F-14D. FLIGHT MANUAL (2004)

NAVAIR 01-F14AAD-1

3F6

26 VAC BUS FDR

8E6

APN-154

3F7

AC ESS BUS NO. 2 FDR PH A

4D6

APX-100 AC

4F1

AC ESS BUS NO. 2 FDR PH B

7F7

APX-100 DC

4F2

AC ESS BUS NO. 2 FDR PH C

7C2

ARC-182 NO. 1

2I4

ACM LT/SEAT ADJ/STEADY POS LT

7C1

ARC-182 NO. 2

3C6

ADF AC

8A5

ARMT GAS/L ENG AFT CONT/RAT IND

8D6

ADF DC

9A4

ASC

RC2

AFCS/NOSE WHEEL STEER

1D2

ASC PH A

LF1

AICS L

1D5

ASC PH B

2I5

AICS L HTR

1D6

ASC PH C

8E2

AICS L LKUP PWR/EMER GEN TST

1F1

ASPJ AUG PH A

7A6

AICS L RAMP STOW

1F3

ASPJ AUG PH B

LG1

AICS R

1F6

ASPJ AUG PH C

2I8

AICS R HTR

1F2

ASPJ BASIC PH A

8E1

AICS R LKUP PWR/ANTI SKID

1F4

ASPJ BASIC PH B

7A5

AICS R RAMP STOW

1F5

ASPJ BASIC PH C

RD2

AIR SOURCE CONTROL

9G5

ASPJ DC

8C2

AIR/ANTI ICE CONTR HOOK CONT/WSHLD

9G6

ASW-27

9B6

ALE-47 CHAFF/FLARE DISP

1J2

ASW-27 AC

9B5

ALE-47 SEQ 1 & 2 SQUIBS

RB1

ALPHA COMP/PEDAL SHAKER

LA3

AUTO PITCH DRIVE TRIM

4F4

ALPHA HTR

1J1

AUTO THROT AC

2H1

ALR-67 CMPTR

9B7

AUTO THROT DC

9F5

ALR-67 CONTR

8G3

AUX FLAP/FLAP CONTR

2H3

ALR-67 RCVR PH A

2H6

ALR-67 RCVR PH B

7D3

BARO ALT/TURN SLIP

2H9

ALR-67 RCVR PH C

3D4

BDHI INST PWR/JTIDS/DPG

7B6

ALT LOW WARN

8E7

BDHI/JTIDS DPG

9D1

AMC BIT/R DC, TEST

1B3

BEAM PS

7A3

ANGLE OF ATTACK IND DC

8F6

BINGO CAUTION

3F3

ANGLE OF ATTK IND AC

8F2

BLEED AIR/L OIL HOT

4F5

ANL ATTK/TOTAL TEMP HTR

4B4

BLEED DUCT AC

9C2

ANN PNL DIM CONTR

7A4

BOS CONTR/B/U OXY LOW

8C1

ANN PNL PWR

9D5

BRAKE ACCUM SOV

9A6

ANT LOCK EXCIT

7A2

B/U OXY PRESS IND

1C2

ANT SVO HYD PH A

1C4

ANT SVO HYD PH B

8A1

CABIN PRESS

1C6

ANT SVO HYD PH C

8C5

CAN/LAD CAUTION/EJECT CMD IND

8C2

ANTI-ICE CONTR HOOK CONT/

LA2

CHAN 1 CADC PH A

WSHLD/AIR

LB2

CHAN 1 CADC PH B

8E1

ANTI SKID/R AICS LKUP PWR

LC2

CHAN 1 CADC PH C

2I1

ANTICOLL/SUPP POS/POS LT

LD2

CHAN 2 CADC

RG2

ANTI-ICE/ENG/PROBE

3E7

CIU PH A

6C3

AN/AWW 4 PH A

4E1

CIU PH B

6C2

AN/AWW 4 PH B

4E2

CIU PH C

6C1

AN/AWW 4 PH C

3B3

COMB HYD PRESS IND

9A2

APG-71 ANT

9F4

COOLING INTLK/GND PWR

2G3

APG-71 PUMP PH A

8D8

CURSOR CONT/SNSR

2G6

APG-71 PUMP PH B

2G7

APG-71 PUMP PH C

7B5

DC ESS NO. 1 FDR

1I1

APG-71 XMTR AC

8A2

DC ESS NO. 2 FDR

9A3

APG-71 XMTR DC

Figure 2-38. Circuit Breaker Alphanumeric Index (Sheet 1 of 5)

CHANGE 1

2-64

NAVAIR 01−F14AAD−1

7A7

DC L TEST/RUDDER TRIM

RF2

FLT CONTR AUTH DC

9D1

DC R TEST/AMC BIT

2A1

FLT HYD BACKUP PH A

9B2

DD ENABLE/RDP

2C1

FLT HYD BACKUP PH B

9I6

DEKI

2E1

FLT HYD BACKUP PH C

8B1

DFCS BUS FDR

3B4

FLT HYD PRESS IND

3F4

DP 1 PH A

3A2

FORM LT/TAXI

4F3

DP 1 PH B

RE1

FUEL FEED/DUMP

4F6

DP 1 PH C

8F7

FUEL LOW CAUTION

1G2

DP 2 PH A

RD1

FUEL MGT PNL

1G4

DP 2 PH B

8F1

FUEL PRESS ADVSY

RG1

FUEL P/MOTIVE FLOW ISOL V

1G6

DP 2 PH C

3C3

FUEL QTY IND AC

9G2

DSS

7D1

FUEL QTY IND DC

RE1

DUMP/FUEL FEED

8E4

FUEL TRANS ORIDE

9E7

DYHR UNIT

8F9

FUEL VENT VALVE

8D4

ECS TEMP CONTR DC

8F5

GEN L CAUTION

9H3

ELECT COOLING

8F4

GEN R CAUTION

7B2

EMER FLT HYD AUTO

9F4

GND PWR/COOLING INTLK

7B1

EMER FLT HYD MAN

8G1

GND ROLL BRAKING/SPOILER POS IND

9I2

EMER GEN CONTR

9D6

GND TEST

8E2

EMER GEN TEST/L AICS LKUP PWR

3E2

GPS

7E3

EMER JETT #1

8E9

GPS ANT AMPL

7E2

EMER JETT #2

6A1

GUN CONTRL PWR AC

RC1

ENG ANTI−ICE VALVES

5D2

GUN PWR NO. 1

7D5

ENG INST NO. 1

5C2

GUN PWR NO. 2

7D4

ENG INST NO. 2

8A5

ENG L AFT CONT/ARMT GAS/RATS IND

1H1

HUD CAMERA PH A

3A3

ENG L BACKUP IGN

1H5

HUD CAMERA PH B

3B1

ENG L OIL PRESS

1H7

HUD CAMERA PH C

8A4

ENG R AFT CONT/EXHAUST NOZZLE

3C1

HUD PH A/MFD 1

3A4

ENG R BACKUP IGN

4C5

HUD PH B/MFD 1

3B2

ENG R OIL PRESS

4C6

HUD PH C/MFD 1

8D1

ENG OIL COOL

1A1

HV PWR SUP PH A

8D3

ENG SEC

1A3

HV PWR SUP PH B

8F10

ENG STALL TONE

1A5

HV PWR SUP PH C

RF1

ENG START

7B3

HYD PRESS IND

RG2

ENG/PROBE/ANTI−ICE

8G11

HYD PUMP SPOILER CONTR

8A4

EXHAUST NOZZLE/R ENG AFT CONT

8E5

HYD VALVE CONTR

8G10

EXT LT CONTR

LE3

ICE DET

9D2

FEMS

7F3

ICS NFO

7C7

FIRE L DET LT

7F2

ICS PILOT

7C5

FIRE L EXT

1J7

IFF A/A AC

7C6

FIRE R DET LT

9F6

IFF A/A DC

8C7

ILS ARA−63 DC

7C4

FIRE R EXT

3E5

ILS ARA−63 PH A

8G3

FLAP CONTR/AUX FLAP

4E3

ILS ARA−63 PH B

3D6

FLAP IND/TAIL/RUDDER

4E4

ILS ARA−63 PH C

RA2

FLAP/SLAT CONTR SHUT−OFF

8G9

INBD SPOILER CONTR

9B6

FLARE DISP/ALE−39 CHAFF

1I7

INS BATT PWR

LC1

FLT CONTR AUTH AC

Figure 2−38. Circuit Breaker Alphanumeric Index (Sheet 2 of 5)

2−65

ORIGINAL

NAVAIR 01−F14AAD−1

3C7

INS PH A

LE2

MACH TRIM AC

4C1

INS PH B

RE2

MACH TRIM DC

4C2

INS PH C

1A8

MAIN L XFMR RECT

3E4

INS SYNC

2F4

MAIN R XFMR RECT

3A1

INST LTS

LE1

MANUV FLAP/WG SWP DR NO. 2

3F5

INSTR BUS FDR

5A2

MASTER ARM

1I2

INTEG TRIM AC

9H4

MASTER TEST

9F3

INTEG TRIM DC

9G3

MDL

9I5

INTRF BLANKER

9G4

MFA

9I1

INTRPT FREE DC BUS FDR NO. 1

3C1

MFD 1/HUD PH A

9C6

INTRPT FREE DC BUS FDR NO. 2

4C5

MFD 1/HUD PH B

9D4

IRST DC

4C6

MFD 1/HUD PH C

2G2

IRST PH A

1G1

MFD 2/MFD 3 PH A

2G5

IRST PH B

1G3

MFD 2/MFD 3 PH B

2G8

IRST PH C

1G5

MFD 2/MFD 3 PH C

1G1

MFD 3 PH A/MFD 2

1J4

JTIDS BATT HEATER

1G3

MFD 3 PH B/MFD 2

1J3

JTIDS DPG PH A

1G5

MFD 3 PH C/MFD 2

1J5

JTIDS DPG PH B

8G5

MLG HANDLE RLY NO. 1

1J6

JTIDS DPG PH C

8G4

MLG HANDLE RLY NO. 2

8E7

JTIDS DPG/BDHI

7F5

MLG SAFETY RLY NO. 1

3D5

JTIDS RT PH A

7F4

MLG SAFETY RLY NO. 2

4D3

JTIDS RT PH B

9D3

MONITOR BUS CONTR

4D4

JTIDS RT PH C

RG1

MOTIVE FLOW ISOL V/FUEL P

3D4

JTIDS/DPG/BDHI INST PWR

5B2

MPRU DC PWR

7C3

KY−58/Z−AHP

6D3

MPRU PH A/SMP

LF1

L AICS

6D2

MPRU PH B/SMP

2I5

L AICS HTR

6D1

MPRU PH C/SMP

8E2

L AICS LKUP PWR/EMER GEN TST

8C8

MSL PWR HUD TEST

7A6

L AICS RAMP STOW

6B3

MSL PWR SUP PH A

7A7

L DC TEST/RUDDER TRIM

6B2

MSL PWR SUP PH B

8A5

L ENG AFT CONT/ARMT GAS/RATS IND

6B1

MSL PWR SUP PH C

3A3

L ENG BACKUP IGN

3C5

MSN CMPTR NO. 2 PH A

3B1

L ENG OIL PRESS

4C3

MSN CMPTR NO. 2 PH B

7C7

L FIRE DET LT

4C4

MSN CMPTR NO. 2 PH C

7C5

L FIRE EXT

1D1

MSN CMPTR NO. 1 PH A

8F5

L GEN CAUTION

1D3

MSN CMPTR NO. 1 PH B

1A8

L MAIN XFMR RECT

1D7

MSN CMPTR NO. 1 PH C

8F2

L OIL HOT/BLEED AIR

3B7

L PH A TEST/P−ROLL TRIM

8A3

NLG STRUT LCH BAR ADVSY

4B1

L PH B TEST/P−ROLL TRIM

RC2

NOSE WHEEL STEER/AFCS

4B2

L PH C TEST/P−ROLL TRIM

2I2

NFO CONSOLE LT

4E5

L PITOT STATIC HTR

3C4

OBOGS CONC

8C5

LAD CAUTION/EJECT CMD IND/CAN

7A1

OBOGS CONTR

9E1

LANTIRN POD CONT/RECON ECS

8F2

OIL L HOT/BLEED AIR

9E2

LANTIRN POD PWR/RECON CONTR

8D2

OIL R HOT

2C3

LANTIRN PWR/RECON HTR 3 PH

2H10

LIQUID COOLING CONTR AC

9C5

OUTBD SPOILER CONTR

9B4

LIQUID COOLING CONTR DC

2B3

OUTBD SPOILER PUMP

Figure 2−38. Circuit Breaker Alphanumeric Index (Sheet 3 of 5)

ORIGINAL

2−66

NAVAIR 01−F14AAD−1

2H5

OXY CONC HTR

9B1

RDP

8F6

OXY/BINGO CAUTION

1E4

RDP PH A

1E5

RDP PH B

3B7

P−ROLL TRIM/L PH A TEST

1E6

RDP PH C

4B1

P−ROLL TRIM/L PH B TEST

9B2

RDP/DD ENABLE

4B2

P−ROLL TRIM/L PH C TEST

2G4

RECON ECS CONT AC

4A6

PANEL FLOOD LTS

9E1

RECON ECS/LANTIRN POD CONT

RB1

PEDAL SHAKER/ALPHA COMP

9E2

RECON CONTR/LANTIRN POD PWR

3B7

PH A L TEST/P−ROLL TRIM

2C3

RECON HTR/LANTIRN PWR 3 PH

4B1

PH B L TEST/P−ROLL TRIM

1E2

RECON POD

4B2

PH C L TEST/P−ROLL TRIM

9E4

RECON POD DC PWR NO. 1

2H2

PH A R TEST

9E3

RECON POD DC PWR NO. 2

2H4

PH B R TEST

1A8

RECT/L MAIN XFMR

2H8

PH C R TEST

5H1

REL PWR/STA 1 TYPE I DCDR

4A5

PILOT CONSOLE LTS

5G1

REL PWR/STA 1 TYPE II DCDR

4A3

PILOT LCD INST LTS

5F1

REL PWR/STA 3 DCDR

LB1

PITCH A AC

5E1

REL PWR/STA 4 DCDR

8B7

PITCH A DC

5D1

REL PWR/STA 5 DCDR

LH1

PITCH B AC

5C1

REL PWR/STA 6 DCDR

8B3

PITCH B DC

5B1

REL PWR/STA 8 TYPE I DCDR

4E5

PITOT STATIC HTR L

5A1

REL PWR/STA 8 TYPE II DCDR

4E6

PITOT STATIC HTR R

LB3

ROLL A/YAW M

8F3

PLT ANN PNL AUX PWR/TR ADVSY

8B4

ROLL A DC

2I1

POSLT/ANTICOLL/SUPP POS

LA1

ROLL B AC

4A4

PROBE LT

8B2

ROLL B DC

RG2

PROBE/ANTI−ICE/ENG

9A7

RSP

1B2

RSP PH A

LG1

R AICS

1B5

RSP PH B

2I8

R AICS HTR

1B8

RSP PH C

8E1

R AICS LKUP PWR/ANTI SKID

3D7

RUDDER TRIM PH A

7A5

R AICS RAMP STOW

4D1

RUDDER TRIM PH B

8A5

RAT IND/L ENG AFT CONT/ARMT GAS

4D2

RUDDER TRIM PH C

9D1

R DC TEST/AMC BIT

7A7

RUDDER TRIM/L DC TEST

8A4

R ENG AFT CONT/EXHAUST NOZZLE

3D6

RUDDER/TAIL/FLAP IND

3A4

R ENG BACKUP IGN

3B2

R ENG OIL PRESS

9I3

SAHRS DC

7C6

R FIRE DET LT

1I3

SAHRS A

7C4

R FIRE EXT

1I5

SAHRS B

1I6

SAHRS C

8F4

R GEN CAUTION

2E4

R MAIN XFMR RECT

2I4

SEAT ADJ/STDY POS LT

8D2

R OIL HOT

RA2

SLAT CONTR SHUT−OFF/FLAP

7E5

SMP ESS

2H2

R PH A TEST

2H4

R PH B TEST

6D3

SMP/MPRU PH A

6D2

SMP/MPRU PH B

2H8

R PH C TEST

4E6

R PITOT STATIC HTR

6D1

SMP/MPRU PH C

8D8

SNSR/CURSOR CONT

4B3

RADAR ALTM

1B1

SOL PWR SUP PH A

1C3

RADAR DD PH A

1B4

SOL PWR SUP PH B

1C5

RADAR DD PH B

1B7

SOL PWR SUP PH C

1C7

RADAR DD PH C

Figure 2−38. Circuit Breaker Alphanumeric Index (Sheet 4 of 5)

2−67

ORIGINAL

NAVAIR 01−F14AAD−1

RB2

SPD BK P−ROLL TRIM ENABLE

6F6

STA 8B PWR PH A

8G1

SPOILER POS IND/GND ROLL BRAKING

6F5

STA 8B PWR PH B

5D3

STA 1 AIM−9 COOL

6F4

STA 8B PWR PH C

2I10

STA 1 BOL PWR

8F8

STARTER VALVE LT

5I3

STA 1 IFOL

3A7

STBY ATTD IND PH A

5H1

STA 1 TYPE I DCDR/REL PWR

4A1

STBY ATTD IND PH B

5G1

STA 1 TYPE II DCDR/REL PWR

4A2

STBY ATTD IND PH C

6A3

STA 1A AIM−9 PWR AC

2I1

SUPP POS/ANTICOLL/POS LT

5B3

STA 1A AIM−9 PWR DC

2I6

STORM FLOOD LTS

5J2

STA 1B NO. 1/2 DC

3D6

TAIL/RUDDER/FLAP IND

6A6

STA 1B PWR PH A

3A2

TAXI/FORM LT

6A5

STA 1B PWR PH B

1H2

TCS PH A

6A4

STA 1B PWR PH C

1H3

TCS PH B

5F1

STA 3 DCDR/REL PWR

5I2

STA 3 NO. 1/2 DC

1H6

TCS PH C

9C3

TCS SEL

6B6

STA 3 PWR PH A

6B5

STA 3 PWR PH B

4B5

TEMP CONT AC

4F5

TOTAL TEMP HTR/ANL ATTK

6B4

STA 3 PWR PH C

5H3

STA 3/6 IFOL

8F3

TR ADVSY/PLT ANN PNL AUX PWR

7D3

TURN SLIP/BARO ALT

5E1

STA 4 DCDR/REL PWR

5H2

STA 4 NO. 1/2 DC

7F6

UHF CONTR/VHF

6C6

STA 4 PWR PH A

3A6

UTILITY LTS

6C5

STA 4 PWR PH B

6C4

STA 4 PWR PH C

7F6

VHF/UHF CONTR

5G3

STA 4/5 IFOL

5D1

STA 5 DCDR/REL PWR

LE1

WG SWP DR NO. 2/MANUV FLAP

5G2

STA 5 NO. 1/2 DC

7D6

WHEELS POS IND

6D6

STA 5 PWR PH A

3F2

WING POS IND AC

6D5

STA 5 PWR PH B

7D2

WING POS IND DC

6D4

STA 5 PWR PH C

LD1

WING SWEEP DRIVE NO. 1

5C1

STA 6 DCDR/REL PWR

8C3

WSHLD DEFOG CONTR

5F2

STA 6 NO. 1/2 DC

8C2

WSHLDAIR/ANTI ICE CONTR/

6E6

STA 6 PWR PH A

HOOK CONT

6E5

STA 6 PWR PH B

6E4

STA 6 PWR PH C

8B6

YAW A DC

5C3

STA 8 AIM−9 COOL

8B5

YAW B DC

2I9

STA 8 BOL PWR

LD3

YAW A AC

5F3

STA 8 IFOL

LC3

YAW B AC

5B1

STA 8 TYPE 1 DCDR/REL PWR

LB3

YAW M/ROLL A

5A1

STA 8 TYPE II DCDR/REL PWR

6F3

STA 8A AIM−9 PWR AC

3F6

26 VAC BUS FDR

5A3

STA 8A AIM−9 PWR DC

5E2

STA 8B NO. 1/2 DC

Figure 2−38. Circuit Breaker Alphanumeric Index (Sheet 5 of 5)

ORIGINAL

2−68

NAVAIR 01−F14AAD−1

2.16 HYDRAULIC POWER SUPPLY SYSTEMS

transfer pump. This unit consists of two hydraulic pumps, one

in each of the main hydraulic systems, interconnected by a

The aircraft employs two main, independent, engine−

common mechanical shaft. Thus, a pressure deficiency in

powered hydraulic systems, supplemented by two electroĆ

one system is automatically augmented using pressure in the

hydraulic power modules, a bi−directional transfer unit, and

other system as the motive power. The result is bi−directional

a cockpit handpump. The systems are pressurized to 3,000 psi

transfer of energy without an interchange of system fluid.

and use MIL−H−83282 hydraulic fluid circulated through

The efficiency of the pump is such that a 3,000 psi system on

stainless steel and titanium lines. Hydraulic fluid is cooled by

one side will pressurize the other system to approximately

heat exchangers that use ejector air on deck. Hydraulic power

2,400 to 2,600 psi.

system controls and indicators are shown in Figure 2Ć39. The

components serviced by each hydraulic power system are

To prevent damage to the hydraulic transfer pump with

shown in FO−10.

the loss of system fluid on one side and to conserve hydraulic

power in the remaining good system, the pump is automatiĆ

2.16.1

Flight and Combined Systems

cally secured when pressure less than 500 psi is detected on

either side of the pump for 10 seconds. In addition, the pilot

2.16.1.1

Engine−Driven Pumps

can manually shut off the hydraulic transfer pump by lifting

the guarded HYD TRANSFER PUMP switch, located aft on

the right outboard console.

The flight and the combined systems are each pressurĆ

ized by engine−driven pumps. The flight hydraulic system

pump is driven by the right engine and the combined hydrauĆ

lic system pump by the left engine. Each of the main systems

is normally pressurized to 3,000

$ 100 psi at any time the

respective engine is operating.

If pressure in either system remains below 500

2.16.1.2

Hydraulic Pressure Light

psi for 5 seconds, immediately lift the guard and

select SHUTOFF with the HYD TRANSFER

A HYD PRESS caution light illuminates when the disĆ

PUMP switch. Failure of the hydraulic transfer

charge pressure from either engine−driven hydraulic pump

pump to automatically shut off after 10 seconds

falls below 2,100 psi; thereafter, the light goes out when

below 500 psi may cause the driving system to

pressure in both systems via the engine−driven pumps

cavitate and overheat.

exceeds 2,400 psi. If the HYD PRESS caution light has been

illuminated by low pressure in one main system, pressure

With ground electrical power connected to the aircraft,

failure in the other system will not cause the MASTER

the hydraulic transfer pump is deactivated and can only be

CAUTION light to illuminate again. The COMB and FLT

energized by a switch on the ground check panel. Normally,

gauges on the hydraulic pressure indicator reflect system

with both engines running, the hydraulic transfer pump is off.

pressure provided by either the engine−driven pumps or the

However, with less than 2,100 psi hydraulic pump discharge

hydraulic transfer pump. With both systems normally presĆ

pressure from either system, the pump will automatically

surized to 3,000 psi, the gauge needles form a horizontal line.

come on and supply hydraulic power to the faulty system. In

addition, the HYD PRESS caution light will also illuminate.

Note

The pilot has no direct control over the direction of pump

flow, the system automatically shifts in the direction that

High−rate lateral movements may illuminate the

supplemental power is required. Because of the location of

HYD PRESS light when engines are at idle

the flight and combined system pressure switches, the presĆ

power.

surization contribution of the hydraulic transfer pump is

reflected on the hydraulic pressure indicator but the HYD

2.16.1.3

Hydraulic Transfer Pump

PRESS caution light will remain illuminated. Operation on

(Bi−Directional Pump)

the hydraulic transfer pump may produce slight pressure

fluctuations. If the failed system discharge pressure is

To assure the continuance of main system hydraulic

restored to normal operating pressure (>2,400 psi) by the

pressure with an engine or engine−driven pump inoperative,

engine−driven pump, this HYD PRESS light will go out and

a second source of pressure is provided by the hydraulic

the hydraulic transfer pump will shut off.

2−69

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

HYD PRESS indicator

COMB and FLT Ċ Indicates pump discharge pressure on each engine,

normally 3,000 psi, or hydraulic transfer pressure

approximately 2,400 psi.

SPOIL Ċ

When the outboard spoiler hydraulic module is pressurized

(1,950 to 2,050 psi) the ON flag appears. If pressure drops

below 1,900 to 1,800 psi, the OFF flag appears.

EMER FLT Ċ When pressure from the backup flight control hydraulic

module reaches 500

$ 50 psi the ON flag appears.

Pressure below 350

$ 50 psi: the OFF flag appears.

HYD ISOL switch

FLT Ċ

Combined system hydraulic pressure is shutoff to landing

gear, nosewheel steering, antiskid, and wheel brakes.

T.O./LDG Ċ

Hydraulic pressure is available to all combined system

components.

BRAKE PRESSURE

AUX Ċ

Green segment indicates hydraulic pressure (2,150

$ 50

gage

to 3,000 psi) in the auxiliary brake accumulator; auxiliary

braking may be applied by rudder toe pedals (approximately

13 to 14 applications available). Red segment indicates

1,900 to 2,150 psi (approximately 5 applications available).

PARK Ċ

Green segment indicates hydraulic pressure (2,150

$ 50

to 3,000 psi) in the parking brake accumulator. The

parking/emergency brake handle must be pulled to apply

emergency braking (approximately 3 applications available).

Red segment indicates 1,900 to 2,150 psi.

HYD PRESS caution

Illuminates when hydraulic pressure from either engine−driven pump is below

light

2,100 psi. It will go out with pressure in both systems at 2,400 psi or above, if

pressure is provided by engine−driven pumps.

HYD TRANSFER

SHUTOFF Ċ Guard must be lifted. Shuts off hydraulic transfer pump.

PUMP switch

The pump should be secured when hydraulic pressure

drops below 500 psi and does not rise again within

5 seconds.

NORMAL Ċ Safety guard down. Pressure loss below 2,100 psi in one

(Guarded)

hydraulic system activates hydraulic transfer pump to

supply pressure from the other system.

EMERG FLT HYD

HIGH Ċ

Guard must be lifted. Activates the power module

switch

(high speed mode) bypassing flight and combined

2,100−psi switches.

LOW Ċ

Guard must be lifted. Activates the backup power module

(low−speed mode) bypassing flight and combined

2,100−psi switches.

AUTO (LOW) Ċ Safety guard down. The backup flight control system is

automatically activated (low−speed mode) when pressure

in both the flight and combined systems is less than

2,100 psi.

Figure 2−39. Hydraulic System Controls and Indication (Sheet 2 of 2)

2−71

ORIGINAL

NAVAIR 01−F14AAD−1

2.16.1.4

Cockpit Handpump

Major components in the combined and flight hydrauĆ

lic power supply systems are shown on FO−10. Each system

A manually operated pump handle is provided as a

has a piston−type reservoir and filter module in the sponson

supplementary source of power for ground operations with

aft of the main landing gear strut on the respective side

engines shut down and as a backup for the loss of combined

(combined−left; flight−right). Protrusion of mechanical pins

system pressure to operate the in−flight refueling probe or

on each filter module indicates a clogged filter.

charge the brake accumulator. It is an extendable handle in

the pilot cockpit between the left console and ejection seat.

2.16.2.1

Hydraulic Priority Valves

Forward and aft stroking of the handpump operates a double−

acting wobble pump. The pump, which draws fluid from the

The combined and flight hydraulic systems each incorĆ

combined system return line, recharges wheelbrake accumuĆ

porate two priority valves (1,800 psi and 2,400 psi) shown on

lator pressure when the landing gear handle is down. With the

FO−10. Hydraulic fluid will not pass through the one−way

gear handle up, it also serves as a backup means of extending

priority valves unless the input pressure exceeds the cracking

or retracting the in−flight refueling probe by placing

threshold of the valve. Basically, the 2,400 psi priority valves

the REFUEL PROBE switch in the desired position (EXT

give priority of the individual engine−driven pump discharge

or RET).

pressure to the primary flight controls

(horizontal tails,

rudders, inboard spoilers) and stability augmentation actuaĆ

The handpump is the only means of pressurizing the

tors. Conversely, the 1,800 psi priority valves give priority to

radome fold actuator, an operation that must be manually

the remaining systems on the other side (inlet ramps, wing

selected and the radome unlocked on deck from the nose

sweep, etc.) with pressure supplied by the hydraulic transfer

wheelwell. The recommended rate of operation is approxiĆ

pump. Under such circumstances, the pilot should be aware

mately 12 cycles per minute (a cycle is a complete forward

of the hydraulic energy available and demands of the various

and aft movement of the pump handle).

system components. Large and abrupt control commands can

rapidly consume total energy with the engine(s) at IDLE

2.16.2

Hydraulic Power Distribution

speed. For example, during a single−engine landing rollout,

if excessive horizontal tail movements are commanded, the

The distribution of hydraulic power in the flight and

nosewheel steering and wheelbrake operation could be

combined systems is shown on FO−10. Except for the left

temporarily lost.

empennage control surfaces, the flight system services only

those components on the right side of the aircraft and does not

2.16.2.2

Normal Hydraulic Isolation

penetrate into the wings. The combined system distribution

is more extensive throughout the aircraft, yet its services are

The combined system incorporates isolation circuits to

predominantly concentrated to the left side and extend to the

limit distribution to flight essential components. With the

inboard sections of the movable wing panels and to the

LDG GEAR handle UP, normal isolation may be selected by

landing gear. Although the flight and combined systems are

the pilot to prevent loss of hydraulic fluid in the event of

completely independent of each other, in certain components

material failure or combat damage to the isolated systems.

both pressure sources are used without an interchange of

Normal isolation electrically shuts off hydraulic pressure to

fluid. Both systems operate in parallel to supply power for

wheelbrakes, antiskid, landing gear and nosewheel steering.

operation of the primary flight control surfaces

(except

It is activated by placement of the HYD ISOL switch to FLT

spoilers) and stability augmentation actuators; if one system

on the landing gear panel. Placement of the gear handle to DN

fails, the other can continue to supply pressure for operation

mechanically cams the HYD ISOL switch to T.O./LDG or the

(with reduced power capability of such components). If

pilot can manually select it before lowering the landing gear.

either or both main hydraulic systems should fail, backup

Such action returns all combined−system components to norĆ

sources provide the capability for safe return flight and

mal operation.

landing.

ORIGINAL

2−72

NAVAIR 01−F14AAD−1

2.16.3

Outboard Spoiler System

2.16.4

Backup Flight Control System

The outboard spoilers are powered by a separate

The backup flight control system consists of a two−

closed−loop system, independent of the main hydraulic sysĆ

speed electrohydraulic power module known as the backup

tems (see Figure 2Ć40). An electrohydraulic power module

flight control module. The BFCM provides fluid energy to

supplies hydraulic pressure for outboard spoiler deflection

operate the horizontal tails and rudders at a reduced rate

and provides a backup power source for the main flaps and

(see Figure 2Ć41). Emergency power provides sufficient

slats. Outboard spoiler operation is electrically inhibited at

pitch, roll, and yaw control for return flight and landing with

wing−sweep angles greater than 62_ and the power module

both main hydraulic power systems inoperative.

is deactivated at wing−sweep angles greater than 65.

Return flow from the combined side of the rudder and

stabilizer actuators is first used to ensure the BFCM reservoir

A thermal cutout circuit secures the system in the event

is filled. When filled, a reservoir bypass valve opens, which

of overheating. Normal operation is automatically restored

allows return flow to the combined system. A priority valve

when fluid temperature falls below the prescribed limit. The

connects the BFCM return to the aircraft’s combined sys−

thermal cutout circuit is disabled with the gear handle down

tem return. When the combined system pressure falls below

and weight off wheels to prevent overtemperature shutdowns

300 psi, the priority valve closes, isolating the BFCM return

during takeoff or landing. To avoid overheating due to proĆ

from the combined system return. When the combined presĆ

longed ground operations, the outboard power module is

sure exceeds 500 psi, the priority valve opens allowing the

deactivated with the flap handle up when on internal electriĆ

backup system return to flow into the combined system

cal power with weight on wheels. Electrical power for the

return. A check valve isolates backup system pressure from

outboard spoiler system motor is supplied from the right main

the combined system when the BFCM is energized.

ac bus. The module can be activated using external ac electriĆ

cal power. With the module pressurized, the ON flag appears

2.16.4.1

Backup Flight Control Operation

in the SPOIL window at the bottom of the hydraulic pressure

indicator; otherwise, an OFF indication is displayed in the

window.

The BFCM may be operated in two modes: emergency

and ground test. In the emergency mode, the BFCM is conĆ

trolled by the EMERG FLT HYD switch, on the MASTER

Reservoir servicing level is shown by an indicator rod

TEST panel. The switch has three positions:(AUTO) LOW,

protruding from the integral power package. A fluid temperaĆ

LOW, and HIGH mode. Electric power to the motor is supĆ

ture gauge that registers current and retained peak system

plied by the right main ac electrical bus through the FLT

temperatures is on the power module. Protrusion of a red−

HYD BACKUP PH A (2A1), PH B (2C1), and PH C (2E1)

tipped pin on the integrated filter package is an indication of

circuit breakers located on right main ac circuit breaker panel

a clogged filter.

(No. 2) in the rear cockpit. Loss of both engine−driven electriĆ

cal generators eliminates in−flight use of the BFCM.

2.16.3.1

Flap and Slat Backup Operation

Although normal operation of the main flap and slat

segments is powered by a combined system motor on the flap

and slat gearbox, an auxiliary motor powered by the outboard

spoiler system is geared to the same shaft to provide for

Never use the three−phase circuit breakers (PH A,

emergency operation (retraction and extension) of the main

PH B, and PH C) to start or shut off the BFCM as

flaps and slats at a reduced rate. Failure of combined system

damage to the motor may result. These circuits

pressure activates the auxiliary motor to drive the flap and

must be engaged prior to any system test.

slat gearbox when selected by the normal flap handle or

maneuvering flap thumbwheel.

Automatic control of the BFCM is provided by the

closing of both flight and combined hydraulic system presĆ

sure switches. Since the switches are set at 2,100 psi, both

flight and combined hydraulic system pressures must drop

below 2,100 psi before the BFCM is turned on in the

2−73

ORIGINAL

NAVAIR 01−F14AAD−1

automatic low mode. Once in this automatic mode of

Ground checks of the BFCM are performed by the pilot

operation, the BFCM cannot be turned off unless either or

using the EMER FLT HYD switch. Before performing

both flight and combined systems are pressurized above

ground checks, the combined and brake system accumulators

2,400 psi. The EMERG FLT HYD switch is used to select the

must be charged. The BFCM has a small volume capacity,

low or high mode. Either of these positions overrides the

1,000 cc (@ 1 qt.) when full, but will decrease in volume

circuitry of the automatic low mode and the BFCM will

500 cc (@ ½ qt.) when the aircraft is not in use.

remain on even if either or both system pressures become

Below 500 cc (@ ½ qt.), cavitation of the pump and overheatĆ

pressurized above 2,400 psi. When the BFCM pump reaches

ing of the motor may occur. If the accumulators are not

500 psi, the ON flag appears in the selected window at the

charged prior to starting the BFCM, depletion of the reservoir

bottom of the hydraulic pressure indicator.

hydraulic fluid will occur. If this occurs too frequently, sysĆ

tem damage and failure may result. Both hydraulic system

In the low−speed mode, the system can operate indefiĆ

pressures should indicate zero in order to fully test indepenĆ

nitely and should be used for maximum range and endurance.

dent operation of the BFCM.

Emergency power

(high mode) provides a maximum

unloaded horizontal tail deflection rate approximately one

fourth of that available from a full powered hydraulic system

(10_ per second vice 36_ per second). The maximum deflecĆ

tion rate available will decrease as airloads increase.

D A 180_F thermal cutoff switch is bypassed

when the BFCM is selected on with the

EMERG FLT HYD switch. Prolonged ground

operation in the emergency mode will result

in BFCM burnout.

When operated in conjunction with zero comĆ

D Since flight control demands can exceed

bined system pressure, some BFCM hydraulic

BFCM capability, all surface demands must

fluid will be forced out by thermal expansion.

be performed slowly and cautiously in order

The BFCM will remain fully serviced and will

not to exceed the output rate of the system.

operate normally as long as the elevated temperĆ

Excessive system demands will cause the

atures are maintained. Once operating, the

pump to cavitate and the motor to overheat.

BFCM should not be turned off in flight without

Checks should be made slowly enough to

combined system pressure available to reservice

ensure continuous on indication in the

it. Doing so would result in fluid contraction and

hydraulic pressure indicator.

an underserviced condition that could prevent

subsequent pump operation.

2.16.4.2.1

Ground Test Mode

The ground test mode of operation is controlled by the

AUX HYD CONT switch on the ground test panel in the rear

cockpit. In this mode, the BFCM operates in the high mode

only. Ground test from the rear cockpit is electrically inhibĆ

D If either the flight or the combined hydraulic

ited when the aircraft is on internal electrical power.

system pressure drops below 2,100 psi withĆ

out illuminating the HYD PRESS caution

light, the automatic low mode of the backup

flight control system may be inoperative.

D Prolonged use

(approximately

8 minutes

cumulative time) of the BFCM in the high

The ground test mode incorporates a solenoid

mode may result in a failure of the BFCM.

valve that allows the BFCM to pressurize the

entire combined hydraulic system. If the comĆ

2.16.4.2

Ground Operations

bined and brake accumulators are not fully

charged

(brake pressure indicator at top of

For ground inspection purposes, protrusion of a red−

green), or if the combined system is not fully

tipped button on either the inlet or outlet filter cases is a

serviced, the reservoir will be depleted and the

positive indication of a dirty filter. Both such indications may

motor will cavitate and overheat. This could

be observed through an access door on the underside of the

result in motor failure prior to activation of the

aft fuselage.

thermal cutoff switch.

ORIGINAL

2−76

NAVAIR 01−F14AAD−1

2.17 PNEUMATIC POWER SUPPLY SYSTEMS

side of the nose wheelwell. Expenditure of bottle pressure

is controlled by a twist−pull operation of the landing

The pneumatic power supply systems consist of three

gear handle. Minimum bottle pressure for accomplishing

independent, stored pneumatic pressure sources for normal

emergency extension of the gear to the down−and−locked

and auxiliary operation of the canopy and for emergency

condition is 1,800 psi. Normal preflight bottle pressure is

extension of the landing gear. The high−pressure bottles for

3,000 psi at 70_ F.

normal canopy operation and emergency landing gear extenĆ

Note

sion are ground−charged through a common filter connection

in the nose wheelwell to 3,000 psi at 70_ F ambient temperaĆ

D Emergency extension of the landing gear shall

ture. Individual bottle pressure is registered on separate

be logged in the Maintenance Action Form

gauges on the right side of the nose wheelwell. An auxiliary

(OPNAV Form 3760−2).

canopy−open N₂ bottle, filter valve, and gauge is on the

D Once the landing gear is extended by emerĆ

turtleback behind the cockpit to allow opening the canopy

gency means, it cannot be retracted while

from the cockpit or ground. Charges may be compressed air;

airborne and must be reset by maintenance

however, pressurized dry nitrogen is preferred because of its

personnel.

low moisture content and inert properties.

D Use of emergency gear extension results in

2.17.1

Normal Canopy Control

loss of nosewheel steering.

The bottle that supplies a pressurized charge for normal

2.18 MISSION COMPUTER SYSTEM

operation of the canopy is on the right side of the forward

fuselage, inboard of the air refuel probe cavity. Expenditure

The MCS consists of two AN/AYK−14 digital comĆ

of bottle pressure for normal operation of the canopy is

puter

(MC1 and MC2) and the dual redundant MIL−

controlled by three (pilot, RIO, and ground) canopy control

STD−1553B buses. The MCS is operated at 16 MHz clock

handles. A fully charged bottle provides approximately 10

speed to perform 1 million instructions per second using up

complete cycles (open and close) of the canopy before

to 1 megabyte of memory. The 1553B bus system in the

reaching the minimum operating pressure of 225 psi.

F−14D uses time division multiplexing (TDM) with informaĆ

tion coded into 20−bit words.

2.17.2

Auxiliary Canopy Open Control

Communication protocol is established by a command

The auxiliary canopy air bottle supplies a pneumatic

response system in which all bus transmissions occur under

charge to translate the canopy aft so that the counter−poise

command of a bus controller or, in case of failure of primary

action of the canopy actuator facilitates opening. It is on the

bus controller, a backup bus controller. Each bus is capable

turtleback behind the canopy hinge line.

of addressing up to 31 remote terminals; however, address 31

is not used in the aircraft. Figure 2Ć42 depicts the physical

Activation of the auxiliary mode can be effected from

connection of the WRAs in the MCS data bus system.

either of the three (pilot, RIO, or ground) canopy control

Remote terminals incapable of communicating directly with

handles. After activation of the auxiliary open mode, the

the MCS on the 1553 data buses are routed through the

control system will not return to the normal mode of operaĆ

converter interface unit for required analog−to−digital and

tion (canopy will lower but will not translate forward) until

digital−to−analog conversion.

the auxiliary selector valve on the aft canopy deck is manuĆ

ally reset (lever in vertical position). Servicing of the auxilĆ

2.18.1

Aircrew Interface

iary canopy air bottle is through the small access panel immeĆ

diately behind the canopy on the turtleback. The reservoir is

The principle aircrew interface with the MCS is

normally serviced to 3,000 psi at 70_ F ambient temperature.

accomplished through the pushbuttons on each MFD. The

A fully charged bottle provides more than 20 operations

RIO has an additional interface through the DEU communiĆ

in the auxiliary open mode. Minimum preflight pressure is

cating directly with the MCS as a remote terminal. The RIO

800 psi.

can also interface indirectly with the MCS through the radar

system digital display.

2.17.3

Emergency Gear Extension

The bottle that supplies the pneumatic force for a

single emergency extension of the landing gear is on the right

2−77

ORIGINAL

NAVAIR 01−F14AAD−1

2.18.2

Operational States

2.19 STANDARD CENTRAL AIR DATA COMPUTER

The MCS has three operational states: startup, full up,

Note

and backup. These states are mutually exclusive and are

The acronyms SCADC and CADC are used

determined automatically based on aircraft operation and

interchangeably throughout this manual.

MC1/MC2 condition.

The SCADC CPU−175/A is installed in F−14D aircraft

The SYS RESET button on the NAV MODE panel

incorporating AFC 793. The SCADC is functionally interĆ

forces both mission computers to transition to the startup

state and execute cold start logic. It can be used to assure the

changeable with the CADC 1166B/A with one difference,

the SCADC software incorporates the static−error source−

aircrew that the MCS is functioning properly and/or to

correction curve required for the true values of Mach

reinitialize the MCS by restarting the OFP. When SYS

number,ăairspeed, and altitude. In aircraft prior to AFC 793

RESET is pressed, the following events occur:

(CADC 1166B/A), aircrew should refer to performance

1. The MCS immediately stops executing the OFP.

charts, NAVAIR 01−F14AAP−1.1 for HUD−displayed altiĆ

tude and Mach number correction curves.

2. The mission computers go off line and run software

BIT.

Note

3. The OFP is automatically restarted.

The standby airspeed indicator is not corrected

for position error.

4. The aircraft goes into the TLN master mode.

The SCADC is a single−processor digital computer

5. Displays revert to defaults.

with a separate, independent, analog, backup wing−sweep

channel. It is capable of making yes and no decisions, solving

Recycling power (by cycling circuit breakers) to the

mathematical problems, and converting outputs to either digĆ

MCS has the same effect as pressing the SYS RESET except

ital or analog form as required by each aircraft system. The

that both hardware and software BIT is performed.

SCADC gathers, stores, and processes pitot pressure, static

pressure, total temperature and AOA data from the aircraft

Note

airstream sensors (see Figure 2Ć43). It performs wing−sweep

and flap and slat schedule computations, limit control and

Cycling subsystem circuit breakers initiates a

electrical interlocks, failure detection, and systems test logic.

cold start for that subsystem. A system reset may

Major systems that depend on all or part of these CADC

be required to resynchronize the MCS and the

functions are shown in Figure 2Ć44.

restarted subsystem.

The following legends appear on the MFD when actiĆ

Refer to the Supplemental NATOPS, NAVAIR

vated by the CADC:

01−F14AAD−1A for a complete description of the MCS

architecture, operational states, and backup operation.

1. RDC SPD (warning legend)  (REDUCE SPEED)

 Indicates flaps down above 225 knots; maximum

2.18.3

Aircraft Master Modes

safe Mach exceeded (2.4 Mach/total temperature

above 388_ F).

There are three aircraft master modes of operation:

takeoff−landing−navigation (TLN), air−to−air (A/A), and air−

2. W/S

(caution legend)

 (WING SWEEP) 

to−ground (A/G). The controls, displays, and avionics equipĆ

Indicates dual wing−sweep channel failure or wing−

ment are tailored as a function of the master mode selected

sweep detent disengaged.

by the pilot. The TLN master mode is entered automatically

when power is applied to the aircraft, when the landing gear

2.19.1

Standard Central Air Data Computer Tests

is down, or when the TLN master mode pushbutton is

selected on the PDCP. The A/A master mode is entered by

2.19.1.1

Built−In Test

pressing the A/A master mode pushbutton on the PDCP,

selecting an air−to−air weapon with the weapon select switch

BIT capabilities provide continuous monitoring of the

on the pilot control stick, or by commanding a radar dogfight

SCADC and its inputs and outputs. The failure indicator

mode. The A/G master mode is entered by pressing the A/G

matrix (Figure 2Ć45) tabulates the functions that are moniĆ

master mode pushbutton on the PDCP.

tored and associated fail indications.

2−79

ORIGINAL

NAVAIR 01−F14AAD−1

INDICATION

REMARKS

FAILURE

CADC − P_S/P_t SENSOR COMPARE DIGITAL

WING−SWEEP LIMIT BUG MAY BE INACCURATE, MAXIMUM

PROCESSOR

SAFE MACH INDICATOR MAY BE INACCURATE, MANEUVER

FLAPS.

CADC WING−SWEEP COMMAND

(SINGLE FAILURE)

CADC WING−SWEEP COMMAND

AUTO WING SWEEP INOPERATIVE

(DUAL FAILURE)

WING SWEEP (SINGLE FAILURE)

WING SWEEP (DUAL FAILURE)

AUTO WING SWEEP INOPERATIVE

CADC MANEUVER FLAP COMMAND

MANEUVER FLAPS VIA THUMBWHEEL INOPERATIVE

MANEUVER FLAP − COMMAND AND

MANEUVER FLAPS VIA THUMBWHEEL INOPERATIVE

[

SERVO MISCOMPARE

MANEUVER FLAP − HYDRAULIC VALVE

[

AND/OR ACTUATOR MISCOMPARE

MANEUVER FLAP − HANDLE AND/OR

HYDRAULIC VALVE MISCOMPARE

AUXILIARY FLAP AND MANEUVER FLAP

MISCOMPARE

AUXILIARY FLAP ASYMMETRY

CADC RUDDER OR STABILIZER

COMMAND AUTHORITY

ANGLE−OF−ATTACK SIGNAL

ANGLE−OF−ATTACK DISPLAY NOT PRESENT ON HUD DURING

LANDING MODE

TOTAL TEMPERATURE SIGNAL

AUTO PILOT CAUTION LIGHT ILLUMINATES IF IN ALTITUDE

HOLD. ALTITUDE HOLD WILL BE DISENGAGED. VERTICAL

SPEED NOT PRESENT ON HUD DURING TAKEOFF AND

LANDING MODE

CADC WING SWEEP INDICATOR OUTPUT

WING SWEEP INDICATOR INACCURATE

ECS FAILURE AND MACH >0.25.

CABIN TEMPERATURE MAY RISE AFTER LANDING. COOLING

AIR ADVISORY LIGHT MAY ILLUMINATE

ECS FAILURE AND MACH >0.4.

CADC− DIGITAL DATA TO CSDC

ALTITUDE AND MACH NOT DISPLAYED ON HUD. ANGLE−OF−

ATTACK DURING LANDING DISPLAY NOT ON HUD. DURING

TAKE OFF AND LANDING VERTICAL SPEED NOT ON HUD.

ALTITUDE HOLD OUTPUT

AUTO PILOT CAUTION LIGHT ILLUMINATES IF IN ALTITUDE

HOLD. ALTITUDE HOLD WILL BE DISENGAGED

ALTITUDE RATE OUTPUT

MACH TRIM OUTPUT

* REDUCE SPEED LEGEND WILL APPEAR IF AIRSPEED >225 KNOTS AND FLAPS ARE DOWN OR MAXIMUM SAFE MACH EXCEEDED

2.4 IMN/TOTAL TEMPERATURE ABOVE 388_ F.

Figure 2Ć45.ĄCADC Processor Indicators

2−81

ORIGINAL

NAVAIR 01−F14AAD−1

2.19.1.2

On−Board Checkout

Under normal operating conditions, the wings are

automatically positioned to the optimum sweep angle for

The CADC performs a self−test during OBC only with

maximum maneuvering performance. The pilot can selecĆ

weight on wheels. When OBC is initiated, normal air data

tively position the wings at sweep angles aft of optimum.

inputs are locked out and in their place constants from the

computer memory are received. Self−test detected failures

A mechanical backup control system is provided for

may be manually reset by pressing the MASTER RESET

emergency and oversweep operations. Details of the wing−

pushbutton.

sweep system are shown in FO−11.

Pressing the MASTER RESET pushbutton for 1 second

The outboard location of the wing pivot reduces the

resets transient failures in the CADC. Activating the master

change in longitudinal stability as a function of wing−sweep

reset circuit recycles the failure detection process in the

angle. Two independently powered, hydromechanical screwĆ

CADC. This recycling process puts off the caution and

jack actuators, mechanically interconnected for synchroĆ

advisory light(s) and may take as long as 10 seconds to check

nization, position the wings in response to pilot or CADC

out the status of the system. If a failure exists, the light(s) will

commands. In flight, the wings can be positioned between

illuminate again. If a transient failure existed, the light(s)

20_ and 68_ wing sweep angle. On the deck, the range is

will remain off.

extended aft to 75_ (oversweep position) to reduce the span

for spotting. Such authority results in a variation of wing span

The following caution and advisory lights are activated

from approximately 64 to 33 feet.

by the SCADC:

Cavities above the engine nacelles and the midfuselage

1. CADC

accommodate the inboard portions of the wing panels as they

2. FLAP

sweep aft. Sealing of the underside is by a wiper seal and

airbag. The bag is pressurized by engine bleed air. Airbag

3. WING SWEEP (advisory)  If the WING SWEEP

pressure is released during oversweep to avoid overloading

advisory light does not recycle when MASTER

of the flap mechanism. An overwing fairing encloses the wing

RESET pushbutton is depressed, the light is actiĆ

cavity and provides a contoured seal along the upper surface

vated by the wing flap controller.

of the wing for the normal range of in−flight sweep angles. The

Three independent SCADC fail signals drive the DFCS

left and right overwing fairing actuators are pressurized by

failure detection circuits. If these signals exist, the DFCS will

the combined and flight hydraulic systems, respectively.

illuminate the following lights:

2.20.1

Wing−Sweep Performance

1. CADC fail signal pitch computer  No Light

Maximum wing−sweep rate (approximately 15_ per

2. CADC fail signal to yaw computer  RUDDER

second) is adequate for most transient flight conditions;

AUTH and HZ TAIL AUTH

however, wing−sweep rate can be significantly reduced or

3. CADC fail signal to roll computer  MACH TRIM.

stalled by negative−g or large positive−g excursions. SuffiĆ

(FCS CAUTION and ARI DGR lights will also be

cient capability has been provided in the system, consistent

illuminated.)

with the sustained performance capabilities of the aircraft.

With a failure of either the combined or flight hydraulic

Note

systems, the wings will move at a reduced rate.

If autopilot is engaged the AUTOPILOT light

will illuminate when any of the three fail signals

exist. If ACL is engaged the AUTOPILOT and

ACLS/AP lights will illuminate when any of the

three fail signals exist.

Slower than normal wing sweep cycling times

Pressing MASTER RESET pushbutton will also

may also be indicative of a failed hydraulic wing

update the wing−sweep and flap commands to their respecĆ

sweep motor or an impending failure. With airĆ

tive feedback signals. As a result, there may be movement in

craft on the ground and both FLT and COMB

the wings and maneuver flaps when MASTER RESET

hydraulic power, the time to sweep the wings

pushbutton is depressed.

from 68° to 20° should not exceed 9 seconds.

Note

2.20 WING−SWEEP SYSTEM

D The overwing fairings and flaps are susceptiĆ

The variable geometry of the wing−sweep system

ble to a high frequency (60 cycles per second),

provides the pilot with considerable latitude for controlling

low−amplitude oscillation that can be felt in

wing lift and drag characteristics to optimize aircraft

the cockpit. This overwing fairing and flap

performance over a broad flight spectrum.

ORIGINAL

2−82

NAVAIR 01−F14AAD−1

buzz is normal and is influenced by the rigging

optimum sweep angle for developing maximum maneuverĆ

of the fairings and air in the hydraulic systems.

ing performance. In addition to providing an automatic wing

positioning function, the programmer also defines the

D Overwing fairing and flap buzz is usually

forward sweep limit that cannot be penetrated using any of

encountered between 0.9 and 1.4 Mach.

the other electrical (manual or bomb) modes. The forward

sweep limiter prevents electrical mispositioning of the wings

2.20.2

Wing−Sweep Modes

from a wing structure standpoint.

Normal control of the wing−sweep position in AUTO,

AFT, FWD, and BOMB modes is by the four−way wing−

Pilot selection of the AUTO mode or automatic transfer

sweep switch on the inboard side of the right throttle grip

from the manual mode causes the AUTO flag to appear in the

(Figure 2Ć46). As an emergency mode of control, changes in

wing−sweep indicator. Once in the AUTO mode, the four−

wing−sweep position can be selected manually with the

way wing−sweep switch can be in the center position without

emergency WING SWEEP handle on the inboard side of the

changing the command mode.

throttle quadrant. The handle is connected directly to the

wing−sweep hydraulic valves. The command source for

2.20.2.2

Manual Mode

positioning the wings depends upon the mode selected by the

The manual wing−sweep mode is commanded by

pilot or, in certain cases, is automatically selected. Electrical

selecting AFT or FWD from the neutral position of the

and mechanical wing−sweep command paths are shown on

wing−sweep switch, driving the wings in the commanded

FO−11. Wing−sweep modes are shown in Figure 2Ć47.

direction to any wing−sweep position aft of the automatic

program. The switch is spring−loaded to return to the center

position. Manual command mode exist unless the wing−sweep

program is intercepted, at which point transfer to the AUTO

mode is automatic. Indication of the existing mode is provided

The emergency wingsweep handle can be moved

by the AUTO and MAN flags in the wing−sweep indicator.

independent of the wings and wingsweep indicaĆ

tors when no hydraulic power and/or electrical

2.20.2.3

Bomb Mode

power are on the aircraft. Care must be taken to

accurately determine the position of the emerĆ

Bomb mode is selected by moving the wing−sweep

gency wingsweep handle prior to application of

switch to the down (BOMB) position. With the switch in

hydraulic power. Inadvertent wingsweep to the

BOMB, the following occurs:

position selected by the emergency wingsweep

1. Wing SWEEP indicator shows MAN flag.

handle may occur anytime hydraulic power is on,

resulting in potential damage to the aircraft.

2. If wing sweep is less than 55_, wings will drive

When positioning the wings during ground

to 55_.

operation other than pilot poststart or postĆ

3. If wing sweep is greater than 55_, wings will not

landing checklist procedures, use the emergency

move.

wingsweep handle to minimize the possibility of

moving the wings inadvertently.

4. If maneuver flaps are extended, they will retract and

wings will sweep to 55_.

Note

As the aircraft accelerates and the AUTO wing−sweep

D When positioning the wings, do not command

schedule is intercepted, the wings will follow the AUTO

opposite direction until wings have stopped in

schedule even though the switch remains in BOMB mode.

original commanded position

(all sweep

Upon decelerating, the wings will sweep forward to 55_ and

modes) to increase motor life.

stop.

D The optimum wing position (triangular index)

and the AUTO/MAN flags may be unreliable

2.20.2.4

Emergency Mode

when the CADC caution light is illuminated.

During normal mode operation of the wing−sweep

system, the wing−sweep control drive servo drives the

2.20.2.1

AUTO Mode

hydraulic valve command input through a spider detent

Selection of the AUTO mode is made by placing the

mechanism. The emergency handle under a transparent

four−way wing−sweep switch in the upper detented position

guard is moved in parallel with the servo output. The

(AUTO) permitting the CADC wing−sweep program to

emergency mode provides an emergency method of controlĆ

position the wings automatically. The program positions the

ling wing sweep. It bypasses the normal command path of the

wings primarily as a function of Mach number but includes

system (CADC and control drive servo loop).

pressure altitude biasing. Wing position is scheduled to the

2−83

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

Wing sweep switch

AUTO Ċ Wing sweep angles are determined by CADC according to wing

sweep program. Detented switch position.

BOMB Ċ Wings are positioned at 55_ or further aft if commanded by the

CADC program. Detented switch position.

AFT/FWD Ċ The pilot can select AFT or FWD wing positions within limits

imposed by the wing sweep program. Switch is spring−loaded to

the center position. When the forward limit is intercepted, the

mode is transferred to AUTO.

Emergency WING

Provides a mechanical means of wing sweep control that overrides the

SWEEP handle

CADC program commands. Wing sweep angles between 20_ and 68_ are

unrestricted except for flap interlocks. Oversweep 75_ is provided with

weight−on−wheels, horizontal stabilizer authority restricter in reduced range,

and air bag pressure dumped.

Wing SWEEP

Displays (from right to left) actual wing sweep position, commanded position

indicator

and wing sweep program position, which is the maximum forward angle at

present airspeed and attitude. Indicator windows show the operating mode.

W/S caution legend

Indicates failure of both wing sweep channels and/or disengagement of spider

on MFD

detent. Wing sweep positioning requires using the emergency wing sweep

handle.

Figure 2Ć46.ĄWing−Sweep Controls and Indicators (Sheet 1 of 2)

ORIGINAL

2−84

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

CADC caution light

Indicates hardware failure and/or that certain computations of the air data

computer are unreliable. Illumination of WING SWEEP advisory light and/or

W/S caution legend on MFD determines pilot action.

WING SWEEP

Indicates failure of a single channel in the system. Illumination of both

advisory light

WING SWEEP advisory and CADC caution light indicates failure of one

channel in CADC.

Figure 2−46. Wing−Sweep Controls and Indicators (Sheet 2 of 2)

To select emergency mode, the handle must be

D If operating in the emergency wing−sweep

extended vertically. The guard should be moved out of the

mode, positively confirm all flaps are

way before the handle is operated. Vertical extension of

retracted prior to attempting AFT wing

the emergency handle provides for better accessibility and

sweep.

leverage. The detent is not disengaged by raising the handle

Note

vertically. An initial fore or aft force of up to 30 pounds

In certain failure modes, the flap indicator may

breakout and 13 pounds maximum is necessary for operation.

not accurately reflect the position of all flaps.

The spider detent is reengaged if the handle is

Since the wing−sweep program acts as a forward limiter

repositioned to the detent (servo) position.

only for the normal modes of operation, the pilot must follow

The emergency WING SWEEP handle incorporates

the following schedule in the emergency mode:

locks at approximately 4_ increments between 20_ and 68_.

1. 0.4 Mach  20_

These locks are provided to eliminate random wing moveĆ

ment in the emergency mode should electrical system

2. 0.7 Mach  25_

transients be experienced. When the locks are engaged, wing

3. 0.8 Mach  50_

movement is inhibited provided that the wings match handle

position. The wing−sweep locks eliminate the need for the

4. 0.9 Mach  60_

installation of wing−sweep servo cutout switches. Locks are

5. 1.0 Mach  68_.

engaged by raising the handle

1 inch from the stowed

position. In order to bypass the locks and select a wing

When operating in the emergency mode, pulling the

position, the handle is raised an additional 1 inch (2 inches

WING SWEEP DRIVE NO. 1 (LD1) and WG SWP DR

from stowed) and moved to the desired position. The handle

NO.Ă2/MANUV FLAP (LE1) circuit breakers on the pilot left

is spring−loaded to return to the lock position when released.

knee panel assures that the electrical command path cannot

The handle can be raised from 20_ to 68_ and oversweep, but

interfere with the emergency mode.

can only be returned to the stowed position at 20_ and

oversweep. This feature is intended to prevent inadvertent

2.20.2.5

Oversweep Mode (75_)

engagement of the AUTO MODE, commanding the wings to

spread causing possible damage to the aircraft or injury to

The wing oversweep mode allows sweeping the wings

aft of 68_ to 75_ during on−deck operation only, thereby

personnel in a confined area. The handle is spring−loaded

toward the stowed position, but requires depressing the

reducing the overall width of the aircraft for deck spotting.

At 75_, the wing trailing edge is over the horizontal tail surface.

release button on the inboard side of the lever in order to

return the handle to the stowed position.

If the over" flag is not displayed in the wingĆ

sweep indicator with the wings in oversweep, the

D Except for wing flap (main and auxiliary) and

stick should remain centered.

oversweep interlocks in the control box, the

emergency mode does not prevent pilot misĆ

With the wings at 68_, oversweep can be initiated by

positioning of the wings from a structural

raising the emergency WING SWEEP handle to its full

standpoint.

extension and holding. Raising the handle releases air pressure

2−85

ORIGINAL

NAVAIR 01−F14AAD−1

from the wing−seal airbags and activates the horizontal tail

2.20.3

Wing−Sweep Interlocks

authority system, restricting the surface deflections to 18_

Automatic limiting of wing−sweep authority is proĆ

trailing edge up and 12_ trailing edge down. During motion

vided under normal in−flight control modes to prevent

of the horizontal stabilizer restrictors, the HZ TAIL AUTH

mispositioning of the wings at conditions that could result in

caution light is illuminated. When the horizontal tail

the penetration of structural boundaries. Wing−sweep interĆ

authority restriction is accomplished

(approximately

locks within the CADC are shown in Figure 2Ć48. Wing

seconds), the HZ TAIL AUTH caution light will go off and

sweep is also electrically inhibited at normal accelerations

the OVER flag on the wing−sweep indicator will be visible.

less than −0.5 g.

This advises the pilot that the oversweep interlocks are free,

allowing movement of the emergency WING SWEEP handle

2.20.3.1

Flap and Slat Wing−Sweep Control Box

75_ and stow. The EMER and OVER flags on the

wing−sweep indicator will be visible.

Electromechanical (auxiliary flaps, oversweep enable)

and mechanical (main flap) interlocks in the control box limit

aft wing−sweep commands at 21_15′, 50_, and 68_. InterĆ

locks in the control box are shown in Figure 2Ć48.. These

interlocks, which serve as a backup to the electronic

interlocks in the CADC, are imposed on both the normal and

D Failure of the oversweep interlocks while

the emergency inputs to the control box and assure nonĆ

trying to achieve oversweep may result in

interference between movable surfaces and the fuselage.

damage to the wingtip and horizontal tail

trailing edges, and the maneuver flap actuator.

2.20.4

Wing−Sweep System Test

D If unusual resistance is encountered while

2.20.4.1

Continuous Monitor

attempting to put the wings into oversweep,

continued aft pressure on the WING SWEEP

The command and execution of the wing−sweep

handle may cause failure of the wing−sweep

system is continually monitored by a failure detection

actuator.

system. The failure detection system in the CADC governs

the change from wing−sweep channel 1 to channel 2 or the

The reverse process takes place when sweeping forĆ

disabling of wing−sweep channel 1 or 2 by switching the

ward from oversweep. However, there is no need to hold the

respective control drive servo off. A single channel failure in

emergency handle in the raised position at 68_. Motion out

the wing−sweep electrical command path is indicated by

of oversweep is completed (wing−seal airbag pressure estabĆ

illumination of the WING SWEEP advisory light followed

lished and horizontal tail authority restriction removed)

by normal operation on the remaining channel. Failure of the

when both the OVER flag and the HZ TAIL AUTH caution

remaining channel is indicated by a W/S caution legend on

lights are off. Six seconds later the WING SWEEP advisory

the MFD and requires that wing−sweep control be exercised

light will illuminate. Upon engagement of the spider detent

through the emergency WING SWEEP handle. Transient

by further unsweeping the emergency handle, MASTER

failures in the CADC can be reset by pressing the MASTER

RESET pushbutton is pressed to clear the WING SWEEP

RESET pushbutton, which recycles the failure detection

advisory light, thus activating the electrical command cirĆ

system.

cuits of the wing−sweep system.

2.20.4.2

Preflight Check

A preflight check of the wing−sweep system to assure

proper operation of the electrical command circuits without

moving the wings should be accomplished after starting

engines while the wings are in oversweep (75_).

When coming out of oversweep and a 68_ wing

position is desired, the wings should be moved

1. Set wing−sweep mode switch to AUTO.

forward to approximately 60_ and then back

to 68_.

Note

The CADC caution light will illuminate and test

will not run if AUTO is not selected on the wing−

sweep switch.

2−87

ORIGINAL

NAVAIR 01−F14AAD−1

2. Press MASTER RESET pushbutton.

2.21.1.1.1

Emergency Flaps

EMER UP enables the pilot to override any electroĆ

3. Set MASTER TEST switch to WG SWP.

mechanical commands that may exist because of malĆ

4. Monitor test by observing:

function of the CADC. To position the flaps, move the FLAP

handle to the end of the normal travel range; then, move the

a. Wing−sweep limit pointer drives to 44°.

handle outboard and continue moving to extreme EMER UP.

While moving the handle, forces may be higher than normal.

b. Illumination of the WING SWEEP advisory light

EMER DN has no function.

and FLAP caution light.

Note

The WING SWEEP advisory light will illumiĆ

nate 3 seconds after test starts, then go off and

A slip clutch assembly is installed between the

illuminate again at 8 seconds into test.

combined system forward flap hydraulic motor

and the center gearbox assembly. While this will

c. RDC SPD warning legend on MFD.

relieve some stall torque on the hydraulic motor,

extremely fast reversals of flap direction while

d. At end of test (approximately 25 seconds) the

flaps are in motion may result in eventual failure

limit pointer will drive to 20_ and the above

of the flap and slat flexible driveshaft.

lights will go off.

2.21.1.2

Maneuver Flap and Slat Thumbwheel

5. Set MASTER TEST switch to OFF.

The maneuver flap and slat thumbwheel is located on

Note

the left side of the stick grip and is spring loaded to the center

position. With LDG GEAR and FLAP handles up, automatic

Ignore illumination of RUDDER AUTH caution

CADC flap and slat positioning can be overridden with pilot

or MACH TRIM advisory lights and motion of

thumbwheel inputs to partially or fully extend or retract the

the control stick if they occur during the test.

maneuvering flaps and slats; however, the next time angle of

attack crosses an extension or retraction threshold, the

2.21

FLAPS AND SLATS

automatic command will again take precedence, unless

The flaps and slats form the high−lift system, which

manually overridden again. Manual thumbwheel command

provides the aircraft with augmented lift during the two

is a proportional command.

modes of operation: takeoff or landing, and maneuvering

2.21.1.3

Main Flaps

flight. The flaps are of the single−slotted type, sectioned into

three panels on each wing. The two outboard sections are the

The main flaps on each wing consist of two sections

main flaps utilized during both modes of operation. The

simultaneously driven by four mechanical actuators geared

inboard section (auxiliary flap) is commanded only during

to a common flap driveshaft. Each wing incorporates a flap

takeoff or landing. The slats consist of two sections per wing

asymmetry sensor and flap overtravel switches for both the

mechanically linked to the main flaps. Flaps down greater

extension and retraction cycles.

than 10_ enables the wheels warning light interlock, and

Cove doors, spoilers, eyebrow doors, and gusses

greater than

25_ enables direct lift control and power

operate with the flaps to form a slot to optimize airflow

approach spoiler gearing.

over the deflected flap. The cove doors are secondary

surfaces along the underside of the wing forward of the flap

2.21.1

Flap and Slat Controls

(Figure 2Ć50.) As the flaps pass 25_ deflection, a negative

Pilot controls for flap and slat takeoff, landing, and

command received from the DFCS depresses the spoilers to

maneuvering modes are illustrated in Figure 2Ć49.

− 4½_ to meet with the cove doors. Because the spoilers do

not span the entire wing as do the flaps, gusses inboard and

2.21.1.1

FLAP Handle

outboard of the spoilers perform the flaps−down function of

the spoilers. With the flaps retracted, the eyebrow doors,

The FLAP handle, located outboard of the throttles, is

which are the forward upper surface of the flaps, are

used to manually command flaps and slats to the takeoff and

spring−loaded in the up position to close the gap between

landing position. Flap handle commands are transmitted by

the trailing edge of the spoiler or guss and the leading edge

control cable to the flap and slat and wing−sweep control box

of the flaps. Mechanical linkage retracts the eyebrow door

where they are integrated with CADC electromechanical

when the flaps are lowered to provide a smooth contour over

inputs to command proper flap and slat position.

the upper surface of the deflected flap.

2−89

ORIGINAL

NAVAIR 01−F14AAD−1

2.21.1.4

Auxiliary Flaps

2.21.2.3

Flap Wing Interlocks

The auxiliary flaps are inboard of the main flaps and are

The main flap and auxiliary flap commands are

interlocked electrically and mechanically with the wing

designed to mechanically lock the auxiliary flaps when in the

sweep to prevent flap/fuselage interference. An electrical

up position. In the event of high dynamic pressure conditions,

interlock in the CADC and a mechanical command in the

a bypass valve within each control valve opens causing the

wing−sweep control box prevent wing sweep aft of 22_ with

auxiliary flap to be blown back, thus avoiding possible

auxiliary flaps extended. In a similar manner, upon extension

structural damage. During loss of electrical power, the

of the main flaps, the wings are electrically and mechanically

control valve is spring−loaded to retract, retracting the

limited to wing−sweep angles less than 50_. The FLAP

auxiliary flaps within 1 minute. The auxiliary flaps use cove

handle is mechanically prevented from moving to the down

doors, eyebrow doors, and gusses identical in purpose and

position if wing position is aft of 50_. If flaps are lowered

operation with those associated with the main flaps.

with wings between 21_ and 50_, main flaps will extend but

auxiliary flaps will remain retracted.

2.21.1.5

Slats

The slats on each wing are divided into two sections,

both of which are driven simultaneously by a single−slat

driveshaft. The slats are supported and guided by seven

curved tracks.

D If flaps are extended with wings between 21_

and 50_, auxiliary flap extension is inhibited and

2.21.2

Flap and Slat Operation

a large nosedown pitch trim change will occur.

D Pulling the FLAP/SLAT CONTR SHUT−OFF

Note

circuit breaker (RA2) will eliminate flap overĆ

travel protection, could eliminate mechanical or

D There is no automatic flap/slat retraction.

electrical main and auxiliary flap interlocks and

may allow the wings to be swept with the flaps

D With flaps extended by the FLAP handle and

partially or fully down in the wing−sweep

an airspeed of 225 knots or greater, the RDC

emergency mode.

SPD legend appears on the MFD and HUD.

2.21.2.4

Maneuver Flap and Slat Mode

2.21.2.1

Normal Operation

The main flaps can be extended to 10_ with the slats

The main flap and slat portion of the high−lift system

extended to 7_ within the altitude and Mach envelope shown

is positioned with a dual redundant hydromechanical servoĆ

in Figure 2Ć51.

loop in response to the FLAP handle command. The auxiliary

flap is a two−position control surface powered by the

Maneuver flaps and slats are automatically extended

combined hydraulic system. With the FLAP handle exceedĆ

and retracted by the CADC as a function of angle of attack

ing 5_ deflection, the auxiliary flaps fully extend. ConĆ

and Mach number (Figure 2Ć52). The schedule commands

versely, they retract for a FLAP handle position equal to or

full maneuver flaps and slats as soon as the slatted wing

less than 5_. The torque of the flap and slat drive hydraulic

maneuvering efficiency exceeds that of the clean wing.

motor is transmitted by flexible driveshafts to each wing.

Note

2.21.2.2

Degraded Operation

CADC maneuver flap commands are automatiĆ

cally reset when the flap handle is placed down

In the event of a combined hydraulic system failure,

greater than 2_, wing−sweep BOMB mode is

outboard spoiler module fluid is automatically directed to a

selected, or maneuver flaps are commanded to less

backup hydraulic motor to lower main flaps and slats only. In

than 1_ by the CADC because of dynamic pressure.

the event of main flap asymmetry greater than 3_, slat

asymmetry greater than 4_, or flap surface overtravel, the

The angle−of−attack input to the CADC from the alpha

flap and slat system is disabled. Flaps and slats will remain

computer is inhibited and will retract the maneuver devices

in the position they were in when failure or malfunction

if they are extended when the LDG GEAR handle is lowered.

occurred. The auxiliary flaps are automatically commanded

This is to ensure that the maneuver devices are retracted

to retract. There is no asymmetry protection for the auxiliary

before lowering the FLAP handle. Maneuver devices

flaps.

extended condition is indicated by a SLATS barberpole and

an intermediate (10_) flap position.

2−93

ORIGINAL

NAVAIR 01−F14AAD−1

To avoid fuel impingement on the fuselage boattail and

nozzles, fuel dump operations are prevented with the speedĆ

brakes extended.

If maneuver devices are not retracted prior to lowĆ

Note

ering the FLAP handle, a rapid reversal of the flaps

D Loss of combined hydraulic pressure with the

will occur with possible damage to the flap system.

speedbrakes retracted or extended will cause

the speedbrakes to move to a floating position.

2.22 SPEEDBRAKES

D The speedbrake/fuel dump interlock is elecĆ

The speedbrakes consist of three individual surfaces,

trically bypassed during a combined hydrauĆ

one upper and two lower panels on the aft fuselage between

lic system failure, enabling the pilot to dump

the engine nacelles (Figure 2Ć53). The speedbrakes may be

fuel when the speedbrakes are floating or

infinitely modulated on the extension and retraction cycle.

modulating. The electrical bypass is enabled

Operating time for full deflection is approximately 2 seconds.

whenever the combined pressure falls below

Hydraulic power is supplied by the combined hydraulic

500 psi.

system (nonisolation circuit), and electrical power is through

the essential No. 2 dc bus with circuit overload protection

D Do not extend the speedbrakes in flight within

on the pilot right circuit breaker panel (SPD BK P−ROLL

1 minute

(nominal) after terminating fuel

TRIM ENABLE) (RB2).

dump operations to allow residual fuel in the

dump mast to drain.

2.22.1

Speedbrake Operation

D A throttle must be held in MIL (or greater) for

Pilot control of the speedbrakes is effected by use of the

approximately

3 seconds in order for the

three−position speedbrake switch on the inboard side of the

automatic function to completely retract the

right throttle grip

(Figure 2Ć54). Automatic retraction of

speedbrake. Anything less will cause partial

the speedbrakes occurs with placement of either or both

retraction.

throttles at MIL or loss of electrical power.

Figure 2Ć53.ĄSpeedbrakes

2−95

ORIGINAL

NAVAIR 01−F14AAD−1

The speedbrakes will start to blowback

(close) at

and pitch acceleration are produced by fore and aft bobĆ

approximately 400 knots and will continue toward the closed

weights. Aircraft overstresses from abrupt stick inputs are

position as airspeed increases to prevent structural damage.

minimized by an eddy current damper that resists large, rapid

A reduction in airspeed will not automatically cause the

control deflections.

speedbrakes to extend to the originally commanded position.

2.23.1.2

Longitudinal Trim

2.23 FLIGHT CONTROL SYSTEMS

Longitudinal trim is provided by varying the neutral

Flight control is achieved through an irreversible,

position of cam and roller feel assembly with an electroĆ

hydraulic power system operated by a control stick and

mechanical screwjack actuator. The manual pitch trim button

rudder pedals. Aircraft pitch is controlled by symmetrical

on the stick is a five−position switch, spring loaded to the

deflection of the horizontal stabilizers. Roll control is

center (off) position (Figure 2Ć58). The fore and aft switch

effected by differential stabilizer deflections and augmented

positions produce corresponding nosedown and noseup trim,

by spoilers at wing−sweep positions less than 62_. DirecĆ

respectively. The manual trim switch is deactivated when the

tional control is provided by dual rudders. During power

autopilot is engaged.

approach maneuvers, the aircraft flightpath can be controlled

through symmetric spoiler displacement by the pilot selectĆ

2.23.1.3

Mach Trim

ing direct lift control. Control surface indicators are shown

in Figure 2Ć55.

Mach trim control is provided by the DFCS and is

continuously engaged to provide automatic Mach trim

The horizontal stabilizer and rudders are powered by

compensation during transonic and supersonic flight. A

the flight and combined hydraulic systems and controlled by

failure of Mach trim compensation is indicated by the

pushrods and bellcranks. A third independent flight control

MACH TRIM advisory light. Transient failures can be reset

hydraulic power source is provided by the backup module.

by depressing the MASTER RESET pushbutton.

Spoiler control is effected by an electrohydraulic, fly−by−

wire system and powered by the combined hydraulic system

The manual and DFCS automatic trim and Mach trim

(inboard spoilers) and outboard spoiler module (outboard

actuator is installed in parallel with the flight control system.

spoilers).

Trim actuation produces a corresponding stick and control

surface movement.

The DFCS includes a stability augmentation system,

an autopilot and auxiliary control functions for spoiler

2.23.2

Integrated Trim System

control, rudder authority control, lateral stick authority

control, and Mach trim compensation.

The ITS is incorporated to reduce longitudinal trim

changes because of the extension and retraction of flaps and

2.23.1

Longitudinal Control

speedbrakes. Disagreement of command position removes

power from the motor and illuminates the INTEG TRIM

Longitudinal control

(Figure 2Ć56) is provided by

advisory light. Transient failures can be reset by pressing the

symmetric deflection of independently actuated horizontal

MASTER RESET pushbutton. ITS schedules are shown in

stabilizers. Control stick motion is transmitted to the

Figure 2Ć59.

stabilizer power actuators by pushrods and bellcranks to dual

tandem actuators independently powered by the flight and

combined hydraulic systems. The power actuators control

the stabilizers symmetrically for longitudinal control and

differentially for lateral control. This is accomplished by

mechanically summing pitch and roll commands at the

When the AIM−54 weapon rail pallet(s) is

pitch−roll mixer assembly. Nonlinear stick−to−stabilizer gearĆ

installed, the speedbrake compensation schedule

ing provides appropriate stick sensitivity for responsive and

in the integrated trim computer changes. If less

smooth control. Longitudinal system authority is shown in

than four AIM−54 missiles are carried on the

Figure 2Ć57.

weapon rails, the ITS may overcompensate for

the speedbrake trim change. In the worst case

2.23.1.1

Longitudinal Feel

(low altitude, between 0.7 and 0.8 Mach, PITCH

SAS OFF, and weapon rails without AIM−54 misĆ

Artificial feel devices in the control system provide the

siles), the ITS can cause an incremental 2 g noseĆ

pilot with force cues and feedback. A spring−loaded cam and

down trim change when the speedbrake is

roller assembly produces breakout force when the stick is

extended. Under these conditions with the

displaced from neutral trim and provides increasing stick

PITCH SAS engaged, maximum trim change is

forces proportional to control stick displacement. Control

reduced to approximately 1 g.

stick forces, proportional to normal acceleration (g forces)

2−97

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

HORIZONTAL tail

Indicated by two pointers marked R (right) and L (left) on a scale

stabilizers position

35_ up and 15_ down. Scale is graduated in 2_ increments.

The inner pointer indicates left wing down or right wing down

(differential stabilizer position).

ANTI SKID SPOILER

BOTH Ċ Antiskid activated. Spoiler brakes operate with weight on

BK switch

wheels and throttle at IDLE.

OFF Ċ Antiskid and spoiler brakes inoperative with weight on wheels.

SPOILER BK Ċ Spoiler brakes operate with weight on wheels and

both throttles at IDLE. Antiskid is deactivated.

HZ TAIL AUTH

Failure of lateral tail authority actuator to follow schedule or CADC

caution light

failure.

RUDDER AUTH

Disagreement between command and position, failure of rudder

caution light

authority actuators to follow schedule, or CADC failure.

Note

The RUDDER AUTH caution light may illuminate when the

in−flight refueling probe is extended. Press the MASTER

RESET button to reset the light.

SPOILERS caution

Spoiler system failure, causing a set of spoilers to be locked down.

light

Note

SPOILERS caution light will not illuminate with SPOILER

FLR ORIDE switches in ORIDE position.

INTEG TRIM advisory

Discrepancy between input command signal and actuator position

light

or an electrical power loss within the computer.

MACH TRIM advisory

Failure of Mach trim actuator to follow schedule.

light

Note

Transient failures involving HZ TAIL AUTH, RUDDER

AUTH, or SPOILERS caution lights and INTEG TRIM and

MACH TRIM advisory lights can be reset by pressing the

MASTER RESET pushbutton.

RUDDER TRIM switch

Controls the electromechanical actuator that varies the neutral position

of the mechanical linkage for rudder trim.

Figure 2−55. Control Surface Indicators (Sheet 2 of 2)

2−99

ORIGINAL

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

DLC engage, disengage,

Momentary depression of the switch with flaps greater than 25 degrees

and chaff switch

down, throttle less than MIL, and no failures in spoiler system engages

DLC. With flaps up, switch will dispense chaff or flares. DLC is

disengaged by momentarily pressing the switch, raising the flaps,

or advancing either throttle to MIL.

Autopilot reference and

With weight on wheels, nosewheel steering can be engaged by

nosewheel steering

depressing switch momentarily. Weight off wheels and autopilot engaged;

pushbutton

switch engages compatible autopilot modes. The switch also disengages

ACL mode.

Autopilot emergency

Disengages all autopilot modes and DLC. Releases all autopilot switches.

disengage paddle

Depressing the paddle switch reverts throttle system from AUTO or

BOOST mode to MAN mode and reverts engines to SEC MODE only

while depressed and with weight on wheels.

Camera and forward

Pilot control of CCTVS, gun camera, and/or forward firing weapons.

weapon firing trigger

First detent of trigger starts gun camera and color cockpit television

sensor (CCTVS).

Figure 2−58. Control Stick and Trim (Sheet 2 of 2)

2.23.2.1

Preflight

2.23.3.2

Lateral Trim

Lateral trim is by differential deflection of the

The ITS is automatically energized with hydraulic and

horizontal stabilizers. The wing spoilers are not actuated for

electrical power applied. It can be checked by operating flaps

lateral trim control. Trim is provided by adjusting the neutral

or speedbrakes and observing a change in indicated stabilizer

position of the spring roller−cam−feel assembly with an

position.

electromechanical screwjack. Left or right deflection of the

roll trim button on the stick grip produces corresponding stick

2.23.3

Lateral Control

movement and left or right wing−down trim, respectively.

The normal stick grip trim switch is inoperative when the

Lateral control (Figure 2Ć60) is effected by differential

autopilot is engaged.

displacement of the horizontal stabilizers and augmented by

Note

wing spoilers at wingsweep positions of less than 62_. With

gear handle up, a ±½−inch stick deadband is provided to

With lateral trim set at other than 0_, maximum

preclude spoiler actuation with small lateral stick commands.

spoiler deflection is reduced in the direction of

The spoilers are commanded to the flush−down (0_) position

applied trim.

at wing−sweep angles of greater than 62_, and roll control is

provided entirely by differential stabilizer. At wing−sweep

2.23.3.3

Lateral Control Stops

angles of 65_ and greater, the hydraulic power to the spoiler

To limit the torsional fuselage loads, variable lateral

actuators is cut off, locking the spoiler in the 0_ position.

control authority stops are installed. The lateral stick stops

Lateral stick commands are transmitted by pushrods and bellĆ

vary according to dynamic pressure airloads from full stick

cranks to the independent stabilizer power actuators and

authority at low Q, to one−half−stick throw limits at high−Q

electrically to the spoiler actuators. Lateral system authority

conditions. Failure of the lateral stick stops is indicated by the

is tabulated in Figure 2Ć61.

HZ TAIL AUTH caution light. Transient failures can be reset

with the MASTER RESET pushbutton. Failure of the stops

2.23.3.1

Lateral Feel

in the one−half−stick position limits low−Q rolling perforĆ

mance. However, ample roll control is available for all

An artificial feel system provides the pilot with force

landing conditions and configurations. Failure in the open

cues and feedback. The lateral feel mechanism is a spring

condition with SAS ON requires the pilot to manually limit

roller−cam assembly with a neutral stick position detent and

stick deflection at higher speeds to avoid exceeding fuselage

a constant stick deflection force gradient.

torsional load limits, as lateral stops do not limit SAS

authority.

ORIGINAL

2−102

NAVAIR 01−F14AAD−1

COCKPIT CONTROL

SURFACE

PARALLEL TRIM

CONTROL

SURFACE

ACTUATION

MODE

MOTION

AUTHORITY

RATE

AUTHORITY

RATE

Differential

Control Stick

Manual

3.5 inches

$7_

36_ per

3/8_ per

Stabilizer

left

second

3.5 inches

right

DFCS

Series

None

$5_

33_ per

second

Inboard

Control Stick

Manual

3.5 inches

$55_

250_ per

None

and

for

left

second

Outboard

Ω v62_

3.5 inches

Spoilers

right

DFCS (ACL)

Series

None

15_ maximum

250_ per

None

(inboard only)

8_ neutral

second

DLC/

Manual

DLC/

inbd only

125_ per

None

Maneuver

17.5_ neutral

second

Flap

−Ă4.5_ down

(minimum)

(inboard only)

Command

+55_ up

Thumb−

wheel

$45_

Ground Roll

Series

None

55_ up

250_ per

None

Braking

second

Armed,

Weight−on−

Wheels

*Lateral

Control Stick

Manual

1.75 inches

$3½_ Diff.

36_ per

Stops

Restricted

left

Stabilizer

second

1.75 inches

**28_ĄSpoiler

250_ per

right

second

* Programmed by CADC (Horizontal Tail Authority) as a function of dynamic pressure.

** Maximum SAS off deflection limits with full lateral stops engaged.

Figure 2Ć61.ĄLateral System Authority

ORIGINAL

2−104

NAVAIR 01−F14AAD−1

Figure 2Ć62.ĄSpoiler Control System

2.23.4

Spoiler Control

for all landing configurations, the DFCS uses the power

approach spoiler gearing whenever the landing gear or main

Four spoiler control surfaces (Figure 2Ć62) on the upper

flaps are down. With DLC engaged in the power approach

surface of each wing augment roll control power and

mode, the inboard spoilers are positioned from normal − 4.5_

implement aerodynamic ground−roll braking. The inboard

to +17.5_ position. Lateral stick inputs result in the spoilers

spoilers also provide DLC. The inboard and outboard spoilers

extending on one side in the direction of stick displacement

are powered and controlled by separate hydraulic and

and depressing toward the landing flaps down drooped

electrical command systems. The DFCS monitors each spoiler

(− 4.5_) stowed position on the other side. This is the primary

panel individually. The pitch computer and outboard spoiler

reason for better roll response in the landing configuration

module control the outboard spoilers; and the roll computer

with DLC engaged.

and the combined hydraulic system control the inboard

spoilers. (Refer to digital flight control system in FO−12.)

2.23.4.1

Lateral Trim and Spoiler Deflection

The inboard spoilers are controlled and monitored by

As mentioned earlier, lateral trim is provided by

the ROLL A and PITCH A computer segregations respecĆ

adjusting the neutral position of the stick. This movement of

tively. The outboard spoilers are controlled and monitored by

the neutral position has an effect on the amount of spoiler

the PITCH B and YAW B computer segregations respecĆ

deflection available. That is, as lateral trim is applied away

tively. Hydraulic actuation of the servo actuators is conĆ

from the neutral trim position, maximum spoiler deflection

trolled by electric servo valves at the actuator and comĆ

is reduced in the same direction (right trim ć less right wing

manded by control stick displacement. The aircraft has two

spoiler deflection).

spoiler gearing curves called cruise and power approach.

Cruise spoiler gearing is the schedule that spoilers follow in

the clean configuration and is shown in Figure 2Ć63. Power

approach is the schedule that spoilers follow with the flaps

down greater than 25° and is shown in Figure 2Ć63 (DLC

Full slat asymmetry (17_) can result in an out−of−

engaged). The power approach spoiler gearing schedule is

control situation at 15 units AOA or greater even

modified to provide predictable roll response with lateral

with 55_ of spoilers available.

stick deflection. To provide the appropriate spoiler gearing

2−105

ORIGINAL

NAVAIR 01−F14AAD−1

Full lateral trim in the same direction as lateral stick

signal from the DFCS is lost (i.e., DFCS power

displacement will still provide approximately 25_ to 35° of

failure). If this bias is reversed, the affected

spoiler deflection to counteract an asymmetric flap and slat

spoiler will extend instead of retracting when the

condition (see Figure 2Ć63). This is sufficient to control full−

command signal is lost. A DFCS power failure

flap asymmetry with symmetrically down slats.

coupled with a reverse spoiler bias will result in

a fully deployed spoiler. All unaffected spoilers

2.23.4.2

Ground−Roll Braking

will remain retracted and will not respond to

flight control inputs until the DFCS command

Aerodynamic ground−roll braking is provided by

signals are restored.

symmetric deflection of all spoilers to +55_. Ground−roll

braking is controlled by the ANTI SKID SPOILER BK

Note

switch on the pilot left vertical console. The three−position

D DFCS synchronization can take up to two

switch allows optional selection of BOTH (spoiler brake and

seconds following a power interrupt. If the

wheel antiskid braking), SPOILER BK (spoiler brake only),

MASTER RESET pushbutton is depressed

or OFF where neither spoilers nor antiskid is armed. With

during the synchronization time, an addiĆ

SPOILER BK or BOTH selected, two conditions are required

tional depression of the MASTER RESET

to actuate the spoilers:

pushbutton will be required to restore spoiler

functionality.

1. Weight on wheels

D Do not press and hold the MASTER RESET

2. Both throttles at idle.

pushbutton. Pressing and holding the

Failure to satisfy any one of the above conditions will

MASTER RESET pushbutton during the

cause the spoilers to return to the down position.

synchronization time will have no effect since

the DFCS computers only recognize the

leading edge of the pulse from the MASTER

RESET pushbutton, and not the fact that the

button is continuously depressed.

Ground−roll braking may fail to extend spoilers

D On deck, when the flap handle is cycled to UP,

on touchdown due to a momentary miscompare

the outboard spoiler module is shut down.

of the weight−on−wheels switches. MASTER

This will cause the outboard spoilers to

RESET should restore normal ground−roll

remain extended if activated. If this occurs,

braking operation.

position the flap handle to DN and deactivate

Note

the spoilers. This may also cause the spoiler

indicators to inaccurately indicate a droop or

During initial spoiler brake operation, it is

down position. If this occurs, position the flap

normal for the indicators in the SPOILER

handle to DN and move the control stick

window to momentarily flip−flop.

laterally to correct spoiler indicators.

2.23.4.3

Spoiler Failure

2.23.4.4

Spoiler Test

Spoiler monitoring is accomplished by directly

Proper spoiler operation is verified when IBIT is run

comparing the commanded spoiler position with the actual

during startup if wings are at 20 degrees and flaps are down.

spoiler position. When a miscompare is detected, the affected

See Chapters 7 and 38.

individual spoiler panel and the corresponding spoiler panel

on the opposite wing are commanded to − 4½_ and the

2.23.5

Yaw Control

SPOILERS caution light illuminated. Transient spoiler

failures can be reset by depressing the MASTER RESET

Yaw control (Figure 2Ć64) is effected by twin rudders,

pushbutton. If the affected spoiler panel is mechanically

one on each vertical tail. The rudder pedals adjust through a

stuck−up, the DFCS automatically restores normal operation

10−inch range in 1−inch increments with the adjust control on

of the opposite spoiler panel within 2 seconds.

the lower center pedestal, forward of the control stick.

Yaw commands are transmitted mechanically from the

rudder pedals to the rudder power actuators by pushrods and

bellcranks. Tandem power actuators are powered indepenĆ

dently by the flight and combined hydraulic systems. Yaw

The spoiler actuators are mechanically biased to

system authority is tabulated in Figure 2Ć65.

the retracted position in order to cause the

spoilers to retract in the event that the command

2−107

ORIGINAL

NAVAIR 01−F14AAD−1

2.23.5.1

Rudder Feel

Before DLC can be engaged, the following conditions

are required:

Artificial feel is provided with a spring roller−cam

mechanism similar to the longitudinal and lateral feel systems.

1. Flaps down greater than 25_.

Rudder force with pedal deflection is nonlinear with a

2. Throttles less than MIL power.

relatively steep gradient about the neutral detent and

gradually decreasing with increased pedal travel.

3. Inboard spoilers operational.

4. Pitch B, and yaw B computers operational.

2.23.5.2

Rudder Trim

Rudder trim is effected by varying the neutral position

5. Operable combined hydraulic pump.

of the feel assembly with an electromechanical screwjack

actuator. Rudder trim control is actuated by a three−position

2.23.6.1

DLC Operation

switch on the left console outboard of the throttle quadrant.

DLC is engaged with the control stick DLC switch and

Left (L) and right (R) lateral switch movement commands

commanded by the thumbwheel. The thumbwheel is spring−

left and right rudder trim, respectively. The switch is

loaded to a neutral position. Forward rotation of the wheel

spring−loaded to the center off position. Trim actuation

extends spoilers and aft rotation retracts them proportionally

produces an associated movement of the rudder pedals,

to the degree of thumbwheel rotation. Absolute spoiler

rudders, and rudder indicator.

deflection is dependent upon lateral stick position

(see

Figure 2Ć63). DLC is provided by the yaw computer.

2.23.5.3

Rudder Authority Stops

Upon engagement of DLC, the roll computer extends

Rudder authority control stops limit rudder throws in

the inboard spoilers from the landing flaps down drooped

the high−Q flight environment. Rudder deflection limits are

(− 4½_) position to +17.5_ above the flush (0_) position. The

scheduled by the CADC, commencing at about 250 knots.

pitch computer displaces the trailing edges of the horizontal

Above approximately 400 knots, the stops are fully engaged,

stabilizers

2.75_ down from their trim position. If the

restricting manual rudder deflection to 9.5_. Disagreement

thumbwheel control is rotated fully forward, the spoilers

between command and position removes power from the

extend to their 55_ position and the stabilizer trailing edges

motor and illuminates the RUDDER AUTH caution light.

remain at 2.75_. This increases the rate of descent. If the

thumbwheel control is rotated fully aft, the spoilers retract to

their − 4.5_ position and the stabilizer trailing edges return

to the trim position. This decreases the rate of descent.

2.24 DIGITAL FLIGHT CONTROL SYSTEM

A CADC failure may drive the rudder authority

stops to 9.5_. This condition should be deterĆ

The DFCS (FO−12) augments the aircraft natural

mined prior to making a single−engine or crossĆ

damping characteristics and provides automatic commands

wind landing. With the 9.5_ stops in, rudder

for control of attitude, altitude, heading, and approach modes

control may be insufficient to maintain direcĆ

selected by the pilot. All DFCS functions are integrated into

tional control with single−engine afterburner

the primary flight control system.

operation or during crosswind conditions. NoseĆ

wheel steering authority is greatly reduced with

The DFCS also provides an Up and Away Automatic

the 9.5_ stops engaged.

Rudder Interconnect (UA−ARI) to enhance departure resistĆ

ance, spin recovery and high angle of attack flying qualities,

2.23.5.4

Rudder Pedal Shaker

and a Power Approach Automatic Rudder Interconnect

(PA−ARI) to enhance the landing approach flying qualities.

The rudder pedal shaker operates during IBIT and in

The DFCS consists of three computers, one computer for

flight when the landing gear is extended and angle of attack

each axis (pitch, roll, and yaw). Each computer has two

is above ~20 units. Rudder pedal shaker will deactivate once

distinct and independent processors called channels or

the angle of attack is reduced below ~19 units.

segregations (one A" and one B" channel per axis), each

controlling one of the dual series servoactuators. All

2.23.6

Direct Lift Control

channels share data through cross channel data links.

During landing approaches, the inboard spoilers and

A BIT capability is provided to exercise in−flight

horizontal stabilizers can be controlled simultaneously to

monitoring and to conduct an automatic operational readiĆ

provide vertical glidepath correction without changing

ness test for preflight checks. DFCS rates and authorities are

engine power setting or angle of attack. Only the inboard

tabulated in Figure 2Ć66.

spoilers are used for DLC.

2−109

ORIGINAL

NAVAIR 01−F14AAD−1

AXIS

ACTUATOR

SURFACE

AUTHORITY

SURFACE RATE

Pitch

Dual Series SAS

Stabilizer

± 3_

20_ per second

Dual Series ITS

Stabilizer

± 3_

3_ per second

Parallel (ACL only)

Stabilizer

10_ TED

36_ per second

33_ TEU

Parallel Trim

Stabilizer

10_ TED

0.1_ per second

18_ TEU

Roll

Dual Series

Differential Stabilizer

± 5_

33_ per second

Spoilers (ACL)

15_ maximum

250_ per second

Yaw

Dual Series

Rudder

± 19_

80_ per second

Figure 2Ć66.ĄDFCS Rates and Authorities

2.24.1

Stability Augmentation System

Note

Depressing the paddle switch does not disable

Stability augmentation is provided for all three aircraft

the PITCH and ROLL SAS. If problems are

axes (pitch, roll, and yaw) and is controlled by the three

suspected with any SAS axis, the appropriate

STAB AUG switches on the upper half of the DFCS control

STAB AUG switch must be manually selected

panel (DCP) (Figure 2Ć67). SAS is engaged by placing these

OFF. Depressing the paddle switch will disĆ

switches to ON during normal poststart procedures. The

engage the autopilot and DLC inflight, revert the

PITCH, ROLL, and YAW STAB AUG switches are manually

throttles to MANUAL mode on deck, and revert

operated toggle switches mechanically held in the selected

fuel control to SEC mode.

ON or OFF position.

2.24.1.1

DFCS Control Panel/Fault Display

The PITCH SAS incorporates a pitch rate feedback

function that is reduced as airspeed is increased above 650

The DFCS control panel (Figure 2Ć67), located on the

KIAS. This is necessary to maintain adequate control system

pilot’s left side console, includes all the controls for the

stability and is not noticeable. The ROLL SAS is independent

DFCS and an LED alphanumeric fault display with the

with the landing gear up, at low angle of attack (less than 15

associated INC and DEC pushbuttons to control display

units), and at supersonic flight conditions. At all other condiĆ

operation. This fault display is intended for ground use only

tions, the ROLL SAS is part of the UA−ARI and PA−ARI.

to assist in the troubleshooting and repair of the DFCS and

Similar to the PITCH SAS, the roll rate feedback is reduced

related components.

as airspeed is increased above 300 KIAS. With the landing

gear down, the YAW SAS becomes part of the PA−ARI.

2.24.1.2

Up and Away Automatic Rudder

Interconnect (UA−ARI)

All SAS switches should remain ON during flight.

Deselection of either the ROLL or YAW SAS switch will

The UA−ARI is selected when the landing gear handle

disable the affected SAS axis and all ARI functions, and

is up and provides several functions designed to improve high

illuminate the ARI/SAS OUT caution light. Deselection of

angle of attack flying qualities and departure resistance

the PITCH SAS switch will disable the PITCH SAS, but no

(Figure 2Ć67). These include:

caution light will illuminate since no restriction exists for

D Differential Stabilizer Fadeout

PITCH SAS OFF.

D Lateral Stick−to−Rudder Interconnect (LSRI)

D Low airspeed/high angle of attack cross

control (LSXC)

D Wing Rock Suppression

Maneuvering with YAW SAS OFF or inoperative

shall not be conducted above 15 units AOA with

D Spin Recovery Function.

landing gear retracted.

ORIGINAL

2−110

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

PITCH STAB AUG

Engages dual−channel pitch stability augmentation.

engage switch

ROLL STAB AUG

Engages dual−channel roll stability augmentation.

engage switch

YAW STAB AUG

Engages dual−channel yaw stability augmentation.

engage switch

DFCS fault display

With weight on wheels, the DFCS fault display yields three categories of fault

codes including: currently existing failures (FAIL), in−flight detected failures

(FLT), and IBIT detected failures (IBIT).

INC pushbutton

With weight on wheels, depressing the INC pushbutton scrolls forward through

logged DFCS fault codes.

DEC pushbutton

With weight on wheels, depressing the DEC pushbutton scrolls backward

through logged DFCS fault codes.

AUTOPILOT ENGAGE−

ENGAGE Ċ Engages autopilot. PITCH, ROLL, and YAW SAS switches

OFF switch

must be engaged. No warmup required. Engages attitude hold.

Requires weight off wheels.

OFF Ċ

Disengages autopilot.

HDG−OFF−GT switch

HDG Ċ

Autopilot will lock on constant aircraft heading when aircraft is

less than "5 roll.

OFF Ċ

Disengages heading hold and ground track.

GT Ċ

Selects autopilot ground tracking computed at time of

engagement using inertia navigation system (INS) data.

Engaged by nosewheel steering pushbutton.

ALT−OFF switch

ALT Ċ

Autopilot will maintain barometric altitude. Engaged by

nosewheel steering pushbutton.

OFF Ċ

Disengages altitude mode.

VEC/PCD−OFF−ACL

VEC/PCD Ċ Autopilot roll axis commands steer aircraft using data link

switch

signals for vectoring. If the precision course direction (PCD)

discrete is present, both roll and pitch axis commands are used.

Engaged by nosewheel steering pushbutton.

OFF Ċ

Disengages VEC/PCD and ACL modes.

ACL Ċ

Autopilot will accept data link signals for carrier landing, using

spoilers for roll and parallel servo for pitch. Only pitch

commands are transmitted to stick movement. Engaged and

disengaged by nosewheel steering pushbutton.

ACLS/AP caution light

Autopilot and automatic carrier landing system (ACLS) mode disengaged.

A/P CPLR advisory

Indicates the aircraft can be coupled to the ACL system for a mode I or

legend on MFD

mode IA approach. A/P CPLR legend remains displayed in conjunction

with the CMD CTRL legend after coupling is accomplished.

A/P REF advisory

Autopilot mode is selected but is not engaged. (Except attitude and

legend on MFD

heading hold.)

Figure 2−67. DFCS Controls and Indicators (Sheet 2 of 3)

ORIGINAL

2−112

NAVAIR 01−F14AAD−1

NOMENCLATURE

FUNCTION

PITCH SAS caution light

Indicates inoperative pitch channel or PITCH SAS failure.

FCS CAUTION light

Indicates DFCS failure has occurred. If no other lights are illuminated, indicates

loss of redundancy only (subsequent failure may result in loss of significant

DFCS functionality).

ROLL DGR caution light

Indicates inoperative roll channel and degraded roll authority.

ARI DGR caution light

Indicates degraded ARI performance. If caused by loss of a Mach number

signal, LSXC and wing rock suppression functions will be inoperative.

YAW DGR caution light

Indicates inoperative yaw channel and degraded yaw authority.

ARI/SAS OUT caution

Indicates loss of either ROLL or YAW SAS and all ARI functions.

light

Will be illuminated if either the ROLL STAB AUG or YAW STAB AUG switches

are selected off.

AUTOPILOT caution light

Indicates failure of one or more of pilot relief modes.

Autopilot reference and

Engages the ALT, GT, ACL or VEC/PCD autopilot mode selected.

nosewheel steering

Autopilot must be engaged and compatible autopilot modes selected.

pushbutton

Also disengages ACL mode. Requires weight off wheels.

Autopilot emergency

Disengages all autopilot modes and releases all autopilot switches.

disengage paddle

PLM pushbutton

With the A/P CPLR legend and the VEC/PCD ACL switch latched in the

ACL position, depressing the PLM pushbutton engages and disengages

the ACL mode and autopilot.

Figure 2−67. DFCS Controls and Indicators (Sheet 3 of 3)

These functions are active throughout the subsonic

by activation of LSXC or spin recovery functions). As Mach

flight envelope and are scheduled with Mach number and

number is increased, the differential stabilizer is faded out at

angle of attack (Figure 2Ć68). The effects of these functions

a lower AOA. This reduces the effects of kinematic coupling

on flight characteristics are discussed in Chapter 11.

and results in less adverse sideslip with lateral stick deflecĆ

tion as angle of attack is increased and reduces the tendency

Note

for lateral control induced departures.

The primary AOA input for control law schedulĆ

The LSRI function gradually applies coordinating

ing is based on degrees AOA provided by the ARI

rudder with lateral stick as angle of attack is increased above

alpha nose−probe vice units AOA as displayed on

~15 units AOA. A maximum of +/−19 deg coordinating

the cockpit AOA indicator provided by the ADD

rudder is provided by the LSRI above ~23 units AOA. This

AOA side−probe. Descriptions of control law

results in the desired roll response with lateral stick input

functions are written in units AOA, but it should

alone at elevated AOA.

be noted that the correlation between units and

degrees AOA is a function of Mach number.

The LSXC function provides a means to override the

differential stabilizer fadeout and LSRI functions when

The differential stabilizer fadeout function reduces the

CADC Mach number is below 0.4 and angle of attack is

amount of differential stabilizer the pilot can command as

above ~30 units. At these conditions, the pilot can command

angle of attack and Mach number are increased. Below ~15

up to +/−10 deg differential stabilizer deflection in the direcĆ

units angle of attack, the pilot can command up to the maxiĆ

tion of lateral stick and up to +/−30 deg rudder deflection in

mum +/−12 deg differential stabilizer authority. Above ~30

the direction of commanded rudder. Because rudder

units angle of attack, the differential stabilizer is limited to

a maximum of +/−2 deg deflection (except when overridden

2−113

ORIGINAL

NAVAIR 01−F14AAD−1

effectiveness is diminished above 30 units angle of attack, the

roll rate. Because the pilot’s roll rate command

LSXC function is necessary to provide adequate roll/ yaw

is based upon lateral stick position from a

maneuvering capability at extreme angles of attack. When

trimmed position, it may be necessary to select

LSRI is no longer effective, the aircraft can still be rolled

ROLL SAS OFF, trim the airplane, and then reseĆ

through LSXC by manually applying greater than 1.75 inch

lect the ROLL SAS ON. The same procedure

of rudder pedal in the direction of the desired roll, and greater

may be necessary to trim an airplane with a latĆ

than 1 inch lateral stick in the opposite direction. This will

eral store asymmetry, trapped wing fuel, etc.

create an adverse yaw response causing the aircraft to

The LSRI function gradually fades in coordinating

roll/yaw in the same direction as the rudder input.

rudder with lateral stick as angle of attack is increased above

The wing rock suppression function uses roll rate

~10 units AOA. This minimizes adverse yaw response from

feedback to command the differential stabilizer and rudder to

lateral stick only inputs, greatly enhancing heading and

damp lateral−directional oscillations from between 20 to 30

centerline capture during lineup corrections. At normal

units AOA. This results in smoother tracking capability for

approach conditions (15 units angle of attack, flaps down), up

the majority of the maneuvering flight envelope. Wing rock

to +/− 19 deg coordinating rudder is provided by the LSRI.

suppression is disabled when a pedal input greater than 2

Raising the flaps decreases the amount of coordinating

inches is applied or CADC Mach number is above 0.77 to

rudder available.

prevent the system from applying inadvertent cross control

The Dutch roll damping function provides sideslip rate

inputs and allow the pilot to roll the aircraft with rudder

feedback to the rudder to reduce directional nose wandering

inputs alone.

during approach. Airplane roll rate, yaw rate, lateral

The spin recovery function applies full SAS authority

acceleration, Mach number, and angle of attack are used to

of up to +/−19 deg rudder and up to +/−5 deg differential

calculate sideslip rate. At normal approach conditions, the

stabilizer to oppose yaw rate during a departure. The spin

sideslip rate feedback to rudder provides a deadbeat Dutch

recovery function is activated when angle of attack is above

roll response.

~30 units and yaw rate is above 20 deg/sec. The spin recovery

The spiral mode stabilization function provides yaw

inputs are in addition to the pilot’s mechanical inputs,

rate feedback to the differential stabilizer to reduce bank

providing full control surface authority for departure recovĆ

angle excursions during stabilized turns. At normal approach

ery

(+/−30 deg rudder opposite roll/yaw and

+/−12 deg

conditions, the yaw rate feedback to differential stabilizer

differential stabilizer into roll/yaw). Misapplied pilot

provides an essentially neutral spiral mode.

recovery inputs are limited to +/−11 deg pro−spin rudder and

+/−2 deg pro−spin differential stabilizer.

2.24.1.4

Aircraft Sensors

2.24.1.3

Power Approach ARI

The DFCS uses the aircraft sensor inputs distributed to

the various computer channels as shown in Figure 2Ć69. The

The PA−ARI is selected when the landing gear handle

DFCS distributes sensor inputs to all computer channels

is down and provides roll rate command, LSRI, Dutch roll

through cross channel data link (CCDL) communication.

damping, and spiral mode stabilization functions. In addiĆ

Each computer channel compares like sensor data

(for

tion, the spoiler gearing has been modified for 0.1 inch lateral

example, yaw rate A, B, and M) to determine validity of each

stick spoiler breakout to improve roll sensitivity and

input and then consolidates the good inputs. The consoliĆ

predictability.

dated sensor inputs are then used to generate output

The roll rate command function tailors differential

commands. This provides an additional level of monitoring

stabilizer to maintain a constant lateral stick to roll rate

and redundancy.

relationship. This is achieved by comparing the roll rate

Note

command (based upon lateral stick position as measured

from trimmed position) to the actual roll rate, then increasing

Loss of a computer segregation or individual

or decreasing the differential stabilizer deflection to maintain

cross channel data link will result in loss of

the commanded roll rate. The control gains are designed to

sensor information provided by the affected segĆ

provide approximately 20 deg/sec roll rate per inch of lateral

regation or link and illuminate the appropriate

stick deflection from trim.

caution/advisory lights.

Note

The aircraft sensors are supplemented by a pitch/roll

voter monitor and air data redundancy management algoĆ

The PA−ARI will perceive a lateral trim offset as

rithm to provide a fail−operational capability following a

an uncommanded roll rate and will attempt to

single sensor failure. Following a second sensor failure, the

reduce roll rate to zero with the stick in the

DFCS reverts to a fail−safe configuration.

trimmed position. As a result, it is possible to

have the stick offset due to lateral trim with zero

2−115

ORIGINAL

NAVAIR 01−F14AAD−1

2.24.1.4.1

Pitch/Roll Voter Monitor

The pitch/roll voter monitor (PQVM) algorithm proĆ

vides triple redundancy for the existing duplex pitch and roll

rate gyro sensors. The PQVM operates by calculating aircraft

Pulling the Alpha computer circuit breaker will

pitch and roll attitudes from three axis rate information and

result in loss of the primary AOA source. This

comparing this against pitch and roll attitudes supplied by the

may degrade DFCS performance as the backup

IMU/INS. If a pitch or roll rate sensor miscompare occurs,

ADD AOA is subject to sideslip−induced errors.

the PQVM algorithm is used to select the remaining good

For the DFCS to operate properly, the Alpha

sensor signal. The monitor does not provide additional rate

computer should not be disabled.

information for averaging with the sensor signals. Since only

pitch and roll attitude are available from the IMU/INS, the

The primary Mach number input is provided by the

effectiveness of the monitor depends on aircraft attitude. The

CADC calculated from the left and right pitot−static probe

monitor is incapable of detecting pitch rate failures with the

inputs. Total and static pressure data (Mach number) inputs

wings near vertical (±90 deg bank angle) or roll rate failures

from the left and right AICS programmers are used for

with the fuselage near vertical (±90 deg pitch angle).

comparison. Similar to the AOA implementation, the upwind

AICS Mach number is selected based upon estimated sideslip

Note

and is compared with the CADC Mach number. When the

estimated sideslip angle is less than

2 deg, the Mach

Transient IMU/INS attitude failures will result in

calculations from both the left and right AICS programmers

a PQVM miscompare. This will be indicated by

are cross checked to determine pressure validity. An AICS

an FCS CAUTION light accompanied by PS and

cross check miscompare (1^st Mach fault) results in a loss of

RS acronyms. Depressing the MASTER RESET

Mach redundancy and is indicated by an FCS CAUTION

pushbutton will restore normal PQVM operation

light and PS acronym. Mach scheduling functions are still

and clear the failure indications. Failed velocity

being performed using the CADC Mach number. A single

information from the IMU/INS does not affect

Mach miscompare between the CADC Mach number and the

operation of the PQVM function.

upwind AICS Mach number results in a default Mach number

being set and is indicated by FCS CAUTION, ARI DGR

2.24.1.4.2

Angle of Attack / Mach Redundancy

caution lights and PS acronym. Any subsequent Mach failure

Management

(2^nd Mach fault) results in the UA control laws configuring

to a fail−safe degraded mode and is indicated by PITCH SAS,

The primary angle of attack (AOA) input is provided

FCS CAUTION, ROLL DGR, ARI DGR, and ARI/SAS

by the ARI alpha nose−probe (AOA range 0° to +37°). The

OUT caution lights and PS acronym with the landing gear

fuselage−mounted ADD AOA side−probe (AOA range −4° to

handle up. Subsequent Mach failures do not further degrade

+24° for Mach <0.4) is used for comparison and as a backup

the PA control laws and are indicated by an FCS CAUTION,

AOA source in the event of a detected failure of the ARI

ARI DGR lights and PS acronym with the landing gear

AOA input. Although two AICS AOA inputs are available

handle down.

(AOA range

0° to

+30° for Mach

>0.5), significant

sideslip−induced errors distort these measurements as a

function of their locations on either side of the fuselage.

Location of sensor probes is discussed in paragraph 2.32

(Pitot Static System). A sideslip estimation routine, based

Pulling either AICS programmer circuit breakers

upon lateral acceleration, differential stabilizer, and rudder

(LF1 or LG1) will result in DFCS air data failures

position, is implemented to select the upwind AICS AOA

and degraded control system capability. This will

source. This AICS AOA is utilized as a triplex monitor to

be indicated by one or more of the following

vote out a failed ARI or ADD AOA (1^st AOA fault), but is not

caution lights depending on flight condition:

of sufficient accuracy to be used as a primary source for

FCS CAUTION, PITCH SAS, ROLL DGR, and

control law gain scheduling. In the event of a subsequent

ARI DGR. Once the AICS circuit breaker is

miscompare between the remaining AOA sources (2^nd AOA

reset, depressing MASTER RESET should

fault), the ARI control laws are reverted to fail−safe fixed

restore normal operation and clear the failure

values for AOA. First AOA miscompares are indicated by an

indications.

FCS CAUTION light and PS acronym. Second AOA

Note

miscompare are indicated by FCS CAUTION, ROLL DGR,

ARI DGR, and ARI/SAS OUT caution lights and PS acronym

D Extremely aggressive maneuvering at high

with the landing gear handle up and FCS CAUTION, ARI

AOA or intermittent CADC problems may

DGR caution lights and PS acronym with the landing gear

result in a transient AOA or Mach misĆ

handle down.

compare. Depressing the MASTER RESET

ORIGINAL

2−116

NAVAIR 01−F14AAD−1

pushbutton will restore normal operation and

2.24.1.6

Digital Flight Control Computers

clear the failure indications for all single

The DFCS consists of three computers, one for each

failure situations. Dual failure of either the

axis (pitch, roll, and yaw) (Figure 2Ć69). Each computer has

Mach or AOA inputs is not resettable with

two distinct channels/segregations (A and B), each in turn

MASTER RESET.

controlling one of the dual series servoactuators in the

D With a dual AOA or Mach failure set, lowerĆ

respective axis. The DFCS uses cross channel data link

ing the landing gear handle will result in the

communication to provide redundancy management so that

PITCH SAS (Mach only), ROLL DGR, and

a miscompare between the A and B channels of a given axis

ARI/SAS OUT lights extinguishing. This

cannot result in a loss of the entire axis. Each A channel (for

does not imply the dual failure is no longer

example, pitch A) is monitored by every B channel (pitch B,

present, only that the impact of the dual failĆ

roll B, and yaw B) and conversely each B channel is

ure on PA control functions are less severe

monitored by every A channel. As a result of this monitoring

than on the UA control functions due to the

structure, loss of a single channel will only result in loss of

more confined operational flight envelope.

functions controlled by that channel and the corresponding

series servoactuator, rather than the entire axis

(sensor

2.24.1.5

Sensor Failures

information provided by the affected channel will also be

lost). Loss of a second channel will significantly degrade

First sensor input failures are indicated by the FCS

DFCS performance or result in a complete loss of all DFCS

CAUTION light and the appropriate acronym (PS, RS, YS,

functions. Computer failure effects and cockpit indications

or AM). An FCS CAUTION light with no other lights indiĆ

are summarized in Figure 2Ć70.

cates a loss of sensor redundancy with no loss of functionĆ

ality. The only single sensor failure to light more than the

2.24.1.6.1

Flight Control Computer Reset

FCS CAUTION light is a single Mach miscompare that is

Each computer channel is independently powered and

indicated by FCS CAUTION, ARI DGR caution lights and

can be disabled by pulling the appropriate DC or AC circuit

PS acronym. With a single Mach miscompare, the LSXC and

breaker. In general, there should be no need to cycle DFCS

wing rock suppression functions will be lost. Depressing the

circuit breakers. If power is removed from a channel and then

MASTER RESET pushbutton should clear transient first

restored, the pilot must depress MASTER RESET to restore

sensor failures and failure indications. Second sensor failures

full DFCS functionality. If power is removed from two or

result in loss of functions in the affected axis. The affected

more of the same letter" channels/segregations, most or all

axis will be indicated by a PITCH SAS, ROLL DGR, or YAW

of the DFCS functions will be lost. Restoring power will

DGR caution light. For failures affecting either the roll or

cause the DFCS to execute a power on reset" (POR) and the

yaw axis, all ARI functions may be lost. Partial loss or

system will re−initialize interpreting the current sensor

degrade of the ARI is indicated by an ARI DGR caution light.

information as valid. This can create a potentially hazardous

Complete loss of the ARI is indicated by the ARI/SAS OUT

situation under conditions where a dual sensor failure

light. For second AOA, Mach, or roll rate sensor failures the

occurred prior to restoring power. When the DFCS re−

spin recovery function is retained even though the ARI/SAS

initializes it is possible for the failed signals to be interpreted

OUT light is illuminated.

as valid and the remaining good signal to be interpreted as

invalid. Therefore, careful consideration should be given

DFCS failure indications and effects are summarized

before executing a POR airborne, since it can result in

in Figure 2Ć70. The DCP fault codes are listed in alpha−

erroneous DFCS commanded control deflections.

numeric order followed by their classification to the IBIT

and/or OFP fault group(s). The IBIT fault group includes

fault codes generated by the IBIT self−test and will be identiĆ

fied in the DCP FAULT DISPLAY following the IBIT"

group header. The OFP fault group includes fault codes genĆ

erated by the PBIT and/or ABIT self−test(s) and will be idenĆ

If a dual failure has been declared, performing a

tified in the DCP FAULT DISPLAY following the FAIL"

power on reset" to clear the failure can result in

and/or FLT" group header. The failure indication and potenĆ

erroneous DFCS commanded control deflections.

tial functions lost are listed for each individual fault code.

The associated caution lights and maintenance file acronyms

When performed on deck, the DC breakers (ROLL A

are listed for each fault code in addition to any applicable

DC, YAW B DC, and YAW A DC − 8B4, 8B5, 8B6) are

notes. In certain cases, potential functions lost for multiple

generally used to avoid inadvertently inducing additional

failures (also known as a dual faults or 2^nd faults) have been

faults. A MASTER RESET is required to extinguish the

identified.

resulting DFCS caution lights following restoration of power.

ORIGINAL

2−120

NAVAIR 01-F14AAD-1

CAUTION LIGHTS

FAILURE

FUNCTIONS LOST

SENSOR FAILURES

PITCH RATE 1F

PGYx

REDUNDANCY LOSS ONLY.

PITCH RATE 2F

PGY7

PITCH SAS OFF.

ROLL RATE 1F

RGYx

REDUNDANCY LOSS ONLY.

ROLL RATE 2F

RGY7

ROLL SAS/ARI OFF. SPIN

RECOVERY RETAINED.

YAW RATE 1F

YGYx

REDUNDANCY LOSS ONLY.

YAW RATE 2F

YGYx

YAW SAS/ARI OFF. ALSO

(PA MODE)

GENERATES PQVM FAULT.

YAW RATE 2F

YGYx

ROLL AND YAW SAS/ARI

(UA MODE)

OFF. ALSO GENERATES

PQVM FAULT.

LATERAL

LATx

REDUNDANCY LOSS ONLY.

ACCELERATION 1F

LATERAL

LATx

YAW SAS/ARI OFF.

ACCELERATION 2F

(PA MODE)

LATERAL

LATx

ROLL AND YAW SAS/ARI

ACCELERATION 2F

OFF.

(UA MODE)

LATERAL STICK

RCPx

REDUNDANCY LOSS ONLY.

POSITION 1F

LATERAL STICK

RCPx

ROLL SAS/ARI OFF.

POSITION 2F

NO SPOILERS.

RUDDER PEDAL

RPPx

REDUNDANCY LOSS ONLY.

POSITION 1F

RUDDER PEDAL

RPPx

NO PEDAL FADEOUT

POSITION 2F

(NO SLIPPED

(PA MODE)

APPROACHES).

RUDDER PEDAL

RPPx

NO LSXC OR WING ROCK

POSITION 2F

SUPPRESSION.

(UA MODE)

ANGLE OF ATTACK

NOTE(1)

REDUNDANCY LOSS ONLY.

ANGLE OF ATTACK

NOTE(1)

SLIGHTLY DEGRADED

2F (PA MODE)

PA-ARI FUNCTIONS.

Figure

2-70.

DFCS Failure Modes and Indications (Sheet

1 of 4)

2-121

CHANGE 1

NAVAIR 01-F14AAD-1

CAUTION LIGHTS

FAILURE

FUNCTIONS LOST

ANGLE OF ATTACK

NOTE(1)

ROLL SAS/ARI OFF. SPIN

2F (UA MODE)

RECOVERY RETAINED.

MACH NUMBER 1F

AD02

NO LSXC OR WING ROCK

SUPPRESSION.

MACH NUMBER 1F

NOTE(2)

SLIGHTLY DEGRADED

OR 2F (PA MODE)

DUTCH ROLL DAMPING.

MACH NUMBER 2F

NOTE(2)

PITCH AND ROLL SAS/ARI

(UA MODE)

OFF. SPIN REC. RETAINED.

AICS MACH

AICX

REDUNDANCY LOSS ONLY.

CROSS CHECK

AICS MACH

AICX

PITCH SAS OFF. SLIGHTLY

CROSS CHECK

DEGRADED ROLL SAS.

(QBAR>1230)

INS ATTITUDE

PS,

IMUx

REDUNDANCY LOSS ONLY.

(PQVM ATTITUDE

MONITOR)

SCADC VALID TO

AD03

NO AUTOPILOT.

PITCH COMPUTER

SCADC VALID TO

AD04

NO AUTOPILOT OR MACH

ROLL COMPUTER

TRIM. MACH FAULT SET.

SCADC VALID TO

AD05

NO AUTOPILOT OR

YAW COMPUTER

AUTHORITY STOPS.

MAIN LANDING

LDGx

REDUNDANCY LOSS ONLY.

GEAR 1F

MAIN LANDING

LDGx

DEFAULTS TO UA-ARI

GEAR 2F

(SAME AS AFCS AT 15U).

(PA MODE)

MAIN LANDING

LDGx

REMAINS IN UA-ARI.

GEAR 2F

(UA MODE)

Figure 2-70. DFCS Failure Modes and Indications (Sheet 2 of 4)

CHANGE 1

2-122

NAVAIR 01-F14AAD-1

CAUTION LIGHTS

FAILURE

FUNCTIONS LOST

COMPUTER FAILURES

PITCH A CHANNEL

PS,

PC05

HALF AUTHORITY PITCH

PC,PA

SAS. NO INBOARD SPOIL-

RA, RS

ERS,

DLC, OR AUTOPILOT.

PITCH B CHANNEL

PS, PC,

PC06

HALF AUTHORITY PITCH

PA,

SAS. NO OUTBOARD SPOIL-

ERS, DLC, OR AUTOPILOT.

MACH FAULT SET.

ROLL A CHANNEL

PS, RS,

RC05

HALF AUTHORITY ROLL

RC, RA

SAS/ARI. NO INBOARD

SPOILERS, DLC, MACH

TRIM, OR AUTOPILOT.

ROLL B CHANNEL

RS, RC,

RC06

HALF AUTHORITY ROLL

SAS/ARI. NO MACH TRIM

OR AUTOPILOT.

YAW A CHANNEL

PS, YS,

YC05

HALF AUTHORITY YAW

YC, YA,

SAS/ARI. NO AUTHORITY

STOPS.

YAW B CHANNEL

PS, PA,

YC06

HALF AUTHORITY YAW

YS,

SAS/ARI. NO OUTBOARD

YA(4),

SPOILERS, DLC, OR

YC, AM

AUTHORITY STOPS.

ACTUATOR FAILURES

ROLL OR YAW

—

INDICATES ROLL OR YAW

STAB AUG

SAS/ARI SELECTED OFF.

SWITCH OFF

PITCH SERIES

PSA1

HALF AUTHORITY PITCH

SERVO 1F

SAS.

PSA2

PITCH SERIES

PSA1

PITCH SAS OFF.

SERVO 2F

AND

PSA2

ROLL SERIES

RSA1

HALF AUTHORITY ROLL

SERVO 1F

SAS/ARI.

RSA2

Figure 2-70. DFCS Failure Modes and Indications (Sheet 3 of 4)

2-123

CHANGE 1

NAVAIR 01-F14AAD-1

CAUTION LIGHTS

FAILURE

FUNCTIONS LOST

ACTUATOR FAILURES (continued)

ROLL SERIES

RSA1

RSAA/B

ROLL SAS/ARI OFF

SERVO 2F

AND

RSA2

YAW SERIES SERVO

YSA1

YSAA/B

HALF AUTHORITY YAW

SAS/ARI (GAINS DOUBLED).

YSA2

YAW SERIES SERVO

YSA1

YSAA/B

YAW SAS/ARI OFF.

AND

(PA MODE)

YSA2

YAW SERIES SERVO

YSAA/B

ROLL AND YAW SAS/ARI OFF.

(UA MODE)

INBOARD SPOILER

SP1/2

SP1/2/

AFFECTED SPOILER PANEL

ACTUATOR

L/R

INOPERATIVE.

OUTBOARD

SP3/4

SP3/4/

AFFECTED SPOILER PANEL

SPOILER

L/R

INOPERATIVE.

ACTUATOR

RUDDER

RUDA

FROZEN AT LAST

AUTHORITY ACTUA-

COMMANDED POSITION.

TOR

HORIZONTAL TAIL

HZTA

FROZEN AT LAST

AUTHORITY ACTUA-

COMMANDED POSITION.

TOR

MACH TRIM ACTUA-

MTRM

FROZEN AT LAST

TOR

COMMANDED POSITION.

PITCH PARALLEL

PPA

NO ACLS.

ACTUATOR

(ACL ENGAGED)

DLC TRIM

DLT1

NO DLC.

ACTUATOR

PITCH AUTOTRIM

PTRM

NO AUTOPILOT.

ACTUATOR

+ Lights will illuminate only if autopilot or ACL mode engaged when failure occurs.

“x” or “xx” in fault code depicts multiple numeral possibilities.

NOTE: 1. DCP CODES MAY INCLUDE AOAC, AOAT, DPSR, DPSL, AOAL, AOAR, AC28.

2. DCP CODES MAY INCLUDE AD01, AD02, AD04, CAD7, SPSR, SPSL, TPSR, TPSL, MACL,

MACR, AICX.

3. DCP CODES MAY INCLUDE SPSR, SPSL, TPSR, TPSL, LAIC, RAIC

Figure 2-70. DFCS Failure Modes and Indications (Sheet 4 of 4)

CHANGE 1

2-124

NAVAIR 01-F14AAD-1

Note

2.24.3

Autopilot

z DFCS synchronization can take up to two

The autopilot is controlled by four switches on the

seconds following a power interrupt. If the

lower half of the DFCS control panel (Figure 2-67) and the

MASTER RESET pushbutton is depressed

autopilot reference and nosewheel steering pushbutton on the

during the synchronization time, an addi-

stick grip. With all three SAS axes engaged, autopilot opera-

tional depression of the MASTER RESET

tion is commanded by placing the ENGAGE/OFF switch to

pushbutton will be required to restore spoiler

ENGAGE. Nowarmup period is required. The autopilot may

functionality.

be engaged with the aircraft in any attitude. If, however,

z Do not press and hold the MASTER RESET

aircraft attitude exceeds ±30

in pitch and ±60

in roll, the

pushbutton. Pressing and holding the

autopilot will automatically return the aircraft to these limits.

MASTER RESET pushbutton during the

Normally, IMU/INS is the prime reference and SAHRS a

synchronization time will have no effect since

backup.

the DFCS computers only recognize the

leading edge of the pulse from the MASTER

2.24.3.1

DFCS Series Actuator

RESET pushbutton, and not the fact that the

button is continuously depressed.

The series actuator is a dual-channel servoactuator that

An IBIT must always be run following POR on deck to

is controlled and commanded by the DFCS computers to

ensure full system capability is restored.

provide a low-authority input that can be mechanically over-

ridden by the pilot. Each servo of the dual actuator is moni-

Note

tored to provide failure detection and automatic shutdown of

a malfunctioning actuator channel. The remaining functional

Both pitch rate gyros are powered through the

channel will continue to provide half authority in the affected

Pitch A AC circuit breaker (LB1), while both roll

axis. In the yaw axis, the output command is doubled to

rate gyros are powered through the ROLL B AC

provide normal YAW SAS response up to the authority of the

circuit breaker (LA1). Removing DFCS Pitch A

remaining yaw series servoactuator (+/-9.5 deg). Autopilot

or Roll B computer channel AC power via one of

modes may be engageable but will have reduced authority. A

these circuit breakers for more than ~15 seconds

dual pitch series servoactuator failure results in loss of the

either on deck or airborne can result in latched

PITCH SAS. A dual roll or yaw series servoactuator failure

dual rate sensor faults requiring a POR to clear.

results in loss of the affected SAS and all ARI functions.

With landing gear retracted, the ROLL SAS will be disabled

The only situation in which an airborne power on reset

whenever ARI functions are lost. This is to minimize risk

should be considered is when a DFCS failure has resulted in

of departure if a dual failure occurs during aggressive

unsuitable controllability for landing approach (i.e., com-

maneuvering.

plete loss of spoilers and asymmetric load condition). In this

case, the risk of potential uncommanded SAS inputs is

Pitch series servoactuator failures are indicated by the

outweighed by the risk of attempting landing with marginal

PITCH SAScaution light and PA acronym. Roll series servo-

or uncontrollable flying qualities at approach speed. Refer to

actuator failures are indicated by the ROLL DGR, ARI DGR

Chapter 14.12.1 - Controllability Check.

caution lights and RA acronym for a first failure and ROLL

DGR, ARI DGR, ARI/SAS OUT caution lights and RA acro-

2.24.2

Voltage Monitoring

nym for a second failure. Yaw series servoactuator failures

The DFCS monitors voltage levels internally and no

are indicated by the YAW DGR, ARI DGR caution lights and

longer requires the VMCU. When a low voltage condition of

YA acronym for a first failure and YAW DGR, ARI DGR,

88.5 volts ac or less or a high voltage condition of 126.5 volts

ARI/SAS OUT caution lights and YA acronym for a second

ac or more is detected, the computer channel detecting the

failure. With landing gear handle up, the ROLL DGRcaution

abnormal voltage will be isolated andits associated functions

light will also be illuminated indicating inhibited UA ROLL

will be lost. Failure indications for an isolated computer

SAS. Series servoactuator failure effects and cockpit indica-

channel are shown in Figure 2-70. Depressing MASTER

tions are summarized in Figure 2-70.

RESET will restore normal system operation once the

voltage is within allowable limits.

Note

Similar to the ac power, when dc voltage drops below

Taxiing with one engine shut down and the

11.0 volts, the affected computer channel will be isolated and

HYD TRANSFER PUMP off may illuminate

the associated functions will be lost. Depressing MASTER

the PITCH SAS, ROLL DGR, ARI DGR, and

RESET will restore normal system operation once the

YAW DGR caution lights.

voltage is within allowable limits.

2-125

ORIGINAL

NAVAIR 01-F14AAD-1

2.24.3.2

DFCS Pitch Parallel Actuator

2.24.3.4

Autopilot Emergency Disengage

The DFCS pitch parallel actuator is a single-channel

Operation of the autopilot emergency disengage

electrohydraulic servoactuator that provides automatic longi-

paddle on the control stick (Figure 2-58), disengages the

tudinal control during mode I and mode IA ACLS approaches.

autopilot and DLC. Depressing the paddle with weight on

Pitch commands received by the data link are supplied to the

wheels reverts throttle system from the auto or boost mode

parallel actuator via the DFCS pitch computer. As a safety

to the manual mode; reverts the engines to SEC mode only

feature, the parallel actuator system contains a mechanical

while depressed.

force link that is designed to disconnect the actuator from the

Note

control system whenexcessive force (greater than 90pounds)

The AUTO PILOT light may or may not illumi-

is encountered at the actuator control rod, thus uncoupling the

nate when the autopilot is disengaged with the

autopilot from the ACL system. Upon ACL engagement, the

autopilot emergency disengage paddle.

parallel actuator centers itself automatically as a function of

stick position, pitch rate, and pitch attitude. Coupling with the

2.24.4

Pilot Relief and Guidance Modes

aircraft out of trim or in a climb or descent will result in

improper centering of the parallel actuator and decreased

2.24.4.1

Control Stick Steering

actuator authority in one direction. This will greatly increase

the probability of uncoupling during the approach since the

With the autopilot engaged, the aircraft may be

actuatormaycommandthecontrolsystemagainstthephysical

maneuvered using control stick steering. In control stick

stop in the direction of reduced authority and disconnect the

steering mode, the DFCS automatically synchronizes to the

force link. Similarly, it is possible for the force link to

new attitude.

disconnect during pilot OBC if longitudinal trim is not

properlysetpriortoOBCcommencement. Oncetheforcelink

2.24.4.2

Attitude Hold

is disconnected, further mode I or mode IA approaches will

Attitude hold is selected by setting the AUTOPILOT

be impossible until the force link is reset by maintenance.

ENGAGE switch to ENGAGE. To change attitude, use

control stick steering. Reengagement is achieved by releas-

ing pressure on the stick. The autopilot will hold pitch

attitudes up to ±30

and bank angles up to ±60 . Inertial

measurement unit failure will cause mode disagreement and

z It is absolutely imperative that the aircraft be

the engage switch will return off. The mode may be

trimmed hands-off in level, on-speed, wings-

reengaged using SAHRS as a reference.

level flight with landing checks complete prior

to coupling in order toachieve propercentering

2.24.4.3

Heading Hold

of the pitch parallel actuator. Engagement of

Heading

(HDG) hold is engaged by setting the

ACL in anyother flight condition will seriously

HDG-OFF-GT switch to HDG. After maneuvering the

degrade mode I/IA flight characteristics and

aircraft to the desired reference heading, release the control

may result in a force link disconnect. The

stick at a bank angle of less than ±5

. The autopilot will then

recommended method for coupling is to

hold the aircraft on the desired heading. Heading reference

engage ACL after 15 to 30 seconds of flight in

is obtained from the SAHRS via the CIU.

the landing configuration with DFCS attitude

and altitude hold engaged to utilize the DFCS

2.24.4.4

Ground Track

automatic pitch trim system.

z Commencement of OBC with longitudinal

To engage ground track, set the HDG-OFF-GT switch

trim set below zero units with flaps up or

to GT. When the A/P REF legend appears, press the

3 units nose up with the flaps down will likely

nosewheel steering pushbutton on the control stick grip.

result in a force link disconnect when the

When the A/P REF legend goes out, the mode is engaged.

control stick hits the forward stop during

Disengagement will occur if more than 1½ pounds

DFCS pitch parallel actuator checks.

lateral stick force is applied and will be indicated by the A/P

REF legend. The ground-track mode may be reengaged by

2.24.3.3

Automatic Pitch Trim

releasing the stick force and pressing the nosewheel steering

Automatic pitch trim is used in all autopilot pitch modes

pushbutton.

to trim the aircraft in order to minimize pitch transients when

Ground-track steering computations are performed by

disengaging autopilot functions. The pitch servo position is

the weapon system computer, based on inputs from the CIU,

monitored to drive the aircraft pitch trim motor at one-tenth

IMU, and SAHRS. The computer output, in the form of

manual trim rate. The pilot manual trim button on the control

ground-track error signals, is processed in the CIU, which

stick is inoperative during all autopilot operations.

ORIGINAL

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NAVAIR 01-F14AAD-1

generates steering commands to the autopilot roll axis. Bank

roll computers in response to inputs from the data-link

angles are limited to ± 30°. Failure of the INS or SAHRS will

converter and IMU. If the data-link vector discrete is present,

cause loss of ground-track steering.

the autopilot roll axis will respond to data-link heading

commands and bank angle authority will be limited to ± 30 .

Note

When the PCD discrete is present, the autopilot roll and

Performingasystemreset withgroundtrack(GT)

pitch axes will respond to data-link commands.

engaged will cause the DFCS AUTO PILOT

caution light to illuminate, which may cause the

ground-track mode to disengage.

2.24.4.7

Automatic Carrier Landing

The DFCS incorporates ACLS software with control

2.24.4.5

Altitude Hold

laws provide a vertical rate (h-dot) command system with

integrated direct lift control

(DLC). These control laws

Altitude hold mode is engaged by setting the ALT-OFF

provide corrections for glidepath deviations commanded

switch to ALT. When the A/P REF legend appears, press the

directly by horizontal stabilizer and DLC through altitude

nosewheel steering pushbutton when at the desired altitude.

rate “h-dot” feedback. Since the F-14 DFCS does not have a

This will engage the altitude hold mode and the A/P REF

direct altitude rate input from the IMU/INS, the DFCS has

legend to go out. Applying 10 pounds longitudinal stick force

incorporated the normal accelerometer (Nz) sensor to derive

will cause the A/P REF legend to appear. The mode may be

a pseudo vertical rate feedback signal by sensing motion in

reengaged by depressing the nosewheel steering pushbutton

the vertical axis. This normal accelerometer is the same

on the stick grip, when at the desired altitude, and observing

sensor used in the autothrottle approach power compensator

that the A/P REF legend goes out. Altitude hold should not

(APC) system and the autopilot altitude hold mode.

be engaged during any maneuvers requiring large, rapid,

pitch trim changes because of limited servo authority and

ACLS control of the aircraft is achieved through the

slow automatic trim rate. Disengagement of altitude hold is

autopilot by pitch parallel servo actuator and DLC com-

accomplished by applying 10 pounds or more longitudinal

mands in pitch and spoiler commands in roll. The pitch

stick force or by placing the ALT-OFF switch to OFF.

parallel actuator is utilized to command the control stick and

horizontal stabilizers to provide a large amplitude, low

Note

frequency control response. The integrated “blended” DLC

z Do not actuate in-flight refueling probe with

is utilized to provide a small amplitude, high frequency

altitude hold engaged because of large tran-

control response. This system is significantly more capable

sients in pitot-static systems sensed by the

than previous versions of compensating for varying engine

CADC.

response, winds, and/or deck motion. The DFCScontinues to

z Altitude hold performance in the landing con-

provide roll control through the spoilers only and does not

figuration with cg forward of 12% will be

capitalize on the full benefits of the automatic rudder

degraded due to rapid limiting of servoactua-

interconnect (ARI). The lateral axis is the primary limitation

tor authority. Aircrew should avoid aggressive

of the F-14 DFCS ACLS and must be closely monitored for

power or bank angle changes in this condition

any unacceptable course deviations during the approach.

or undesirable pitch attitudes may result (with-

out decoupling of AUTOPILOT switch).

Note

If the pitch parallel actuator force link is mechan-

2.24.4.6

Data-Link Vector —

ically disconnected, the A/P REF legend indicat-

Precision Course Direction

ing ACL mode engagement may go out when

coupling is attempted, but the aircraft will not

This mode is engaged by placing the VEC/PCD switch

respond to SPN-46 commands and the autopilot

to VEC/PCD and pressing the nosewheel steering push-

will then uncouple from the ACLS when the first

button. Mode engagement is evidenced by the A/P REF

pitch commands are received.

legend going out.

The F-14

DFCS ACLS control laws require the

Disengagement of the mode is accomplished by

incorporation of a software upgrade in the AN/SPN-46

application of stick forces of 7½ pounds lateral or 10 pounds

Automatic Carrier Landing System.

longitudinal, or by placing the VEC/PCD switch to OFF. If

the switch is left in VEC/PCD, the A/P REF legend will

Note

appear and the mode may be reengaged by depressing the

ACLS mode I/IA approaches are authorized for

autopilot reference and nosewheel steering pushbutton.

F-14 DFCS aircraft incorporating OFP 4.4 or

Determination of whether data link or precision course

subsequent only.

direction signals are present is made in the DFCS pitch and

2-127

ORIGINAL

NAVAIR 01-F14AAD-1

the ACL mode. DLC disengagement during an

approach will result in automatic downgrade.

When the ACL mode is engaged, the DLC

neutral spoiler position is shifted from 17.5°to

DFCS software OFP 4.4 is not compatible with

8°. In the event of a downgrade, the DLC

AN/SPN-42

systems. ACLS mode I/IA ap-

neutral spoiler position will return to 17.5°.

proaches are only authorized with AN/SPN-46

This slight transient will occur over a 1 sec

systems.

fade-in schedule so as not to result in any

perceptible change in aircraft energy and/or

2.24.4.7.1

ACL Operation

rate of descent during the approach.

PriortoACLSengagement, theaircraft shouldbe inthe

Between the time the autopilot ACL mode is

landing gear down, full flaps, speedbrakes extended

engaged (A/PREFlegend goesout) andtransi-

approach configuration with direct lift control

(DLC),

tion to command control (CMD CONTROL

autothrottle approach power compensator (APC), and auto-

legend appears), the aircraft may experience a

pilot altitude hold mode engaged.

slight altitude deviation of less than 100 feet.

Normal system operation should correct for

this deviation prior to tip-over.

Care should be taken not to couple above glide-

slope. If above glideslope or reference altitude

ACLS mode I/IA approaches are not authorized

wheninitial pitchcommands aresent, theresult-

with the THROTTLE MODEswitch inMANUAL.

ing nose down correction may cause a force link

disconnect resulting in automatic decouple and

Note

an inability to perform mode I/IA approaches

The APG-71 should be in STBY to avoid beacon

until maintenance action is performed.

interference problems.

Care should be taken not to couple after tip

With a valid ACLS coupler discrete (A/P CPLR legend),

over or prior to tip over with greater than 500

the autopilot can be armed in the ACL mode with the A/P REF

foot per minute rate of climb or descent. If

advisory legend displayed, indicating that a pilot relief mode

coupling is attempted after tip over, degraded

(in this case, ACL) has been selected, but not engaged (altitude

system performance should be expected,

hold mode will automatically disengage). The pilot can then

possibly requiring a PTOno later than 200 feet

couple the autopilot ACL mode to the data link by means of

or ½ mile on final. If excessive climb/descent

the autopilot reference pushbutton on the control stick, at

rate is established prior to coupling, system

which time, if the DFCS is functioning properly and the ACL

control authority may be insufficient to arrest

mode interlocks are satisfied, the AP REF light will be

the trend and capture reference altitude.

extinguished. The pilot should report coupled and the

The ACL mode (and autopilot) will be automatically

controller will then sendadiscrete command controlmessage

disengaged by loss of any aircraft autopilot or ACL mode

that illuminates the CMD CONTROL light. The Naval

interlock requirement, if the information stored in the data

Tactical Data System (NTDS) begins transmitting ACLS

link is not updated within any 2-second period, or the aircraft

data-link pitch and bank commands to the aircraft. The

exceeds the flightpath control envelope. The DFCS will

autopilot actuates the appropriate control surface to execute

revert to basic stability augmentation and the pilot can

the desired command, while the autothrottle APC maintains

continue the descent in mode II or mode III.

approach angle of attack by controlling the throttle setting.

Note

z Application of more than 2 to 3 pounds of

stick force while attempting to couple will

If the autopilot/ACLS uncouples after approach

cause the AUTOPILOT caution light to illu-

commencement, donotattempttorecouplewiththe

minate and coupling cannot be accomplished.

CMD CONTROL light illuminated. To do so could

It is imperative that any stick force be avoided

cause abrupt attitude changes and a possible force

while depressing the autopilot reference

link disconnect. The pilot should verbally instruct

pushbutton to preclude illumination of the

the approach controller, “downgrade to mode II.”

AUTOPILOT caution light.

Upon downgrading, the CMD CONTROL

z In the autopilot ACL mode, the ACLS control

lawsutilize DLCtoaugment glideslope control.

DLC engaged is an interlock requirement for

ORIGINAL

2-128

NAVAIR 01-F14AAD-1

light should go out. The ACL RDY and A/P

aircraft bolters or if the pilot decides to go around, the

CPLR legends must be displayed prior to any

autopilot/ACL mode is disengaged automatically by weight-

attempt at recoupling.

on-wheels or overriding the control stick, as the pilot enters

the bolter/waveoff pattern. If the aft longitudinal stick force

Until

12.5

seconds from touchdown, the landing

method is used at or inside the in-close position, the pilot

system commands the aircraft to follow a stabilized glide-

must avoid over-rotation. The waveoff technique described

slope. Inside of 12.5 seconds, the landing system commands

in Chapter 8 applies.

the aircraft to follow the vertical movement of the intended

touchdown point. As a result, some deviations from the

SLOLS glideslope will be noted with large pitching deck

motions.

Between 12.5 and 1.5 seconds from touchdown, the

A PTO initiated by autothrottle APC disengage-

approachcontrollersendsanautomatic waveoffdiscreteifany

ment with large power additions prior to uncou-

part of the carrier-based equipment fails and up to 5 seconds

pling from ACLS will result in large nose down

from touchdown if the aircraft exceeds the AN/SPN-46

commands. A force link disconnect may occur if

flightpath control envelope. Waveoff signals may also be

the control stick hits the forward stop.

issued by the final controller between lock-on and touchdown

and the landing signal officer between 1mile andtouchdown.

Approaches must be wavedoff atprecision approach weather

minimums if the pilot cannot see the meatball.

At 1.5 seconds from touchdown, the landing system

The paddle switch will disengage the autopilot.

freezes the vertical rate command and sends a bank command

Use of the paddle switch to disengage DFCS for

to return the aircraft to a wings-level attitude. The DFCS

mode IAlanding is not recommended since DLC

followsthese commands totouchdown, unlessthe pilot elects

will also be disengaged. The PITCH and ROLL

to disengage from the ACL mode via pilot takeover.

SAS switches will remain engaged.

Note

z The paddle switch will revert throttles to

MANUAL mode with weight-on-wheels.

z The paddle switch, control stick forces, orloss

If the pilot and/or LSO recognizes a course drift

of any aircraft ACL mode interlock will

immediately prior to or at command freeze, the

illuminate the MASTER CAUTION light,

pilot will be required to make a lateral correction

AUTOPILOT caution light, and ACLS/AP

toprevent unacceptable deviation fromcenterline.

caution ladder light.

Pilot takeovers (PTO) may be desired/required during

ACL mode disengagement via the autopilot reference

ACLS approaches. In the case of an ACLS mode IA approach

pushbutton or PLM pushbutton will illuminate the ACLS/AP

the PTO shall be executed prior to 200-feet altitude and

caution ladder light, but not the MASTER CAUTION and

½ -mile. All approaches must be waved off at precision

AUTOPILOT caution lights. The PLM pushbutton com-

approach weather minimums if the pilot cannot see the

mands the radar to pilot lock-on mode when the A/P CPLR

meatball. The recommended method for a PTO is via the

legend is not displayed. However, when the A/P CPLR

autopilot reference pushbutton located on the control stick to

legend is displayed, selection of the PLM pushbutton

disengage the ACL mode and the CAGE/SEAM pushbutton

disengages the ACL mode. ACL mode disengagement via

on the throttle to disengage autothrottle APC. An alternative

control stick forces or the emergency disengage paddle will

method to disengage the ACL mode is via the PLM

illuminate the MASTER CAUTION light, the AUTOPILOT

pushbutton. Another method to disengage the ACL mode is

caution light, and the ACLS/AP caution ladder light.

via the manual deselection of the ACL or AUTOPILOT

Manually disengaging the ACL mode and/or AUTOPILOT

switches. Manual deselection of the THROTTLE MODE

switches will illuminate the ACLS/AP caution ladder light,

switch will disengage the autothrottle APC. Manual deselec-

but not the MASTER CAUTION and AUTOPILOT caution

tion of these switches may be difficult to accomplish,

lights.

especially during the final stages of the approach. The paddle

switch will disengage the ACL mode and autopilot, but will

2.24.5

DFCS Test

also disengage DLC. As a last resort, overriding the control

stick with 10 pounds longitudinal or 7 pounds lateral control

The DFCS has several self-test modes. These include

stick force will disengage the ACL mode and 11 pounds of

power-up BIT (PBIT), initiated BIT (IBIT), and automatic

force per throttle will disengage the autothrottle APC. If the

BIT (ABIT). The results of these tests are indicated by the

2-129

ORIGINAL

NAVAIR 01-F14AAD-1

illumination of applicable caution lights, maintenance file

serve as an indication that IBIT is running. The IBIT

acronyms, and DFCS control panel (DCP) fault display

sequence will continue with the pitch trim check (slow longi-

codes.

tudinal stick motion), PITCH SAS actuator check (no longi-

tudinal stick motion), the pitch parallel actuator check (rapid

2.24.5.1

DFCS Power-up BIT (PBIT)

longitudinal stick motion), the individual spoiler operation

check (from right to left), and the ROLL and YAW SAS

A DFCS power-up BIT is an automatic function of

actuator checks. DFCS IBIT concludes with disengagement

the DFCS that is initiated when power is initially applied

of the AUTOPILOT switch, activation of the rudder pedal

to the aircraft. Power-up BIT is completed in approximately

shaker check, and illumination of an alternating test pattern

2 seconds. Following a successful power-up BIT, the flight

to test all pixels of the DCP LED fault display.

control computers will synchronize and enter the operational

flight program mode following depression of the MASTER

Premature termination of the IBIT sequence will cause

RESET pushbutton. Failure of power-up BIT will result in

the ACLS/AP and AP REF lights to stop flashing and leave

illumination of caution lights and DCP fault display codes

all other DFCS caution lights illuminated. The DCP fault

(Figure 2-71) associated with the failed computer(s) that will

display will indicate

“ABRT” when IBIT is terminated

not reset with MASTER RESET. The failed computer(s)

prematurely.

will remain isolated and will not enter the operational flight

program mode.

Following completion of a successful IBIT, all DFCS

caution lights will be extinguished, the AUTOPILOT switch

2.24.5.2

DFCS Initiated BIT (IBIT)

will be OFF, and the DFCS will automatically enter the

operational flight program mode. The DCP will display a

“PASS” indication in the DCP fault display.

ADFCSInitiated BITis athorough preflight indication

of DFCS performance and can be obtained during poststart

OBC or a DFCS BIT. All SAS switches must be engaged,

Following an IBIT with one or more failures, caution

weight-on-wheels, flaps extended greater than 25°or wings

lights and acronyms for the detected failures will be dis-

swept aft of 62°, and ANTI-SKID SPOILER BK switch OFF.

played. The DCP fault display will indicate “NO GO” and

If one of these interlocks is not satisfied the DFCS will not

fault codes for the specific failed WRAs can be viewed using

enter the IBIT ARM state. The AUTOPILOT switch must be

the INC and DEC pushbuttons. The DFCS will automatically

engaged to test autopilot functions and can only be engaged

enter the operational flight program, even though IBIT has

in the IBIT ARM state.

detected failures. Depression of the MASTER RESET push-

button will extinguish caution lights and acronyms, but will

Longitudinal trim should be greater than 0°for flaps up

not clear DCP IBIT fault codes. Any discrepancies detected

and greater than 3° for flaps down. The MASTER TEST

by IBIT may still exist even though caution lights have been

switch must be selected to “IBIT ARM” by raising and rotat-

extinguished with MASTER RESET. Another IBIT must be

ing to the “OBC” or “DFCS BIT” position. The DCP fault

completed to ensure proper system operation.

display will alternate between an “IBIT” and “ARM” indica-

tion to confirm that IBIT is in the armed state. The AUTO-

PILOT switch can be engaged at this time in order to test

autopilot functions during IBIT. If the INC or DEC push-

buttons are depressed in the IBIT ARM state the DCP will

indicate any existing fault display codes. In this case, the

Following an IBIT, a MASTER RESET will

DFCS is still in the IBIT ARM mode and the depression of

clear the IBIT caution/advisory light failure indi-

a MASTER RESET will restore the IBIT ARM codes to the

cations, but will not clear the FAULT DISPLAY

fault display, but is not required. A DFCS IBIT test sequence

IBIT codes. This does not indicate that the fail-

will commence upon depression of the MASTER TEST

ures detected during IBIT are resolved. The

switch in the “DFCS BIT” position, or a complete OBC

DFCS should not be considered fully opera-

encompassing all aircraft functions may be subsequently

tional. Only the successful completion of another

initiated by the RIO with the MASTER TEST switch in the

IBIT can verify proper system operation.

“OBC” position.

Note

The DFCS IBIT sequence will commence with the

following: The DCP fault display will alternate between an

Spoiler actuator IBIT tests are run only with the

“IBIT” and “RUN” during the entire IBIT run sequence. All

wings forward and flaps down. During IBIT,

ten DFCS caution lights (including the HZ TAIL AUTH,

spoilers are deflected individually, one at a time

RUD AUTH, and SPOILERS lights) will illuminate and the

starting with the right no. 4 spoiler.

ACLS/AP and AP REF lights will flash once per second to

ORIGINAL

2-130

NAVAIR 01-F14AAD-1

CODE

REMARKS

ARM

IBIT is armed awaiting BIT initiate (alternates between “IBIT” and “ARM”).

RUN

Indication that IBIT is running (alternates between “IBIT” and “RUN”).

ABRT

IBIT has aborted before it completed.

PASS

IBIT passed without any failures.

NOGO

IBIT completed with failures.

No current, in-flight, or IBIT failures have been logged.

FAIL

Failures following this code are current fault indications.

FLT

Failures following this code were logged during flight.

IBIT

Failures following this code were logged during the most recent IBIT run.

END

Failure list end has been reached.

CLR

Failure clearing sequence has been started.

----

Indicates completion of clearing sequence.

Figure 2-71. DFCS DCP System Display Codes

If the pitch parallel actuator is functioning properly,

2.24.6

DFCS Control Panel Fault Reporting

large

longitudinal control stick deflections should be

observed during IBIT. An IBIT with the flaps down requires

The DFCS control panel (DCP) incorporates an LED

a longitudinal trim of 3

or more noseup; an IBIT with the

alphanumeric fault display. This fault display is intended for

wings at 68

requires notless than 0

noseup.Apitch parallel

ground use only to assist in the troubleshooting and repair of

actuator force link disconnect during IBIT is indicated by

the DFCS and related components. The DCP will not display

illumination of the AUTOPILOT caution light, a PA

anyfault data with weight off wheels. DFCSoperational DCP

acronym, and the absence of large control stick deflections.

system display codes are listed in Figure 2-71. The fault dis-

It is possible for the force link to be partially disconnected;

play will group faults into three categories: currently existing

that is, disconnected mechanically while electrical

faults (FAIL), faults detected in flight (FLT), and faults that

continuity is maintained. If this has occurred, the

are detected during initiated BIT (IBIT). Fault codes will be

AUTOPILOT caution light or PA acronym may be absent

displayed in order by repeated depression of the INC push-

after IBIT, but no large stick deflection will be observed. The

button. Current failures will be displayed first followed by

implications of this condition are the same as for a total

in-flight detected failures, and any IBIT detected failures.

disconnect (no ACL capability).

This will be indicated by “FAIL” followed by any current

failures, then “FLT” followed by any in-flight logged fail-

2.24.5.3

DFCS Automatic BIT (ABIT)

ures, and finally by “IBIT” followed by failures detected

during the last executed IBIT. If there are no failures in a

A DFCS Automatic BIT provides continuous failure

particular group that group’s header will not be displayed.

monitoring of the DFCS. Test coverage for ABIT is not as

When all faults have been displayed, “END” will be dis-

extensive as IBIT and should not be used as a replacement for

played. The INC or DEC pushbuttons may be used to scroll

performing a pre-flight IBIT. ABIT failures will be recorded

forward or backward through the fault codes. If no failures

in a maintenance data store and are listed on the DCP fault

have been logged, depression of the INC or DEC pushbutton

display following the “FAIL” and “FLT” headers. Depending

will display a “GO” indication.

on the severity of the problem detected, functionality may be

lost and the appropriate caution/advisory lights illuminated

and acronyms displayed.

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F-14D. FLIGHT MANUAL (2004) - page 3