the aneroid will delay automatic opening of the parachute until the occupant free-falls to the preset altitude. At or below the preset altitude, only the timer function is required to deploy the parachute.
Canopy jettison followed by seat ejection is initiated by raising handgrips. This action exposes the catapult firing triggers and automatically locks the shoulder harness inertial reel. Squeezing either or both triggers jettisons the canopy, and seat ejection occurs .3 second later. Accompanying this action, the seat adjuster power cable and personal leads are disconnected, the calf guard is lowered into position, and the automatic safety belt 1-second delay initiator is activated. Following the 1-second delay the initiator fires; subsequent pressure buildup opens the safety belt and also actuates the man-seat separator, forcing the crewmember from the ejection seat. The open safety belt releases the shoulder harness straps but retains the parachute arming lanyard. With the zero delay lanyard hook stowed, the parachute arming lanyard arms the parachute aneroid and timer device as the crewmember separates from the seat. Above a preset altitude, the aneroid will delay automatic opening of the parachute until the crewmember free falls to the preset altitude. At or below the preset altitude, only the timer function is required to deploy the parachute. With the zero delay lanyard hook attached to the parachute ripcord handle, the parachute arming lanyard and zero-delay lanyard pull the parachute ripcord. See section II for proper connection of the zero-delay lanyard and to section III for the proper use of ejection equipment.
The zero-delay lanyard must be disconnected and stowed when operating at high altitudes to permit the automatic parachute aneroid and timer to function.
The Improved seat ejection sequence functions in basically the same manner as the Standard seat, except that the automatic safety belt .65-second delay initiator is activated during seat/aircraft separation. After the seat has left the cockpit, the drogue chute deploys to stabilize the seat, and the safety belt initiator fires, opening the safety belt and actuating the man-seat separator. As the crewmember separates from the seat, the parachute arming lanyard arms the parachute aneroid and timer device. Above a preset altitude, the aneroid will delay automatic opening of the parachute until the crewmember free-falls to the preset altitude. At or below the preset altitude, only the timer function is required to deploy the parachute. When the BA-22 parachute is used and with the zero delay lanyard hook attached to the parachute ripcord handle, the parachute arming lanyard and zero-delay lanyard pull the parachute ripcord. See section II for proper connection of the zero-delay lanyard and to section III for the proper use of ejection equipment.
The zero-delay lanyard must be disconnected and stowed when operating at high altitudes to permit the automatic parachute aneroid and timer to function.
The F is equipped with a sequenced seat ejection system for automatic or manual ejection of either the front or rear ejection seat, independently or to sequence. Seat ejection sequence is determined by the positioning of an ejection sequence selector on the rear cockpit pedestal (figure 1-68) and whether the ejection is initial in the front or rear cockpit. A forcible pull of the selector is required to select either of three positions: SOLO, NORMAL, or DUAL.
With the sequence selector at SOLO, no automatic ejections sequencing is provided. The ejection must be initiated separately for each seat. Squeezing the firing trigger(s) jettisons the canopy and retracts the shoulder harness. The seat will eject .3 second after firing trigger squeeze. With SOSO selected, the two crewmembers in the aircraft, the rear seat should eject first. The front seat should initiate ejection 1 second after rear seal ejection.
With the ejection sequence selector in SOLO position and both cockpits occupied, intercockpit coordination is required to avoid seat collision after ejection.
With the selector at NORMAL, ejection sequence is determined by the crewmember initiating the ejection when the firing trigger(s) are squeezed. If rear seat ejects first, both seats must eject independently. If the ejection is initiated in the front cockpit, the rear cockpit canopy will be jettisoned and the shoulder harness will retract, .3 seconds later, the rear sweat will eject. The front cockpit canopy will be jettisoned and the shoulder harness of the front seat will retract .45 second after the rear seat ejects. The front seat will eject .3 second after the shoulder harness retracts. If the ejection is initiated in the rear cockpit, only the rear seat will eject. The front cockpit crewmember must eject independently.
With the selector at DUAL, when ejection is initiated in either cockpit by raising the handgrips and squeezing the firing trigger(s), the rear cockpit canopy will be jettisoned and the shoulder harness will retract, .3 second later, the rear seat will eject. The front cockpit canopy will be jettisoned and the shoulder harness of the front ejection seat will retract .45 second after rear seat ejects. The front cockpit seat ejects .3 second after shoulder harness retracts.
To ensure positive selection of SOLO or DUAL positions, pull selector full aft and rotate beyond detent positions (override marking provided) and push selector full forward. Selector will automatically detent in selected position.
When ejection occurs, the seat adjuster power cable, the personal leads, and the sequenced ejection dual gas-coupling are disconnected, the calf guard is lowered into position, and the automatic safety belt .65-second delay initiator is activated. After the seat leaves the cockpit, the drogue chute deploys to stabilize the seat, and the safety belt initiator fires, opening the safety belt and actuating the man-seat separator. The open safety belt releases the shoulder harness straps but retains the parachute arming lanyard. The man-seat separator strap assembly is drawn ??, separating the crewmember from the seat. As the crewmember separates from the seat, the parachute arming lanyard arms the parachute aneroid and timer device. Above a preset altitude, the aneroid will delay automatic opening of the parachute until the crewmember free-falls to the preset altitude. At or below the preset altitude, only the timer function is required to deploy the parachute. With the zero delay lanyard hook attached to the parachute ripcord handle, the parachute arming lanyard and zero-delay lanyard pull the parachute ripcord. See section II for proper connection of the zero-delay lanyard and to section III for the proper use of ejection equipment.
The zero-delay lanyard must be disconnected and stowed when operating at high altitudes to permit the automatic parachute aneroid and timer to function.
A personnel locator beacon in the parachute harness, if installed, is used to locate a pilot who has ejected. The beacon transmits a signal on 243.0 MHz. Upon parachute deployment, the beacon will operate automatically when the actuator tab is snapped to the stud tab on the right main lift web of the harness.
The survival kit fits in the ejection seat and is attached to the parachute harness by web straps and quick-disconnect buckles. The forward section of the kit top is equipped with a seat cushion and the rear section provides support for a back type parachute. Depending on the local command desires, kit contents will vary and may include a life raft.
The standard survival kit must be manually released from the parachute harness following ejection or during emergency exit on the ground. After ejection from the aircraft, the survival kit is deployed by pulling the yellow emergency release handle on the right side of the kit. Pulling the handle up and backward releases the kit from the parachute harness, the kit will open, and the life raft, if installed, deploys and automatically inflates when the survival kit lanyard attached to the harness reaches full length. For emergency exit on the ground, pulling the yellow emergency release handle, with pilot’s weight on seat, releases the kit and lanyard from the parachute harness. Normal ground egress from the cockpit should be accomplished by manually disconnecting the two quick-disconnect buckles from the parachute harness.
The improved survival kit incorporates an automatic deployment feature which may be selected by the AUTO/MANUAL selector. This survival kit is for use with the BA-25 or modified BA-22 parachute. The kit will be automatically released during the ejection sequence or retained for manual release, depending upon the selected position of the survival kit AUTO/MANUAL selector. During parachute deployment, the parachute shroud lines pull the kit auto-release cable. If the AUTO/MANUAL selector is at AUTO, the kit auto-release cable pull causes an initiator cartridge to fire, and after a 4-second delay, the survival kit is automatically released. If the selector is at MANUAL, the cartridge is safetied and the kit must then be released manually by pulling the emergency release handle. When the kit is released, either automatically or manually, the quick-disconnect buckle/web assemblies separate from the kit, permitting it to open and fall away from the crewmember until the lanyard, attached to the parachute harness, is full extended. The lift raft, if included in the kit, automatically deploys and inflates.
For emergency exit on the ground, pulling the emergency release handle, with pilot’s weight on seat, releases the kit and lanyard from the parachute harness, regardless of the position of the AUTO/MANUAL selector. Normal ground egress from the cockpit should be accomplished by manually disconnecting the two quick-disconnect buckles from the parachute harness.
The environmental control system consists of the following: air-conditioning, pressurization, canopy and windshield defog, anti-g, air distribution systems, windshield rain removal (E EXCEPT E-1 only). All systems except anti-g suit, canopy and windshield seal system, hydraulic reservoirs, external fuel tanks, and radar waveguide pressurization are controlled by controls on the right vertical panel (F front cockpit). Air from the ninth stage of the compressor section of each engine is used to perform cooling, heating, conditioning, and pressurization functions. Either engine will provide sufficient air to operate the system in the event of engine failure. Check valves prevent air bleedoff to an inoperative engine.
Air is routed thru a heat exchanger, cooling turbine, and water separator before entering the cockpit area. Cockpit temperature is automatically or manually selected by a temperature switch. In the automatic mode, a temperature control valve automatically maintains the temperature level selected by the temperature knob. In manual mode, the temperature controller is inactive. Temperature is controlled by manual operation of the temperature switch until desired temperature is achieved. Manual mode should be used only if a malfunction occurs in automatic mode.
A pressure regulator automatically maintains the cockpit pressure differential schedule. Cockpit pressure altitude is indicated on the cabin pressure altimeter (F front cockpit). Static pressure ports on each side of the fuselage below the windshield area provide a static air pressure source reference for the regulator and safety valve. A pressure safety valve incorporated in the system automatically protects the cockpit from excessive high or low pressure and depressurizes the cockpit when the cockpit pressurization switch placarded CABIN PRESS is in the RAM DUMP position. Pressurizing air is supplied to the external tank system, anti-g suit system, canopy and windshield seal system, hydraulic reservoirs and radar waveguide.
The canopy and windshield are defogged by a mixture of bleed air and partially cooled package heat exchanger air that is directed thru ducting to the canopy and windshield surfaces. Defogging air temperature is independent of the temperature selected by the cockpit temperature knob, but is maintained with temperature limits by the defog temperature control valve and the defog temperature sensor.
Anti-G suit air pressure is routed thru a regulating valve to the anti-G suit. A flexible hose from the regulating valve to the anti-G suit passes thru a quick-disconnect fitting on the left side of the ejection seat to allow automatic disconnection upon ejection. The anti-G suit valve is on the left floor aft of the seat. The valve regulates air pressure to the anti-G suit to inflate the suit when positive G is encountered. The valve operates automatically and begins to function at about 1.75G exerting an increasing pressure as the G-load is increased. When the acceleration decreases below the valve opening G-setting, the valve closes and the suit deflates.
The cockpit air distribution subsystem provides air conditioning and pressurization airflow, and routes cooling air to two torso outlets on the left canopy frame. An additional torso outlet is provided on the right vertical panel of E-1. Airflow volume of the outlets can be adjusted or shut off by turning the outer opening and can also be adjusted directionally by tilting the outlet left-right or up-down. An outlet in the right lower rear bulkhead of the cockpit provides conditioned air to the floor area and is stationary and permanently open. In an emergency, the pilot can shut down the cabin conditioning and pressurization system by selecting the RAM DUMP position of the cockpit pressurization switch. The RAM DUMP position fully opens the pressure safety valve, opens a small ram air door in the left side of the cockpit to provide ambient airflow (E-1 opens a ram air valve to provide ambient airflow from an opening behind the right engine air inlet duct), and closes the air-conditioning shutoff valve.
The ram air door can be closed at any airspeed but cannot be closed at airspeeds above 400 KIAS.
The cockpit air distribution subsystem provides distribution of tire air-conditioning and pressurization airflow. conditioned air is delivered to each cockpit thru an air-conditioning torso outlet on the right vertical control panel and an outlet on the left canopy frame. Airflow from the torso outlet on the right vertical panel can be adjusted or shut off by rotating and/or turning. The canopy outlet can only be rotated for directional airflow. outlets at the forward end of the left and right consoles of each cockpit provide conditioned air to the floor area of the cockpits and are stationary and permanently open. In an emergency, the pilot in the front cockpit can shut down the cabin conditioning and pressurization system by selecting the RAM DUMP position of the cockpit pressurization switch. The RAM DUMP position fully opens the pressure safety valve, opens a ram air valve to provide ambient airflow from an opening behind the right engine air inlet duct, and closes the air-conditioning shutoff valve. A ram air valve in the forward avionics bay opens to provide cooling air as cabin air is discharged overboard thru the safety valve. This valve also automatically opens to supply cooling air whenever the aircraft is at or above an altitude of 40,000 feet.
On the ground, two ac-powered blowers circulate ambient air within the forward avionics bay when electrical power is on. When the canopy is closed, conditioned air from the cockpit area is discharged thru the cabin pressure regulator to the forward avionics bay. This conditioning maintains temperature limits in flight. The aft electrical bay is cooled by circulating conditioned air.
The windshield rain removal system is provided to improve forward visibility in rain. The system consists of a rain removal switch outboard of the throttles, windshield spray nozzles at the exterior base of the windshield, a pressurized rain repellent fluid container, timer, and solenoid valve in the nose compartment, and a system pressure gage in the nosewheel well.
Holding the rain removal switch momentarily at RAIN REMOVAL will provide approximately 1/2 second of system operation. Rain repellent fluid will squirt from the nozzles at the base of the windshield and react with the rain, spreading a transparent water-repellent film over the face of the windshield. One 1/2 second application will last approximately 10 minutes. More than one application may be required initially if windshield is dirty or rain intensity is excessive; thereafter, application is made as necessary to maintain clear visibility. Rewetting will start to occur at the lower outer corners of the windshield. If the rewetted area is allowed to advance toward the center of the windshield, subsequent application of rain repellent fluid may not allow recleaning of the rewetted area. Applications of fluid should be repeated as necessary to prevent the rewetted area from advancing toward the center of the windshield.
The system anti-ice system directs engine ninth-gage compressor hot air to the engine inlet guide vanes (IGV) T2 sensor, and the bullet nose of each engine. An electrically controlled engine anti-ice valve controls the flow of hot air to each engine. Both anti-ice valves are activated by an anti-ice switch on the right vertical panel (F front cockpit) and actuated by engine compressor discharge pressure. The switch has two positions: ENGINE and OFF. A caution light placarded ENGINE ANTI-ICE ON on the caution light panel illuminates when the switch is at ENGINE.
The engine anti-ice valves are normally closed until electrically energized and sufficient air pressure is received from the engine to open them. The valves open when the engine anti-ice switch is positioned to ENGINE. At high engine rpm, (below T5 modulation), a slight increase in egt can be expected when the system is operating. Thrust loss during system operation is approximately 9% at MIL power and 6.5% at MAX power. At MIL power, the opening of the anti-ice valve may produce an approximate 100 lb/hr decrease in fuel flow and a 2% increase in nozzle opening indication. The engine anti-ice valve will fail to the closed position of dc power is lost.
To check engine anti-ice system operation prior to flight, with throttle at 75% rpm, position ENGINE ANTI-ICE switch to ENGINE, and check for a slight rise in egt. Also check that ENGINE ANTI-ICE ON caution light comes on when switch is actuated.
The pitot boom, total temperature probe, and AOA vane contain electric heating elements for anti-icing. The pitot heater is powered by the right ac bus; the AOA vane and total temperature probe elements are powered by the left ac bus. Positioning the two-position pitot anti-ice switch on the right vertical panel (F front cockpit) to PITOT activates all heating elements.
For detailed description and operation of fire control radar, lead-computing optical sight, sight camera, gun and missile system, and armament controls, refer to the Aircrew Nonnuclear Weapons Delivery Manual, T.O.1F-5E-34-1-1. See Jettison System, this section, for description and operation of stores jettison controls. See section V for authorized store configurations and limitations.
The A/A37U-15 (Dart) tow target system can be carried for aerial gunnery. The system consists of an RMU-10/A tow reel pod on the centerline pylon and an adapter and launcher assembly on the left outboard pylon to carry, launch, and tow a TDU-10/B Dart target. A nylon rope is routed under the aft fuselage and the left horizontal stabilizer. The rope is suspended forward to the target and attached to the aircraft with cloth tape. Armament circuitry and switches provide controls for launching, towing, and freeing the target. Cable cutters in the tow reel can be electrically actuated to cut the tow cable. The tow reel pod and target carrier are not jettisonable. See Section V for limitations and the appendix for performance.
Refers to T.O.1F-5E-34-1-1 for operation of controls.
The rear cockpit may be equipped with an instrument hood for simulated instrument training flights. The hood is positioned on guides and is stowed behind the ejection seat when not in use.
The MXU-648 baggage/cargo pod is a modification of the BLU-1 unfinned and BLU-27 unfinned series fire bombs. Each pod has a lunged access door on the left side. Some pods have a removable tail cone for loading a variety of cargo in size and length. The cargo compartment contains a metal floor and a cargo tie-down system, which consists of straps and/or netting secured to permanent hooks installed in the floor.
The photoreconnaissance camera system is integrally mounted in the nose section. The system consists of four KS-121A 70mm cameras, a computer-junction box, camera cooling and a camera window defog lens, a camera control panel, and camera operate lights. The system provides high-resolution aerial photographic coverage of ground targets at a full range of speeds and at low to medium altitudes.
The camera compartment environmental control system controls the compartment temperature and directs defog air to the camera windows. The system draws cooling and defog air from the cockpit air-conditioning unit. The compartment is automatically cooled and the camera windows defogged when the cockpit air-conditioning system is operating and the camera mode selector is at TEST, RMT, or OPR. Temperature in the compartment is maintained between 80° and 90°, except for occasional transients to 120°F on hot days at maximum airspeed, low-level missions.
The cameras may be arranged in six basic arrays, depending on specific target and mission requirements. Camera usage in the basic arrangements is dictated by scale, ground coverage, environment (hostile), and type of target. The arrangements provide two trimetrogon arrays, two split-vertical arrays, one split-oblique array, and one left oblique array. Camera and lens usage is restricted to the six basic arrangements.
The KS-121A camera is an aerial photographic sequential, pulse-operated still picture type with three alternate local length lenses and shutter speeds of 1/250 to 1/4000 second, which are infinitely variable within that range. The film format is 70mm (2.25 inches) square. A light filter, integral light sensor, and a 200-foot film magazine with a capacity of 916 exposures are included. Lenses are 1.5-inch, 3-inch, and 6-inch focal length. Exposure is automatically controlled by the light sensor and automatic exposure control computer thru adjustment of lens aperture and shutter speed. Shutter speeds are automatically set by the automatic exposure control circuits. The shutter operates at 1/4000 second until the lens is completely open and then automatically adjusts down to the speed required, with 1/250 second the minimum shutter speed. Each camera is electrically connected to the computer-junction box. The computer-junction box and the cameras operate on 28-volt dc power. The computer-junction box controls and coordinates camera operation.
The camera control panel has a camera selector switch for each camera, a mode selector, an interval selector, a built-in-test (BIT) button with GO and NO-GO lights, a camera override switch, and four frames-remaining counters with reset controls.
A camera operate light for each camera on the instrument panel provides monitoring of camera operation head up. The green lights are numbered 1 thru 4 to correspond to cameras, camera selector switches, and frames-remaining counters. Each light comes on while the corresponding camera is operating. If the selected exposure interval is 1 second or less, the light will be on steady. If the interval exceeds 1 second, the light will pulse on with the camera for approximately 1 second each cycle.
If vertical stereo coverage is required, use the Vertical Stereo--60 Percent Overlap Coverage chart as a guide. Vertical stereo coverage is vertical photographs of the same target area taken from slightly different angles. When the stereo (overlap) area is viewed thru special stereo viewing equipment, targets show vertical development permitting more effective analysis. Determine the size of the target and the scale required in the photography. Enter the chart with the scale and largest size and determine first an altitude and lens focal length which will provide the required scale; second, the number of exposures required to cover the longest dimension of the target; and third, the interval setting required between exposures at your planned groundspeed. The length of your flight line to cover the long dimension required and the other dimension of the target will determine whether additional flight lines are required. Generally, a 20 percent side-overlap between flight lines is considered satisfactory.
Stereo coverage requires 60 percent overlap from the exposure to the next, so that only 40 percent of each exposure is actual “ground advance”. The intervals in figure 1-77 are based on the formula, so the intervals recommended will provide optimum stereo overlap coverage.
To provide stereo coverage of an industrial target area approximately 15,000 feet by 7500 feet at a desired scale of 1:10,000 go to figure 1-77. Determine that a 1.5 inch focal length lens at 1250 feet altitude, a 3-inch focal length lens at 2500 feet altitude, or a 6-inch focal length lens at 5000 feet altitude will satisfy your scale requirements.
Considering tactical and other requirements, you select the 6-inch lens. In the GROUND COVERAGE (SINGLE FRAME) column, you find that each exposure with this lens at this altitude will cover 1875 feet. With 60% overlap, each exposure will advance only 40% of the coverage, so that 1875 X .40= 750 feet is the ground advance for each exposure. Dividing 750 into 15,000 (the longest dimension of the target) discloses that 20 exposures will cover the target lengthwise. Add 1 exposure at each end of each flight line to allow for turning error and lineup. Allowing 20% sidelap (side overlap) for each flight line (80% of 1875) discloses that each flight line covers 1500 feet across the target. Dividing 7500 by 1500 shows that 5 flight lines are required. Selecting 420 knots ground speed gives an INTVL-SEC switch setting of 1.0 second and the complete result is:
|ABSOLUTE ALTITUDE||-5000 feet|
|INTVL-SEC Setting||-1.0 second|
|EXP PER FLT LINES||-2.2|
|NO. OF FLT LINES||-5|
Note: most figures/tables omitted by author because of time constraints. They will be added as time permits.
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