F-14D. FLIGHT MANUAL (2004)

NAVAIR 01−F14AAD−1

CHAPTER 1

Aircraft and Engine

1.1

AIRCRAFT

areăcarried on the movable portion of the wing. The fuel

system incorporates both in−flight and single−point ground

The F−14D aircraft is a supersonic, two−place, twin−

refueling capabilities. Aircraft general dimensions are shown

engine, swing−wing, air−superiority fighter designed and

in Figure 1Ć1. FO−1 shows the general placement of comĆ

manufactured by Grumman Aerospace Corporation. In addiĆ

ponents within the aircraft. A summary of aircraft limitaĆ

tion to its primary fighter role, carrying missiles (Sparrow

tionsăand characteristics are shown in Figure 1Ć2. Refer to

and/or Sidewinder) and an internal 20−mm gun, the aircraft

Chapter 4 for detailed information on operating limitations.

is designed for fleet air defense

(Phoenix missiles) and

ground attack

(general purpose and precision ordnance)

1.1.1

Aircraft Weight

missions. Armament and peculiar auxiliaries used only durĆ

ing secondary missions are installed in low−drag, external

The basic weight of the aircraft is approximately

configurations. Mission versatility and tactical flexibility are

43,735 pounds, which includes trapped fuel, oil, gun, and

enhanced through independent operational capability or

pylons. Consult the applicable Handbook of Weight and

integration under existing tactical data systems.

Balance (NAVAIR 01−1B−40) for the exact weight of any

series aircraft.

The forward fuselage, containing the flightcrew and

electronic equipment, projects forward from midfuselage

1.1.2

Cockpit

and wing glove. Outboard pivots in the highly swept wing

glove support the movable wing panels, which incorporate

The aircraft accommodates a pilot and RIO in a tandem

integral fuel cells and full−span leading−edge slats and

seating arrangement. To maximize external field of view,

trailing−edge flaps for supplemental lift control. In flight, the

stations within the tandem cockpit are prominently located

wings may be varied in sweep, area, camber, and aspect

atop the forward fuselage and enclosed by a single clamshell

ratioăby selection of any leading−edge sweep angle between

canopy. Integral boarding provisions to the cockpit and

20_ and 68_. Wing sweep can be automatically or manually

aircraft top deck are on the left side of the fuselage. Each

controlled to optimize performance and thereby enhance airĆ

crewăstation incorporates a rocket ejection seat that is

craft versatility. Separate variable−geometry air inlets, offset

vertically adjustable. A single environmental control system

from the fuselage in the glove, direct primary airflow to two

provides conditioned air to the cockpit and electronic bays

F110−GE−400 dual−axial compressor, turbofan engines

for pressurization and air−conditioning. Oxygen for breathĆ

equipped with afterburners for thrust augmentation. The disĆ

ing is supplied to the crew under pressure from an on−board

placed engine nacelles extend rearward to the tail section,

oxygen generating system with stored gaseous oxygen as

supporting the twin vertical tails, horizontal tails, and venĆ

backup. The cockpit arrangement provides minimum dupliĆ

tralăfins. The middle and aft fuselage, which contains the

cation of control capability, thereby necessitating two airĆ

main fuel cells, tapers off in depth to the rear where it

crew for safe flight.

accommodates the speedbrake surfaces and arresting hook.

All control surfaces are positioned by irreversible hydraulic

1.1.2.1

Pilot Cockpit

actuators to provide desired control effectiveness throughout

the flight envelope. Stability augmentation features in the

The forward station of the cockpit is arranged and

flight control system enhance flight characteristics and

equipped for the pilot (FO−3). In addition to three electronic

thereby provide a more stable and maneuverable weapons

displays for viewing tactical, flight, navigational, and ECM

delivery platform. The tricycle−type, forward−retracting

data, the pilot instrument panel also contains armament conĆ

landing gear is designed for nosegear catapult launch and

trols, flight and engine instruments. Engine controls, fuel

carrier landings. Missiles and external stores are carried from

management, auxiliary devices, autopilot, and communicaĆ

eight hardpoint stations on the center fuselage between the

tion control panels are on the left console. Display, power,

nacelles and under the nacelles and wing glove; no stores

lighting, and environmental controls are on the right console.

1−1

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 1Ć1.ĄAircraft Dimensions

1.1.2.2

RIO Cockpit

tions. The following are types of technical directives used in

this manual:

The aft station of the cockpit is equipped for the RIO

AAC

Aviation Armament Change

(FO−4). This instrument panel contains controls and three

electronic displays for the radar, tactical, and navigational

ACC

Aircrew System Change

flight instruments. Armament controls, sensor controls, and

AFC

Airframe Change

communication panels are on the left console. The right

console contains a navigational display, data entry unit,

AVB

Avionics Bulletin

ECMăcontrols, data−link controls, lighting controls, and the

AVC

Avionics Change

IFF panel.

AYC

Accessories Change.

1.1.3

Electronic Nomenclature

1.1.5

Block Numbers

Figure 1Ć3 is an alphabetical listing of the tactical,

The following production block numbers include the

communication, navigation, flight control, and instruments

indicated aircraft serial numbers (BuNo). Selected aircraft in

in the aircraft.

blocks 85 and 110 have been updated to create the F−14D/

block 170 configuration.

1.1.4

Technical Directives

Block No.

Serial No. (BuNo)

As technical changes are made to the aircraft, those that

160

163412 − 163418

affect aircraft operation or pilot and RIO need−to−know

165

163893 − 163904

operation will be incorporated in the appropriate sections and

listed in the Summary of Applicable Technical Directives in

170

164341 − 164351 and 164599 − 164604

the front of this manual. In some instances, technical

159592, 159595, 159600, 159603,

directives may be incorporated in the aircraft while it is still

159610, 159613, 159618, 159619,

on the production line. Check the Technical Directives

159628 − 159630

Section of the Aircraft Log Book for applicable modificaĆ

110

161159, 161163, 161166.

ORIGINAL

1−2

NAVAIR 01−F14AAD−1

F−14D AIRCRAFT CHARACTERISTICS AND LIMITATIONS

Aircraft Dimensions

Starter Cranking Limits

Length

62′

8.5″

Cross bleed

2 min continuous then 10 min off

Height (Tail)

16′

0″

Start Cart

5 min continuous then 10 min off

Wingspan @ 20_ wingsweep

64′

1.5″

Wingspan @ 68_ wingsweep

38′

2.5″

Idle

Wingspan in oversweep

33′

3.5″

Wing Area

565 sq/ft

RPM = 62

− 78%

OIL PRESS = 15

− 45 psi

EGT = 350

− 650_C

NOZ POS = 100%

Gross Weights

FF = 950 − 1400 PPH

HYD PRESS = 3000 psi

Empty A/C (w/crew & trapped fuel)

43,735 lbs

MIL (and above)

Catapult

76,000 lbs

Field takeoff

72,000 lbs

RPM = 95 − 104%

OIL PRESS = 25 − 65 psi

Min descent rate landing (350 fpm)

72,000 lbs

EGT = 780− 935_C

NOZ POS INFLIGHT:

Field landing (max 500 fpm ROD)

60,000 lbs

MIL = 3 − 10%

Carrier/FCLP land

54,000 lbs

Min AB = 5 − 12%

Max AB = 60 − 90%

T/O & Land Flap/Slat Transition Limits

Ground Start Malfunctions

AOB < 45_

HOT START:

890_C (will normally peak

ROLL SAS ON

@ 30 − 40% RPM)

Min 200′ AGL for flaps/slats up on takeoff

HUNG START: Hung RPM below 63% with rising EGT

Min 800′ AGL for dirty−up

WET START: No lightoff within 20 sec of

Min 180 kts for retraction of flaps

throttle to IDLE

12 units AOA for all transitions

225 kts

max flap/slat speed

INST Test

280 kts

max landing gear speed

RPM

96%

Landing & Braking

EGT

950 ± 10_C

(initiates engine overtemp alarm)

15 kts

min speed for antiskid operation

10,500 pph

145 kts

max wheel braking (51.0 k A/C)

AOA

18 ± 0.5 UNITS

165 kts

max wheel braking (46 k A/C)

W/S

45 ± 2.5

190 kts

max tire speed

FUEL

2,000 ± 200# (all tapes and windows)

145 kts

max E 5 engagement speed (69.8 k A/C)

LIGHTS:

L and R FUEL LOW

175 kts

max E−28 engagement speed

20 kts

max 90_

crosswind

component

60 kts

max canopy open

Oil System

Engine Speeds (RPM)

Normal Range

25− 65 psi

10% Ignition system becomes operative

Min (stabilized idle)

15 psi

30 − 46% Engine crank switch will not engage

Oil pressure light

11 psi decreasing

20% Min RPM with throttle at idle

off @ 14 psi increasing

30% Max RPM with pneumatic starter

30% RPM must fall below for generator

Pneumatic Pressure

auto−reset

50% Engine crank switch automatically shuts off

60% Generator comes on line

Min auxiliary canopy (3000 psi)

800 psi

50% Generator light illuminates (if RPM falling)

Min normal canopy (3000 psi)

1200 psi

60% Ensure engine crank off (osp/valve)

CHS

accumulator

1800 psi

− 78% Normal idle

FHS accumulator

1800 psi

Min emergency gear preflight

3000

psi

75 − 90% Auto throttle range

− 104% MIL & above

Min emergency gear extension

1800 psi

Wheel brake accumulat

or (2

gages)

1900

± 50

psi

107.7% Overspeed (chevrons flash)

Arresting hook dashpot

800 ± 10 psi

110.0% Engine secures (f

uel shutoff)

Figure 1Ć2.ĄCharacteristics and Limitations (Sheet 1 of 2)

1−3

ORIGINAL

NAVAIR 01−F14AAD−1

F−14D AIRCRAFT CHARACTERISTICS AND LIMITATIONS (CONTINUED)

Fuel System

In−Flight Refueling

Aft−Left Tank Group Capacity

5900 − 6200 lbs

170 − 200

Approach configuration

Fwd−Right Tank Group Capacity

6300 − 6600 lbs

200 − 300/0.8 TMN

Cruise configuration

Max split between cockpit totals

300 lbs

400/0.8 TMN

Max IFR probe speed

Fuel dump rate

1500 lbs/min

Fuel dump au

to shutoff with

4000 lbs remaini

Ground refuel rate @ 50 psi

450 gal/min

SAS Stability Augmentation System

In−flight refuel rate @ 57 psi

475 gal/min

FCS CAUTION:

Airspee

< 600 kts/1.3 TMN

Hydraulic System

> 0.5 TMN/10 units No cross controls

> 0.6/15 units AOA coordinate all

lateral stick

Normal Hyd. system operating press

3000 ± 100 psi

BIDI activates when one system is

< 2100 psi

PITCH SAS:

No Limitations

BIDI output w/3000 psi on good side

2400 − 2600 psi

BIDI shuts off when failed system is

< 500 psi for

ROLL DGR/YAW DGR Airspeed <1.0 TMN

sec

and/or ARI DGR

Emer. Flt. Hyd. on if both systems

< 2100 psi

Outb’d Spoiler Module

ARI/SAS OUT:

Airspeed <1.0 TMN,

electrically inhibited @

62_ W/S

AOA

− max 15

units

Outb’d Spoiler Module

No aggressive maneuvering

depressurized @

65_ W/S

Miscellaneous

Prohibited Maneuvers

450

Windmill airstart airspeed required

1. Intentional Spins.

300

Spooldown airstart airspeed required

2. During AB operations; sustained 0 to −0.5 g flight;

250

With RUDDER AUTH, limit inputs to <10_

−0.5 g to −2.4 g’s for more than 10 seconds.

400

Rudder authority limits inputs to < 9.5_

3. At MIL power or less; zero or negative g flight for more

400/0.9

With HZ TAIL AUTH, limit lateral stick<1/4 throw

than 20 seconds.

400

Speedbrake blowback

4. AIM−9 launch with flaps/slats extended.

300

Hook blowback in transit

5. Fuel dump while in AB or with S/Bs extended.

300/0.8 TMN Max speed w/ airsource off (overwing fairing)

350/1.5 TMN Max ramdoor open airspeed − heat from

6. Dual eng AB takeoff, waveoffs, bolters or cat launches.

friction

7. Single eng MAX AB takeoff, waveoff, bolter, or

> 0.7 TMN Coordinate stick and rudder input

cat launches.

> 0.5 TMN No cross controls above

8. Rolling maneuvers with bank angle changes in excess

2.4 TMN

Reduce speed light (airframe limit)

of 360 degrees.

Ejection

G Limits

Zero − 250 KIAS

Ejection Safe

250

− 600 KIAS

Ejection Hazardous

> 600 KIAS

Ejection Extremely Hazardous

Gear Down Symmetric Limit

0 − 2.0

GearDown Rolling

External Tank Limits

(coordinated turns only 225 − 280 KTS)

0 − 2.0

Flaps/Slats Down

− 2.0

68,000 lb aircraft symmetric limit

4.6

650

Max below 12,000 ft

58,000 lb aircraft symmetric limit

5.5

700

Max 12

− 25,000 ft

53,000 lb aircraft

symmetric limit

6.0

650

Max 25 − 34,000 ft

50,000 lb aircraft symmetric limit

6.5

1.75 TMN

Max above 34,000 ft

1.6 TMN

Max with ROLL SAS OFF

Figure 1−2. Characteristics and Limitations (Sheet 2 of 2)

ORIGINAL

1−4

NAVAIR 01−F14AAD−1

TACTICAL

CHAFF DISPENSING SET

AN/ALE−39

DIGITAL DATA LINK

AN/ASW−27B/C

ELECTRONIC COUNTERMEASURES SET

AN/ALQ−165

FUZE FUNCTION CONTROL SET

AN/AWW−4

GUN CONTROL UNIT

C−11414/AYQ−15

IFF INTERROGATOR SET

AN/APX−76C

IFF TRANSPONDER SET

AN/APX−100(V)

INTERFERENCE BLANKER

MX−10666/A

IRSTS

AN/AAS−429XN−1

JOINT TACTICAL INFORMATION DISTRIBUTION SYSTEM

AN/URC−107

MISSILE LAUNCHER/BOL CHAFF DISPENSER

LAU−138A/A

MISSILE POWER SUPPLY

PP−8043/A

MISSION COMPUTERS

AN/AYK−14 9XN−60 PMM

PANORAMIC CAMERA

KA−99A

RADAR SYSTEM

AN/APG−71 (XN−1)

RADAR WARNING SET

AN/ALR−67(V)

SERIAL FRAME CAMERA

KS−87B

STANDARD CENTRAL AIR DATA COMPUTER

CPU−175/A

STORES MANAGEMENT SET

AN/AYQ−15

TARPS POD

LA−610

TELEVISION CAMERA SET

AN/AXX−1

COMMUNICATION

CRYPTOGRAPHIC SYSTEM

TSEC/KY−58

INTERCOMMUNICATIONS SYSTEM

LS−460B/AIC

VHF/UHF COMMUNICATIONS SET

AN/ARC−182

NAVIGATION

AUTOMATIC DIRECTION FINDER

OA−8697/ARD

INERTIAL NAVIGATION SYSTEM

AN/ASN−139

MAGNETIC AZIMUTH DETECTOR SET

DSU−4A/A

MINIATURIZED AIRBORNE GPS RECEIVER

R−2512

MISSION DATA LOADER

MU−1053/A

RADAR ALTIMETER

AN/APN−194(V)

RADAR BEACON AND AUGMENTOR SET

AN/APN−154(V) and R−1623/APN

RECEIVER DECODER GROUP

AN/ARA−63

STANDARD ATTITUDE HEADING REFERENCE SYSTEM

AN/USN−2(V)

TACTICAL NAVIGATION SET

AN/ARN−118(V)

FLIGHT CONTROL AND INSTRUMENTS

AIR INLET CONTROL PROGRAMMER

C−8684B/A

APPROACH POWER COMPENSATOR SET

AN/ASN−146

DIGITAL FLIGHT CONTROL SET

AN/ASW−59

BEARING DISTANCE HEADING INDICATOR

ID−663D/U

STANDBY AIRSPEED INDICATOR

AVU−30/A

STANDBY COCKPIT ALTIMETER

AAU−39/A

STANDBY COMPASS

AQU−5/A

VERTICAL VELOCITY INDICATOR

AAU−8/A

Figure 1Ć3.ĄElectronic Nomenclature

1−5 (Reverse Blank)

ORIGINAL

NAVAIR 01−F14AAD−1

CHAPTER 2

Systems

2.1

AIR INLET CONTROL SYSTEM (AICS)

2.1.1.1

Ground and Low−Speed Operation

The purpose of the AICS is to decelerate supersonic air

During ground static and low−speed (Mach < 0.35)

and to provide even, subsonic airflow to the engine

operation, the inlet ramps are mechanically restrained in the

throughout the aircraft flight envelope. The AICS consists of

stowed

(retracted) position. The predominant airflow is

two variable−geometry intakes, one on each side of the

concentrated about the lower lip of the inlet duct and is

fuselage at the intersection of the wing glove and fuselage.

supplemented by reverse airflow through the fixed bleed

Intake inlet geometry is varied by three automatically

door, around the forward lip of the third ramp. As flight speed

controlled hinged ramps on the upper side of the intakes. The

is increased to 0.35 Mach, hydraulic power is ported to the

ramps are positioned to decelerate supersonic air by creating

ramp actuators but the ramps are not scheduled out of the

a compression field outside the inlet and to regulate the

stowed position until Mach 0.5 (see Figure 2Ć4). The fixed

amount and quality of air going to the engine. The

bleed door bleeds low−energy, boundary layer air from the

rectangular intakes are separated from the fuselage to

movable ramps.

minimize boundary layer ingestion and are highly raked to

optimize operation at high angle of attack.

2.1.1.2

Subsonic and Transonic Speeds

Inlet ramps are positioned by electrohydraulic actuaĆ

tors that respond to fixed schedules in the AICS programĆ

At airspeeds greater than 0.5 Mach, the ramps program

mers. Separate programmers, probes, sensors, actuators, and

primarily as a function of Mach for optimum AICS perforĆ

hydraulic power systems provide completely independent

mance. At transonic speeds, a normal shock wave attaches to

operation of the left and right air inlet control systems.

the second movable ramp. The third ramp deflects above

Figure 2Ć1 shows the basic elements of AICS mechanization.

0.9ĂMach to maintain proper bleed slot height (∆h) for

transonic and low−supersonic flight.

Electrical power for the AICS programmers is proĆ

vided by the ac and dc essential No. 2 buses. The ramp stow

At supersonic speeds, four shock waves compress and

function is powered by the dc essential No. 1 bus. Hydraulic

decelerate the inlet air. The bleed slot removes boundary

power is supplied individually to the left AICS from the

layer air and stabilizes the shock waves. This design results

combined hydraulic system and to the right AICS from the

in substantially higher performance above Mach

2 than

flight hydraulic system. The left AICS programmer also

simpler inlet designs.

functions as a wing−sweep backup computer.

2.1.2

AICS Test

2.1.1

Normal AICS Operations

No pilot control is required during the normal (AUTO)

Two types of AICS tests are provided to check the

mode of operation. Electronic monitoring in the AICS

general condition of the AICS and to pinpoint system comĆ

detects failures that would degrade system operation and

ponents causing detected failures: AICS built−in test and

performance (refer to AICS BIT). AICS caution lights (L and

on−board check.

R INLET, L and R RAMPS) and INLET RAMPS switches

are shown in Figure 2Ć2.

2.1.2.1

AICS Built−In Test

Sectional side views of representative variable geomeĆ

try inlet configurations scheduled by AICS programmers and

BIT in the AICS computer programmer is automatiĆ

cally and continually initiated within the programmer to

descriptive nomenclature are shown in Figure 2Ć3.

check AICS components when the programmer is energized.

2−1

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

INLET RAMPS

AUTO − Inlet ramp position is determined by the AICS programmer.

switches

STOW − Electrically commands the respective inlet ramp actuator to the stow

position; opens the appropriate hydraulic shutoff valve.

WARNING

• DO NOT take off with the INLET RAMPS switches in STOW.

Hydraulic power is on and may drive the ramps out of the stow locks

during certain servocylinder failure modes causing an engine stall.

• If wing−sweep advisory light illuminates, cycling L AICS circuit

breaker (LF1) may cause unintentional wing sweep unless WING

SWEEP DRIVE NO. 1 (LD1) and WG SWP DR NO. 2/ MANUV

FLAP (LE1) circuit breakers are pulled.

Note

Cycling either AICS circuit breaker while airborne may result in

DFCS air data failures, illuminating the FCS CAUTION and ARI

DGR lights, and cause degraded control system capability.

MASTER RESET should restore normal operation once the

cb(s) are reset.

RAMPS caution

Indicates ramps not positioned in either stow or trail locks during critical flight

light

conditions (see Figure 2Ć5).

Figure 2Ć2.ĄAICS Control and Indicators (Sheet 1 of 2)

2−3

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

INLET caution light

Indicates AICS programmer/system failure: Reduce airspeed to Mach 1.2 and

check AICS acronym for failure indication.

AICS failure

Less than Mach 0.5: Ramps should be restrained by actuator stow locks.

Greater than Mach 0.5: Ramp movement is restrained by trapped hydraulic

pressure and mechanical locks, depending on Mach when INLET light illuminates.

Greater than Mach 0.9: Ramp movement is minimized by actuator spool valves and

the aerodynamic load profile in this Mach range and a RAMP light should illuminate.

Figure 2−2. AICS Control and Indicators (Sheet 2 of 2)

The operational status of the AICS depends on

hydraulic power. Detected AICS failures are indicated by

BIT−detected failures in AICS components. Failures of static

AICS acronyms or AIC acronym(s) with associated INLET

or total pressure sensors; ramp No. 1, 2, or 3 positioning;

caution light(s) displayed after completion of OBC.

programmer continuous end−to−end BIT; or hydraulic presĆ

sure to any of the ramp actuators would seriously degrade

Note

AICS performance. Detected failures of these items cause the

D With INLET RAMP switches in STOW, AICS

AICS to automatically transfer to a significantly degraded

OBC will fail test and INLET lights will illuĆ

fail−safe mode of operation, indicated by illumination of an

minate.

INLET caution light.

D If the engine enters secondary mode during

Detected failures of angle of attack, engine fan speed,

OBC, the ramps will stow and fail OBC. To

or out−of−calibration detection of the difference between P1

reinitiate OBC, select primary mode and reset

and P2 (∆P), P_s or P_t sensors will cause the AICS to revert to

the AICS.

the slightly degraded fail−operational mode of operation.

D An S4 acronym indicates the AICS programĆ

Nominal values of angle of attack, total temperature, or

mers may be operating on the REV 4 (TF−30/

engine fan speed are substituted for the failed values in the

F14A) schedule. As a result, below 25,000 feet

AICS programmer, without illumination of an INLET

at airspeeds greater than 1.1 Mach, unloading

caution light.

the aircraft to less than 1g will reduce inlet staĆ

In both fail modes of operation, detected failures are

bility and may result in inlet buzz and possibly

continuously registered by the in−flight performance moniĆ

an engine stall. Cycling AICS circuit breakers

toring system and displayed with air inlet control acronyms

at a constant subsonic Mach number should

on the multifunction display and the programmable tactical

eliminate the S4 acronym and reset the proĆ

information display (Figure 2Ć5 and Figure 2Ć6).

grammer to the REV 5 (F−110) schedules.

2.1.2.2

AICS On−Board Check

2.1.3

AICS Failure Modes of Operation

OBC, initiated by the pilot during poststart or ground

AICS mode of operation following a BIT−detected

maintenance checks, performs a dynamic check of the left

failure may be either fail−operational mode (Figure 2Ć5) or

and right AICS. In addition to the regular AICS BIT program,

fail−safe mode (Figure 2Ć6).

sensor calibration checks are made. The status of the proĆ

grammer electronics and the ramp actuators are checked

2.1.3.1

Fail−Operational

throughout an altitude and airspeed schedule as psuedoĆ

pneumatic inputs to the programmer are varied to simulate a

Failures in the AICS are detected by the AICS proĆ

flight sequence of maximum airspeed condition and back to

grammer, which automatically initiates appropriate corĆ

static sea level conditions within 65 seconds. This cycles the

rective action. Mode entry is indicated by the display of a

ramp actuators through their full range, illuminates the ramp

fail−operational AIC acronym. The fail−operational mode

lights, exercises the complete AICS for preflight failure

results in no significant degradation in AICS operation, and

detection, and ensures the ramps are in their stow locks. OBC

the mission can be continued without any flight restrictions

is the only way to ensure stow lock integrity since it verifies

or corrective action by the pilot.

the ramps are in the stowed position and then removes ramp

ORIGINAL

2−4

NAVAIR 01−F14AAD−1

ACTUATOR POSITION

HYDRAULIC POWER

FLIGHT CONDITION

RAMP ACTUATORS

RAMP NO. 1

RAMP NO. 2

RAMP NO. 3

M < 0.35

OFF

Mechanically restrained by stow locks in stowed position;

electrical stow commands output from AICS programmer.

M > 0.35 to < 0.5

Electrical stow commands output from AICS programmer.

M > 0.5 to < 2.2

Variable position scheduled by AICS programmer as a

function of mach number, corrected engine fan speed, and

angle−of−attack. Ramps no. 1 and no. 2 begin positioning

at Mach 0.5; ramp no. 3 begins at Mach 0.9.

M > 2.2

Variable position scheduled by AICS programmer as a

function of Mach number.

Figure 2Ć4.ĄAICS Normal Operating Mode

PTID FAILURE

MAINTENANCE

DETECTED

READOUT ACRONYM

FAILURE

CAUSE

RESULT

AIC S1

P_s, P_t or

Limits exceeded.

Ramps may not program during

(Possible only

programmer out of

OBC. Reset AICS L and R circuit

during OBC)

calibration

breakers (LF1, LG1) prior to

attempting another OBC.

NONE

Engine fan speed

Loss of engine fan

Substitutes 7,300 rpm. Ramps do not

rpm from AFTC.

speed signal.

program during OBC.

AIC S3

None

Mask continuous monitor (CM) so

that subsequent AIC acronyms are

displayed.

AIC S4

Angle−of−attack or

Limits exceeded.

IN FLIGHT: Substitutes +2_

engine fan speed.

angle−of−attack or 7,300 rpm.

AIC S4

Alpha delta pressure

Limits exceeded.

• Substitutes +2_ angle−of−attack

(During OBC)

sensor out of

Augmenter fan

value until reset.

calibration or engine

temperature

fan speed.

controller (AFTC) may

• Substitutes 7,300 rpm.

be in secondary mode.

AIC A4

Open wire

None

Note

AIC symbol has L or R appended (AICL, AICR) to

identify on which side failure was detected.

Figure 2Ć5.ĄFail−Operational Mode  No INLET Light

ORIGINAL

2−6

NAVAIR 01−F14AAD−1

RESULT

FAILURE MAINTENANCE

DETECTED

MACH < 0.5

MACH > 0.5

READOUT ACRONYM

FAILURE

CAUSE

AIC P

AICS programmer

Failed end−to−end

Hydraulic shutoff

Ramp movement is

(P)

BIT

valve remains closed.

restrained by actuator

Ramp actuators

mechanical locks if

AIC S1

remain mechanically

failure occurred with

Static pressure (P_s)

Minimum or

restrained within stow

ramps within locks.

maximum limits

AIC S2

locks, provided they

Otherwise ramp(s)

Total pressure (Pt )

exceeded

failed within stow

move slowly with

locks.

aerodynamic loads.

AIC A1

Ramp No. 1

Sustained command

and feedback

error

AIC A2

Ramp No. 2

AIC S3

Ramp No. 3

AIC A1,

Hydraulic pressure

Sustained error due

Ramp(s) may move

AIC A2, or

loss of ramp No. 1,

to loss of hydraulic

if failure occurred

AIC A3

No. 2, or No. 3

pressure

with ramp(s) out of

(INLET caution light

mechanical locks.

eventually illuminates

RAMP light will

> 0.5 Mach)

illuminate.

NONE (No INLET caution

Loss of hydraulic

light < 0.5 Mach)

pressure

Note

AIC symbol has L or R appended (AICL, AICR) to identify on which side failure was detected.

Figure 2Ć6.ĄFail−Safe Mode  INLET Light Illuminated

Note

Transferring to SEC mode will revert the AICS

Fail−safe operations result in a slight degradation

programmers to the REV 4 (TF−30/F14A) schedĆ

of cruise and excess thrust performance because

ule because of the loss of the AFTC N₁ speed

of the off−optimum configuration.

signal and will display an S4 acronym. Below

If the hydraulic shutoff valve closes above Mach 0.9,

25,000 feet and at airspeeds greater than 1.1

the ramps are normally in an unsafe configuration and the

Mach, unloading the aircraft to less than 1g will

appropriate RAMPS caution light will accompany the

reduce inlet stability and may result in inlet buzz

INLET caution light (Figure 2Ć7). Above Mach 0.9, the

and possible engine stall. To restore full REV 5

No. 3 ramp normally begins programming below the actuator

(F110/F14B/D) schedule and eliminate S4 acroĆ

stow lock. When the fail−safe mode is entered above Mach

nym following an airborne engine mode reset to

0.9, the unpowered No. 3 ramp will eventually move and may

PRI, recycle AICS circuit breakers at constant

cause compressor stalls at higher power settings. The aircraft

subsonic Mach number.

shall be decelerated below 1.2 Mach, and the appropriate

2.1.3.2

Fail−Safe Mode

INLET RAMPS switch shall be selected to STOW.

The fail−safe mode results in significantly degraded

AICS operation. Mode entry is indicated by the display of a

fail−safe AIC acronym and illumination of the appropriate

INLET caution light. Under these conditions, the AICS

Do not select STOW at speeds greater than

programmer provides a shutoff signal to close the ramps’

1.2 Mach. Compressor stalls may occur because

hydraulic shutoff valve. If the hydraulic shutoff valve closes

of ramp mispositioning.

below Mach 0.9, the ramps are normally in a safe configuraĆ

tion (No. 1 and No. 2 ramp within trail locks and No. 3 ramp

2.1.3.3

Stow Mode of Operation

in stow locks are restrained by trapped hydraulic pressure).

The STOW position of the INLET RAMPS switch

Engine operations may be successful below 1.2 Mach in this

commands the appropriate hydraulic shutoff valve to open

configuration; however, corrective procedures shall be

performed.

2−7

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć7.ĄRamp Monitor Logic

and provides a direct electrical signal to the ramp actuators,

Whenever the hydraulic shutoff valve closes (i.e., fail−

porting hydraulic pressure directly to the retract side of the

safe mode entry), hydraulic spool valves in the ramp actuaĆ

actuator. When the ramps are retracted to the stow position,

tors sense the absence of pressure and block the actuator

the RAMPS light will extinguish and the stow locks should

pressure and return ports, causing a hydraulic lock (dump

remain engaged even if hydraulic power is subsequently lost.

inhibit). This feature reduces ramp movement when an AICS

Once in STOW, AICS programmer−detected electronic

failure occurs and the ramps are not being restrained by

failures may be reset below Mach 0.5.

mechanical actuator locks. Although dump inhibit prevents

the ramp from rapidly extending and causing an engine stall,

2.1.3.4

Hydraulic Shutoff and Dump Inhibit

the ramps will still slowly move. Under normal circumĆ

The AICS hydraulic systems include a hydraulic shutoff

stances, the pilot will have sufficient time to select STOW

valve to control hydraulic system pressure. The hydraulic

and prevent an engine stall. F−14A flight test results show that

shutoff valve is normally controlled by the AICS programmer,

with dump inhibit, the time interval between illumination of

which removes the hydraulic−on signal below 0.35M or in the

a RAMPS caution light and engine stall following an AICS

event of a programmer failure. The STOW position of the

failure is 15 to 40 seconds on the ground at military power,

INLET RAMPS switch bypasses the AICS programmer to

and approximately 50 seconds at 10,000 feet at military

energize the hydraulic shutoff valve, providing pressure for

power.

ramp motion. To ensure hydraulic pressure is shut off, the

respective AICS programmer must be deenergized by pulling

2.1.3.5

Ramp Actuator Mechanical

the circuit breaker (LF1, left or LG1, right) and the INLET

Locks/Positioning

RAMPS switch placed in the AUTO position.

In addition to the actuator stow locks, the first and

Note

second ramp actuators have another set of latches (trail locks)

Cycling either AICS circuit breaker while airĆ

that prevent further ramp actuator extension after a failure

borne may result in DFCS air data failures, illuĆ

within these trail locks. The actuator stow and trail locks

minating the FCS CAUTION and ARI DGR

restrain actuator movement in tension only. Hydraulic

lights, and cause degraded control system capaĆ

pressure (500 psi) is required to disengage the lock finger

bility. MASTER RESET should restore normal

latches.

operation once the cb(s) are reset.

ORIGINAL

2−8

NAVAIR 01−F14AAD−1

Safe positioning of the ramp actuators is monitored

fail−safe operation. In flight, the No. 1 and 2 ramps tend to

by the ramp monitor logic shown in Figure 2Ć7. A RAMPS

blow back to the stow position or are restrained within the

light should always be accompanied by an INLET light with

trail locks because of aerodynamic loads. The hydraulic

the landing gear handle UP. With the landing gear handle

restriction of all ramps during loss of hydraulic power and

DOWN, a RAMPS light can be illuminated without an

after fail−safe mode entry should prevent rapid ramp

INLET light. The emergency procedure in any case is the

movement. Internal failure of an actuator especially the

same. RAMPS lights will extinguish when a safe configuraĆ

No. 3 ramp actuator, may allow rapid ramp extension and

tion is attained.

cause engine stall. Additionally, failure to stow the ramps in

a reasonable amount of time after INLET light illumination

Note

or inability to stow following a hydraulic system failure may

D Following an AICS programmer/ramps failĆ

result in compressor stalls at high power settings. Engine start

ure, the safest configuration results when the

attempts may not be successful unless the ramps are stowed

(RAMPS caution light extinguished).

ramps are in the stowed position. The proĆ

grammers are disabled by pulling the affected

AICS circuit breaker and returning the INLET

2.1.4

AICS Anti−Ice

RAMPS switch to AUTO.

AICS anti−ice is activated only by selecting ORIDE/

D Cycling either AICS circuit breaker while airĆ

ON with the AICS ANTI−ICE switch and airspeed between

borne may result in DFCS air data failures, illuĆ

0.35 to 0.9 Mach (hydraulic power is available at 0.3 Mach).

minating the FCS CAUTION and ARI DGR

Above and below these airspeeds the AICS anti−ice is

lights, and cause degraded control system

disabled. When the ENG/PROBE anti−ice switch is in

capability. MASTER RESET should restore

AUTO, the AICS anti−ice is off. When AICS anti−ice is

normal operation once the cb(s) are reset.

activated, the AICS programmer repositions the No. 1 and

No. 2 ramps to positions below the No. 3 ramp (Figure 2Ć8)

In the event of an engine or hydraulic failure, the

so that ice will not form above the No. 3 ramp.

following conditions exist with respect to AICS reset:

2.2

ENGINE

If hydraulic pressure is zero, there is no need to safe

the ramps (by stowing ramps, pulling AICS circuit

The aircraft is powered by two F110−GE−400 turbofan

breakers, and returning to AUTO) since selecting

engines (Figure 2−9) with variable exhaust nozzles and AB

STOW will have no effect without hydraulic

augmentation. They are dual−rotor engines consisting of a

pressure.

three−stage fan driven by a two−stage, low−pressure turbine

and a mechanically independent, aerodynamically balanced,

If airspeed is less than .35 Mach, there is no need to

nine−stage, high−pressure compressor driven by a single−

safe the ramps since hydraulic pressure has already

stage, air−cooled high−pressure turbine. Engine operation is

been removed and ramps should be in the stow

automatically regulated and maintained electrically by the

locks. If the ramps are not in the stow locks, the

augmenter fan temperature control unit and by throttle inputs

RAMPS light will illuminate when the landing gear

to the main engine control.

handle is lowered. If the RAMPS light does illumiĆ

nate, then the ramps should be stowed and the AICS

Each engine is slung in a nacelle with the thrust axis

programmer reset. (Depressing MASTER RESET

laterally offset approximately

4½ feet from the aircraft

following an AICS programmer reset should restore

centerline. The installed static engine thrust at military power

normal DFCS operation.)

13,800 pounds and, at maximum AB power, thrust is

If hydraulic pressure is greater than zero and airĆ

23,600 pounds. Installed engine thrust at maximum AB at

speed is greater than .35 Mach, then the ramps

0.9M at sea level is 30,200 pounds. Acceleration time from

should be stowed and, if time allows, the programĆ

idle to military power is approximately 4 seconds.

mer reset after engine failure or a low hydraulic

pressure situation. This will ensure that if the ramp

During operation, air entering the engine is directed

is out of the stow lock (as is normal above .5 Mach),

into the fan, which initially compresses the air and directs

it will be returned to the stow lock and kept there for

its flow into the engine core compressor and fan bypass duct.

landing regardless of subsequent hydraulic or

Direction of airflow into the fan is optimized by variable−

electrical malfunctions.

geometry inlet guide vanes (IGV) and into the compressor by

variable geometry stator vanes. The high−pressure compresĆ

2.1.3.6

AICS Failure In−Flight Operation

sor further compresses the air through the nine−stage comĆ

pressor before discharging it into the annular combustion

Most AICS failures occurring in flight do not require

chamber to mix with fuel from the fuel nozzles. This fuel−air

rapid pilot response because of system design features for

2−9

ORIGINAL

NAVAIR 01−F14AAD−1

mixture is initially ignited by the main spark igniter in the

Note

combustion chamber. As a result of this combustion,

SEC mode transfer while in AB may result in pop

expanding gases drive the high− and low−pressure turbines.

stalls. Nonemergency manual selection of SEC

Power to drive the two accessory gearboxes is obtained from

mode on the ground should be performed in basic

the high−pressure rotor.

engine. Nonemergency manual selection of SEC

mode airborne should be performed in basic

From the turbine section, the exhaust gases pass into

engine with power set above 85−percent rpm.

the afterburner section and are mixed with air from the fan

bypass duct. During AB operation, fuel is sprayed into this

Transferring to SEC mode will revert the AICS proĆ

mixed airflow and ignited for additional thrust.

grammers to the REV 4 (TF−30/F14A) schedule because of

the loss of the AFTC N1 speed signal and will display an

OBC AICS − LEFT (RIGHT) and ANGLE OF ATTACK

acronym. Below 25,000 feet and at airspeeds greater than 1.1

TMN, unloading the aircraft to less than 1g will reduce inlet

stability and may result in inlet buzz and possible engine

During night and/or IFR conditions, the

stall. To restore the full REV 5 (F110/F14B/D) schedule and

increased acceleration during AB use will result

eliminate the OBC acronym following an airborne engine

in inner ear disturbances. In addition, the large

mode reset to PRI, cycle AICS circuit breakers at constant

amount of light generated by the AB exhaust

subsonic Mach number.

reflecting around the aircraft will compound this

condition. These factors may result in severe

2.2.1.1

Main Engine Control

aircrew disorientation/vertigo.

The MEC is a hydromechanical control that provides

2.2.1

Engine Control

fuel shutoff, variable stator vane scheduling, and main fuel

metering in both primary and secondary modes. The MEC

The engine is controlled by three units: the hydroĆ

controls fuel flow until

59−percent rpm and provides

mechanical main engine control, the electronic augmenter

high−pressure compressor rotor overspeed protection autoĆ

fan temperature control, and the AB fuel control. There are

matically by securing fuel flow to the engine when an

two modes of operation: primary (electronic) and secondary

overspeed condition of 110 percent is reached.

(mechanical), with provisions for automatic and manual

switchover to secondary. Manual selection is controlled

Note

through the ENG MODE SELECT panel (Figure 2Ć10).

D To regain engine operation following an autoĆ

Automatic or manual selection of the secondary mode illuĆ

matic engine overspeed shutdown, the throttle

minates an ENG SEC caution light. When one engine reverts

must be cycled to OFF then IDLE.

to secondary mode, the other engine continues in primary

mode. Cycling the ENG MODE SELECT switch may reset

D An overspeed condition in excess of

110

the AFTC if the faults are temporary. If the change back to

percent will result in momentary loss of rpm

primary mode is successful, the ENG SEC light will go out.

indication until N₂ rpm falls below 110 ± .5

Automatic or manual selection of secondary mode is possible

percent. EGT and FF indicators will continue

throughout the flight envelope. Selection of secondary mode

to function normally.

will cause a loss of fan speed signal to the AICS.

2.2.1.2

Augmenter Fan Temperature Control

The AFTC is a modular solid−state electronic device

that performs control schedule computations, integration and

logic functions, limit control, failure detection, and provides

engine core speed (N₂) signal for instrument display and

SEC mode transfers with throttles in AB above

engine fan speed (N₁) signal to the AICS. It also controls the

450 KCAS could result in pop stalls and damage

distribution of electrical power to the entire engine electrical

to the IGV linkage.

and monitoring systems. Figure 2Ć11 shows the various

interface signals used by the AFTC. Normally the CADC

supplies Mach number value to the AFTC. If this signal is

erroneous, the AFTC assumes a default Mach number value

in order to continue operation.

2−11

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć11.ĄAFTC Functional Relationships

engine operation with unrestricted throttle movement

throughout the flight envelope. The AFTC computations are

used to control basic engine and AB fuel flow, IGV, and AB

nozzle positioning; VSV positioning is controlled by the

The loss of Mach number signal from the CADC

MEC. The AFTC incorporates independent control schedĆ

results in the loss of both airflow limiting and idle

ules that are prioritized so that the optimum amount of fuel

lockup functions of the AFTC. This may result in

flow is provided to the main combustor. At any given time,

pop stalls while supersonic on a cold day, or at

only one of these schedules is actually in control of fuel flow.

high power settings, or at idle. Inlet buzzing may

The remaining schedules are always active and are calculatĆ

also result at high power settings. If occurring

ing the change in fuel flow required (if any) to attain the

while supersonic and at high power settings,

desired value of their assigned parameter. The selection of

decelerate at military power until subsonic.

the schedule in control is accomplished by a series of

minimum and maximum selectors. These selectors control

2.2.1.3

Afterburner Fuel Control

scheduling of the following:

The AFC is controlled by the AFTC for afterburner

1. Acceleration/deceleration

operation. The AFTC computes AB fuel flow ratios and

provides them to the AFC. The AFC converts ratio comĆ

2. Minimum/maximum compressor discharge

mands to metered fuel flows into local, core, and fan AB fuel

pressure

manifolds. When staging up the AB, local fuel flow is

initiated first, followed by core and fan flow last. When

3. Minimum/maximum rpm

staging down, the reverse sequence occurs. Thrust changes

are smooth when staging up or down.

4. Fan speed limiting

2.2.1.4

Primary Mode

5. Maximum turbine blade temperature limiting

In the primary mode, the AFTC controls the MEC,

6. Idle lockup speed.

AFC, and AB nozzle hydraulic pump to provide optimum

2−13

ORIGINAL

NAVAIR 01−F14AAD−1

Other AFTC functions include engine start control,

asymmetric thrust limiting, reduced arrestment thrust, autoĆ

matic relight, and fault detection. Fault detection automatiĆ

cally switches the engine control to the secondary mode in

the event of core overspeed, fan speed signal loss, nozzle full

SEC mode transfers with throttles in AB above

open when engine is not at idle or maximum AB, AFTC

450 KCAS could result in pop stalls and damage

power deviations, fuel flow demand mismatch with throttle

to the IGV linkage.

settings, fan speed greater than 800 rpm and not accelerating,

or throttle signal error.

Note

D SEC mode transfer from AB may result in pop

2.2.1.4.1

AB Operation (Primary Mode)

stalls. Nonemergency manual selection of

SEC mode on the ground should be performed

For AB operational characteristics, refer to Figure 2Ć12.

in basic engine. Nonemergency manual selecĆ

Unrestricted throttle operation into and out of AB is per−

tion of SEC mode airborne should be perĆ

mitted throughout the flight envelope. During AB operation,

formed in basic engine with power set above

rpm, EGT, fuel flow, and nozzle position vary with altitude

85−percent rpm.

and airspeed. The nozzle position will also increase as the

throttle is transitioned from minimum AB to maximum AB.

D If the fan speed limiter circuit has failed, engine

If an AB blowout occurs, the autorelight feature attempts to

rollback may occur with selection of SEC

reinitiate AB without throttle movement. The engine has

mode. In the event of engine rollback, PRI

reduced AB region of operation at high altitudes and low airĆ

mode must be reselected above 59−percent rpm

speeds. An AFTC automatic rich stability cutback" feature

or flameout will occur and airstart will not be

reduces or limits maximum AB fuel flow at high

possible.

altitudes and low airspeeds to prevent AB instabilities

(Figure 2Ć12). Indication of rich stability cutback is a nozzle

2.2.1.6

Engine Alternator

position of approximately 30 to 50 percent at maximum AB

rather than the normal 60 to 70 percent. Also, because of airĆ

Each engine’s electrical system is powered by an alterĆ

flow and temperature characteristics, AB light−off characterĆ

nator mounted on the engine aft gearbox. The alternator

istics are slower at high altitudes and low airspeeds.

consists of four windings. Two windings are redundant in

providing power to the AFTC and its components. A third

2.2.1.5

Secondary Mode

winding provides power for both main high−energy ignition

and AB ignition. The fourth winding provides power to the

Basic engine operation in SEC mode is extremely reliĆ

engine monitoring system processor (for FEMS), and a signal

able. In the secondary mode, the electronic functions perĆ

for the rpm gauge. The last winding is also an alternate source

formed by the AFTC are eliminated. The MEC provides

of power for the fan speed limiting circuit. The fan speed

complete control of the engine with the exception of fan

limiting circuit may be powered by either the essential

speed limiting. SEC mode is manually selected via the

No. 2 dc bus or the engine−driven alternator winding, dependĆ

ENG MODE SELECT switch or the autopilot emergency

ing on which source has the highest stable output.

disengage paddle switch, or via automatic default. In

SEC MODE, the exhaust nozzle is commanded full closed,

If engine alternator power output drops below a preset

the nozzle position indicator goes to the not−powered

value, engine control will automatically transfer to SEC

position (subzero indication), the IGVs are fixed full open,

mode, illuminating the respective engine SEC light. If the

high−energy ignition is continuously energized, AB is

engine reverts to SEC mode as a result of a sheared alternator

inhibited, and idle lockup protection is lost.

shaft, engine high−energy ignition will not be available and

the engine SEC light will not illuminate. Cockpit indications

In SEC mode, engine stall margin is decreased at low

are loss of engine rpm and nozzle position indicating below

rpm because of IGV positioning. The FEMS engine stall

zero. In failure modes, redundant aircraft electrical power

detection circuit is inoperative, but overtemperature warning

will be available for fan overspeed protection.

is still available. A low−level vibration/rumble may be sensed

in ground idle operation when in secondary mode. This

The engine is completely operable should the aircraft

vibration/rumble has no adverse affect on the engine and

experience a complete electrical failure. The engine operates

disappears when the throttle is advanced slightly (5−percent

in either PRI or SEC mode, which can be selected automatiĆ

rpm increase or less). Maximum thrust available at military

cally or manually. In case of a complete electrical failure all

power in SEC mode is depicted in Chapter 14, Figure 14Ć4.

engine lights and indicators are inoperative.

ORIGINAL

2−14

NAVAIR 01−F14AAD−1

2.2.1.7

Turbine Blade Temperature (Pyrometer)

gear with the hook handle down or the hook out of the stowed

position. The RATS light, located on the pilot’s advisory

The pyrometer is a fuel−cooled, photodiode, optical

panel, illuminates when the aircraft circuit is activated but it

unit that measures infrared radiation from the metal surface

is not an indication that the engines are operating at reduced

of the high−pressure turbine blades. This temperature signal

thrust.

is transmitted to the AFTC and is used to regulate engine fuel

flow, which maintains turbine blade temperature within

Note

limits. Cockpit indications of turbine blade temperature

The RATS light will be illuminated anytime the

appear on the MFD.

aircraft circuit is enabled, even if the engines are

operating in SEC mode or the engine circuit has

2.2.1.8

Flame Sensor

been overridden by selection of AB.

The flame sensor is an ultraviolet radiation sensing unit

in the AB duct. During AB operation, ultraviolet rays

2.2.2

Variable Exhaust Nozzle

detected through a quartz window activate a gas filled sensor

Engine exhaust gases at higher thrust settings are

that electrically transmits a flame−present signal to the

discharged through the nozzle throat at sonic velocity and

AFTC. Without this signal, only minimum AB fuel flow is

are accelerated to supersonic velocity by the controlled

available. AB will be inhibited if the flame sensor fails on.

expansion of the gases. Varying nozzle throat area controls

A L/R AUG acronym is displayed in the ENGINE FAULTS

fan stall margin, which optimizes performance.

block of the MFD engine page.

The variable exhaust nozzle is a three−flap, convergent−

2.2.1.9

Asymmetric Thrust Limiting

divergent−type nozzle. Nozzle variation is accomplished by

The asymmetric thrust limiting circuit is designed to

axial movement of four hydraulic actuators mechanically

hold both engines to minimum AB until both ABs are lit off.

synchronized for geometric stability. These hydraulic actuaĆ

The AFTC releases the hold on the AB when both engine AB

tors use oil from a separate compartment in the engine oil

pumps are on and both engine flame sensors are on. Selecting

storage tank and are operated by a hydraulic pump that

the ASYM LIMITER switch to OFF (guard cover up)

responds to AFTC signals. A failed open nozzle may be

overrides the comparison of left and right AB status and

caused by an oil leak, but if the leak is in the nozzle system,

allows each AB to operate independently.

only a portion of the main engine lube oil will be lost. During

basic engine operation, the nozzle area is modulated to a near−

closed position, and, in AB, the nozzle area is infinitely variĆ

able to a full−open position. The nozzle will go full open

airborne with the throttle at IDLE at low altitude and airspeeds

(Figure 2Ć13). A gauge for each engine on the pilot instrument

A malfunctioning or deselected ATLS can

panel next to the engine instruments indicates nozzle position

greatly increase the magnitude of asymmetric

in percentage from 0 to 100. Normal indication for maximum

thrust because of engine stall or failure.

AB is approximately 70 percent.

2.2.1.10

Reduced Arrestment Thrust System

Note

The RATS is a feature of the AFTC provided to reduce

When AFTC is operating in secondary mode, the

thrust for carrier landings to a level consistent with carrier

nozzle is commanded closed and the exhaust

(CV) wind−over−deck operations. When activated, the AFTC

nozzle indicator is inoperative.

automatically reduces the military power core speed (N₂)

by approximately 4.5 percent. This results in an approximate

With the landing gear handle down, engine at IDLE,

20 to 25−percent decrease in thrust.

and weight off wheels, the nozzle is restricted to a near closed

position (maximum 26 percent) to prevent exhaust nozzle

RATS employs two enabling circuits: an engine circuit

flap contact with the deck/hook during landing. Five seconds

incorporated within each engine’s AFTC, and an aircraft

after weight on wheels, the nozzle resets to full open to

circuit. The engine circuit is enabled by the aircraft circuit via

reduce idle power during landing rollout and while taxiing.

switch closure. Since the engine circuit is a function of the

On deck in PRI mode with throttle above IDLE detent, nozzle

AFTC, it is not available in SEC mode and can be overridden

position varies linearly with throttle position.

in PRI mode with selection of AB. The aircraft circuit is

enabled when weight is placed on either or both main landing

ORIGINAL

2−16

NAVAIR 01−F14AAD−1

Figure 2Ć14.ĄFEMS Multifunction Display Configuration

2.3

FATIGUE ENGINE MONITORING SYSTEM

The FEMS consists of the following components (see

Figure 2Ć15).

2.3.1

FEMS Functional Description

2.3.1.1

Engine Monitoring System Processor

The FEMS is a solid−state electronic system that

The EMSP is engine−mounted, engine−powered, and

provides data acquisition, processing, and storage. FEMS

converts control system electrical signals from the AFTC

information is displayed on the MFDs (Figure 2Ć14). The

into digital format for transmission to the ADAC. It also

system accumulates airframe stress and fatigue data and

receives and digitizes other noncontrol system−related data

relevant engine performance data, both in flight and on deck,

such as anti−icing system status, lube oil level, and lube temĆ

from the engine monitoring system processors. Engine faults

perature data for transmission to the ADAC. In addition, the

are isolated to the appropriate WRA or combinations of

EMSP calculates and stores engine cycle count data, making

WRAs and recorded for later transfer to the DPGS for

this data readily available for each serial−numbered engine

diagnostic analysis, troubleshooting, and appropriate mainĆ

even when the engine is not installed in an aircraft.

tenance. The DPGS also computes and stores engine parts

life tracking and failure−trending data. This tracking of

Note

engine data extends the life and safety of fleet aircraft by

EMSP is only operational with the engines in

permitting maintenance routines at periodic intervals. FEMS

primary mode.

also provides a signal to the stall warning system that initiates

2.3.1.2

Airborne Data Acquisition Computer

a 10−second warning tone (identical to overtemperature tone)

and illuminates the L or R STALL warning legend on the

The ADAC is the central processor of FEMS and

MFD/HUD indicating an engine stall. FEMS will record

executes airframe and engine fatigue algorithms. The ADAC

aircraft overstress when it determines that normal acceleraĆ

acquires aircraft data by direct analog and digital inputs.

tion has exceeded:

Additional aircraft data received by the ADAC from the CIU

1. 7.5g with landing gear UP and Mach greater

to be stored as a result of structural, engine, or other mission

events are transferred to the DSS for postflight analysis. In

than .24

addition, ADAC stores fault code messages, in nonvolatile

2. 4.5g with landing gear DOWN (as in hard landing)

memory, for display on the FMI. ADAC is powered by the

3. 4.5g when Mach is .24 or less.

28−Vdc right main bus.

ORIGINAL

2−18

NAVAIR 01−F14AAD−1

2.3.1.3

Data Storage Set

The DSS, located in the nosewheel well has a removĆ

able DSU that provides in−flight recording of engine and CSS

data for analysis. In flight, the MCS transfers engine−related

data via the 1553 bus to the DSU for postflight analysis. This

data is collected for engine diagnostic purposes and compiled

for long−term maintenance records. A fault code on the FMI

will alert the maintenance crew when the DSU has reached

80 percent of its capacity for engine data recording. If the

DSS is inoperative or is not loaded with a DSU, engine part−

life tracking data is maintained only by the EMSP.

2.3.1.4

Flight Maintenance Indicator

The FMI (Figure 2Ć16) displays to the maintenance

crew ADAC data for engine/airframe status. It is mounted in

an easily accessible location on the forward bulkhead in

the nose wheelwell. After each flight, the FMI FAIL,

CAUTION, and/or FLUIDS fault trip indicators will be

either black, signifying the absence of a FEMS−detected failĆ

ure, or white, indicating FEMS detected a failure. The indicaĆ

tors should normally be reset by maintenance personnel prior

to flight. With electrical power applied to the aircraft, pressĆ

ing the STATUS SWITCH button displays either a fault code

(if a fault is present) or NONE in the STATUS window. All

fault codes may be scrolled line by line by pressing the

STATUS SWITCH button once for each line. When no more

Figure 2Ć16.ĄFlight Maintenance Indicator

fault codes are displayed, the display will read END*. When

END* is displayed, pressing and holding the CLEAR button

changes the display from END* to CLR for approximately

Exhaust nozzle off schedule or signal out of range

5 seconds followed by NONE, erasing all fault codes.

Fan inlet guide vanes off schedule or signal out of

Note

range

The FMI is designed to be a maintenance tool

AB fuel valve operation (dry power)

only and should not be used as a go/no−go device

by aircrew on preflight. Likewise, aircrew

AB fuel schedule fault or signal out of range

should not take it upon themselves to reset the

device. Do not press both CLEAR and STATUS

AB signal on but not selected

SWITCH at the same time. Failure to comply

will result in the FEMS onboard clock being

10. No AB light−off signal

altered.

11. AB blowout

The following is a composite listing of the data autoĆ

12. Secondary mode operation

matically recorded in memory for maintenance and disĆ

played in numeric code on the FMI:

13. Pilot−initiated EMS data

1. Fan/core overspeed

14. Anti−icing fault

2. Decay in core speed or signal out of range

15. Low oil quantity or signal out of range

3. Compressor stall

16. Oil overtemperature

4. Turbine blade temperature limit exceeded or signal

17. AFTC power out of limits

out of range

ORIGINAL

2−20

NAVAIR 01−F14AAD−1

18. Throttle/AFTC signal fault

2.4.1

Motive Flow Fuel Pump

The motive flow fuel pump is a gear−driven centrifugal

19. Mach signal to AFTC fault

pump on each engine accessory gearbox that returns

20. Aircraft 28−volt supply to AFTC fault

high−pressure fuel to the fuselage and wing tanks to effect

normal fuel transfer. Motive flow is used to power the boost

21. EMSP fault

pump in the respective sump tank. This fuel continues

through control valves to ejector pumps in the fuselage and

22. ADAC/EMSP interface fault

wing fuel tanks. There is no cockpit control for the motive

flow fuel pumps. Failure of one pump illuminates the R or L

23. ADAC BIT fault and system failure

FUEL PRESS caution light and reduces the rate of fuel

transfer but does not inhibit the transfer of fuel from any tank.

24. ADAC battery low

Motive flow pump failures cause the engine to draw fuel

through suction feed. Higher altitudes and decreased ambient

25. Data storage set memory full and requires service

pressure result in reduced fuel flow, which may cause engine

flameout because of fuel starvation. With a single motive

26. Aircraft overstress

flow fuel pump failure, AB selection above 15,000 feet MSL

may cause engine flameout. With failure of both motive flow

27. System DSS

fuel pumps, high power settings in basic engine may cause

28. ADAC A−6 failure

flameout above 25,000 feet MSL. If a dual motive flow fuel

pump failure occurs, wing fuel will not be available.

29. RATS failure.

2.4.2

Engine Fuel Boost Pump

2.3.2

FEMS Operation

The engine

(total flow) fuel boost pump is an

FEMS data acquisition for monitoring engine perforĆ

engine−driven centrifugal pump on the aft accessory gearbox

mance is automatic. However, the pilot may encounter

that provides boosted pressure and flow from the fuel supply

unusual engine behavior of a nature that does not automatiĆ

system to meet main and AB fuel requirements. The pump

cally initiate data recording. This data is valuable for

receives fuel at aircraft boost pressure and boosts fuel

diagnosis of the cause of unusual behavior and should be

pressure to levels adequate to operate the engine at all power

recorded by the pilot by depressing the ENG RCD button on

settings

(maximum 40−psi pressure rise). During non−AB

the fuel management panel. Pressing the ENG RCD button

operation, some fuel is circulated between AB fuel control

momentarily causes

21 seconds of engine data to be

and the engine fuel boost pump so that fuel pressure is readily

recorded:

6 seconds before and 15 seconds after switch

available to the spray bars for AB light−off.

initiation. It is important to remember that if a transient

problem is to be recorded by FEMS, the ENG RCD button

2.4.3

Main Fuel Pump

must be activated quickly so the actual event is not missed.

The main fuel pump is a two−stage pump that receives

Manual recording will not interfere with data automatically

fuel flow from the engine fuel boost pump. It provides

saved by the FEMS.

additional fuel pressurization and transmits mechanical−

gear−driven power to the MEC from the gearbox.

2.3.3

FEMS and OBC

FEMS is checked during OBC preflight and in flight

2.4.4

Main Engine Control

(Class III). It is designated by a FEM acronym. This acronym

The MEC is a fuel−operated, hydromechanical fuel

is displayed at the completion of OBC if FEMS fails its BIT

flow regulator that operates in tandem with the main fuel

during OBC. Engine−life tracking data is still available

pump and is capable of operating in two modes. In the

through EMSP if FEMS is lost.

primary mode, it meters main fuel flow as commanded by the

AFTC and provides VSV scheduling. The secondary mode

2.4

ENGINE FUEL SYSTEM

hydromechanically meters main fuel flow to govern N₂ speed

The engine fuel system, which is identical for each

based on pilot throttle commands and provides basic engine

engine, provides motive flow fuel to effect fuel transfer and

control except for AFTC fan speed limiting.

metered fuel for combustion as a function of pilot throttle

commands and numerous engine parameters (Figure 2Ć17).

2−21

ORIGINAL

NAVAIR 01−F14AAD−1

VSVs aerodynamically match high− and low−pressure

2.5

THROTTLES

compressor stages by changing the angle at which airflow

enters the compressor rotor blades. The MEC contains the

Two throttle levers for regulating engine thrust are on

scheduling mechanism and provides fuel pressure to vary

the left console of the forward cockpit. Unrestricted engine

VSV positioning. A flexible mechanical cable provides feedĆ

operation under independent control is afforded; however,

back from the compressor stator to the MEC.

normal symmetric thrust control is provided by collective

movement of the throttle levers. Numerous engine control

2.4.5

Afterburner Fuel Pump

and subsidiary functions are performed by movement of the

throttle levers within the full range of travel as shown in

The AB fuel pump is a centrifugal gear−driven pump

Figure 2Ć19. The forward and aft throw of each throttle lever

that receives fuel from the engine boost pump, increases

in the quadrant is restricted by hard detents at the OFF, IDLE,

pressure, and delivers fuel to the AB fuel control. During

MIL, and MAX (AB) positions. At the OFF and IDLE detĆ

non−AB operation, fuel is circulated between the AB fuel

ents, the throttles are spring loaded to the inboard position.

control and the engine boost pump; the AB fuel pump

At the MIL detent, the throttles can be shifted outboard to the

impeller runs dry with the bearings lubricated by the engine

AB sector or inboard to the basic engine sector of operation

oil system. Failure of an AB fuel pump will result in an AB

by merely overcoming a lateral breakout force. Lateral shiftĆ

blowout.

ing of the throttles at the MIL detent does not affect engine

2.4.6

Afterburner Fuel Control

control. Thus, placement of the throttle outboard at MIL

provides a natural catapult detent to prevent unintentional

The AB fuel control is a fuel−operated, electrohydroĆ

retarding of the throttles during the launch. This, however,

mechanical unit that regulates fuel flow in response to AFTC

does not inhibit the selection of afterburner. The friction

scheduling and compressor discharge pressure. Fuel pressure

control lever on the outboard side of the quadrant permits

from the AB fuel control provides on−off signals to the AB

adjustment of throttle friction to suit individual requireĆ

fuel pump.

ments. With the friction lever in the full aft position, no

throttle friction is applied at the quadrant; increased throttle

The AB fuel control splits fuel flow into three metered

friction is obtained by forward movement of the lever.

streams (local, core, and fan) on a sequential basis into the

AB manifolds for distribution through spraybars in the AB

A locking pin device prevents the left throttle from

duct. Throttle commands initiate local fuel flow and AB

moving into the cutoff position when the right throttle is

ignition (minimum AB). Once local fuel flow and flame are

either traversing or at rest on the face of the right−hand idle

established, core fuel flow commences. As maximum core

stop block.

fuel flow is established, fan fuel flow commences and

increases until maximum AB is achieved. The transitions

between local, core, and fan fuel flow are smooth and

2.5.1

Throttle Control Modes

unnoticed

(Figure 2Ć18). During non−AB operation, fuel

flow is circulated through the AB manifolds to prevent thrust

Manual, boost, and automatic are the three modes of

lags and surges when AB is initiated.

throttle control over engine operation selectable by the

THROTTLE MODE switch located outboard of the quadrant

on the pilot console. The toggle switch must be lifted out of

a detent to select MAN from BOOST or BOOST from MAN.

The switch is solenoid held in AUTO upon successful

D Zero− or negative−g flight longer than 10 secĆ

engagement of the automatic mode. A functional schematic

onds in AB or 20 seconds in MIL or less will

of throttle control modes, including system major compoĆ

deplete the fuel sump tanks (cells 3 and 4),

nents, is shown in Figure 2Ć20. Except for the autothrottle

resulting in flameout of both engines.

computer and mode control switch, the throttle control

system for each engine is completely redundant. Independent

D To prevent engine instability and/or flameout,

engine operation is possible in the manual or boost mode of

avoid holding zero or negative g when doing

throttle control; however, full system operation is necessary

a low−altitude, maximum−thrust acceleration.

in the automatic mode since operation under single−engine

D With fuel in feed group below 1,000 pounds,

control is impracticable because of asymmetric thrust

AB operation could result in AB blowout.

considerations.

Note

Fuel dump operations with either engine in AB

are prohibited. The fuel dump mast can be

torched.

2−23

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć19.ĄThrottle Interlocks

2.5.1.1

Manual Throttle Mode

electric clutch in the throttle servoactuator, which is also

mounted to the power lever shaft, is disengaged in the manual

The manual throttle is a degraded mode of operation

mode to reduce operating forces.

and was designed as a backup system. Because of hysteresis

With the throttle friction lever in OFF, approximately

and friction in the manual system, engine rpm may vary from

8 pounds of force per throttle must be applied at the grip to

the boost mode at a given throttle position. If an engine fails

operate the throttles in the IDLE to MAX range.

to secure when the throttle is moved to the OFF position, the

throttles have probably reverted to the manual mode and are

2.5.1.2

Boost Throttle Mode

slightly out of rig. Cycling the throttle switch to MAN and

back to BOOST may allow engine shutdown. If shutdown is

The boost mode of throttle is used for normal

unsuccessful, then the engine may be secured with the FUEL

operations. A force of 2 to 3 pounds at the grip is required to

SHUTOFF handle.

move each throttle throughout its range with the throttle

friction lever off. Essentially, the boost mode provides

electric throttle operation, with the push−pull cables serving

as a backup control path. Throttle movement is detected by

the throttle position sensor. The signal is resolved in the

amplifier to provide positional followup commands to the

D Engine shutdown at high power settings using

actuator. Movement of the actuator rotates the engine power

the FUEL SHUTOFF handle may result in

lever shaft, which drives the push−pull cable.

damage to the aircraft fuel system.

If a boost system malfunctions, applying approxiĆ

D Engine startup in manual mode may cause

mately 17 pounds at the throttle grip automatically reverts the

tailpipe fires as fuel flow may not be secured.

throttle control to the manual mode by disengaging the

In the manual mode of operation, movement of each

actuator electric clutch. The throttle control reverts to

throttle is mechanically transmitted to the respective engine

manual mode in 0.25 second. In the event of a boost system

by a push−pull cable and a rack and sector mechanism

malfunction, the throttle mode switch will remain in the

mounted to the main engine control power lever shaft. An

BOOST detent. By manually placing the throttle mode

2−25

ORIGINAL

NAVAIR 01-F14AAD-1

switch in MAN and then back to BOOST, transient failures

selecting AUTO to ensure a valid test. Once AUTO is

in the boost mode can be reset. Additionally, if an actuator

engaged, the control stick should be programmed fore and aft

seizes, a mechanical clutch in the actuator will slip when a

to check for the appropriate power response.

force of approximately 50 pounds is applied at the throttle

grip. This permits the pilot to override an actuator seizure.

There is no visible warning of these anomalies only the

noticeable increase in the forces required to manipulate the

affected throttle.

High-power settings may result during aft stick

2.5.1.3

Approach Power Compensator

deflection.

(Automatic Throttle Mode)

If the THROTTLE MODE switch does not remain

The automatic mode of throttle control is a closed-loop

engaged or the APC does not respond properly to indicated

system that automatically regulates basic engine thrust to

AOA and longitudinal stick movements, a malfunction exists

maintain the aircraft at 15 units angle of attack for landing.

in the autothrottle system.

All components of the throttle control system except the

throttle position sensor are used in the automatic mode of

Depressing and holding the autopilot emergency dis-

control. The AOA signal from the AOA probe on the left side

engage paddle switch with weight on wheels causes the

of the forward fuselage is the controlling parameter within

throttle control system to be placed in the manual mode. If the

the autothrottle computer. Additional parameters are inte-

automatic mode was selected before depressing the paddle

grated within the computer to improve response. The air

switch, the THROTTLE MODE switch will automatically

temperature switch on the pilot left console effects a

move to BOOST. The THROTTLE MODE switch must be

computer gain change to compensate for pilot-preferred

moved from BOOST to MAN while holding the paddle

reaction rate. In order to engage the autothrottle, throttles

switch depressed if the manual mode is desired after the

must be between 75 to 90-percent rpm with weight off

paddle switch is released.

wheels, gear handle down, and throttle friction off. With all

conditions met, the throttle mode switch will be held by an

2.6

ENGINE BLEED AIR

electrical solenoid when placed in AUTO. The throttle

control mode automatically reverts to the boost mode upon

Bleed air is extracted from the high-pressure compres-

interruption of any interlock in the system or by manually

sor to perform engine-associated services and to supply hot,

overriding the throttles with a force of approximately

high pressure air for operation of auxiliary equipment.

11 pounds per throttle in either direction. The throttle

Fifth-stage bleed air supplies hot air for the engine anti-icing

mode switch automatically returns to BOOST and the

system and is used to draw cooling air through the aircraft

AUTO THROT caution light illuminates for 10 seconds.

hydraulic heat exchangers to cool flight and combined

See Figure 2-21 for autothrottle controls.

fluids and to ventilate the nacelle when weight is on wheels

(Figure 2-22). Ninth-stage bleed air supplies hot air to the

The pilot can revert from automatic to boost mode by

environmental control system, provides air for crossbleed

selecting the CAGE/BRST (UP) position on the CAGE/

engine starts, and draws air through the integrated drive

SEAM switch located on the inboard throttle grip. This pro-

generator heat exchanger (ventral fin) when weight is on

vides a smooth throttle override for an automatic-to-boost

wheels.

mode approach, while maintaining a grip on both throttles.

2.6.1

Engine Anti-Ice

2.5.1.3.1

Autothrottle Test

The fan IGV and nosedome are susceptible to icing

An automatic check of the autothrottle control system

undera widerrange ofconditions,particularlyatstatic orlow

while on deck is accomplished during OBC. Signals to the

speed with high engine rpm, than that which cause ice to form

servoactuators are inhibited during the OBC autothrottle test

on external surfaces of the airframe. Ice formation at the fan

so that the engines remain at idle thrust. A malfunction is

face can restrict engine maximum airflow, which results in a

indicated by an APC acronym at the conclusion of OBC.

thrust loss, decreased stall margin, and dislodgment of ice,

which can damage the compressor. The engine anti-icing

Rotating the MASTER TEST switch to FLT GR DN

system is designed to prevent the formation of ice rather than

and depressingit bypassesthe autothrottle weight-on-wheels

de-ice the IGV and nose dome. Hot bleed air (5th stage) is

interlock and an end-to-end check of the autothrottles may be

passed through the hollow IGV to the nose dome and is

performed on deck. The throttles should be placed at about

discharged into the engine along the vanes and at the rotor

80-percent rpm and the throttle mode switch placed in

hub. Cockpit control of the engine anti-icing system is

AUTO. The throttles must be positioned above idle before

effected through the ANTI-ICE switch (Figure 2-23).

2-27

CHANGE 1

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

THROTTLE MODE switch

AUTO Ċ Engine thrust is automatically regulated by the throttle control computer

to maintain optimum angle of attack for landing.

BOOST Ċ Normal operating mode. Reduces effort required to move throttles

manually with friction control aft.

MAN Ċ Movement of each throttle is mechanically transmitted to the respective

engine cross−shaft by a push−pull cable.

THROTTLE TEMP switch

Used with the AUTO throttle mode to effect throttle computer gain changes to

compensate for air temperature.

HOT Ċ Increases normal throttle computer gain.

NORM Ċ Normal throttle computer gain.

COLD Ċ Decreases normal throttle computer gain.

AUTO THROT caution light

Auto throttle mode is disengaged. During preflight check, remains illuminated for

10 seconds, then goes off and throttle mode switch automatically returns to BOOST.

Note

If the auto throttle is disengaged by deselecting the throttle

mode switch, the AUTO THROT light will not illuminate.

CAGE/SEAM switch

When in TLN master mode with the throttle mode switch in AUTO, selecting the

CAGE/BRST position on the CAGE/SEAM switch reverts the throttles to the

BOOST mode.

Autopilot emergency paddle

Reverts throttle system from AUTO or BOOST mode to MAN mode only while

disengage depressed and with weight on wheels.

Figure 2−21. Autothrottle Controls and Indicators (Sheet 2 of 2)

Note

from 520° to 1,180_ F inside the bleed air portion of the

ECS, and from 400_ to 500_ F inside the hot air portion

Because of its adverse effects on engine perforĆ

(400_ F manifold).

mance, the engine anti−icing system should be

used only when icing conditions exist or are

The entire bleed air portion of the ECS, from engine

anticipated.

bleed air shutoff valves to the primary heat exchanger, is monĆ

During engine start, the engine anti−ice valve remains

itored by two detection systems. Fire detection circuits moniĆ

open to bleed the compressor to prevent engine stall. The

tor the bleed air system from each engine to its respective

valve closes when the engine approaches idle rpm. In flight,

firewall. When a fire detection circuit in an engine compartĆ

the valve is normally closed unless the ANTI−ICE switch is

ment senses temperatures above threshold, the appropriate

in ORIDE/ON, or AUTO/OFF, when the ice detector probe

L or R FIRE warning light illuminates (refer to fire detection

in the left inlet is activated. Ice accumulation on the ice

system). The remainder of the bleed air system, from engine

detector illuminates the INLET ICE caution light. The

firewalls to the primary heat exchanger, is monitored by bleed

engine anti−icing control valve on the engine is powered

air leak−sensing elements. When the bleed air leak−detection

closed (fails open) from the essential dc No. 2 bus through the

circuit detects temperatures in excess of 575_ F, the BLEED

ENG/PROBE/ANTI−ICE circuit breaker (RG2).

DUCT caution light illuminates.

2.6.2

Environmental Control System

The hot air portion of the ECS is monitored by hot air

Leak Detection

leak−sensing elements. The hot air system extends from the

primary heat exchanger through the 400_ manifold to the

Thermal detection circuits are routed in proximity to

cockpit floor. When the hot air detection circuit detects

ECS ducts and components to provide cockpit indications of

temperatures in excess of 255_ F, the BLEED DUCT caution

high−temperature air leaks. Normal air temperatures range

light illuminates.

2−29

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć22.ĄEngine Bleed Air/Compartment Ventilation

2.7

ENGINE COMPARTMENT VENTILATION

2.7.2

Engine Ground Ventilation

With weight on wheels, cooling airflow through the

Each engine compartment is completely isolated from

engine compartment is induced by the hydraulic heat

the primary air inlet, and the efficiency and cooling of the

exchanger ejector in the forward end of the compartment. Air

variable−area exhaust nozzle are not dependent upon nacelle

enters through the nacelle ram−air scoop on the left side, passes

airflow. Therefore, within the bounds of the forward firewall

through the hydraulic heat exchanger and is discharged into

(landing gear bulkhead) and the nozzle shroud, the cooling

the engine compartment. The air flows through the full length

system for each engine compartment is a separate entity.

of the nacelle to discharge overboard through a louvered port

Cooling requirements for the turbofan engine are mini−

atop the nacelle on the outboard side of the vertical tail.

mized by the annular fan bypass duct. Figure 2Ć22 shows

cooling airflow patterns through the engine compartment

2.8

ENGINE IGNITION SYSTEM

during ground and flight operations. Two air−cooled heat

exchangers are also shown; however, only the hydraulic

There are three electrical ignition circuits, each utilizing

heat exchanger cooling airflow is associated with engine

a dedicated igniter, for each engine: main high energy, afterĆ

nacelle cooling. Fire access doors are on the outboard side

burner, and backup.

of the nacelles at the forward end to permit application of

fire suppressing agents by ground personnel in event of an

2.8.1

Main High−Energy Ignition

engine compartment fire.

The main high−energy ignition provides ignition in the

2.7.1

Engine In−Flight Ventilation

combustion chamber for ground and air starts. It is powered

by one of the four windings in the engine−driven ac alternator.

In−flight cooling of the engine compartment is accomĆ

The AFTC provides logic to control main high−energy igniĆ

plished by nacelle ram−air scoops, circulating boundary−

tion automatically. Ignition is available when N2 rpm is 10

layer air through the length of the compartment and expelling

percent or greater and is automatically provided from 10−to

the air overboard through louvered exits, just forward of the

59−percent rpm when the throttle is above cutoff. Ignition

engine nozzle shroud.

ORIGINAL

2−30

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

ANTI−ICE switch

ORIDE/ON Ċ Overrides ice detector system to turn on INLET ICE caution

light, and activate external probe heaters and engine anti−ice.

Commands the anti−ice mode to the AICS programmers.

AUTO/OFF Ċ When icing is sensed, ice detector activates engine anti−ice

system, turns on INLET ICE caution light, activates external

probe heaters with weight off wheels, and disables anti−ice

mode to AICS programmers.

OFF/OFF Ċ Engine anti−ice system and probe heaters shut off. INLET

ICE caution light disabled. Disables anti−ice mode to AICS

programmers.

INLET ICE caution light

Illuminates when ice accumulates on ice detector with ANTI−ICE switch in

AUTO/OFF or if ORIDE/ON is selected. Does not illuminate with switch

in OFF/OFF.

BLEED DUCT caution light

Illuminates when bleed air leak sensing elements detect temperatures

greater than 575_ F between the left and right firewalls, past the primary

heat exchanger and up to the right diverter area. Also illuminates when an

additional sensor detects temperatures of 255_ or greater from the right

diverter area, along the 400_ manifold and into the bootstrap turbine

compartment.

Figure 2Ć23.ĄAnti−Ice Control

2−31

ORIGINAL

NAVAIR 01−F14AAD−1

is secured 0.5 second after N₂ rpm rises above 59 percent. At

The air hose is connected to the aircraft fitting in the

rpm above 59 percent, ignition is provided if N₂ deceleration

left sponson area, behind the main gear strut. Ground start air

exceeds a 5 percent rpm per second rate. Ignition continues

is ducted into a central bleed air (9th stage) manifold, which

for 20 seconds after N₂ deceleration falls below the 5 percent

interconnects the air turbine starters on both engines.

rpm per second rate. Main high−energy ignition is provided

The air supply to each air turbine starter is pressure reguĆ

continuously when the engine is in the secondary (SEC) mode.

latedĂ(52.5 psi) and controlled by a shutoff and regulating

valve at the turbine. Each pneumatic starter is composed of

2.8.2

Afterburner Ignition

a turbine, gear train, sprag clutch with a speed−sensing

device, and an overspeed disengagement mechanism with a

The AB ignition ignites AB fuel flow for AB light−offs

shear section. Shutoff valves in the bleed air manifold selecĆ

and relights (in the event of an AB blowout). The AB ignition

tively isolate the other starter, subsidiary bleed lines, and the

is powered by the same winding in the engine−driven alternaĆ

environmental control system air supply. Maximum engine

tor that powers the main energy ignition. The AFTC provides

motoring speed with the pneumatic starter is approximately

logic to control AB ignition automatically and prevents

30−percent rpm.

simultaneous powering of the main high−energy and AB igniĆ

tions. In the event of an AB blowout, relight is normally

2.9.2

Engine Crank

provided within 1.5 seconds. AB ignition is not powered if

the engine is in SEC mode.

Placing the ENG CRANK switch in either L or R opens

the corresponding starter pressure shutoff valve to allow

2.8.3

Backup Ignition

pressurized air to drive the turbine. The ENG CRANK switch

energizes the appropriate shutoff valves to configure the

The backup ignition provides ignition in the combusĆ

bleed manifold for starting.

tion chamber for ground and air starts when the BACK UP

IGNITION switch on the THROTTLE CONTROL panel is

2.9.2.1

Engine Crank Switch

set to ON. It is powered by the essential No. 1 ac bus and

provides less power than main high−energy ignition. After

The ENG CRANK switch is held in L or R by a holding

use, the BACK UP IGNITION switch should be set to OFF.

coil. At approximately 50−percent rpm, a centrifugal cutoff

To allow ground checkout of backup ignition, main high−

switch closes the turbine shutoff valve and returns the ENG

energy ignition is disabled when the BACK UP IGNITION

CRANK switch to the center or off position. A START

switch is ON and weight is on wheels.

VALVE caution light illuminates if the starter valve remains

in the open position after the ENG CRANK switch automatiĆ

cally returns to the center (off) position.

The BACK UP IGNITION switch shall be

selected to OFF prior to applying external elecĆ

trical power to prevent ignition of fuel puddled

D If the starter valve does not close during

in the engine.

engine acceleration to idle rpm, continued airĆ

flow through the air turbine starter could

2.9

ENGINE STARTING SYSTEM

result in catastrophic failure of the starter

turbine.

Each engine is provided with an air turbine starter that

D If the START VALVE caution light illumiĆ

may be pressurized from an external ground starting cart or

nates after the ENG CRANK switch is off,

by crossbleeding high−pressure bleed air from the other

select AIR SOURCE to OFF to preclude

engine. Figure 2Ć24 shows the components associated with

starter overspeed.

the engine start system.

D If the ENG CRANK switch does not automatiĆ

cally return to the OFF position by 50 percent,

2.9.1

External Airstart

ensure that the ENG CRANK switch is off by

60−percent rpm to avoid starter turbine failure

A high−pressure (75 psi) air source and 115 volt, 400 Hz

as a result of an inoperative automatic starter

ac power are required for engine start on the deck.

cutout.

ORIGINAL

2−32

NAVAIR 01−F14AAD−1

Return of the ENG CRANK switch to the center or off

D When attempting a crossbleed or normal

position resets the bleed air manifold valves to permit 9th−

ground start, do not attempt to reengage the

stage bleed air to flow to the environmental cooling system

ENG CRANK switch if the engine is spooling

and ejectors in the engine compartment.

down and rpm is greater than 46 percent. At

rpm’s of 30 to 46−percent rpm, the ENG

Starter cranking limits:

CRANK switch may not stay engaged

because of normal variations in starter cutout

1. Crossbleed

 2 minutes continuous then

speed.

minutes OFF.

D The ENG CRANK switch should automatiĆ

cally disengage between 49 to 51−percent rpm

2. Start cart  5 minutes continuous then 10 minutes

OFF.

during a crossbleed or normal ground start.

2.9.4

Airstarts

2.9.3

Crossbleed Start

AFTC logic provides main high−energy ignition autoĆ

Engine cranking procedures during a crossbleed start

matically during automatic and manual spooldown, crossĆ

are the same as with a ground start cart. Engine crossbleed

bleed, and windmill airstarts. Selecting the BACKUP

start on the ground can be accomplished with the throttle on

IGNITION switch to ON provides continuous backup

the operating engine at or above idle rpm. When high−

ignition to both engines, and backs up main high−energy

residual EGT (remains from a hot start) and/or throttles are

ignition during manual spooldown, crossbleed, and windmill

advanced from OFF to IDLE prior to 20−percent rpm, higher

airstarts.

than normal EGT readings may occur.

2.10 ENGINE OIL SYSTEM

When initiating crossbleed starts with ambient temperĆ

ature less than

40° F (4° C), the starter torque load is

increased. Above 80° F (27° C), engine bleed air provides

Each engine has a self−contained, dry sump non−

less energy potential to the starter turbine. Either extreme can

pressure regulated oil system that provides filtered oil for

affect engine starting acceleration rates, resulting in hotter−

lubricating and cooling engine main shaft bearings, oil seals,

than−normal starts. When crossbleed starting with an operatĆ

gearboxes, accessories, and provides a hydraulic medium to

ing engine at idle, the operator should be aware of either

operate the engine exhaust nozzles (FO−5).

condition and increase the operating engine rpm in 5−percent

increments until normal starting acceleration rate is

A storage tank feeds oil to an oil pump that supplies oil

achieved. Low percentage rpm−to−EGT ratio can increase

under pressure to the forward sump in the engine front hub,

turbine distress without necessarily exceeding the EGT limit.

the mid sump in the fan hub, the aft sump in the turbine hub,

and the inlet and accessory gearboxes. Oil is recovered from

When performing an idle crossbleed start, advance the

the sumps and accessory gearboxes, pumped past a chip

throttles from OFF to IDLE at 20−percent rpm or greater

detector, and cooled in a fuel/oil heat exchanger before

while monitoring EGT. If EGT rises rapidly, advance the

returning to the storage tank.

operating engine rpm to slightly above idle. The exhaust

nozzles start to close when rpm is slightly above idle.

A separate compartment in the storage tank provides

oil to the exhaust nozzle hydraulic system. Oil returning from

Note

the nozzle to the tank provides auxiliary oil supply to the

No. 3 bearing when normal supply is interrupted or during

engine spooldown.

D To prevent possible engine overtemperature

during crossbleed and backup ignition start

The oil system permits engine operation under all flight

attempts, select AIR SOURCE for the operĆ

conditions. During zero− or negative−g flight, oil pressure may

ating engine and return to BOTH after rpm

decrease to zero but will return to normal when positive−g

stabilizes at idle or above.

flight is resumed. Normal oil consumption is 0.03 gallon per

D If attempting a ground restart after a hot start,

operating hour with the maximum being 0.1 gallon per operĆ

windmill the engine until EGT is below

ating hour. Capacity of the oil storage tank is 3.7 gallons,

250° C prior to advancing the throttle from

with 2.9 gallons usable. A sight gauge on the side of the

OFF to IDLE to avoid a subsequent hot start.

ORIGINAL

2−34

NAVAIR 01−F14AAD−1

storage tank indicates down to a 2−quart−low oil level. The

300° F during a temperature increase and go out at 280° F

protrusion of a bypass indicator underneath the oil scavenge

minimum during a temperature decrease. The caution lights

pump indicates a clogged filter element.

also illuminate when respective forward engine gearbox

scavenge temperature exceeds 375° F during a temperature

Note

increase, and go out at 345° F minimum during a temperature

decrease.

D Engine oil level must be checked within

2.11 ENGINE INSTRUMENTS

30Ăminutes of engine shutdown, otherwise the

engine must be run at 80−percent rpm or greater

Instruments for monitoring engine operation are on the

for 10 minutes to ensure proper servicing.

pilot left knee panel (Figure 2Ć25). Engine operating parameĆ

D A failed−open nozzle may indicate an oil leak;

ters are displayed on the engine instrument group which is a

however, if the leak is in the nozzle hydraulic

single WRA with LCD readouts. The display provides white

circuit, only that portion of the main engine

readout segments and scales on a dark backg0round and is red

lube oil will be lost.

backlighted for night operations. Left and right engine comĆ

pressor speed (rpm), EGT, and fuel flow are displayed on the

2.10.1

Oil Cooling

EIG. Adjacent to the EIG are circular instruments for both

engines’ oil pressure and nozzle position. Takeoff checks at

Filtered and scavenged oil is cooled by a fuel/oil heat

military (MIL) thrust should display evenly matched tapes on

exchanger. This oil is then used in a heat exchanger to cool

corresponding vertical scale instruments and all pointers on

the exhaust nozzle oil. A cold−oil bypass valve opens when

the circular instruments should be at the 9−o’clock position.

the heat exchanger pressure differential is 44 psi, because of

Data on engine operating limits are provided in Chapter 4.

reduced oil temperature or exchanger blockage, allowing oil

flow to bypass the heat exchanger (for example, during cold

2.11.1

Engine RPM Indicator

engine starts).

The RPM indicators (Figure 2Ć25) have a range of 0 to

2.10.2

Oil Pressure Indicators

110 percent. The tape display steps in 5−percent increments

and the upper segment flashes to indicate rpm increasing at

An oil pressure transducer in each engine’s oil supply

more than 0.4 percent per second from 0 to 60−percent rpm.

line provides a continuous signal to the oil pressure indicator.

The tape steps in 1−percent increments when greater than

Another, independent oil pressure switch in each oil supply

60−percent rpm. Nominal indications are 62 to 78 percent at

line activates the OIL PRESS light when either engine’s oil

idle and 95 to 104 percent at military and above. At 107.7

pressure decreases to 11 psi. The oil pressure switches and

percent and above, the affected engine’s exceeded portions

lights receive electrical power from the essential No. 2 ac

of the chevrons will flash at a rate of 2 to 3 flashes per second.

bus. The OIL PRESS light and oil pressure indicator are

At 20−percent rpm a horizontal segment will illuminate

independent of each other.

giving an indication of proper motoring speed to start the

engine. There is an rpm reading for each engine.

Note

D During cold starts, oil pressure may exceed

An overspeed condition in excess of 110 percent

65 psi. The 65 psi oil pressure limit should not

will result in momentary loss of rpm indication

be exceeded for more than 1 minute.

until N₂ rpm falls below 110 ± 0.5 percent. EGT

D Maneuvers that result in zero or negative−g on

and fuel flow indicators will continue to function

the engine (such as rapid rolls, pushovers, or

normally.

bugout maneuvers) may cause oil pressure

fluctuations and momentary illumination of

2.11.2

Exhaust Gas Temperature Indicator

the low oil−pressure light.

The EGT indicators (Figure 2Ć25) provide a nonlinear

2.10.3

OIL HOT Caution Lights

vertical scale with a range of 0 to 1,100° C. The compressed

lower portion has a range of 0 to 600° C. The expanded

The L or R OIL HOT caution light may be illuminated

upper portion of the scale has a range of 600 to 1,100° C.

by either high engine oil temperature or by high forward−

The display moves in 50° increments in the compressed

engine gearbox scavenge oil temperature. The caution lights

portion and 10° increments in the expanded portion of the

illuminate when respective engine oil temperature exceeds

2−35

ORIGINAL

NAVAIR 01−F14AAD−1

display. The normal indications are 350 to 650° C at idle and

2.11.6

Engine Oil Pressure Indicator

780 to 935° C at MIL and above. Above 940° C, the affected

engine’s exceeded portions of the chevrons flash. With a

The engine oil pressure indicators display oil pressure

reading of 940° C, the stall warning light and the aural

from 0 to 100 psi. Normal oil pressure is 25 to 65 psi and

warning tone will be activated signifying an engine overĆ

increases in proportion to engine rpm within the pressure

temperature condition. The tone is present for a maximum of

limit range. Stabilized idle oil pressure may be a minimum

10 seconds unless the fault clears sooner. There is an EGT

of 15 psi. The OIL PRESS caution light illuminates at 11 psi

reading for each engine.

with decreasing oil pressure and extinguishes at 14 psi with

increasing oil pressure. Maximum allowable oil pressure

2.11.3

Fuel Flow Indicator

fluctuation is ± 5 psi.

The fuel flow indicators have a nonlinear vertical scale,

2.11.7

Exhaust Nozzle Position Indicator

with a range of 0 to 17,000 pph. The expanded lower portion

of the scale has a range of 0 to 5,000 pph. The compressed

The nozzle position indicators

(Figure 2Ć25) have

upper portion of the scale ranges from 5,000 to 17,000 pph.

a range of

0 to

100−percent open. Normal indications

The display moves in 100 pph increments in the expanded

(Figure 2Ć13) are l00 percent at idle with WOW and vary in

portion and in 500 pph increments in the compressed portion

flight:

3 to 10 percent at MIL thrust, 5 to 12 percent at

of the display. Normal indications on deck are 350 pph

MIN AB, and 60 to 90 percent at MAX AB.

starting, 950 to 1,400 pph at idle, and approximately 10,100

pph at military and above. The fuel flow reading for each

Note

engine indicates only basic engine consumption and does not

indicate AB fuel flow.

When operating engine in SEC mode, the nozzle

position indicator is inoperative and indicates

2.11.4

Engine Instrument Group BIT

below zero. No nozzle position indication is

available in SEC mode.

A degraded mode of EIG operation is indicated if the

2.11.8

Engine Monitor Display Format

BIT segment on the top left side of the EGT indicator illumiĆ

nates. This means that either the primary or backup microĆ

processors, or the primary or backup power supply channels

A display of engine parameters (Figure 2Ć26) can be

(internal to the EIG), have failed. An automatic switch to the

selected on the MFD by pressing the pushbutton adjacent to

operative microprocessor/channel takes place if a failure is

the ENG legend on the own−aircraft menu. The display preĆ

detected. The instrument still monitors engine operation and

sents N1 (fan speed), TBT (turbine blade temperature), FF/M

accurately reflects rpm, EGT, and fuel flow. If the input

(fuel flow main engine) or FF/T (fuel flow total, main engine

processing circuit fails, the affected scale reading goes to zero.

and AB), and NP (exhaust nozzle position). FF/M scale indiĆ

cates main engine fuel flow and is similar to the fuel flow

2.11.5

Engine Instrument Group Self−Test

displayed on the EIG. NP is the same as the nozzle position

indicators. Numerical readouts below the N1 and TBT vertiĆ

cal scales digitize the indicated value. The TIME readout

EIG self−test is selected by the MASTER TEST switch

below the FF/M vertical scale indicates time in hours and

in INST. When master test is selected, all display segments

minutes that fuel will last based on current consumption

illuminate, scales drive to maximum readings, and warning

rates. Directly below the TIME readout, engine faults are

chevrons (stripes) flash for 5 seconds. BIT segment on the top

displayed based on current engine operating conditions of

left side of EGT indicator illuminates. L and R STALL

both engines processed by FEMS. If more than three faults

warning acronyms appear on the HUD and MFD and stall

exist at the same time, the acronyms will continuously scroll

warning/overtemp tone is present in pilot earphones for

upward. The ten possible acronyms and their definitions are:

10 seconds. After 5 seconds, all EIG scales decrease to preĆ

determined values of equal height that correspond to an EGT

1. L MACH # or R MACH #  Mach number signal

reading of 950 ± 10° C. If BIT segment remains illuminated,

to designated engine has failed.

EIG has failed self−test and BIT remains illuminated until

self−test is reinitiated. Total self−test time is 15 seconds. If

2. L LO THR or R LO THR  Designated engine may

master test is deactivated prior to this, EIG returns to normal

be producing less than expected thrust.

mode after the 15−second test. If the MASTER TEST switch

remains in INST for more than 15 seconds, the EIG retains

3. L A/ICE or R A/ICE  Designated engine anti−ice

equal height readings until master test is deselected.

is on or anti−ice valve has failed opposite comĆ

manded position.

2−37

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć26.ĄMFD Engine Monitor Display

4. L OIL LO or R OIL LO  Designated engine oil

4. L FLMOUT or R FLMOUT  Designated engine

level is approximately two quarts low. Postflight,

flameout.

engine at idle.

5. L IGV SD or R IGV SD  Designated engine inlet

5. L AUG or R AUG  Designated AB control system

guide vane off schedule.

has failed. AB is not available.

6. L STALL or R STALL  Designated engine stall

Refer to Chapter

12, WARNING/CAUTION/

detected (also on HUD).

ADVISORY LIGHTS/DISPLAY LEGENDS for the

appropriate pilot/RIO response.

7. L FIRE or R FIRE  Designated engine fire/

overheat condition in engine nacelle (also on HUD).

2.11.9

MFD Engine Caution Legends

Refer to Chapter

12, WARNING/CAUTION/

In addition to the engine caution lights on the pilot

ADVISORY LIGHTS/DISPLAY LEGENDS for the apĆ

CAUTION/ADVISORY panel, illumination of the READ

propriate pilot/RIO response.

MFD caution light indicates that one or more of the following

caution legends on the upper left quadrant of the MFD is

2.11.10

Engine Stall/Overtemperature Warning

activated:

An engine stall detection circuit in FEMS monitors

1. L N2 OSP or R N2 OSP  Designated engine N₂

each engine. When a stall condition is detected, a L or R

overspeed condition.

STALL warning legend is displayed on the HUD and MFD

until the condition is cleared. In addition, an aural warning

2. L N1 OSP or R N1 OSP  Designated engine N₁

tone is activated through the pilot ICS for up to 10 seconds.

overspeed condition.

There is no pilot check of the FEMS engine stall detection

3. L TBT OT or R TBT OT  Designated engine

system.

turbine blade overtemperature.

ORIGINAL

2−38

NAVAIR 01−F14AAD−1

Note

the entire length of the engine compartment on both sides

above the nacelle door hinge line. The tube sheath, which is

In SEC mode, FEMS and, therefore, the engine

clamped in grommets to the engine compartment structure,

stall detection circuit, is inoperative. However,

contains a ceramic−like thermistor material in which are

overtemperature warning is still available and

embedded two electrical conductors; one of the conductors

will activate both the STALL warning legends

is grounded at both ends of the loop. Electrical resistance

and aural warning tone.

between two conductors varies inversely as a function of

temperature and length, so that heating of less than the full

When an overtemperature condition occurs, the EGT

length will require a higher temperature for the resistance to

display rises above 940° C, the warning chevrons begin to

decrease to the alarm point. The L or R FIRE warning lights

flash, and a signal from the EGT indicator activates the

in the cockpit illuminate when the respective entire sensing

STALL warning legend and the aural tone. The overtemperaĆ

loop is heated to approximately 600° F or when any 6−inch

ture warning system is checked by the pilot during prestart as

section is heated to approximately 1,000_ F.

part of the MASTER TEST check in INST test.

The fire alarm output relay to the light is a latching type

2.12 FIRE DETECTION SYSTEM

that remains in the last energized position independent of

power interruptions until the fault clears.

The fire detection system provides a cockpit indication

of fire or overheating in either engine compartment. There is

False alarms triggered by moisture in the sensing eleĆ

a separate system for each engine compartment, each consistĆ

ment and connectors or by damage resulting in short circuits

ing of a thermistor−type sensing loop monitored by a transisĆ

or grounds in the sensing element are unlikely because of the

torized control unit. The system is powered by 28 volts from

system design. Additionally, there is no loss or impairment

the essential dc No. 1 bus. Figure 2Ć27 is a functional scheĆ

of fire detector capability from a single break in the sensing

matic of the system.

element as long as there is no electrical short. With two

breaks in the sensing element the section between the breaks

The sensing loop for each engine compartment consists

becomes inactive although the remaining segment ends

of a 45−foot continuous tubular element routed throughout

remain active.

Figure 2Ć27.ĄFire Detection System

2−39

ORIGINAL

NAVAIR 01−F14AAD−1

Fire detection circuits in the engine compartments

be pulled to make the pushbutton for that engine accessible

detect a leak in the high−temperature duct and illuminate the

(see Figure 2Ć28). If the left or right fire extinguishing

appropriate FIRE warning light and activate the L FIRE or

pushbutton is activated, the contents of both extinguishing

R FIRE warning legend on the MFD and HUD. The warning

containers are discharged into the selected engine and its

legend is a repeat of a discrete signal from the fire detect

accessory section. Since it is a one−shot system, both system

system.

advisory lights, ENG FIRE EXT and AUX FIRE EXT, will

illuminate and remain illuminated after container pressures

2.12.1

Fire Detection Test

drop below a preset level.

An integrity test of the fire detection system can be

2.13.2

Fire Extinguisher Advisory Lights

performed by selection of FIRE DET/EXT on the pilot

MASTER TEST switch. The integrity test simultaneously

Two advisory lights are provided to indicate low

checks the sensing element loops of both engine compartĆ

pressure in the fire extinguishing agent containers. The

ments for continuity and freedom from short circuits, and the

lights, ENG FIRE EXT and AUX FIRE EXT, illuminate

fire alarm circuits and FIRE warning lights for proper

when container pressure drops 90 psi below a nominal

functioning. Presence of a short circuit or control unit

pressure of 600 psi at 70° F (see Figure 2Ć28).

malfunction causes the warning light to remain out. Fire

detection test is not available on the emergency generator.

2.13.3

Fire Extinguisher Test

The fire extinguishing system is tested by raising and

2.13 FIRE EXTINGUISHING SYSTEM

rotating the MASTER TEST switch to FIRE DET/EXT and

The fire extinguishing system is capable of discharging

depressing the knob. If the fire extinguisher test passes, the

an extinguishing agent into either engine nacelle and its

GO light illuminates; if the NO GO light illuminates or if both

accessory section. The system consists of two containers of

or neither GO and NO GO lights illuminate, the system has

extinguishing agent, piping and nozzles to route and

not tested properly and a failure exists somewhere in the

discharge the agent, cockpit switches to activate the system,

system.

and advisory lights that alert the flightcrew to a drop in

system pressure below a predetermined level.

2.14 AIRCRAFT FUEL SYSTEM

The aircraft fuel system normally operates as a split

The fire extinguishing agent is a clean, colorless, odorĆ

feed system, with the left and aft tanks feeding to the left

less, and electrically nonconductive gas. It is a low−toxicity

engine and the right and forward tanks feeding the right

vapor that chemically stops the combustion process. It will

engine (refer to FO−6). Except for the external tanks, the

not damage equipment because it leaves no water, foam,

system uses motive flow fuel to transfer fuel. The supply of

powder, or other residue.

high−pressure fuel from engine−driven motive fuel pumps

operates fuel ejector pumps to transfer fuel without the need

The retention time of an adequate concentration of the

of moving parts. The system is not dependent on electrical

extinguishing agent in the engine compartment will deterĆ

power for normal fuel transfer and feed. Total internal and

mine probability of reignition, and, therefore, the probability

external fuel quantity indication is provided, with a selectĆ

of aircraft survival. At high airspeeds, where airflow through

able quantity readout for individual tanks. Fuel system

the engine compartment is increased, agent retention time is

management requirements are minimal under normal operaĆ

reduced.

tion for feed, transfer, dumping, and refueling. Sufficient

cockpit control is provided to manage the system under

The slower the airspeed at the time the extinguisher is

failure conditions. The aircraft fuel system is designed so that

fired, the higher the probability of fire extinction and the

all usable fuel will normally be depleted under two or

lower the probability of reignition.

single−engine operating conditions before an engine flameĆ

out occurs from fuel starvation. However, with complete

Circuit breaker protection is provided on the RIO

motive flow failure, engine fuel starvation can occur with

essential No. 1 circuit breaker panel by the R FIRE EXT

usable fuel aboard.

(7C4) circuit breaker and the L FIRE EXT (7C5) circuit

breaker.

Note

2.13.1

Fire Extinguisher Pushbuttons

All fuel weights in this manual are based on

the use of JP−5 fuel at 6.8 pounds per gallon,

The discharge pushbuttons for the fire extinguishing

JP−4 fuel at 6.5 pounds per gallon, or JP−8 fuel at

system are located behind the FUEL SHUT OFF handles.

6.7 pounds per gallon.

The FUEL SHUT OFF handle for the affected engine must

ORIGINAL

2−40

NAVAIR 01−F14AAD−1

2.14.1

Fuel Tankage

forward−most cell in the aft tank group (cell No. 5) lays

laterally across the center fuselage. Extending aft are two

Figure 2Ć29 shows the general fuel tankage arrangeĆ

coffin−shaped tanks that contain two cells (Nos. 6 and 8) on

ment in the aircraft. The fuel supply is stored in eight separate

the right side and one cell (No. 7) plus an integral fuel vent

fuselage cells, two wing box cells, two integral wing cells,

tank on the left side. The coffin tanks straddle the center

and (optional loading) two external fuel tanks.

trough area, which contains the control rods, Sparrow missile

launchers, and electrical and fluid power lines. All fuel cells

2.14.1.1

Sump Tanks

in the aft tank group are interconnected by one−way flapper

valves at the base for aft−to−forward fuel gravity transfer.

The engine feed group, consisting of the left and right

box−beam tanks and the left and right sump tanks, spans the

2.14.1.4

Wing Tanks

fuselage slightly forward of the mid−center of gravity. Fuel

in each box−beam tank gravity flows to its respective sump

There are integral fuel cells in the movable wing panels

tank. The sump tanks (self−sealing) are directly connected

between the front and aft wing spars. Because of the wing−

to the box−beam tanks and contain the turbine−driven

sweep pivot location and the extensive span (20 feet) of the

boost pumps. The feed tanks supply fuel to the engine.

wing tanks, wing fuel loading provides a variable aft cg conĆ

A negative−g check valve traps fuel in the feed tank during

tribution to the aircraft longitudinal balance as a function of

negative−g flight.

wingsweep angle. Each wing panel consists of the integral

fuel cell, which is designed to withstand loads because of fuel

sloshing during catapulting and extreme rolling maneuvers

with partial or full wing fuel. Fuel system plumbing (transfer

and refuel, motive flow, and vent lines) to the wing tanks

incorporate telescoping sealed joints at the pivot area to proĆ

D Zero− or negative−g flight longer than 10 secĆ

vide normal operation independent of wing−sweep position.

onds in AB or 20 seconds in MIL or less will

deplete the fuel sump tanks (cell Nos. 3 and 4),

2.14.1.5

External Tanks

resulting in flameout of both engines.

D AB operation in the 0 to ć0.5−g regime may

Fuel, air, electrical, and fuel precheck line connectors

result in air ingestion into the fuel boost

are under the engine nacelles for the external carriage of two

pumps, causing possible AB blowout or

fuel tanks. Check valves in the connectors provide an autoĆ

engine flameout.

matic seal with the tank removed. Although the locations are

D With fuel in feed group below 1,000 pounds,

designated as armament stations Nos. 2 and 7, no other store

AB operation could result in AB blowout.

is designed to be suspended there so that the carriage of

external fuel tanks does not limit the weapon−loading capaĆ

Note

bility of the aircraft. Suspension of the drop tanks and their

fuel content has an insignificant effect on the aircraft longituĆ

AB operation with less than 1,000 pounds in

dinal center of gravity, and, even under the most adverse

either feed group may illuminate the FUEL

asymmetric fuel condition, the resultant movement can be

PRESS light because of uncovering of the boost

compensated for by lateral trimming.

pump inlet.

2.14.1.2

Forward Tank

The forward fuselage fuel tank is between the inlet

ducts and immediately ahead of the feed group. The forward

See Chapter 4 for external tank limitations.

tank is partitioned into two bladder cells (Nos. 1 and 2) that

are interconnected by open ports at the top for vent and overĆ

2.14.2

Fuel Quantity System

flow purposes. Flapper valves at the base provide for

forward−to−aft fuel gravity transfer.

The fuel quantity measurement and indication system

2.14.1.3

Aft Tank

provides the flightcrew with a continuous indication of total

internal and external fuel remaining, a selective readout for

all fuel tanks, independent low−fuel detection, and automatic

The aft fuselage fuel tank group is partitioned into four

fuel system control features.

bladder cells (Nos. 5, 6, 7, and 8) and a vent tank. The

ORIGINAL

2−42

NAVAIR 01−F14AAD−1

2.14.2.2

FUEL LOW Caution Lights

A L FUEL LOW or R FUEL LOW caution light

illuminates with 1,000 ± 200 pounds of fuel remaining in the

respective feed group. The RIO is provided with a single

To prevent fuel spills from an overfilled vent tank

FUEL LOW caution light that illuminates with one or both

caused by a failed level−control system, set the

of the pilot FUEL LOW caution lights.

WING/EXT TRANS switch to OFF if the

left tape reading reaches 6,200 pounds or the

Each FUEL LOW caution light is illuminated by two

right tape reading reaches 6,600 pounds. If either

thermistors operating in series. One set of thermistors is in the

fuel tape reading is exceeded, the aircraft shall be

right box−beam tank and cell No. 2. The other set of thermisĆ

downed for maintenance inspection.

tors is in the left box−beam tank and cell No. 5. The FUEL

LOW light illuminates only if both thermistors operating in

Note

series are uncovered.

Fuel in the vent tank is not gauged.

The quantity measurement system uses dual−element,

capacitance−type fuel probes to provide the flightcrew with

a continuous display of fuel quantity remaining. Fuel thermĆ

D If the thermistors in either cell No. 2 or No. 5

istor devices and caution light displays provide a backup fuel

remain covered during a fuel transfer failure,

low indicating system, independent of the capacitance gaugĆ

it is possible to partially deplete the sump tank

ing system. Additionally, the pilot is provided with a BINGO

without illuminating the respective FUEL

set capability on the fuel quantity indicator to preset the total

LOW caution light.

quantity level for activation of a BINGO caution light.

D When both FUEL LOW caution lights illumiĆ

Note

nate, less than 1 minute of fuel is available if

both engines are operating in zone five AB.

Fuel quantity system malfunctions that result in

erroneous totalizer readings will invalidate the

D If the BINGO CAUTION circuit breaker

use of the BINGO caution light.

(8F6) is pulled, the L and R FUEL LOW

caution lights will be disabled.

2.14.2.1

Fuel Quantity Indicators

2.14.2.3

Fuel Quantity Indication Test

The pilot and RIO fuel quantity indicators are shown in

Actuation of the master test switch in INST causes the

Figure 2Ć30 with a definitive breakdown of tape and counter

fuselage tapes and total and feed/wing/external fuel quantity

readings. The white vertical tapes on the pilot indicator show

indicators to drive to 2,000 pounds and illuminates the FUEL

fuselage fuel quantity. The left tape indicates fuel quantity in

LOW caution lights. The test can be performed on the ground

the left feed and aft fuselage; the right tape indicates fuel

or in flight. The test does not check the fuel probes or the

quantity in the right feed and forward fuselage. The L" and

thermistors. A test of the BINGO set device can be obtained

R" labeled counters display either feed group, wing tank, or

concurrently with the INST test by setting the BINGO level

external tank fuel quantity on the side selected using the

at greater than 2,000 pounds. In this case, the BINGO caution

QTY SEL rocker switch on the fuel management panel. The

light will illuminate when the totalizer reading decreases to

rocker switch is spring loaded to FEED. The pilot TOTAL

a value less than the BINGO setting.

quantity display and the RIO display indicate total internal

and external fuel.

2.14.3 Engine Feed

Note

The feed group for each engine is comprised of a

box−beam tank and a sump tank. Each box−beam tank holds

The RIO fuel quantity indicator is a repeater of

approximately 1,300 pounds of fuel and is fed from external

the pilot total fuel indicator. The difference

tank transfer, wing transfer, and fuselage transfer from cell

between the two should not exceed 300 pounds.

No. 2 or 5. When a box−beam tank is full, excess fuel is

returned to the fuselage tanks through an overflow pipe. The

sump tanks, which hold approximately 300 pounds of fuel

2−43

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

QTY SEL switch

WING Ċ

Fuel quantity in each wing is displayed on L and R counter of

pilot’s fuel quantity indicator.

FEED Ċ

Spring−loaded rocker switch returns to FEED when not held in

WING or EXT. FEED group fuel quantity displayed on L and R

counter of pilot’s fuel quantity indicator

EXT Ċ

Fuel quantity in each external tank displayed on L and R counter

of pilot’s fuel quantity indicator.

FEED switch

FWD Ċ

Both engines feed from right and forward tanks. Opens sump

tank interconnect valve, box beam vent valves, fuselage motive

flow isolation valve, defueling and transfer selector valve, and

shuts off motive flow fuel to all aft tank ejector pumps.

NORM Ċ

Right engine feeds from forward and right tanks. Left engine

(guarded

feeds from aft and left tanks

position)

AFT Ċ

Both engines feed from aft and left tanks. Opens sump tank interĆ

connect valve, box beam vent valves, fuselage motive flow isolaĆ

tion valve, defueling and transfer selector valves, and shuts off

motive flow fuel to forward tank ejector pumps.

WING/EXT TRANS

ORIDE Ċ

Airborne − Allows transfer of wing fuel, fuselage tank

switch

pressurization, and pressurization and transfer of external tanks

with landing gear down, and with electrical malfunction in transfer

system. Weight on Wheels −Allows transfer of wing and external

tank fuel.

AUTO Ċ

Airborne − Normal position. Wing fuel is automatically transferred.

Transfer of external fuel and fuselage pressurization is automatic

with landing gear retracted. Automatic shut off of wing and

external tanks when empty. Weight on Wheels −Automatic

transfer of wing and external tank fuel cannot be accomplished;

switch must be set to ORIDE for wing fuel transfer.

OFF Ċ

Closes solenoid operated valve to shut off motive flow fuel to

wing and also inhibits external tank transfer and fuselage

pressurization. Spring return to AUTO when master test switch

is actuated in INST, and when either thermistor in cell 2 and 5 is

uncovered, when DUMP is selected, and when REFUEL PROBE

switch is in ALL EXTD.

In−flight refueling probe

Illuminates whenever probe cavity forward door is open during retraction or

transition light

extension of probe.

DUMP switch

OFF Ċ Dump valve closed.

DUMP Ċ Opens a solenoid operated pilot valve, which ports motive flow

fuel pressure to open the dump valve and allows gravity fuel

dump overboard from cells 2 and 5. Wing and external tank

transfer automatically initiated. Dump electrically inhibited with

weight on wheels or speed brakes not fully retracted.

Figure 2−30. Fuel Controls and Indicators (Sheet 2 of 3)

2−45

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

REFUEL PROBE switch

ALL EXTD Ċ Extends refueling probe. Shuts off wing and external

tank fuel transfer to permit refueling of all tanks.

Returns transfer switch from OFF to AUTO.

FUS EXTD Ċ Extends refueling probe. Normal transfer and feed.

Used for practice plugins, fuselage−only refueling,

or flight with damaged wing tank.

RET Ċ

Retracts refueling probe.

Left and right FUEL SHUT

Pulling respective handle manually shuts off fuel to that engine.

OFF PULL handles

Push forward resets engine fuel feed shutoff valve to open.

L and R FUEL LOW

Fuel thermistors uncovered in aft and left or forward and right feed group.

caution lights

Illuminates with approximately 1,000 pounds remaining in individual feed

(Also single light on RIO

group and the respective fuselage tanks empty.

CAUTION panel.)

BINGO caution light

Illuminates when total fuel quantity indicator reads lower than BINGO

counter value.

L and R FUEL PRESS

Indicates insufficient discharge pressure (less than 9 psi) from respective

caution lights

turbine−driven boost pump.

NOTE

If fuel pressure light is illuminated an engine overtemp warning

tone will sound.

Figure 2−30. Fuel Controls and Indicators (Sheet 3 of 3)

each, are located directly beneath the box−beam tanks and

tank. The pressure head at interconnect A or B created

have three sources of fuel (see Figure 2-31 for identification

by the higher vertical location of the fuel in the box−

of tank interconnects):

beam tank, is greater than that created at interconnect

1. Interconnect A or B provides gravity sump from

C or D by the fuel in either cell No. 2 or 5. Therefore,

the respective box−beam tank.

fuel to replenish the sump tanks will come from the

2. Interconnect C or D connects the sump tank to its

box−beam tanks through interconnects A and B.

respective fuselage tank (cell No. 4 to cell No. 2/

2. Situation 2

cell No. 3 to cell No. 5).

a. Fuel in cell Nos. 2 and 5

3. The sump tank interconnect line and valve E

b. FEED switch in NORM

connect the two sump tanks.

c. High−engine fuel demands (afterburner).

The proportion of fuel supplied to each sump tank

Under these conditions the sump tank interconnect

through the five interconnects (A through E) is a function of

valve is closed and the left and right systems are

the pressure differential existing at each of the interconnects.

isolated. Engine fuel demand can exceed the transfer

The interconnect with the highest pressure differential will

rate into the box−beam tank. If this occurs, the fuel level

provide the most fuel. Valve E is commanded open during

in the box−beam tank will start to drop; however the

low−fuel states and during fuel balancing when the FEED

box−beam tanks are not vented, resulting in a pressure

switch is selected FWD or AFT.

drop above the declining fuel level. This reduced

In a normal sequence, three situations can be defined:

pressure lowers the total pressure at A and B, below the

1. Situation 1

pressure at C and D. Therefore, the majority of the fuel

to replenish the sump tanks comes directly from

a. Fuel in cell Nos. 2 and 5

fuselage cell Nos. 2 and 5 through interconnects C and

b. FEED switch in NORM

D, respectively. The reduction in box−beam tank fuel

c. Normal engine fuel flow (MIL thrust or less).

quantity should not normally result in a feed group

Under these conditions, the sump tank interconnect

quantity indication of less than 1,200 pounds. If the

valve is closed, and the left and right systems are

feed groups drop and then hold in the 1,200−pound

isolated. The transfer capacity into the box−beam tank

range during a high−speed dash, the system is working

exceeds the engine demand, ensuring a full box−beam

normally.

CHANGE 2

2−46

NAVAIR 01−F14AAD−1

3. Situation 3

only in emergencies. Suction feed is drawn from an inlet at

the bottom of the fuel cell that does not incorporate a flexible

a. Fuel in either cell No. 2 or 5 has been depleted

pendulum pickup.

b. FEED switch in NORM

c. Any normal engine demand.

When the low−level thermistor in either cell No. 2 or 5

With a left or right FUEL PRESS light, flight at

is uncovered, both box−beam tanks are vented and the

zero or negative g should be avoided or engine

sump tank interconnect valve is opened. The two

fuel starvation may result.

groups become a common system and will seek a

common level to equalize the static pressure head. Fuel

With both FUEL PRESS caution lights illuminated,

will flow through the open sump tank interconnect

there is a potential that total loss of motive flow pressure has

valve only as a function of differential pressure. With

occurred because both motive flow pumps are not function-

open vent valves, the fuel in both box−beam tanks has

ing. Total loss of motive flow pressure will preclude transfer

a positive vent pressure, forcing the fuel into the

of any remaining wing fuel or fuel dump and result in total

respective sump tank through interconnect A or B.

segregation of the FWD/RIGHT and AFT/LEFT systems

since motive flow provides the force to open the sump tank

Fuel in the sump tank is picked up by the turbine−driven

interconnect valve. Without motive flow pressure, all fuse-

boost pump through a flexible pendulum pickup, boosted to

lage fuel transfer is by gravity, which makes the quantity of

greater than 10 psi, and fed to the engine through the engine

usable fuel a function of aircraft attitude. At cruise attitude,

feed line. Normally the right boost pump only feeds the right

approximately 400 pounds of usable fuel will be trapped in

engine and the left boost pump only feeds the left engine;

the aft fuselage. After illumination of both fuel pressure

however, the boost pump output lines are connected by a

caution lights, any of the following events indicate that some

normally closed engine automatic crossfeed valve. If either

motive flow pressure is available:

boost pump output pressure falls below 9 psi, as indicated by

the illumination of the appropriate FUEL PRESS caution

1. Wing fuel transfer

light, the engine automatic crossfeed valve is commanded

open. The engine automatic crossfeed valve allows fuel from

2. With the FEED switch in FWD or AFT and no

the operating boost pump to supply pressurized fuel to the

transfer of external fuel

engine on the failed side. The engine automatic crossfeed

valve is also opened when either of the low level thermistors

a. The feed group of the selected side remains full.

in cell No. 2 or 5 is uncovered; however if equal boost pump

pressures exist, negligible flow will occur through the valve.

b. Fuel migration from one side to the other.

2.14.3.1

L/R FUEL PRESS Caution Lights

NOTE

If fuel pressure light is illuminated an engine

Illumination of the L or R FUEL PRESS caution light

overtemp warning tone will sound.

results from a malfunction of the boost pump, failure of the

motive flow pump, exhaustion of fuel, or fuel flow interrup-

2.14.3.2

Engine Fuel Feed During Afterburner

tion. With illumination of the caution light, the engine auto-

matic crossfeed valve is commanded open and the fuselage

Operations

motive flow shutoff valve on the failed side is automatically

closed. Because of the reduced pumping and transfer capac-

High AB fuel consumption places extreme demands on

ity while operating on a single boost pump, afterburner

the engine feed system. In addition, the g forces experienced

operation is restricted to altitudes below 15,000 feet. Fuel to

with AB use, especially during unloaded accelerations

both engines is supplied from the side with the operating

( bugouts") and low−g nose−high maneuvering, tend to

boost pump; therefore a fuel quantity imbalance will result.

reduce forward fuel transfer to cell No. 5 and the left engine

Use of the FEED switch to balance fuel quantity will override

sump tank (cell No.3). When these conditions are sustained,

the low−fuel pressure signal to the fuselage motive flow shut-

fuel in cell No. 5 is depleted by both high suction feed through

off valve, allowing normal fuel balancing procedures. Illumi-

the gravity transfer line (C, Figure 2-32), and by reducing

nation of both FUEL PRESS caution lights indicates

gravity fuel transfer from cell Nos. 6 and 7. Zero− or low−g

reduced (< 9 psi) or loss of boosted fuel pressure to both

(less than 0.5) flight tends to force the fuel remaining in cell

engines. Fuel will continue to be supplied by suction feed;

No. 5 toward the aft wall of the tank or, at reduced fuel level,

however, thrust settings should be minimized and AB used

uncovers gravity transfer line

CHANGE 2

2−48

NAVAIR 01−F14AAD−1

Figure 2Ć32.ĄAft Fuselage Fuel Transfer

drawn into the sump tank. Continued zero− or low−g (less

D In the presence of a fuel pressure light, fuel

thanĂ0.5) maneuvers will aggravate this condition and

demand must be reduced and positive g

increase the probability of air ingestion. If air enters the boost

restored to prevent possible engine flameout.

pump and engine feed line, the fuel pressure light will

illuminate. If the maneuver is continued, the left AB will

2.14.3.3

Fuel Shutoff Handles

blow out and subsequent left−engine flameout can occur.

Right−engine flameout can follow after left−engine flameout

Individual engine fuel feed shutoff valves in the left

because engine feed crossfeed operation will reduce the

and right feed lines at the point of nacelle penetration are

effective output of the right boost pump. Aircraft deceleraĆ

connected by control cables to the FUEL SHUT OFF handles

tion can further interrupt fuel transfer from cell No. 2 to the

on the pilot instrument panel. During normal operation, the

right sump through the gravity transfer line (D, Figure 2Ć33).

handles should remain pushed in so that fuel flow to the

Once initiated, this sequence can occur rapidly and is

engine fuel feed system is unrestricted. If a fire is detected in

independent of total fuel state.

the engine nacelle, the pilot should pull (approximately 3 or

4 inches) the FUEL SHUT OFF handle on the affected side

to stop the supply of fuel to the engine.

D During zero− or negative−g flight, the oil

pressure light will normally illuminate and

activate the master caution light. Subsequent

Securing the engine at high power settings using

illumination of a fuel pressure light may go

the FUEL SHUTOFF handles may result in

unnoticed, allowing the pilot to continue the

damage to the aircraft fuel system.

maneuver to the point of AB blowout and

engine flameout.

2−49

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć33.ĄForward Fuselage Fuel Transfer

Note

The path of the motive flow fuel is essentially the same

for either side. Fuel from the engine feed line is pressurized

Engine flameout will occur approximately 4 secĆ

by the engine−driven motive flow pump and initially routed

onds after the FUEL SHUT OFF handle(s) is

through the boost pump turbine. The motive flow fuel is then

pulled with the throttle(s) at MIL. With lower

routed through its respective transfer system. As the presĆ

power settings, time to flameout will increase

surized fuel passes through each ejector pump, it induces

(approximately 30 seconds at IDLE).

transfer fuel to flow along with the motive flow fuel. This

combination of fuel eventually is transferred into the respecĆ

2.14.4

Fuel Transfer

tive wing box−beam tank.

2.14.4.1

Motive Flow Transfer

There are four valves that control motive flow transfer:

With the exception of the external tanks, which utilize

1. Motive flow isolation valve Ċ Normally closed, but

bleed air, all fuel transfer is accomplished by gravity and

when the low−level thermistor in cell Nos. 2 or 5 is

motive flow. In motive flow, a relatively small amount of

uncovered or the FEED switch is out of NORM, the

pressurized fuel moves at high speed through ejector pumps,

valve is commanded open, providing a path for

using the venturi effect to induce flow of the transfer fuel.

motive flow fuel from a normally operating side to

The ejector pumps have no rotating parts or power requireĆ

cross over and power a malfunctioning opposite

ments other than motive flow.

side.

Like other elements of the fuel transfer system, motive

flow transfer is initially segregated to right and left. The

2. Forward fuselage motive flow shutoff valve Ċ

motive flow pump driven by the right engine provides motive

Normally open except when the R FUEL PRESS

flow and pressure to drive the right boost pump and to run the

caution light is illuminated or the FEED switch is in

ejector transfer pumps in the forward fuselage and right wing.

AFT. When the valve is closed, all motive flow

The motive flow pump driven by the left engine provides

transfer in the forward fuselage is shut off. If the

motive flow and pressure to drive the left boost pump and

valve is closed because of the R FUEL PRESS

runs the ejector transfer pumps in the aft fuselage and left

caution light, positioning the FEED switch to FWD

wing.

will open the valve.

ORIGINAL

2−50

NAVAIR 01−F14AAD−1

3. AFT fuselage motive flow shutoff valve Ċ

lage, aft fuselage transfer does not have any high−level control

Normally open except when the L FUEL PRESS

associated with it. Excess fuel in the box−beam tank passes

caution light is illuminated or the FEED switch is in

through an overflow pipe back into cell No. 5. When cell No.Ă5

FWD. When the valve is closed, all motive flow

is full, the fuel cascades into cell Nos. 6, 7, and 8. The aft

transfer in the aft fuselage is shut off. If the valve is

fuselage fuel will continue to circulate until consumed by the

closed because of the L FUEL PRESS caution light,

engine. When their respective cells are empty, the motive flow

positioning the FEED switch to AFT will open the

ejector pumps will be shutoff by their own low−level floats.

valve.

The scavenge ejector pumps do not incorporate shutoff floats.

In the event of loss of aft fuselage motive flow transfer, fuel

4. Wing motive flow shutoff valve Ċ The motive flow

may be gravity fed forward to cell No. 5 and eventually to the

to each wing passes through separate paths in a

left sump tank through interconnect C.

single motive flow shutoff valve. The valve is norĆ

mally open except when:

2.14.4.4

Wing Transfer

a. The WING/EXT TRANS switch is in OFF or in

Wing fuel is transferred by two motive flow ejector

AUTO with both left and right wing thermistors

pumps located in each wing. To prevent overfilling the fuseĆ

dry.

lage, entry of wing fuel into the box−beam tank is controlled

by the refueling/transfer shutoff valve. In the forward fuseĆ

b. Weight is on wheels.

lage, excess fuel overflows through an overflow pipe from

the right box−beam tank into cell No. 2, and then cascades

c. The REFUEL PROBE switch is in ALL EXTD.

into cell No. 1. A high−level pilot valve senses when cell

No.Ă1 is full and sends a signal to close the right refueling/

In any case, the wing motive flow shutoff valve can be

transfer shutoff valve, preventing additional wing fuel from

commanded open by selecting ORIDE on the WING/EXT

entering. When engine fuel consumption provides room in

TRANS switch.

cell No. 1 for additional fuel, the high−level pilot will signal

the refueling/transfer shutoff valve to open. The sequence is

2.14.4.2

Forward Fuselage Transfer

identical for the left box−beam tank and aft fuselage with the

Fuel in cell No. 1 flows by gravity into cell No. 2 where

exception that the high−level pilot valve is located in cell

two motive flow ejector pumps transfer it into the right wing

No.Ă7 and controls the left refueling/transfer shutoff valve

box−beam tank at approximately 18,000 pph. Fuel entering

(see Figure 2Ć34 for wing and external tank fuel transfer).

the box−beam tank beyond engine demands overflows

through an overflow pipe back into cell No. 2. There is no fuel

Normally wing fuel can only transfer to the box−beam

level control associated with fuselage motive flow transfer;

tank on its respective side, except when the thermistor in

therefore, the fuel will continue to circulate from cell No. 2

either cell No. 2 or 5 is uncovered or the FEED switch is

into the right box−beam tank and back through the overflow

selected FWD or AFT. For either condition, the motive flow

pipe. When the fuel in cell Nos. 1 and 2 is depleted, the

isolation valve opens, making motive flow pressure available

motive flow ejector pumps are shut off by their own low−level

to either wing from either engine, and the two defuel/transfer

floats. In the event of failure of the forward fuselage motive

selector valves open, permitting fuel from either wing to

flow, the fuel can reach the right sump tank by gravity flow

transfer to either box−beam tank. Total loss of wing motive

through interconnect D.

flow will preclude transfer of any remaining wing fuel.

Failure of either high−level pilot valve or refueling/transfer

2.14.4.3

Aft Fuselage Transfer

shutoff valve to the closed position could cause a single−wing

transfer failure. Selection of FWD or AFT on the FEED

Fuel in the aft fuselage is transferred forward by

switch opens the defuel/transfer selector valves allowing the

scavenge ejector pumps in cell No. 8 and the vent tank, single

trapped wing fuel to transfer to the opposite box−beam fuel

ejector pumps in cell Nos. 6 and 7, and two ejector pumps in

tank.

cell No. 5. All aft motive flow transfer is into the left box−

Note

beam tank, producing a rate of approximately 36,000 pph.

This flow rate is approximately twice that of the forward

Premature automatic wing motive flow valve

fuselage transfer rate because there are more motive flow

shutoff may occur because of formation of air

ejector pumps in the aft transfer system. More fuel tanks and

bubbles in the wingtip fuel thermistors. Pilot

thus more motive flow ejector pumps are required in the aft

selection of ORIDE with the WING/EXT

transfer system than the forward transfer system because of

TRANS switch will re−enable fuel transfer.

the aircraft structural configuration. Like the forward fuseĆ

2−51

ORIGINAL

NAVAIR 01−F14AAD−1

Note

bypass the landing gear down interlock in the external tank

transfer circuit, the WING/EXT TRANS switch may remain

in the AUTO (normal) position for this check.

D ORIDE transfer should not normally be used

unless AUTO transfer fails to complete transĆ

Note

fer of wing or external tank fuel. ORIDE use

when the wing tanks are dry may allow air to

D Verifying tank operation by observing fuel

enter the box−beam tanks, reducing the effiĆ

transfer is both time consuming with a full fuseĆ

ciency of gravity transfer to the sump tanks.

lage fuel load and aggravates fuel slosh loads

in the external tanks during catapult launch.

D When the thermistor in either cell No. 2 or 5

is uncovered, the WING/EXT TRANS switch

D Engine rpm above idle may be required to

will be deenergized from OFF to AUTO. This

provide sufficient bleed air pressure for a satĆ

automatic feature is to ensure all wing and

isfactory check.

external fuel has been transferred. After 5 secĆ

onds, the pilot may reset this switch to OFF.

2.14.4.6

Vent Valve Failure

A weight−on−wheels inhibit function prevents opening

The vent valves in the right and left box−beam tank are

of the wing motive flow shutoff valve. To transfer wing fuel

always commanded open with the sump tank interconnect

during ground operations, the WING/EXT TRANS switch

valve, making the right and left feed groups a common

must be set to ORIDE to bypass the weight−on−wheels

system. This function occurs when the low−level thermistor

function.

in cell No. 2 or 5 is uncovered. To equalize the static pressure

head at the interconnect valve, the fuel in the sump tanks will

Activation of fuselage fuel dump automatically initiĆ

seek a common level. At matched engine demands, each

ates wing fuel transfer in sequence after external tank transfer

engine will feed from its own side and negligible flow will

by automatically moving the WING/EXT TRANS switch to

occur across the sump tank interconnect valve. If a vent valve

AUTO if in OFF. Positioning the REFUEL PROBE switch to

fails to open, the additional vent pressure on top of the fuel

ALL/EXTD also releases the solenoid holding the WING/

on the vented side creates a pressure differential between the

EXT TRANS switch in OFF.

left and right sump tanks and results in migration through the

interconnect valve to the side with the inoperative vent valve.

2.14.4.5

External Tank Transfer

Therefore, sump tank replenishment of fuel to the side with

the failed vent valve will come primarily from the opposite

External tank transfer is also controlled by the WING/

sump tank because the head pressure at the interconnect

EXT TRANS switch. When external tanks are installed,

valve (E) may be higher than that at interconnects A, B, C,

transfer from the wings and external tanks occurs concurĆ

or D (Figure 2Ć31). A fuel quantity imbalance will occur with

rently. Transfer from the wings and external tanks cannot be

the side of the properly operating vent valve decreasing more

accomplished separately; however, the external tanks should

rapidly than the malfunctioning side. The box−beam tank

complete transfer before the wing tanks. External tank fuel

with the malfunctioning vent valve will eventually vent

is transferred by bleed air pressure regulated to 25 psi. MaxiĆ

through the overflow pipe when the respective fuselage tank

mum transfer rate of each external tank is approximately

(cell No. 2 or 5) is empty. If for any reason the fuel is not

45,000 pph. External tank fuel transfer into the fuselage is

transferred out of the respective fuselage tank, the imbalance

controlled by the same valves that control wing transfer.

will continue until the vented sump tank fuel quantity is

Fuselage level is controlled by the refueling/transfer shutoff

low enough to uncover the interconnect valve and line

valves and, until both the defuel/transfer valves are comĆ

(256 pounds approximately). This permits venting of the

manded open, external tank fuel can only transfer into the

unvented side and permits use of the balance of the fuel in

box−beam tank located on the same side of the aircraft.

the sump tanks.

External tank transfer can be checked on the deck by

Vent valve malfunctions can create disconcerting fuel

placing the WING/EXT TRANS switch to ORIDE, or selectĆ

imbalances. Although engine operation is not affected and all

ing FLT GR UP with the MASTER TEST switch and noting

of the fuel in the aircraft is available, AB use should be

depletion of external tank fuel quantity. In addition, when

avoided when low feed group fuel quantities are indicated. If

FLT GR UP is selected, the GO/NO GO light on the

both engine/boost pumps are operating, there is no advantage

MASTER TEST panel is illuminated by a pressure switch in

in using the cockpit fuel FEED switch to attempt to correct

the aircraft pressure line leading to the external tanks and

the imbalance. Positions other than NORM may simply

indicates status of line pressure. Since FLT GR UP serves to

aggravate the imbalance.

2−53

ORIGINAL

NAVAIR 01−F14AAD−1

2.14.5

Fuel Quantity Balancing

2.14.6

Fuel Transfer/Feed During Single−Engine

Operation

Fuel quantity balancing is not normally required prior

to completion of wing/external tank transfer or until one

Loss of an engine before the low−level thermistor in

fuselage tape drops below 4,500 pounds. The procedure

either cell No. 2 or 5 is uncovered will terminate all motive

requires use of the FEED switch that opens the sump tank

flow transfer on the failed side. External tank fuel will conĆ

interconnect valve, joining the FWD/R and AFT/L systems.

tinue to transfer if room is available in the failed side fuselage

With a high quantity in the FWD/R group, the greater static

tanks. If no pilot action is taken, the operating engine will

head pressure, particularly in noseup attitudes, can cause

feed only from its own side. This will lead to a fuel imbalance

overfilling of the AFT/L group. To prevent this, the FEED

that can normally be corrected through the use of the fuel

switch should be returned to NORM before the AFT/L tape

FEED switch. Selecting the high side (inoperative engine

reaches 6,200 pounds.

side) results in the following:

When the FEED switch is moved to select the high−fuel

Selected side fuselage motive flow shutoff valve is

quantity side, the following occurs:

opened. The valve was commanded closed when the

FUEL PRESS caution light illuminated.

1. Sump tank interconnect valve opens and provides a

fuel path between the right and left tanks.

Operating side fuselage motive flow shutoff valve

2. Both box−beam tank vent valves open and provide

is closed and stops operating side fuselage fuel

equal vent pressure on top of the fuel in each box−

transfer into the box−beam tank.

beam tank, regardless of the fuel level.

Motive flow isolation valve opens. Operating side

3. Fuselage motive flow shutoff valve on the non−

motive flow pressure now powers the inoperative

selected (low−fuel quantity) side closes and termiĆ

side. Failed side fuselage fuel will begin transferĆ

nates the last source of transfer of that fuselage fuel

ring into its respective box−beam tank.

into its respective box−beam tank.

Sump tank interconnect valve opens and provides a

4. Motive flow isolation valve opens and provides a

path for the inoperative side fuel to reach the operĆ

path for the nonselected side motive flow pressure

ating engine.

to reach the opposite side. Thus motive flow transfer

should maintain a full box−beam tank on the

Wing box−beam tank vent valves open and equalize

selected side.

the pressure above the fuel in each wing box−beam

tank, permitting the higher static pressure created

5. Both defuel/transfer selector valves open and

by the full wing box−beam tank on the inoperative

permit either wing/external tank to transfer into

side to induce flow through the open sump tank

either wing box−beam tank.

interconnect valve to the operating engine.

The higher static pressure head created by the full box−

Both defuel/transfer selector valves open and allow

beam tank on the selected side results in the nonselected side

either wing or external tank fuel to transfer into

engine feeding primarily from the sump tank interconnect

either wing box−beam tank.

rather than interconnects A, B, C, or D. With both engines

feeding from the fuel in primarily one side, the correction rate

If no crew action is taken with the FEED switch, the

of the fuel quantity imbalance is essentially a function of

same fuel system functions are automatically provided when

engine demand.

the thermistor in either cell No. 2 or 5 is uncovered. AddiĆ

tional actions that will occur when the cell No. 2 or 5 thermisĆ

tor is uncovered are:

1. Both right and left fuselage motive flow shutoff

valves open, overriding any previous commands to

D During AB operations, NORM shall be

close. Manual override of each valve is still proĆ

selected. FWD or AFT could deplete fuel in

vided through the FEED switch.

sump tanks.

D Aircraft attitude will have a significant influĆ

ence on the direction of fuel movement if

FWD or AFT is selected. Nosedown attitude

will transfer fuel forward, and noseup attitude

will transfer fuel aft.

ORIGINAL

2−54

NAVAIR 01−F14AAD−1

2. Engine crossfeed valve receives a redundant comĆ

2.14.7

Fuel Dump

mand to open. An initial command was provided

when the FUEL PRESS caution light illuminated.

Figure 2Ć35 shows aircraft fuel system components

associated with fuel dump operation. Fuel dump standpipes

3. WING/EXT TRANS switch will automatically go

in the forward (cell No. 2) and aft (cell No. 5) fuselage tanks

to AUTO if originally in OFF. If desired, OFF can

are connected to the fuel dump manifold at the dump shutoff

be reselected after 5 seconds.

valve. The manifold extends aft to the fuselage boattail.

Actuation of the fuel DUMP switch to DUMP supplies power

2.14.6.1

Sump Tank Interconnect Valve Failure

(dc essential No. 2) to open the solenoid−operated pilot valve,

which ports motive flow fuel pressure to open the dump

The major fuel system consideration while operating

shutoff valve with weight off the main landing gear and the

single engine is that the sump tank interconnect valve opens

speedbrakes retracted.

when commanded. This constitutes the only path through

which inoperative side fuselage fuel can reach the operating

The fuel DUMP switch circuit is deactivated on deck

engine. While the probability of an inoperative sump tank

or with speedbrakes extended. Fuel dump with the speedĆ

interconnect valve is very low, the consequences of a malĆ

brakes extended is inhibited because of the resulting flow

function under single−engine conditions are severe, particuĆ

field disturbance, which would result in fuel impingement on

larly at landing fuel weights. With a failed closed sump tank

the fuselage boattail and exhaust nozzles. The speedbrake

interconnect and full fuselage cells on the inoperative side,

switch is electrically bypassed during a combined hydraulic

only the wing fuel on the inoperative side and external fuel

system failure, enabling the pilot to dump fuel when the

can be transferred into the operating side fuselage. Attempts

speedbrakes are floating. The electrical bypass is accomĆ

to transfer the fuel from the inoperative side with the FEED

plished whenever the combined pressure falls below 500 psi.

switch compound the problem when the motive flow isolation

valve and inoperative side motive flow shutoff valve open.

Operating side motive flow fuel, pumped through the open

motive flow isolation valve to permit inoperative side wing

and/or fuselage motive flow transfer cannot be retrieved. Fuel

migration is approximately 100 pounds per minute because

The speedbrake/fuel dump interlock does not

of wing transfer, and approximately 200 pounds per minute

prevent speedbrakes from being deployed if fuel

for fuselage transfer. Coupled with a normal engine demand

dump is activated. It only prevents the dumping

of approximately 100 pounds per minute, a balancing attempt

of fuel if the speedbrakes are already extended.

will result in usable fuel in the operative side being depleted

at approximately 400 pounds per minute.

Note

D The FUEL FEED/DUMP circuit breaker (RE1)

Note

is on the pilot right−knee circuit breaker panel.

Operating side fuel remaining can be protected

D Dump operations with either engine in afterĆ

by pulling the FUEL SHUT OFF handle for the

burner should be avoided since the fuel dump

inoperative side and concurrently selecting the

mast discharge will be torched.

operative side on the FEED switch. This will elimĆ

inate a potential fuel path across the engine autoĆ

D After terminating fuel dump, wait approxiĆ

matic crossfeed valve, through the inoperative

mately 1 minute to allow residual fuel in the

sump tank boost pump into the inoperative side.

fuel dump line to drain before extending

speedbrakes or lighting afterburners.

If the sump tank interconnect is failed closed, the

following additional considerations apply:

Fuel in the wings and external tanks is dumped by

transferring to the fuselage. When the fuselage fuel dump

With the FEED switch selected to the operating side.

circuit is activated, wing and external tank transfer to the

box−beam tanks is automatically initiated. Fuel dump is by

1. Wing and external tank fuel from both sides will

gravity flow with a nominal discharge rate of 1,500 pounds

transfer into the operating side fuselage if the

per minute. The dump rate is affected by aircraft pitch attiĆ

inoperative side fuselage is full.

tude and total fuselage fuel quantity with discharge flow

inhibited at nosedown conditions. The standpipes in the fuel

2. If DUMP is selected, wing motive flow is autoĆ

cells control the minimum fuel dump level in the tanks,

matically activated; therefore, approximately 100

which, under normal operations (feed group full), is approxiĆ

pounds per minute of fuel available to the operating

mately 4,000 pounds.

engine will be lost.

2−55

ORIGINAL

NAVAIR 01−F14AAD−1

Figure 2Ć35.ĄFuel Vent and Dump

2.14.8

Internal Tank Pressurization and Vent

2.14.9

Fueling and Defueling

The internal fuel vent system is shown in Figure 2Ć35.

Figure 2Ć36 shows the refueling system. The aircraft is

It is an open−vent−type system, pressurized by ram air and

equipped with a single−point refueling system, which enables

engine bleed air from the

25−psi external tank pressure

pressure filling of all aircraft fuel tanks from a single recepĆ

system that is reduced to 1.75 psi by a fuselage pressure

tacle. The receptacle is at the recessed ground refuel and

regulator and distributed to all tanks through the fuselage

defuel station, behind a quick−access door on the lower right

vent system. This air is automatically supplied when the

side of the forward fuselage. The maximum refueling rate is

landing gear handle is UP or the WING/EXT TRANS switch

450 gallons per minute at a pressure of 50 psi. Since ground

is in ORIDE. When the WING/EXT TRANS switch is in

and air refueling connections use a common manifold, the

OFF, the low−pressure bleed air is cut off.

refueling sequence is the same.

In flight, the vent tank is maintained at a positive

Standpipes refuel the aft and forward fuselage tanks by

pressure up to 2.5 psi maximum. This pressure is fed by

overflow from the left and right box−beam tanks. A high−

connecting lines to all internal tanks. These connecting lines

level pilot valve at the high point of the forward tank shuts

are routed to provide venting to both the forward and aft end

off the fuselage refueling valve in the right box−beam tank

of each fuselage tank so it can function as both a climb and

when the forward tank group is full. Fuel flows from the left

dive vent. Venting of the box−beam tanks is controlled by

box−beam tank to cell No. 5, after which it overflows to the

solenoid−operating valves, which when closed, provide

right side, then the left side. A high−level pilot valve at the

suction transfer through the gravity flow paths in cell Nos. 2

high point of the left box−beam tank and aft tank (cell No. 7)

and 5 to the sump tanks.

shuts off the fuselage refueling valve in the left box−beam

tank when the aft tank group is full. Individual wing and

external tank filling is accomplished by flow through a

shutoff valve in each tank.

ORIGINAL

2−56

NAVAIR 01−F14AAD−1

Figure 2Ć36.ĄRefueling System

vent pressure indicator monitors pressure in the vent lines.

The gauge consists of a pointer on a scale having two bands,

one green and one red.

The green band indicates a safe pressure range (0 to

Gravity refueling of the aircraft fuel system

4 psi), and the red band indicates an unsafe range (4 to 8 psi).

should be accomplished only under emergency

situations. While performing such an operation,

avoid introducing contaminants into the fuel tanks

or damaging the fuel quantity probes and wiring.

2.14.9.1

Precheck System

During ground refueling operations, the direct−

reading vent pressure indicator shall be observed

Ground refueling control is by two precheck selector

and refueling stopped if pressure indicates in the

valves and a vent pressure gauge adjacent to the refueling

red band (above 4 psi).

receptacle on the ground refuel and defuel panel. The

precheck valves functionally test high−level pilot valve

2.14.10

In−Flight Refueling

operation incident to ground pressure refueling; the valves

Note

separately check the pilot valves in the fuselage tanks and the

See paragraph

9.1 for in−flight refueling

wing and external tanks. In addition to this precheck

procedures.

function, the precheck valves can be used for ground

selective refueling of only the fuselage or all tanks. Since the

The in−flight refueling system permits partial or

precheck valves, which are manually set by the groundcrew,

complete refueling of the aircraft fuel tanks while in flight.

port pressurized servo fuel to the high−level pilot valves and

The retractable refueling probe has an MA−2−type nozzle,

subsequently to the shutoff valves, no electrical power is

which is compatible with any drogue−type refueling system.

necessary on the aircraft to perform ground refueling

A split refueling system is provided with fuel routed into the

operations. Additionally, ground refueling control without

left and right box−beam tanks for initial replenishment of

engines running is completely independent of switch posiĆ

sump tank fuel. Selectable fuel management controls dictate

tioning on the fuel management panel. The direct−reading

the extent of further distribution to the wing tanks, external

2−57

ORIGINAL

NAVAIR 01-F14AAD-1

tanks, and/or fuselage tanks. The maximum refueling rate is

approximately 475 gallons per minute (3,000 lbs per minute

depending on fuel type) at a pressure of 57 psi.

Loss of combined pressure may indicate impend-

ing fluid loss. Without fluid in the combined

system return line, the in-flight refueling probe

will not extend with the handpump. Early exten-

To prevent fuel fumes from entering the cockpit

sion of the refueling probe at the first indication

through the ECS because of possible fuel spill

of a combined system malfunction is recom-

during in-flight refueling, select L ENG air

mended in a carrier environment.

source.

Note

D Extension or retraction of the refueling probe

using the hydraulic hand pump requires the

refuel probe switch to be placed in EXT or RET

57 psi limitation could be exceeded when

(as appropriate), combined system fluid in the

refueling from some aircraft operating more

return line, and essential dc No. 2 electrical

than one transfer pump. Damage to fuel sys-

power. With a total loss of combined hydraulic

tem can result.

pressure in flight, fluid trapped in the return

line/handpump reservoir can be isolated,

D Maximum airspeed for extension or retraction

exclusively for refueling probe extension, by

of the refueling probe is 400 knots (0.8 Mach).

placing the landing gear handle in the up posi-

Note

tion. Extension of the refueling probe requires

approximately 25 cycles of the pump handle.

D With the in-flight refueling probe extended,

D Probe retraction is not available if the FUEL P/

the pilot and RIO altimeter and airspeed and

MOTIVE FLOW ISOL V (P-PUMP) circuit

Mach indicators will show erroneous indica-

breaker (RG1) is pulled.

tionsbecause ofchangesinairflowaroundthe

pitot static probes.

2.14.10.2

Refueling Probe Transition Light

D Flight operations with the in-flight refueling

The red probe transition light immediately above the

probe door removed are not recommended

REFUEL PROBE switch illuminates whenever the probe

because of the effects of water intrusion,

cavity forward door is not in the closed position. Since the

exposure to elements, and structural fatigue to

closed-door position is indicative of both the probe retracted

electrical hydraulic hardware assemblies. If

and extended position, the light serves as a probe transition

operational necessity dictates, the door may

indicator as well as a terminal status indicator. The probe

be removed to prevent damage, loss or engine

external light illuminates automatically upon probe exten-

FOD.

sion with the EXT LTS master switch ON.

D The RUDDER AUTH caution light may

2.14.10.3

In-Flight Refueling Controls

illuminate when the in-flight refueling probe

is extended. Press the MASTER RESET

Regardless of fuel management panel switch position-

button to reset the light.

ing, at low-fuel statesthe initial resupplyoffuel isdischarged

into the left and right box-beam tanks. The split refueling

2.14.10.1

In-Flight Refueling Probe

system to the left and right engine feed group provides for a

relatively balanced cg condition during refueling. Selective

The retractable in-flight refueling probe is in a cavity

refueling of the fuselage or all fuel tanks is provided on the

on the right side of the forward fuselage section, immediately

REFUEL PROBE switch with the probe extended. In

forward of the pilot vertical console panel.

FUS/EXTD normal fuel transfer and feed is unaltered. This

Extension of the refueling probe is provided through

position is used for practice plug-ins, fuselage-only re-

redundant circuits by the REFUEL PROBE switch. A

fueling, or return flight with a damaged wing tank. The

hydraulic actuator within the probe cavity extends and

ALL/EXTD position shuts off wing and external drop tank

retracts the probe. The probe actuator is powered by the

transfer to permit the refueling of all tanks.

combined hydraulic system. It can be extended and retracted

by means of the hydraulic handpump in the event of

2.14.11

Hot Refueling

combined system failure.

Hot refueling can be accomplished with the refueling

probe extended or retracted. If the probe is extended, control

CHANGE 1

2-58

NAVAIR 01−F14AAD−1

of the tanks to be refueled is accomplished in the same

2.15 ELECTRICAL POWER SUPPLY SYSTEM

manner as during in−flight refueling. If the probe is not

extended, select WING/EXT TRANS switch to OFF to refuel

In normal operation, ac power is supplied by the

all tanks. Select ORIDE to refuel the fuselage only.

engine−driven generators. This ac power is converted by two

transformer−rectifiers

(T/R) into dc power (refer to FO−8).

2.14.12

Automatic Fuel Electrical Controls

One generator is capable of assuming the full ac power load

and one T/R is capable of assuming the full dc power load.

2.14.12.1

Automatic Low−Level Wing Transfer

Additionally, a hydraulically driven emergency generator

Shutoff

provides an independent backup supply of both ac and dc

power for electrical operation of essential buses. Ground

A thermistor is located at the low point in each wing

operation of all electrically powered equipment is provided

cell. When both are uncovered, a discrete electrical signal is

through the supply of external ac power to the aircraft.

generated, and through a control, the wing motive flow

Switching between power supply systems is automatically

shutoff valve is energized and closes, terminating all wing

accomplished without pilot action; however, sufficient

transfer. If either or both thermistors are again submerged,

control is provided for the flightcrew to selectively isolate

wing transfer resumes.

power sources and distribution in emergency situations. See

Figure 2Ć37 for a functional description of the control

Failure of this override system could result in a wing

switches. All electrical circuits are protected by circuit

breakers accessible in flight to the pilot and RIO.

transfer failure. Selection of WING/EXT TRANS switch to

ORIDE removes all power from the wing motive flow shutoff

valve, permitting it to open.

2.15.1

Normal Electrical Operation

2.14.12.2

Automatic Fuel Low−Level Override

2.15.1.1

Main Generators

Under normal operating conditions, the forward and

Two engine−driven, oil cooled, integrated drive generĆ

right fuselage tank complex is isolated from the aft and left

ators (IDG) produce the normal 115−volt, 400 Hz, three−

tank. This is necessary for proper longitudinal cg control and

phase ac electrical power. The normal rated output of each

battle damage conditions. However, as fuel depletion progĆ

generator is 75 kVA, which is sufficient to individually

resses to the point of sump tank only remaining, it becomes

assume the complete electrical load of the aircraft. Each main

mandatory that the tanks be connected to maintain an equal

ac generator is controlled by a separate switch on the pilot

balance. To accomplish this, two thermistors are located at

MASTER GEN control panel. Indication of a main power

the low points in cell Nos. 2 and 5, and when either is

supply malfunction is provided by a L GEN and R GEN

uncovered (approximately 1,700 to 2,000 pounds per side)

caution light. The IDG oil system is used for cooling as well

the following operations are electrically performed:

as lubricating the IDG. The oil is normally cooled by the IDG

air/oil cooler and returned to the constant speed drive (CSD)

1. Sump tank interconnect valve is opened.

for recirculation. When AB is used, additional cooling is

provided by the AB fuel/oil cooler before returning to the

2. Motive flow isolation valve is opened.

IDG. Should an excessive amount of heat be developed in an

3. Box−beam vent valves are opened.

IDG, a thermal (390_ F) actuated device automatically

decouples the input shaft from the remainder of the CSD,

4. Engine crossfeed valve is opened.

protecting both the CSD and generator. There are no proviĆ

sions for recoupling the IDG unit in flight.

5. WING/EXT TRANS switch is energized to move

from OFF to AUTO. This signal is maintained for

5 seconds.

6. Defuel transfer selector valves are opened.

Failure of the weight−on−wheels circuit to the in−

flight mode while on the deck will cause the loss

of ECS engine compartment air ejector pumps,

causing a subsequent IDG disconnect and illumiĆ

Uncovering either thermistor in cell No. 2 or 5

nation of the GEN light.

will move the WING/EXT TRANS switch from

OFF to AUTO but under no circumstances will

it override a wing transfer failure.

2−59

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

MASTER GEN

NORMĂĊ

Connects the generator to the main buses through the line contactor.

switch

OFF/RESETĂĊ Disconnects generators from the buses. Resets the generator if tripped by

(lock lever)

an overvoltage, undervoltage, or fault condition.

TESTĂĊ

The generators are energized but are not connected to the buses. Provides

a means to analyze a system malfunction indicated by a generator caution

light when an attempt to reset a generator is unsuccessful.

EMERG

NORMĂĊ

Safety guard down. Electrically controlled by a solenoid control valve

generator

energized by the left main dc bus. Operation is automatically initiated,

switch

connecting to the essential buses, with the loss of the left main dc bus,

regardless of other sources of ac or dc power. Total loss of ac or dc power

will result in the loss of the left main dc bus and, therefore, activation of the

emergency generator.

OFF/RESETĂĊ Safety guard must be lifted. Disconnects the emergency generator from the

essential buses. Resets the generator if tripped by an undervoltage or

under−frequency condition.

L GEN and R

GEN caution lights are on the pilot’s caution/advisory light panel. Each light is tied to its

GEN caution

respective main ac contactor and is powered by the essential bus no. 2. Illumination of the

lights

L GEN or R GEN caution light indicates that the corresponding generator is not supplying

power, due to a fault in the generator, generator control unit, or electrical distribution

system.

TRANS/RECT

A TRANS/RECT advisory light is on the lower half of the pilot’s caution/advisory indicator

advisory light

panel. Illumination of the TRANS/RECT advisory light indicates either a single or dual

transformer−rectifier failure has occurred.

Figure 2Ć37.ĄGenerator Panel

ORIGINAL

2−60

NAVAIR 01−F14AAD−1

2.15.1.1.1

Generator Control Units

An external power monitor prevents application of external

power that is not within tolerances and disconnects external

Generator output voltage and frequency are individu-

power from the buses if undervoltage, overvoltage, under-

ally monitored by GCUs, that prevent application of inter-

frequency, overfrequency, or phase−reversal occurs. Power

nally generated power to the aircraft bus system until the

can be reapplied to the aircraft by pressing the reset button

generator output is within prescribed operating limits. With

adjacent to the receptacle, provided it is within prescribed

the main generator switch in NORM, the applicable genera-

limits. External electrical power is automatically inhibited

tor is self−excited, so that during the engine start cycle, it

from HUD, MFD, AICS, APX−76, CADC, and CIU without

automatically comes on−line at approximately 50−percent

external air−conditioning connected to the aircraft. When the

rpm under normal load conditions. Likewise, during engine

left generator comes on the line during start, it automatically

shutdown, the GCU automatically trips the generator off the

disconnects external power. Although there is no direct

line as the power output decreases below prescribed limits at

cockpit indication of external power being applied after one

approximately 55−percent rpm.

generator is operating, the HYD TRANSFER PUMP will not

operate if the external power plug is still in the aircraft

During normal operations, the generator control

receptacle.

switches remain in NORM continuously. However, subse-

quent to an engine shutdown, stall, or flameout in flight

2.15.2

Electrical Power Distribution

where the GCU has tripped the generator off−line, the relight

of the engine will not automatically reset the generator unless

Electrical power is distributed through a series of

the engine speed decreased below about 30−percent N₂ rpm.

buses. Under normal operation, the ac generator power dis-

If a transient malfunction or condition causes the generator

tribution is split between the left and right main ac buses.

to trip, the generator must be manually reset by cycling the

Failure of either main ac generator trips a tie connector to

applicable generator control switch to OFF/RESET then

connect both buses to the operative generator. If the bus tie

back to NORM.

fails to trip when the generator goes bad, the respective

transformer−rectifier will not be powered and the indication

When normal reset cannot be accomplished, TEST, on

of this double failure will be a L GEN or R GEN caution light

the generator control switch, allows the generator to be

and a TRANS/RECT advisory light. The left and right main

excited but not connected to the aircraft buses. In test, a CSD,

ac buses in turn supply ac power directly to the respective

generator, or GCU failure causes the GEN light to remain

transformer−rectifiers, and the left main ac bus also supplies

illuminated. If the light goes out, the problem is in the dis-

power to both essential ac buses under normal operation.

tribution system.

External power is distributed through the aircraft elec-

2.15.1.2

Transformer−Rectifiers

trical system in the same manner as main generator power.

Like the main ac generators, dc power distribution from the

Two transformer−rectifiers convert internal or external

two transformer−rectifiers under normal operations is split

ac power to 28−Vdc power. A single TRANS/RECT advisory

between the left and right main dc buses. Failure of either

light on the pilot advisory panel provides failure indication

transformer−rectifier trips the respective tie contactor to

for one or both transformer−rectifiers. No flightcrew control

connect both main dc buses to the operative transformer−

is exercised over transformer−rectifier operation aside from

rectifier. The TRANS/RECT advisory light provides a direct

controlling the ac power supply or circuit breakers for the

indication of dc bus tie status. An interruption−free dc bus

power converters. The transformer−rectifiers have a rated

interconnects the left and right main dc buses to provide a

output of 100 amperes each. Each unit is capable of assuming

continuous source of dc power with failure of either main ac

the complete dc electrical load of the aircraft. Forced air

generator and/or transformer−rectifier. The left main dc bus

cooling is provided with engines running to dissipate the heat

additionally supplies power to both essential dc buses under

generated by the power converters.

normal operations. Power to the DFCS bus is normally

supplied from the interruption−free dc bus; however, with an

2.15.1.3

External Power

output failure from both transformer−rectifiers, the DFCS bus

load is automatically transferred to the essential No. 2 bus.

Loss of main dc power automatically activates the emer-

Ground power is applied through a receptacle just aft

gency generator. The emergency generator is electrically

of the nosegear. The pilot has control over external power

application only through hand signals to the plane captain.

2−61

CHANGE 2

NAVAIR 01-F14AAD-1

inhibited by a solenoid control valve energized by the left

5are onthe RIO’sleft andpanels6to9are onthe RIO’sright.

main dc bus. Operation of the generator is automatically

The pilot left and right knee panels are designated L and R,

initiated, connecting to the essential buses, with the loss of

respectively.

the main dc bus, regardless of other sources of ac or dc power.

Total loss of ac or dc power will consequently result in the

The first digit in the three-part locator is the alpha-

loss of the left main dc bus and, therefore, activation of the

numeric that identifies the circuit breaker panel. The second

emergency generator. This, in turn, trips power transfer

part is a letter that designates the row in which the circuit

relays to change essential ac & dc bus loading from the left

breaker will be found. The top row is designated A, the next

main ac and dc buses to the emergency generator, regardless

row lower is B, etc. The third part is a number and designates

of main generator output status.

the column in which the circuit breaker will be found. The

innermost column of each panel 1, 2, 5, 8, and 9 or aft most

2.15.2.1

Circuit Breakers

column on each panel 3, 4, 6, 7 L and R is designated “1,”

the next outboard/forward column is 2, etc. Figure 2-38 is an

Individual circuit protection from an overload condi-

alphanumeric listing of circuit breakers.

tion is provided by circuit breakers, which are all located in

the cockpits for accessibility in flight. The appropriate circuit

Note

breaker will pop out and isolate a circuit that draws too much

current, thus preventing equipment damage and a possible

D Panel No. 1 row A, the column numbering is

fire.

different from rows B to J.

D Panel No. 2 rows A to F, the column number-

ing is different from Rows G to I.

2.15.3

Degraded Electrical Operation

Popped circuit breakers should not be reset more

2.15.3.1

Emergency Generator

than once nor held depressed unless the associ-

ated equipment is absolutely required by opera-

The emergency generator provides a limited but

tional necessity. A popped circuit breaker indi-

independent backup source of ac (5 kVA, 115/200 volts) and

cates an equipment malfunction or an overload

dc (50 amperes, 28 volts) power for flight-essential compo-

condition. Repeated resets or forced depressions

nents. It is driven by combined hydraulic system pressure.

of popped circuit breakers can result in equip-

ment damage and/or serious electrical fire.

With normal combined hydraulic system operation, the

emergency generator powers the essential ac and dc No. 1

Cockpit circuit breaker panels are shown on FO-8 and

and No. 2 buses and the DFCS dc bus in the 5 kVA mode.

FO-9. Circuit breakers in the pilot cockpit comprise the

Operation of the generator is automatically initiated with the

majority of those required for essential aircraft systems. The

loss of dc left main bus even if other dc buses remain ener-

circuit breakers are arranged in rows and are oriented so that

gized. Approximately 1 second elapses from the time of auto-

the white banded shaft of a popped breaker is readily visible

matic initiation before the generator delivers rated power to

to the flightcrew. Panels, rows, and columns of breakers are

flight-essential ac and dc buses. This delay will force the

identified to facilitate breaker location and designation. Plac-

DFCS computers into a power-up BIT sequence, requiring a

ards adjacent to the breakers identify individual circuit

MASTER RESET to regain SAS and ARI functions.

breakers by affected components; amperage ratings are indi-

cated on top of each circuit breaker.

2.15.2.1.1

Circuit Breaker Location

The alphanumeric system for locating circuit breakers

The spoiler actuators are mechanically biased to

in the aircraft is as follows.

the retracted position in order to cause the

spoilers to retract in the event that the command

The panels in the RIO cockpit are labeled 1 through 9

signal from the DFCS is lost (i.e., DFCS power

starting left-aft and proceeding clockwise. Thus, panels 1 to

failure).

CHANGE 1

2-62

NAVAIR 01−F14AAD−1

If this bias is reversed, the affected spoiler will

The exact hydraulic pressure at which the emergency

extend instead of retracting when the command

generator is unable to power all three buses is dependent on

signal is lost. A DFCS power failure coupled

the load placed on the generator and can vary from 2,000 to

with a reverse spoiler bias will result in a fully

1,100 psi indicated. If the emergency generator is required

deployed spoiler. All unaffected spoilers will

and there is a hydraulic emergency that could lower comĆ

remain retracted and will not respond to flight

bined system operating pressure, the ac essential No. 2 and

control inputs until the DFCS command signals

dc essential No. 2 and dc DFCS buses can be powered with

are restored.

lower hydraulic pressure securing nonessential equipment in

order to reduce the electrical load and to maintain DFCS

Note

functionality.

D DFCS synchronization can take up to two

Note

seconds following a power interrupt. If the

MASTER RESET pushbutton is depressed

D When the emergency generator is operating

during the synchronization time, an addiĆ

with one main hydraulic system inoperative,

tional depression of the MASTER RESET

large hydraulic flow requirements for flight

pushbutton will be required to restore spoiler

controls may cause loss of the essential ac and

functionality.

dc No. 2 and DFCS buses. To regain these

D Do not press and hold the MASTER RESET

buses the emergency generator switch must be

pushbutton. Pressing and holding the

cycled through OFF/RESET to NORM after

MASTER RESET pushbutton during the synĆ

the hydraulic pressure recovers. Engine

chronization time will have no effect since the

instruments are powered by essential ac bus

DFCS computers only recognize the leading

No. 1. Engine instruments will be available or

edge of the pulse from the MASTER RESET

restored at lower engine rpm. The airspeed at

pushbutton, and not the fact that the button is

which engine instrumentation is restored

continuously depressed.

(either automatically or by pilot cycling the

emergency generator switch) could be higher

Pilot control of the emergency generator is through the

than the maximum airspeed.

guarded EMERG switch on the MASTER GEN control

panel. The emergency generator is electrically inhibited by

D In the event of L and R generator failure, an

a solenoid control valve energized by the left main dc bus.

attempt should be made to reduce unnecesĆ

sary electrical loads to maintain DFCS SAS

With the switch in NORM, operation of the generator

and ARI functionality.

is automatically initiated, connecting to the essential buses,

with the loss of the left main dc bus, regardless of other

2.15.3.1.2

Emergency Generator Test

sources of ac or dc power. Total loss of ac or dc power will

consequently result in the loss of the left main dc bus and

An operational check of the emergency generator can

activation of the emergency generator. The OFF/RESET

be accomplished anytime the combined system is pressurĆ

switch position provides the pilot with the capability of

ized and at least one main generator is on the line by selecting

isolating emergency electrical power from the aircraft buses

EMERG GEN on the master test switch and depressing the

(as in the case of an electrical fire) or resetting the generator.

switch. This provides 28 Vdc to activate the emergency

generator and checks the tie contactors by connecting

2.15.3.1.1

Emergency Power Distribution

electrical power to the essential ac and dc buses. The GO light

on the MASTER TEST panel indicates a satisfactory check.

An emergency generator control unit monitors the

A malfunction in the emergency generator operation is

emergency generator output. If it senses that the emergency

indicated by the NO GO light.

generator cannot supply power within the proper frequency

and voltage tolerances, the control unit disconnects the

Note

essential ac and dc essential No. 2 and the dc DFCS buses

from the emergency generator (1 kVA mode). It is possible

During the emergency generator test the essenĆ

that this could happen if the combined hydraulic system is not

tial ac No. 2 bus is switched between the left main

operating normally. If combined hydraulic pressure subseĆ

and emergency generator. The DFCS computers

quently recovers, the emergency generator switch must be

detect this as a loss of ac power and perform a

cycled through OFF/RESET and back to NORM to regain the

power−up BIT sequence when the emergency

5 kVA mode, restoring power to the essential No. 2 and DFCS

generator test is completed. A MASTER RESET

buses. The DFCS computers will respond with a power up

is required to regain SAS and ARI functions.

BIT sequence, requiring a MASTER RESET to regain SAS

and ARI functions.

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ORIGINAL

F-14D. FLIGHT MANUAL (2004) - page 2